ntrs.nasa.govSharp Leading-Edge Extension RESE~RC~ MODEL AND LEADING-EDGE DEVICES 60-Deg Delta Wing...

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NASA Contractor Report 165923

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NASA-CR-165923 19830011443

Subsonic Balance and Pressure Investigation of a 60-Degree Delta Wing With Leading-Edge Devices

Stephen A. Tingas Dhanvada M. Rao

MAY 1982 NCCI-46

NJ\SJ\ National Aeronautics and Space Adminisiration

Langley Research Cooter Hamptor., Virginia 23665

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LIST OF TABLES

LI ST OF FIGURES

LI ST OF SYMBOLS

INTRODUCTION

BAC~GROUND .

Chordwise Slots

Fences .. . .

TABLE OF CONTENTS

Pylon Vortex Generators

Leadi~g-Edge Vortex Flaps .

Sharp Leading-Edge Extension

RESE~RC~ MODEL AND LEADING-EDGE DEVICES

60-Deg Delta Wing Model

Slot Contours

Fences

Pylon'Vortex Generators.

Leading-Edge Vortex Flaps


WIND-TUNNEL FACILITY

EXPERIMENTAL PROGRAM

DATA REDUCTION . . .

PRESENTATION OF DATA

Page

iii

iv

. . . . x

1

5 ~";,

5

6

8

9

11

13

13

20

. .~.:~. 22

22

22

29 ~ i

31

32

35

40

~f3-!97/(:#·

RESULTS AND DISCUSSION

60-0eg D~lta Wing

Chordwise Slots .

Fences

Pylon Vortex Gener~tors

Sharp Leading-Edg~ Extension


CONCLUS IONS

REFERENCES

'':

< .:

ii

Page

42

42

60 ~ ..

76

91

121

144

196

200

;,'

,..".,.\

"

I.

II.

III.

LIST OF TABLES

Pressure orifice locations

Test summary

Vortex generator zero~a drag characteristics

iii

Page

17

33

96

1.

2.

3.

4.

5.

6.

7.

8.

9. - - ---

10.

1l.

12.

13.

14.

15.

16.

17.

18.

~ ;

~~ ~; -;:

.~: . '-::- i..:;

:"'1ST OF FIGURES

Photographs of -research model and l~~.di ng-edge devi ces

Drawing of 60-deg delta wing research model

Drawing of slot contours tested

Drawing Of' fences tested . . .

Drawi ng of pylon vortexgenerato,.rs tested

Drawing of pylon vortex generator with missile tested

Drawing of leading-edge vortex flaps tested

Drawing of sharp leading-edge extension tested

Basic wing force and moment characteristics - - .-~ -- ---.- -

Basic wing leading-edge static pressure variations and CPLE derived separation boundary

Static pressure distributions around the basic wing leading edge ........ -.. .

Basic wing spanwjse leading-edge thrust distributions

Basic wing leading-edge thrust characteristics

Basic wing induced drag characteristics

Comparison of;b~~ic.wing experimental data with VLM-SA theory . . . . . . . . . . . . .

Basic wing leading-edge static pressur~-thrust relationship ... '.' .. ; ...... .

Force and moment characteristics of single and multiple chordwi se slot confi gura ti on s . . . . .

Local static pressure effects of the chordwise slot at a = 160 ........ .

19. Local ~uction effects of the chordwise slot

.. :. .

~

iv'

'P~a:ge=-

14

16

21

23

24

25

26

30

43

46

47

51

53

55

57

58

61

63

65

20. Spanwise leading~edge static pressure distributions for single and multiple chordwise slot configurations

21.

22.

23.

Effects of single and multiple chordwise slots on leading-edge thrust ....... .

Effect of multiple chordwise slots on leading-edge static pressure-thrust relationship ..... .

Effects of internal contouring on chordwise slot performance .. . . . . . . . . . . . . . . ;.' .

24. Force and moment characteristics of single fence and slot-fence configurations

25. Local static pressure effects of the fence at a = 160

26. Effects of the fence and slot-fence combination on leading-edge separation (C PLE derived)

27. Effects of the fence on leading-edge thrust

28. Induced drag characteristics of single fence and slot-fence configurations .....

29. Effect of the fence on leading-edge static pressure-

" Page

66

67

70

73

77

79

81.

82

84

thrust relationship ................... ~ 86

. 30. . Effects of slot-fence combi nati ons on 1 eadi ng-edge

31.

32.

33.

34.

thrust . . . . . . . . . . . . . . . . . . . . .

Effects of slot-fence combinations on spanwise leadingedge static pressure distribution at a = 160

....

Force and moment characteristics of single and multiple pylon vortex generator configurations

Effect of the pylon vortex generator on leading-edge separation (C pLE . derived)". . . ...

Local static pressure effects of the pylon vortex generator . . . . . . . . . . . . . . . . . . .

35. Effect~ of single and multiple pylon vortex ~enerators

88

89

92

98

99

on leading-edge thrust. . . . ............ 101

vi

Page

36. Spanwise le"ading-edge static pressure distributions for single and multiple pylon vortex generator configurations at a = 140 . . . . . . . . . . . . . . . . . 103

37. Effect of the pylon vortex generator on leading-edge static pre~sure-thrust relationship ............ 104

38. Performance effects of pylon" vortex gener"ator 1 eadi ng-edge length reduction ., . . . . . . . . . . . ... 108

39. Outboard stati~ pr~ssure effects of pylon vortex generator )eading-edge length reduction . . . . . 110

40.

41.

Performance effects of pylon vortex generator chord reducti on . . . . ." . . . . . . . . ". . . . . . .

Effects of pylon vortex generator chord reduction on leading-edge thrust ........... .

42. Performance effects of py16nvortex generator

112

114

diagonal cutback . . . . . . . . . . . . . . . . . . . 116

43. Effects of pylon vortex generator diagonal cutback oh leading-edge thrust . . . . . . . . . . . . .. . . . . 117

44. Performance of an extend~d chord pylon vortex generator utilized as ~ carrier of slender external stores ...

45. Force and moment characteristics of a sharp leading-edge

46.

47.

48.

extension configuration .......... .

Effects of the sharp leading-edge extension on static pressure distributions around the leading edge at a = 120 . . . . . . . . • . . . . • . • • .

Leading-edge suction effects of the sharp leadingedge extension . . . . . . . . . . . . _ . .

Effects of the sharp leading-edge extension on static pressure distributions around the leading edge at a = 220 • . . . . • . . • • . • • . • • . .

49. Effects of SLEE extension on spanwise distribution of leading-edge thrust ..... .

50. Effects of SLEE extension on performance

1 1 " ll~

122

126

128

130

131

133

~

vii

Page

51. Effects of SLEE extension on static pressure distributions around the leading edge at a = 120 . . . . . 135

52. Effects of SLEE extension on leading-ed~e thrust 136

53. Performance comparison of the fence and chordwise slot at the SLEE apex . . . . . . . . . . . . . . . . . . . . . . 137

54. Performance effects of the fence and chordwise slot in combination with the SLEE . ..... . . -... 139

55. Upper surface tuft photographs of 0.48 cm ext. SLEE with F-2 at n = 0.25 configuration with and without an F-2 fence at n = 0.625 (a = 140 ) ..•...•.

56. Effects of the fence and chordwise slot on leading-edge thrust distribution along the SLEE ...

57. Leading-edge static pressure variation and flow development on a 300 + VF-3 configuration

_ 58. Side tuft photographs of 300 + VF-3 configuration at a = 60 and 100 . • . . • • • . • . . • • • . -. • •

59. LocaJ design point distribution (C PLE 30 + VF-3- ............ . derived) along

60. Side tuft photographs of 600 + VF-3 configuration at

· 140

· 142

· 145

· 147

· 149

a = 140 , 160 , 180 , and 230 . . . . . . . . . . . 150

61. Effects of 300 + VF-3 on wing leading-edge thrust 152

62. Effects of vortex flap deflection angle on force and moment characteristics. . . . . . . . . . . . . . 153

63. Effects of vortex flap deflection angle on longitudinal stability characteristics . . . . . . . . . . . 156

64. Comparison of 30° + VF-3 experimental data with VLM-SA theory . . .. ....... 158

65. Thrust contribution from 30° + VF-3 flap surface 159

66. Upper surface tuft photographs of 300 + VF-3 configuration at a = 140 and 170 .... 161

vii i

Page"

67. Effects of upward vortex flap deflection on force and moment characteristics. . . ..... 164

68. Upper surface tuft photograph of 30° t VF-7 "configuration at a= 6° ••.••. ~. . . . ... 166

69. Effects of "inverse" tapered flap chord distribution and apex shape on vortex flap performance . . . .. . ... 168

70. Effects of inverse tapered vortex flap chord distribution on static pressure distributions around the leading edge at a = 120 . . . . . . . . . . . . . ,.. .. 169

71. Upper surface tuft photographs of 30° f VF-2 and VF-3 configurations at a = 160 .............•... 171

72. Effects of inverse tapered flap chord distribution on local design point distribution (C PLE derived) along the vortex flap .......... ~ ...•....... 172

73. Effects of inverse tapered vortex flap chord distribution and apex shape on flap thrust parameter .......... 173

74.

75.

76.

77.

Effects of segmentation and flap geometry on vortex fl ap performance (6 = 300 ) . . ... . . . . . . . . . . .

Effects of segmentation and flap geometry on vortex flap performance (6 = 450 ) ..•.•..••••.•• .

Effects of segmentation and vortex flap geometry.on flap thrust parameter

Side tuft photographs of 300 f VF-3, VF-4, VF-5, and VF-6 configurations at a = 140 .

· 175

178

· 179

• 180

7S. Upper surface tuft photographs of 300 f VF-3, VF-4, VF-5, and VF-6 configurations at a = lSo .....•.... lS3

79.

so.

Effects of segmentation and flap geometry on local design point distribution (C PLE derived) along the vortex flap ............... .

Effect of differential vortex flap deflection on rolling moment characteristics ....... .

· lS4

· 186"

Sl. Performance effects of the fence and chordwise slot in combination with the vortex flap . . . . . .... lS8

82.

83.

84.

85.

Side tuft ph6tographs of 300 + VF-3 configuration with and without a slot in the flap at n = 0.625 (a = 14°) .... . . . . . . . . . . . . . . . . .

Effect of a chordwise slot in the vortex flap on spanwise local design point distribution (C PLE derived) ....

Effect of a chordwise slot in the vortex flap on flap thrust parameter ...... .

Comparison of sharp leading-edge extension and vortex flap performance . . . . . . . . . . .. ....

ix

Page

. . 190

191

. 192

195

x

LI ST OF SYMBOLS

A - area, cm2

AR aspect ratio

a length of pressure panel along leading edge, em

b span, cm

BW basic wing

CA axial force coefficient, Axial force/qroSref

CAo axial force coefficient at a = 0°, Axial force/qooSref

Co drag coefficient, Drag/qooSref

COo drag coeffieient at a = 0°, Drag/qooSref

CO i induced drag coefficient, Induced drag/qooSref

CL lift coefficient, Lift/qooSref

C1 rolling moment coefficient, Rolling moment/qooSrefb

Cm pitching moment coefficient, Pitching moment/qooSrefc

Cml pitching moment coefficient about original moment reference center indicated in figure 2, Pitching moment/qooSrefc

CN normal force coefficient, Normal force/qroSref

CNo

Cp

normal force coefficient at a = 0°, Normal force/qooSref

static pressure Coefficient, (p - p )/q ro ro

Cs suction force coefficient, Suction force/qooSref

CT leading-edge thrust coefficient, Thrust/qooSref

Cx fl ap thru st pa rameter, /m2

cr root chord, cm

c mean geometric chord, cm

cg center of gravity

"

o drag, N

e spanwise efficiency factor

ext. extension perpendicular to leading edge, em

F fence

K

L

L. E.

9v

PO

P

Poo

qoo

S

SC

SF

SLEE

STA

51

T. E.

Voo

VF

VG

VLM-SA

X

XI

induced drag parameter, lie

1 ift, N

leading edge

semispan leading-edge length, cm

percent drag redu~tion

static pressure, N/m2

free-stream static pressure, N/m2

free-stream dynamic pressure, N/m2

planform area, cm2

slot contour

suction force, N

sharp leading-edge extension

spanwise pressure station

chordwise slot in vortex flap

trailing edge

free-stream velocity. m/sec

leading-edge vortex flap

pylon vortex generator

Vortex Lattice Method with Suction Analogy code

chordwise distance from wing leading edge, cm

distance from vortex flap apex along wing leading edge, em.

xi

xii

x chordwise distance from wing apex, cm

Y' perpendic~lar distance from wing leading edge, cm

y spanwise distance from plane of symmetry, cm

z vertical distance fro~ chordline, cm

a angle of attack, deg,

aD departupeangle of attack, deg

8 ~ortex flap deflection angle, deg

Aleading~edge sweep angle, deg

n spanwise coordinate in fraction of semispan

Subscripts

avg average

cm center of moments

des local design point

front frontal

LE 1 ead i ng edge

loc 1 oca 1

min minimum

ref reference

sep _.leading-edge sepa~atio~

tot total

INTRGDUCTION

Modern aircraft designed for supersonic flight are often fitted

with thin, highly swept delta wings to take advantage of reduced super

sonic wave drag. At high subsonic speeds, where goo~ maneuverability

and high accelerations are desirable, thin delta wings provide added

advantages. These include high drag divergence Mach number and high-g

maneuyerability due to increased lift generated by stable leading-edge

vortices. These vortices, which result from leading-edge flow separa

tion, and subsequent flow reattachment on the upper surface produce the

added advantage of sustained performance capability at high angles of

attack. However, the large induced drag penalty characteristic of delta

wings in high-lift flight limits the aircraft's maneuver capability at

high subsonic and transonic speeds. This increase in the induced drag

is due to early 1eading~edge flow separation, which causes a total loss

of the leading-edge suction associated with attached flow around a blunt

leading edge. Consequently, excess engine thrust must be used to

counteract the drag increase; reducing the acceleration capability of

the aircraft. In addition to the drag penalty, the spread of vortex

origin from the tip to the apex of the wing with increasing angle of

attack produces severe longitudina~ instability in the mid-a range, as

the center of vortex lift moves ahead of the wing center of gravity

(ref. 1).

One method of alleviating the induced drag and stability problems

associated with highly swept delta wings is through leading-edge flow

2

control to delay the onset of leading-edge flow separation. Attached

flow ata blunt leading edge all~ws for partial recovery of leading

edge suction, providing a reduction in drag and alleviation of the

longitudinal instability resulting from leading-edge vortices. The use

of a cambered leading edge has been one method of maintaining attached

flow to higher angles of attack. However, severe drag penalties at

supersonic c~uise speeds and the added weight and mechanical complexity

involved with the use of a variable camber leading edge make this method

less desirable. Leading-edge devices such as flaps and slats have also

been used for the purpose of maintaining attached flow; however, they

have not been overly successful on highly swept wings (ref. 1). Fixed

leading~edge devices such as pylon vortex generators, fences, and chord

wise slots (notches) have previously been used for alleviation of sta

bility problems of highly swept wings. However, recent research has

shown that these devices also have potential for drag reduction by

maintaining attached fl ow to hi gher angl es of attack and,thus, seem

to bea practical solution to the induced drag problem (ref. 2).

An alternative to the conventional approach of drag reduction

through attached flow is accomplished through the use of leading-edge

vortex flaps and sharp leading-edge extensions (SLEE) (refs. 1,2,3,

and 4). These devices utilize the nat0ral tendency toward separation by

forcing vortex formation and using the resulting suction forces for drag

reduction. The downward-deflected vortex flap generates a coiled vortex

whose suction force acts directly on the forward face of the flap, pro

ducing thrust and lift force components. The sharp leading-edge

"

:l.

extension makes use of the same typ\:: of induced vortex; hO\tJf:ver, by

keeping the vortex directly ahead of the leading edge, the suction

effect produces strictly a thrust component and, thus, a reduction in

drag. The flow mechanisms of the vortex flap and sharp leading~edge

extension are shown below:

Leadi ng-edge vortex flap Sharp leading-edge extension

Sketch A

3

This report presents the results of a wind-tunnel investigation

undertaken to examine the potential for further drag reduction through

refined versions of devices such as fences, chordwise slots, pylon

vortex generators, leading-edge vortex flaps, .and sharp leading-edge

extensions. Since previous research had established the effectiveness

of these devices, the present investigation was primarily concerned with

modifications to overcome their limitations. Methods of reducing the

low-angle-of-attack drag penalty while maintaining the effectiveness of

the devices at higher angles of attack were studied. Results of previ

ous research (ref. 5) were used to design and test certain device com

binations which were believed to have drag-reduction capabilities beyond

4

those of the individual devices. In addition, several novel configura

tionswereincluded to determine the feasibility of concepts sucn as the

use of pyl~n vortex " generators as carriers .of slender external stores

(such as air-to-air missiles) and leading-edge vortex flaps for roll

augmentation and as "drag" devices for landing purposes.

Use of commercial products or names of manufacturers in this report

does. not constitute official endorsement of such products or manufac

turers, either expressed or implied, by the National Aeronautics and

Space Administration.

5

BACKGROUND

This section presents a brief discussion of the basic flow

mechanisms of the devices and results of previous research ~h;ch led to

the present investigation.

Chordwise Slots

Typical high-angle-of-attack performance improvements through the

use of chordwise slots cut into swept leading edges are presented in

reference 5. The slot flow mechanism results from the high ve.]ocity jet

sheet emanating from the slot due to the natural flow of air from the

lower to the upper wing surface (sketch B). It is believed that the

~

Spanwise boundary @ layer flow

---------------" I ---------

Root Leading edge

Sketch B

Tip

vortex shed from the outboard edge of the slot acts to obstruct the

span\lJise boundary layer flow on the upper surface, reducing boundary

layer buildup and, thus, separation tendencies near the wing tip. In

addition, the'sense of rotation of the slot vortex opposes that of the

6

primary vortex,; thereby hindering its inboard movement and growth. It

is thi s "compa)~tmentati on" effect of the chbrdwi se slotwhi ch prompted

its use in combina't;ion' withth'e vort~x flaps and sharp leading-edg.e

extension in the pres~nt investigation. The pri nci pa 1 concern , however,

was alleviation of a sudden loss of slot 'effectiveness at the higher_

angles of attack and a low-a drag penalty arising from pressure drag on

the vertical' face and friction on the internal wetted area of the slot.

Through internal contouring of the device, an attempt was made to reduce

this law-a drag and maintain sloteffectiveriess to higher a. In an

attempt at further performance improvements, a verN limited study of the

effects of slot depth and v:Jidthwas also included.

Fences

The use of fences on delta wings has tra-d-itionally been as a fix

faY' longitudinal instability. Previous research, however, ha's shown

these devi'ces to be'effective in the role of drag reducers at high

angles of attack, aiding in the alleviation ·of severe lift-dependent

drag pena Hi es. Typical performance improvements \,Iith si ngl e and

multiple fences similar to those presently tested are given in

reference 5.

Fence flml/.mechanisms are described in detail in reference 6 and

briefly summarized here. First, the fence forces the swept wing upper

surface isobars, normally parallel to th:e leading eoge, to be ut1swept

locally, reducing suction peaks and pressure gradients outboard, with

opposite effects inboard. Sl,I bsequently, outboard staTl and loss of

7

leading-edge suction are delayed. The adverse effects inboard at'e of

minor concern due to lower prevailing upwash. The fence also acts as an

obstruction to the spanwise boundary layer flow on the upper wing sur-

face, further delaying the onset of separation near the tip. However,

a loss of effectiveness is observed at high a, possibly due to an

accumulation of viscous fluid on the inboard side of the device. It

was believed that the flow through an adjacent inboard slot would IIblmv

off" this viscous accumulation (sketch C) and was, thus, tested here.

Boundary

Viscous accumUlatiOn)

----;-

_____ L~ayer ==;---- \ \111- - ---

Root

. ~

Leading edge

Fence only

Tip Root

1~ Slot-Fence combination

Sketch C

Tip

Finally, the fence impedes the inboard movement and growth of the pri-

mary vortex, allowing for the formation of a second, undisturbed

leading-edge vortex inboard, with both acting primarily on the wing

leading edge. For this reason, fences were also selected in the

present investigation ~or use in combination with the vortex flaps and

sharp 1 eadi ng-edgeexte'ns ion.

Pylon Vortex Benerators

The pylon vortex generator relies on the formation of a vortex

8

originating at its swept-forward upper edge as a result of the effective

angle of attack of the device relative to the wing leading-edge cross-

flow. Except at the lowest angles of attack, this vortex travels over

the wing upper surface and rotates in a sense so as to act as a barrier

to the spanwise boundary layer flow, while at the same time inducing a

downwash velocity outboard, with a subsequent delay in outboard stall

(sketch D). In addition, the rotating motion of the vortex promotes a

Root

/

00':1--/8'''''

Sketch 0

(ref. 5)

"

certain degree of boundary layer energization through turbulent mixing

of viscous boundary layer fluid ~.jith high-momentum fluid from the

9

external stream. Again, the result is increased resistance to separa

tion (ref. 7). However, a drag penalty is paid at low angles of attack

since the leading-edge cross-flow is not yet of sufficient magnitude to

cause vortex formation. The present investigation attempts to reduce

the vortex generator low-a drag through systematic size reductions

(basically, lower and aft edge removal), with minimal sacrifice of

high-a performance. The constant 300 forward sweep and 100 toe~in angle . .

of these devices were optimum values selected on the basis of previous

research (ref. 5). In addition, the charactedstic pylon shape of the

vortex generators suggested their possible use also as carriers of

slender external stores (such as air-to-air missiles); therefore, the

effect of a simulated missile attached to the lower edge of a vortex

generator on its aerodynamic effectiveness was investigated.


Results of previous tests on leading-edge vortex flaps and detailed

descriptions of their aerodynamic mechanisms are presented in refer-

erices 1 and 3. In essence, the vortex flap relies on the prevailing

upwash ahead of the wing leading edge to force separation and formation

of a coiled vortex whose suction effect acts on the flap, producing

aerodynamic thrust and lift components (see sketch A in INTRODUCTION).

By maintaining this sweep-stabilized vortex on the flap along its entire

spanwise extent, with the flow reattachment position ideally at the

wing-flap junction (knee), attached flow is maintained on the wing

upper surface. Flow entrainment and increasing upwash, however, cause

~

10

the flap vortex 'to grow and rni grate onto the wing surface wi thi ncreas-

ing angle- of attack and outboard di.stance. In an attempt to maintain

the ideal flow condition along the ~nt~re flap span, an inverse tapered

flap was selected in the present investigation for comparison with a

constant chord flap, under the assumption that increasing chord outboard

cou ldbetter :accominoda te the con i cal fl ap vortex. Segmented fl aps of

various planforni shapes (constant chord, parabolic, and inverse tapered)·

were also included to further assist in this matt'er through the forma

tion of a distinct vortex on· each segmeht, each acting primarily on the

flap surface. The use of fences and chordwise slots in combination with

the vortex flaps was based on 'their compartmentapoh_ effect, again with

the segmehtatian of the primary vortex into two smaller, undisturbed

vortices (sketch E). The ability of the vortex flap to modify the wing

without slot

Sketch E

with slot (or fence)

i':t

~~

~

11

spa~wise lift distribution suggested that sizable rolii~g moments might·

be obtained by means of asymmetric flap deflections at high angles of

attack, when other control surfaces are degt'aded by flow separation.

Thi s concept was tested along with the pass i bil ity of adapt; ng vortex

flaps for aerodynamic braklng at landing through appropriate control of

vortex suction forces (sketch F).

Suction

Sketch F


Upward deflection

Large downward deflection

Results of preliminary research on a sharp leading-edge extension

are given in reference 5. This device, operating on the same principle

of forced separation as the vortex flap, derives its drag-reduction

capabiliti~s from a tightly coil~d vortex maintained ahead of a blunt

12

. leading ,edge, utilizing its 5uction:effect to obtain a thrust force

(ref, 8) .. Ideally, flow reattachment .occurs just aft of the wing

leading~e9ge curvature (see~ketch A in INTRODUCTION). However, a down

stream expansion of the vortex cor~ due to flow entrainment leads to

eventua 1 mi gration of the. reatta·chment poi nt onto the wing upper surface

with increasing,~ngle of attack and. outboard distance. The resulting

upper surface separated .nOVY and asso,ciated drag increase act to

partially nullify th~ thrust derived from the suction effect of the

device. The compartmentationeffect Of fences and chordwise slots

located at various positions along the SLEE was utilized in the present

investigation for the purpose of delaying this growth of the leading

edge vortex. In addition, tests were performed to determine the optimum

SLEE extension producing the ideal vortex size and position just

described.

RESEARCH MODEL AND LEADING-EDGE DEVICES

The following is a description of the research model and leading

edge devices used in the investigation. Actual photographs of the

sting-mounted model and selected devices appear in figure 1.

60-De9 Delta Wing Model

13

Figure 2 shows a drawing of the wooden, 60-deg cropped delta wing

model used in the investigation. The flat-plate wing has semi-elliptic

leading edges (ellipse ratio of 26.7 percent) with uniform leading-edge

radius of 0.231 cm.

The right-hand wing panel was equipped with six chordwise rows of

static pressure orifices around the leading edge at approximately the

20, 33, 45, 57, 70, and 82 percent semispan positions. Each station

consisted of four orifices on the upper surface, four on the lower

surface, and one near the leading edge. Thirteen additional orifices

were provided along the span at the leading edge (X = 0) and one on the

wing upper surface near the trailing edge. The pressure orifices were

fed by 0.10 cm outside diameter metal tubing, which was protected by a

removable metal base plate on the lower surface of the wing. ' The

orifice locations are given in Table I.

The research model also had six 5.08 cm long chordwise slots on

either leading edge at approximately the 25, 37.5, 50, 62.5, 75, and

87.5 percent semispan positions. The slots were meant for holding the

,leading-edge devices but also used as "devices" themselves,.being sealed

when not in use.

14

(a) Basic wing. ( b ) Open slot,.

(c) Vortex generator. (d) Vortex generator wjmissile.

Figure 1.- Photographs of research model and leading-edge devices.

(e) Down-deflected leading-edge vortex flap.

(g) SLEE w/fences.

(f) Up-deflected leading-edge vortex flap.

(h) SLEE w/slots ..

Figure 1.- Concluded.

15

Slot (0.094 wide)

6

Moment reference center

z

CHORDWI5E, SLOTS /

SLOT in-(%)

1 26.40 2. 38.68 3 ',51. 04 l\ 63.36

1 5. 3 5. 61 '" ,6 . 87.89

19.05

PRESSURE OR IFI CE'S

CHORDWI SE ...L ROW ' b/2 ('}'o)

J' 19.97 2 ,32.78 , J ;44.79 4 57.21 ' 5 69; 64 . 6; 81. 75

33.02

'-r--'+--'JJ 1.90

RESEAR CHMODEL

Planform area ........................ 3263.9 em 2

Aspect ratio ........................... 1.60 Taper ratio ............................ 0.18 Mean geometric chord ................. 52.18 em Moment reference center .............. X' 32. 38 cn

(from wi ng apex) y • 0 zoO

Figure 2.- Drawing of 60-deg delta wing research model. Dimensions in centimeters.

/ .

.~~".j

TABLE 1.- PRESSURE ORIFICE LOCATIONS 17

no 109 ----

Station 1 ISlA 1) f

106~ I

,...

105 . . 11M

103 102

ORIFICE NUMBER X(cm) y(cm)

101 0 I 6.03

102 4.07 7.29 103 2.38 104 1.41 105 0.51 106 o· 107 0.57" 108 1.46 109 2.45 110 4.34

111 0 8.38 * 112 0 10.69

113 3.94 11.95 114 2.44 .

1 115 1.52 116 0.69 117 0 118 0.56 11.94

-

X

z(cm)

0 I -0.96 -0.77 -0.61 -0.37 -0.02 0.36 0.61 0.78 0.98

0 0

-0.97 -0.79 -0.62 --OAl -0.05 0.44

~ J>y

I \~ L.E.-l I ;'-.

STAi }\ :\,

~ ~\ , j), >\

1 ~\. 1\

L .: '~r LE.-19

I ,~6 J

DESCRIPTION

L.E.-1

Bottom STA 1

! L. E. - 2, ST A 1

Top STA 1

1 LE.-3 L. E.-4

Bottom STA 2

! L. E . - 5, ST A 2

Top STA 2

TABLE 1.- Continued. 18

ORIFICE -. , NU~1BER'- - '>_: X (em):, -: yCem) 'z(cm) DESCRIPTION, '<

119 1. 57 l 0.64 t *' i~g "~~ ~r"- r'" ~:~~ t

,;1"" ,_','

'122 0 12.88 0 L.E.-6 123 ," 0 ,,15:06 0 L.E.-7

124 4.02' 16.28 -0.99 Bottom STA 3 , 125" 2.56 1 -0.82 ! 1~6 1.48 ,,-0.64 127 0;56 ,':, ',..0.40 128 0 '-0.04 L.E.-8, STA 3

", ,129 0.,67 ",16.32" - ,0.42 Top STA 3 130 1.64 1 0.67 ! 131 2AO ,",' 0.80 ' 132 4.28 1.01

133 0 17.28 0 L.E.-9 * 134 0 19.54 0 L.E.-10

201 3.94 '20. 77 ~ -1. 00 Bottom STA 4

202 2.50. '1 -0~82 ! 203 1..50 " -0.66' ' ' 204 0.61 '-0.42 205 ' 0 ,: .,.O~14 L.E.-ll, STA 4 206 0.78 20.84 0.47 Top STA 4

*207 1.5,8 I"!' 0.67 !' 208 2. 42 " O. 81 209 4.32 ' 1.01 '

210 '021: 77 - 0 L. E..., 12 211 ' 0 23.99 0 L. E.-l3. '

212 4.26 25.27 -,1.01 Bottom STA 5

213 ' 2 A8 -- '1 ' .;. 0 .82 l' 214 1.47 ~0.64 '

215 0.59 -0.40 216 0 "0 L.E.-14, STA 5 217 '0.69 ,'25d5 "0.45 Top STA 5

218 1.62 1 0.68 1 219 2~53 " 0:82 220 4.08 " 1.00

221 0 26.21 0 L.E.~15 222, 0 28.44,0 L.E.-16

TABLE I. - Concluded. 19

NUM!)t.K i I\\wll'j ,)'\ ... 1111 l... L\Cmj I Ut~I...Kl!" ! HJN .

I

223 4.23 29.69 -1.00 Bottom STA 6 I 224 2.36

1 .. -0.8b /

! 225 1.38 -0.61 i

226 0.51 -0.38 227 0 0.04 L.E.-17 STA 6 228 0.55 29.79 0.39 Top STA 6 229 1.43

J 0.65 ! 230 2.43 0.81

231 4.28 1.01

232 0 0 L.E.-18 . 233 0 32.85 0 L.E.-19

234 36.07 20.84 1.25 Upper Surface ---~- -------

*malfunction suspected.

20

. The balance (NASA-LRC model 846) used for force and moment measure

ments was. a six-cof)1ponent internally mOtlntedstrain-gage balance. with

maximum allowable 19adsasfonows (accurate to within 0.5 percent of . . . .

these values):

Component Load

Normal' 3113.6 N Axi al 378 .. 1 N. Side 1334.4 N Pitch 197.8 N-m Roll 36.2 N-m Yaw 84.8 N-m

The balance was located on the upper surface of the wing and shielded by

~,fuselage-li~e aluminUm housing to prevent wind interference.

The two scani -val ves .used for pressure readings were equipped wi th

3.45 N/cm2 pressure transducers, accurate tG within 0.017 N/cm2~ and

were located on the lower surface of the wing. They were a1~0 covered

by an aluminum housing identical to the balance housing.

An accelerometer, l'ocated inside the upper nose cone of the wing.

was used to measure angle of attack. It was a pendulum-type strain-gage

unit, accurate to within 0.2°.

Slot Contours (SC)

Details of the internal slot contours of the open slots are shown

in figure 3. With the exception of SC-4, the slot contours were made

from 0.079 cm thick aluminum stock and were inserted directly into the

21

SC-1 8.26 X 3.00 ellipse

Leading edge



1.78L~ Jr ~ -t- .

1 j ~,--t--1.70

. r----~81~

<~~ ... SC-4 flush with L.E.

SC-4

Figure 3.- Drawing of slot contours tested. Ellipse dimensions given are length of major and minor axes. Dimensions in centimeters.

chordwise slots. 5C-4 was made by forcing 0.159 em thick balsa wood

into the slots and s~nding it flush with the leading edge.

Fences (F)

22

The fences used in the investigation are shown in figure 4. They

were constructed from D.079 cm thick flat~plate aluminum and were also

held in position by the chordwise slots.

Pylon Vortex Generators (VG)

The geometry and dimensions of the pylon vortex generators tested

are given in figure 5. These flat-piate devices were constructed of

0.079 cm thick aluminum and were reinforced by an additional thickness

of 0.159 cm on the outboard side to prevent bending under air loads. An

additional vortex generator, with a'ft extension of the lower edge as a

possible external store-carder (fi"g. 6), was tested with and without a

1.27 cm diameter wooden dowel simulating a sidewinder missile scaled

from the F~4D aircraft~ T~e-~ortex generators were held i~ position by

the chordwise slots.

Leading-Edge Vortex Flaps (VF)

,Undeflected plan views of the vortex flaps, along vlith their corre-:

sponding planform areas, are shown in figure 7. The figure also indi

cates the deflection angles and the semispan ~overage and position of

each flap test configurati6n. The flaps were b~nt from 0.159 cm thick

aluminum and had sharp tapered edges to induce vortex formation. For

mounting, they were bolt~d difectly onto the lower surf~ce of the wing.

1 [ 5.971 +

LV ____ --~--~!~., -1-____ > 7 7 (

L=:60

86 -, t 0.64

==J I . ~3 / t

t-1.27

/' 7 7---r-L

23

~ / (I" -*-0.64

J t l

Figure 4.- Drawing of fences tested. Dimensions in centimeters.

lOa. foeHn

~

2_YQ-~ ___ _

/" VG~2' I· 7 -- ---: - -':"':;-~ ........ --'

VG-l

. 1---:5.59~;

Vortex generator

l-2.79~

I t

L90 ~

100 toe-in

h--4.19 ---l

30 81

1

l 9.14 ~I

24

Figure 5.- Drawing of pylon vortex generators tested. Dimensions . in centimeters.

L.E.

3,81 VG-8 wI missile

VG-9 wI out missile I i

~~ ~~ I. ~ 22.48 ·1

missile

SIDEVIEW FRONT VIEW

Figure 6.- Drawing of pylon vortex generator with missile tested. Dimensions in centimeters.

r

N (J1

Vf-l A =374.00 em2 .

30 - 100% sem i spa n

S "30° t

T ItOO

1-13.87

,--,-----,------'------I' ~

I IS. 62

LJ VF-3

~4S0 ~1 A a 373. 3S em

2 .

?S -100% sell1ispan "~ ~ 3(\0 /,SO 6,,0 I o u, q " v t

~.8l

SO. 52

26

"A" 362.00 em2

30-100% semispan

0"' 30°.

\~1 ~ IljOO

VP'4 2 A = 358.84 em

"25 - 100% sem ,span

8 = 30°, 4So t

Fi gure 7. ~ Drawing of 1 eading-edge . vortex ~ flaps tested. Planform areas given a~efo~ undeflect~d case and include both right arid left flaps. Dimensions in centim~ters~

I 15.62

U VF-5

. 2 A = 312.52 em

25-100% seniispan

0=30°, 45° t

VF-7-A" 329.29 tm

2

25-100% semispan

o • 30° t

T 15.62

U VF-6

A = 206.45 em2

25-100% semispah

8 = 30°, 45° t

VF""'l th rough 6

o .. ==~~/%~ZLL~ ~y .

0.4K' ..

30° t VF-7

r2•621_ --(~

5.0V

Figure 7.- tontinued.

27

I t X'

28

-<

.. YF-5 FlapCoordi nate 5 .

X' y'

0.00 0.00 0.94 0.79 2.92 1.85 7.29 3.00

l3.26 ~.50 1B.90 3.68 27,OB 3~Bl

VF-5 27.0B 0.00 28.45 1.07 31..27 236 36.B3 3.30 44.45 3.73 47.70 ·3.81 . 54.15 Q.OO

.- .

Figure 7.-Concluded.

29

Sha.r.Lle2di ~_-:: EdJL~~!tens tC!~J~lEE)

The sharp leading-edg~ extension tested appears in figure 8: The

SLEE had a variable extension ranging from a to 0.76 em mea~utedperpen

dieular to the wing leading edge. This 0.101 em thick aluminum device

had a sharp leading edge and was bolted directly onto· the lower surface

of the wi ng.

l

· 25-93% semi span

I 13.87

---,-------,I ~ '\ ~·0.094 27.08 ~I SLEE wlSlot at V 62.5% sem i span

SLEE planform

area

cm 2

75.0

'50.0

25.0.

0 0 Q.25 0.50 0.75

SLEE extension (ext.), em

l ext.

Lr .•. C07li 0.1O!< . ~""'''~

SLEE wi Fence (fence extended to SLE~ edge)

Figure 8.-. Drawing of sharp leading"-edge extension tested. Plariform area includes bath right and left SLEE. Dimensions in centimeters ..

30

WIND-TUNNEL FACILITY

The study was conducted in NASA-Langley Research Center's 7- by

10-foot high-speed tunnel. This is a continuous-flow. closed-circuit,

subsonic-transonic atmospheric wind tunnel which operates at ambient

atmospheric conditions.

31

The drive system consists of a motor generator system which powers

a 10.5 megawatt fan motor. The fan motor drives an 18-blade, 9.14 m

diameter fan at a maximum speed of 485 rpm, producing a maximum test

section Mach number of approximately 0.94.

The test section of the tunnel is 2.01 m high and 2.92 m wide, with

a usable length of 3.30 m.

The model support system used in the test is referred to as the

standard angle-of-attack sting. It consists of a vertical strut with

a variable pitch angle sting support system with a range of approxi

mately _1 0 to 230 . In addition to the pitch mode, the standard sting

also has a translation mode which allows the model to be translated

vertically from floor to ceiling, keeping it near the center of the test

section throughout the angle-of-attack range. Reference 9 contains a

detailed description of the tunnel facility.

The data acquisition, display, and control system for the 7- by

10-foot high-speed tunnel is controlled by a dedicated on-site computer.

The system includes a Xerox Sigma. 3 computer, a data acquisition unit, a

line printer, and a Tektronix 4014 graphics terminal. Reference 10

contains a detailed description of the data reduction capabilities of

the system.

32

EXPERIMENTAL PROGRAM

The present investigation was performed at nominal Mach and

Reynolds numbers of 0:16 and 2.0 x 106 (based on a mean geometric chord

of 52.18 cm), respectively~ Force, moment, and surface static pressure

data were taken at angles of attack ranging from _1 0 to 23°. A summ'ary

of the test is presented in Table II. The run number(s) corresponding

to each test configuration is the key to locating the test data pre-

sen ted in reference 11.

Surface flow visualization using a fluorescent tuft technique was

included in the investigation to aid in interpretation of balance and

pressure data. This technique involved the use of 0.02 mm diameter mini-

tufts made of a nylon monofilament material treated with a fluorescent

dye. Approximately 300 mini-tufts of 3.8 em average length were mounted

on the upper surface of the right-hand wing panel and, in some cases, on

the leading-edge devices, using a mixture of three parts Du~o cement and

one part lacquer thinner. The tufts were illuminated by ultraviolet

strobe and photographed through windows in the top and side of the test

section at various angles of attack during a test run. This mini-tuft

technique had been shown in previous testing of this model (ref. 5) to

have a negligible effect on the flow field and, thus, could be performed

simultaneously with force and pressure tests.

In comparison with the model used in the investigation of refer-

ence 5, the present model was basically identical with a few minor

exception~. The most obvious was a reduced fuselage housing in order to

33

TABLE 1I.- TEST SUMMARY

DEVICE IS) SEMISPAN POSITION (%) RUN NUMBER 25 30 ' 50 62.5 75 93 100

Basic Wing 3, 56

Open Slots (no contour) • • • 4 SC-l (short semi-elliQtic) • 47 SC-1 " • • • 36

Slots SC-2 (long semi-elliptic) • • • 7 SC-3 {quarter-elliptic} • • • 8 SC-4 (flush with L.E.) • • • 9 Double Slot Width 10.18 cm) wI SC-1 • 48

rJ~ - • 53

Slot/Fence SC-1 H n' , • 49 50 combination SC-1 F-4 , • 51

SC-1, F-5' ... • 52

VG-1 (baseline) " 40 VG-1 " " " 54 VG-l Ii ~ e e 55 VG-2 (leadi ng-edge Ie ngth reduction) " 43 VG-3 Il " 46 Vortex VG-4 (chord reduction) " 41 generators VG-5 Il " 44 VG-6 (variable chord) • 42 VG"7 Il " 45 VG-8 (extended chord w/ mi ssile) " 38 VG-9 (extended chord wI out missile) 0 39

3no t VF-l' (fu!! 'length, inverse tapered) I

10 -30v t VF-2 Il (tapered apex) 12 30u t VF-3 (full length, constant chord) 11 45

u t VF-3 11 11

60u t VF-3 " 16

30u t VF-4 (segmented, constant chord) 14 Leading- 45u f VE~4 Il 15

edge 30u ~ VF-5 (segmented, parabolic) 17 vortex l4.2". VF-5 "

I

1R 19 flaps 30u t VF-6 (segmented, inverse,tapered) 22

45u J 'VF-6 Il 23 30u ,VF-3 SC-1 SC - 29 300 t VE: 3 £-1 F 30 300 t VF-3 Slot (D. 094 cm) in flap IS') SI" 37 300 t on left' 45u t on rjqht VF-6 57 300 t VH (full length, constant chord) 21 I

0.71 cm ext. SLEE F-2 F ~ '24 I !

0.48 cm ext. SLEE F""2 F ~ 25 Sharp 0.23 cmexl SLEE F-2 F 26

leading- 0.00 em ext. SLEE F-2 F 27 J edge 0.48 em ext. SLEE SC-1 SC~ 28 i

'exten sian s 0.48 em ext. SLEE SC"'l SC~ SC 34 35 0.48 em ext. SLEE F-2 SC-1 F SC- 33 -0.48 cm ext. SLEE, F-2 F ~ F - 32

34

minimize its In{luence on the leading-edge flow development of the basic

delta wing. The accompanying reduction in profile drag acted to better

show up the effect of the leading-edge devices on reduction of the lift-

dependent drag. Indeed, the size of the housing was reduced to a

minimum compatible with the requirement to contain the balance adaptor

and scanning-valves.

In an effort towards economizing wind~tunnel time per configuration,

several pressure orifices on the original research model (ref. 5 investi-

gation) were omitted in the present investigation. The-effect of

eliminating the four most aft pressure orifices (two each on the upper

and lower surfaces at each of the six spanwise leading-edge pressure .-

stations of the original research model) on the accuracy of le~ding-edge

thrust calculations by pressure integration was checked and found to be

within 5 percent of the value obtained with all the original orifices in

use within the a range of interest. This was considered acceptable

since the present investigation was mainly concerned with relative,

rather than absolute, levels of leading-edge thrust. In addition, all

except one of the original upper surface orifices near the trailing edge

were eliminated. However, 13 additional orifices were added along the

wing leading-edge interjacent existing orifices in an attempt to better

define the movem~nt of leading-edge separ~tion in the spanwise di~ec

d~'n'-. It was estimated that, in this mann~r, the test duration per run

_ was reduced by approximately 20 percent without sacrificing the prime

objective of the study.

~

35

DATA REDUCTION

Forces and moments sensed by a wind-tunnel balance must be cor- .

rected for exte~nal interferen~es unrealistic of actual flight. Cali

brationof the wind-tunnel test section in reference 9 shows aconst~~t

streamwise static pressure distribution at the test Mach number and, .

thus, nb correction was needed for longitudinal buoyancy effect. Jet·

boundary corrections were applied to angle of attack to account for the

vertical velocity induced on the model by the test section walls

(ref. 12). To account for initial balance loads due to model weight,

wind-off \'Jeight t~re measurements were taken at various balance atti

tudes and used in the reduction of balance data (ref. 10). Balance

axial force measurements were also corrected to eliminate housing pres

sure drag using chamber (base) pressure measurements. Reference 13 was

used to calculate solid and wake blockage corrections due to the

presence of the model and wake in the test section .. Since the angle

of attack was measured by means of an accelerometer located inside the

model,. no correction for sting bending due to aerodynamic loading was

required.

Once all neces~ary corrections had been applied to the balance

data, the final results were expressed in terms of force, moment, and

pressure coefficients. Force and moment coefficients were computed

based on the basic wing (devices off) planform area. This is quite

legitimate since any addition of planform area from a device is an

essential part of that particular concept. Lift and drag coefficients

eCl' and "CO' 'res'pectivel:y;)an:e or; en ted 'along 'the conventional wind

,axis coordinate system, wi'th t'heaxial and normal force coefficients

{CAand eN' respectively) along thebod:y axis system (sketch G). . ("' ~ (. i

--)10_ ~D

Sketch G

Momentcnef'ficientsrefer to :th:ebodyaxe:s~Pitching momentcoeffi--

36

cientshav.ebeen ·modi'fiedfrom 'those reported in referencell by moving

-the moment ,center further aft ,using the :equa~tion

C m 'c AXc = +~·C

ml 'c ·N' 'r

fix 2m =0. OS for StEEand vortex 'fl aps r

=O.l65 'forother devices'

where em1 i scomputed abouttheorlgina"l moment reference center (see

fig. 2). Definingequati bns for force, moment. and pressure coeffi-, .

dents' u:s~d 'i n theanalysi s are g:;ven in 'the SYMBOLS section .

The effect'·ivenessof the devices under considera'tion depends

;crucially 'on leadi-ng~edge flm'ltontroT,~vhich jnturn is :best observed

37

through improvements in the leading-edge thrust characteristics. The

method of integration of the measured static pressures around the wing

leading edge to obtain the local leading-edge thrust is described with

the aid of the following sketch:

~~ta/CSj I .\/ ~ .Y'

STA j

f

. leading-edge curvature

Sketch H

i-I .........:.=- • . ( , z, - z f /).Z.= 1+1 i-l)

I 2

.. x

(ref. 5)

The suction force developed at the jth spanwise pressure station is

defined as

m SFj = ~(Pij - poo)a ~Zij

;=1

where i, denotes a specific orifice at the jthstation and a is the

length and ~Zij the height of the vertical i-jth pressure panel.

Assuming a constant suction force per unit length of the span (ref. 14),

'SF. J avg

''']

SF. =_J

a

the -totaTsuction-fc:irc.~ ,contr5butedby the jth pressure station is

SF. - ( SF. Jl

oc = 2,Q;)~ a

where ZQ, i sthe "to'tal 'leading..:edge 'length of the wing. Usingthe

defi nitionof ~thepressure toeffi Ci ent"

Cp, .. 'lJ

_Pi} '-POo q

00

ananond'imensi onaltz:ing using the reference force ( q~Sref) yields the

-suc'ti On forte coeffi ciient

c., Sloc

SF· Jlo'c

= Q66Sref

m , " C (2Q,'):2: Pij t;zij

Sref i =1

38

Taking the thrust component of the suction force gives 'the local leading

edge thrustcoeff'i ci ent

'c - ,m (2£)C A 'r _= "" ,p .. DZ'l· J. cos n l~c~' 1~ lL

i=l

39

Averaging the sum of the local thrust contributions and noting that

(21) cos A = b, the total leading-edge thrust coefficient becomes

CTtot

= ~ L: L: .. P~jlJ n (m be !J,z • • )

j=l ;=1 ref

40

PRESENTATION OF DATA

The force; moment,and"staticpres5ure graphs used in support of

discuss'tonof results arepriefly summarized in this section.

Forceslh the body axis coordini:i'te system (CA and CN) were the

main par~meters used for the irtitialperforinance assessment of each

device:' Since the investigation was mainly concerned with drag n~duc

tion through leading-edge flow contrOl, axial force was particularly

well-sU~ted tor this ~Grpose si~c~it providei a sensitive and dir~ct

indication of leading-edge thrust. For demonstration of longitudinal

stabil ityeffects., the pitch; ng-morrient data were transformed to a ref

erence position different from that prevailing during the tests (see

DATAREDUCTTON). This new reference center was chosen s6 as to give

approximately neutral stability at low angles of attack for a closer

approximation to the'condition expected to ,prevail on an actual a i r-

craft. L'ift-lo-drag and drag polar (CD vs CL) curves, conventi ona lly

used fo~descrfptfo~ of the aeroaynainic characteristic~ from a perfor-

, manee point of view~ were given secondary importance in the assessment

of the devices. An advantage of using the lift-to-drag parameter is

elimination of plan form area effects in case of devices such as the

vortex flaps 'and sharpleading..:.edge extension,. Since the basic intent

of this investigation was alleViation of induced drag penalties, the

effect of each devi ceon the; nduced drag of the wi ng was reflected i h

plots of an induced drag parameter; K. Details on the calculation of

this paranieter are given in the Ba'sic Wing section of RESULTS AND

41

DISCUSSION. Additional force plots such as CL, CO' and C1 as func

tions of . a were also included in support of discussion. It should be

noted, however, that the ov~rall force dat~ are representative of the

wind-tunnel model rather than of any actual aircraft. Therefore,

emphasis should be on the relative rather than the absolute magnitudes

of forces and moments and also on the leading-edge static pressure

measurements.

Leading-edge pressure data were basically used as an aid in inter-

pretation of trends in the balance data and also in assessing a device's

ability to favorably modify the leading-edge flow field. Graphs of

leading-edge static pressure (Cp ) and integrated thrust rCT ) as \ LE· '\ . loc

functions of a, at specific spanwiselocations, were useful in detec-

tion of local leading-edge separation and determination of the leading

edge sUction effectiveness of a device throughout the angle-of-attack

range. These figures were also used to plot a boundary between sepa

rated and attached flow and, thus, follow the inboard progression of

leading-edge separation. These same oarameters (C p . , . LE and CTloJ

plotted against spanwise position, at specific angles of attack, also

depicted the effectiveness of a device along the entire span. Static

pressure variations around the leading edge (at the six chordwise rows

of pr,essure. orifices) provided added insight into the specific flow

patterns existing at the leading edge and also reflected the relative

thrust and normal force contributions from various positions on the

leading-edge curvature. For comparison, basic wing (devices off) data

appear as a dashed line on selected balance and pressure data plots~

" RESULT? AND DIS~USSION

,)

60:'D~'g' Del tel ~~rng

The basic '6o'-deg 'delta' vli'ng '(d~\I;;ces'off)served as the baseli'ne

confi gurati on throughout the invest; gati on for compar; sdnand perfor:'

42

manCe assessment of the vari otis leadhig-edge devices. F8rc~a.nd moment

data obtairied for thi s basi c wingconfiguratio'h a.'r'e pres~nteagraphi:" caliy in figure 9. The negative n'ormal force and posi1:1ve pitching

moment at' CJ.~'Od :'are attributed ,to the negative camber eff~ct of an

asymmetri ca lly bevell ed wing trainrlg-edge region (see fig. 2), \<Jhlch

simulated an ,up-deflected trail irig~~dge fl"ap.

BelOw a.'~ 8°, the delta wing isuric:ler the influente of fully

attached 'flow with 100 percent leading~edg'e suction, as evidenced by the

close agreement bf the ax; iil and normal force dirves (corrected for

zero~lift forces)ytith potential flow' th~or.Y (vortex lattiCe m~thb'd with

suction analogj code (~LM-SA); ref. 15) in figure 9. At a~~roximatelj

a. = gO; referred to as the departure a'ngTe of atta.ck (aD)' flow separa':'

bon occur~ ~t the streaniwise tip~ and the subsequent vortex fOrination

spreads toward the wing apex with increasin'g a.. The low pressure pro..:

duced 1 oca lly on the wing upper surfa.c~ by theh; gh rotafiOnal vel Dci-

ties withln the V()rtex generafes additi'dnai 1 itt, which can be noted in

the departUre of the eN curve ~rDm the lrlitially linear theoretical

100 percent suction curve. However; th'ere is an accompanYi ng reducti on

in leading~edge ~ucti6n. The deViation of the 'CA cuHie from the

paraboliC theoretical potential flow <;:urve at approximately gO is

CA

Cm

.02 .~-

-.02

-.04

-.06

~ "',"\.

'~

---("...,-- Basic wing (expt. )

-- - -- Expt. CA - CAo - - - - - - VlM-SA theory

\'\. \ ""-. -- -- .

\ -- ---\ -.... \ \ \ \

Potential flow

I I I L-__ ~ ____ ~ ____ -L _____ ~_~

o 5 I 1 0 15 20 25

.015

.010

aD

.005 f--

1 o !...-- I

0 5

a. deg

Xcm = 0.589 cr

L ____ L 10 15 20

a. deg

25

Figure 9.- Basic wing force and moment characteristics.

43

CN

Cz

1.0 .. ; ;

.8

.6

.4

.2

D06[ 005

I .004

.003

.002

.001

ex.. deg

aD

I-I

Separated flow

Potential

~ Basicwing,(expt)

---Expt C .e.C . . ..• N No - - - -..;.,. - VLM:-SA theory

o I I vi o 5 10 15 20 25

ex. deg

Figure ~.~ Concluded~

44

45

characteristic of this gradual loss of leading-edge suction. This

trade-off between leading-edge suction and vortex-induced normal force

is the basis for the Polhamus suction analogy theory (ref. 15) used in

, this paper for comparison with experifuental data. In addition, the aft

position of the primary vortex at a = 9° locates the center of vortex

lift aft of the wing moment reference center (approximately at the cg),

resulting in a strong pitch-down~ as reflected by the sharp downturn of

the Cm curve (fig. 9). Although not presented here, it is important

to note that the maximum lift-to-drag ratio for th~ basic wing is

attained at approximately a = So. This implies that the additional

vortex lift is insufficient to compensate for the loss of leading-edge

thrust (drag increase) following the onset of separation.

The inboard spread of leading-edge separation is shown in figure 10

to occur very rapidly with increasing a. Plots,of leading-edge pres

sure as a function of angle of attack at selected spanwise positions

indicate a buildup of leadi~g-edge suction (negative Cp)with increasing

a up to the local onset of s~paration, followed by a sharp collapse of

suction at all but the inboard~most station. Plots of this type lead

to the angle for local onset of separation as a function of spanwise

position, shown in the same figure. The locus of data points, which is

effectively a boundary between separated and attached flow, indicates a

movement of leading-edge separation from approximately the 90 to the

30 percent semispan position within only 30 a increment (viz., 11°

to 14°). The corresponding pressute distributions around the leading

edge are presented in figure 11. 'At a = 10°', attached flow,

e", I"'LE

7J

-4

-3 '

-2

-1

Orifice III

232

~\ '2\Q, r \ q

Orifice 111 (T] " O. 232)

I b.. ,I ..' 'o..-D--D-----o ZIDI'IH.6Il2J

~ ''O---~- ---0 232 (7] = 0.847)

o gel

o

a, deg

1.0

0.8

I 'I' c9

LE.

0.6 r d LE.

Separated

Attached

0.4

0: 2,

o 'I "

o 5 10 15 20 25

a sep, deg

46

Figure 10.,.- Bas i ewing- -1 eading-edge, static pressure vari ati o'ns and Cp deri ved separation boundary.

LE

-4r-

-,3 f-

-4r-n= 0.20

-2):, ~,3r-

C Pl

n = 0.33 -2

C P2

-1

2 ,3 4 .5 0 2 ,3 4 5

-4~ X, em

-~ ~l n = 0.57

C P3

Cp 4

1L--1 1 1 2 ,3 4 5 q 1 2 ,3 4 5

X, em X, em

-4i -4r

-,3L n = 0.70 ~3r- ·n = 0.82

-·2

Cps -1l~, CP6

--0 -1~. ~-=: --0----0 °c ·1·1 I I I I I

1 2,3 4 5 1 I ° 1

2 ,3 4 5 0

X, em X, em

(a) 0 ex = 10 .

Figure 11.- Static pressure distributions around the basic wing 1eadingedge.

47

y--

CP1

CP3

c Ps

1 r -I o

-,i

n= O~20

2,,-,. 3 4

-4.~ X, em

n = 0.45

.~~~ -~~I··?

o 2 3 4-

X, em

-4 1

n =0.70 -2

--1

o ~ ~ I ... -

5

5

1 L! ____ ~ ____ L-__ ~ ____ _L __ ~

o 2 3 4 5

X, em

CP2

CP4

, C!'6

'(b) ex = 12°.

-4

-3 n = '0.33

-1

1 0 2 3 4 5

-4~ X, .em

~3 , n = (L'57

-'~ ..•. ~ •..•.•. '," O· "',' . I 1 . o 2 3 4 5

X, em

-4

-3 n =0.82

-:~ 1 LI __ ~L-__ -L __ ~L-__ -L __ ~

o 2 4 5

X, em

Figu're 11. '-Con'ti-nued.

48

49

-4r -4--

I -3,

n = 0.33

cP1 -2~ c

p2 -2~

~--D

-1

1 I I I I I 01' 2 3 4 5 0 2 345

-4 r- X, em -4, X, em

-3f--- -3 r n = 0.45 ,n = 0.57

-2

CP3 lV-----V--O----.:..o CP4 ~-'-O -1 -1~

~~II ~t 1 1 I I o 2 3 4 5 02 3 4 5

~~ ~~

-4~ -4~

-3~ -3[ n = O. 70 n = 0 82

-2~ -2 Cps - , CP6 'I ,

-1 -1~

t~IO ~~I 1° o 2 3 4 5 o 2 3 4 5

X, em X, em

(c) a = 14,0.


50

characterized bya 10caL~uct50n pea.k near·the leading edge, due to flow

acceleration, and subsequent pressure recovery on the upper surface,

persists at each spanwise station considered. Flow visualization indi~

cates separation has begun near th~tip but apparently has not traversed

far enough to be detected. At q, = ,120,. local separation is indicated

at n = 0.82 (STA6) by the constant pressure, stagnated flow region

an the upper surface. The leading~edge flow remains attached inboard.

An increase in a to 14° shows the rapid ..inboard movement of separa

tion, as its apex now appears, to lie between n = 0.33 (STA 2) and

n = 0.45 (STA 3).

The rapid inboard spread of .le,ading-edge separation induces an

erratic wing rolling moment behavior between 90 and 150 a, as shown in

figure 9. This so-called "wi.ng rock,," characteristic of all highly

swept wi ngs, results from asymmetrical ~pam'li se movement of separati on

along the two leading edges. In addition. the forward movement of the

primary vortex leads to pitch-up at approximately nO a, as the center

of vortex lift moves ahead of the wing moment reference center.

The axial force reversal at a ~ 11° is best explained through

inspection of spanwise leadi~g-edge thrust distributions in figure 12.

At lo~."anglesof attack, ,axial force improvement with increasing a is

attributed to local thrust gains all along the span. Increasing thrust

values· toward the tip are due to increasingupwashoutboard, .resulting

in higher local effective angles of attack and, consequently,greater

flow accelerations and.suction forces around the leading edge. In the

mid-a range (11°_14°), however, a balance between thrust loss at the

C Tloe

1.0 r-0.8

0.6

0.4

0.2

a = 22° \

\ \ \

\. \ \ \

° \ 18 \ ~:\

.. \\,\\ / \\\

16° / /~Z~\ .. ,....--~~'X"'" ~\ \' / '" '\ / \s..\ /' \ "

/ ~ ,-- , ----.;:~\ ° I , , 14, ,,'- .~, '(Y = 11 / ' \ ,

\ )~\. " \

8° / --- // '---",-\\_ '~12' /

,-/ ~ "0' 12° ,/,/ >\~", ;3: " / .', "<"'14,

/',- / .... ~" lI°,- /- ....... l~' '/ " 18' 10° /' .///

./ ~·O . ./ -- "'-6 ./

80~ 6°

51

O~I ------~----~------~------~----~ o 0.2 0.4 0.6 0.8 1.0

r;

Figure 12.- Basic wing spanwise leading-edge thrust distributions. Symbols omitted for clarity.

52

outboard stations, due to local leading-edge separation, and continued

thrust gains inboard account for the relative insensitivity of the CA

curve to angl€ of attack. Above' a = 140 , additional thrust gains

inboard with relatively minor further losses outboard account for the .

CA recovery. These same leading-edge thrust data are plotted versus

angle of attack in figure 13~. Here~ .local thrust is compared VJith

balance axial force measurements correCted for profile drag (CA - CAo )'

which should represent strictly leading-edge effects and, thus, an

approximate average leading-edge thrust.. .The close agree~ent between

the local thrust and balance data prior to separation adds credibility

to the pressure integrations. Again, as in figure 12, the axial force

recovery at high a is shown to result from continued thrust gains at . I

the inboard stations as the outboard stations settle at a constant value

below the balance-derived average. Note that in the mid-a range

( 100 to 1<°, the ha1::>·':'r-e d::>+a r-"ns;s+en tl·,,;4=a" ..... '" -'-_" \.oIl ,-,IU,-lIw u.t, yv11 I v II J".I I. below the integrated

, .. , thrust values. The balance is, thus, sensing a source of drag other

than that accbuntable from loss of suttion, possibly from trailing·edge

separation or interference· from the housings.

At high angles of attack, the .rate of inboard spread of separation

diminishes, as reflected by the leveHng off of the separation boundary

in figure 10. Continued axial force improvement is observed up to the

highest angles tested. Previous research has shown that further

increases in angle of attack would eventually lead to vortex burst,

spreading rapidly toward the wing apex with an accompanying loss of

leading-edge suction and pitch~up.

53

----0-- Pressure-integrated ----0----- Sa 1 ance-deri veo -(CA-CA )

0 .10,- n •1O r-.0Sf- n = 0.20 / I

.OS[- n == 0.33

.06f- r! I

C C TI 2.06 Tloc1 CC.,. -

.04 uU·--D .04j P .-D .au

---~ . . 02 .02

0 I~ 0 0 10 15 20 25 0 5 10 15 20 25

IX, deg IX, deg

.10[ . 1°i

n = 0.45 I 11 = 0.57 .os .osl

c- 06r R~ CTloC4 .06 ~ 'Ioc~ .

.04f- Y f""'""t.-n--""'-"" 04~. ~

.02f- J:Ei 02cL . o L .... 1J .. J 1 1 -....J o . _ .. :....._L-.-L_J

0 5 10 15 20 25 0 5 . 10 15 20 25

a, deg a, deg

.10,-.10!

.081- n = 0.70 n == 0.82 .0Sr-

CTIOC5 .06 r' C I T1oc6 .06 r-

c..D 0.-0-".-0 I

.04f- / I .04 1

.02~ .02L

0·' I~ 01 L,· I -....-L ...... J 0 5 10 15 20 25 0 5 10 15 20 25

a. deg a. deg

Figure 13.- Basic wing leading-edge thrust characteristics.

54

It was previously stated that severe induced drag penalties are

characteri sti c of delta wi ngs at hi gh 1 ift. In order to see thee-ffect

of leading-edge separation on the total induced drag of the wing, a plot

of the induced drag fa~tor K, y.Jh~re

c - Co . D mln(?fAR) K = ? . C

L

1 e

i s pr~serJt~d in fi gur~ 14. Thi s formul a is derived from the equati on

for minimum induced drag,

CD = CD. + cL2K

1m n ?fAR

where K = 1 corresponds to fully attach~d flo\f.i and K > 1 to part; a 1

leading-edge separation. Very low angl~s of attack ~v~re ignored since

low lift and drag coefficients reS.l,llt in sporadic K values. Figure 14

depicts attach~d flow at low a,follow~d by a rapid increase in induced

drag as leading-edge separation spreads inboard with its accompanying

loss of leading-edge suction. - The reduction in the rate of spread of

leading-edge separation near the wing apex results in a reduction in

slope of the induc~d drag curve at approximately a =_ 14°. In the case

of an actual aircraft, induced drag would continue to increase at

approxi l11ately the mid-a rate since decreasing leading-edge radius toward

the tip cannot retain the degree of residual suctign of the uniform

radius leading edge of the present research model.

55

2.2

.... :-

2~0 ~C 1

K = _0_ (7r AR) = -C 2 e L

1.8 ~CD = Co - C

Omin

1.6

K

1.4

1.2

1.0

.~ Fully attached flow

0.8

o 5 10 15 20 25

a, deg

Figure 14.- Basic wing induced drag characteristics.

56

A comparison of test data wit~ theory ispresehted in figure 15 in

the form of a drag polar c'orrected for zero-lift drag (CO i = Co - COo),

Experi'mental values are compared with theory for zero and full

(100 percent) leading-edge suction. At low angles of attack, where

fully attached flow prevails, 100 percent leading-edge suction is

realized .. At approximately a = gO (CL ~ 0.25), deviation of the

experimental values from the theoretical 100 percent suction curve

signifies the onset of leading-edge separation. The loss of suction

continues with increasing lift and eventually settles along a constant

percent suction curve, indicating residual leading-edge suction charac-

teristic of blunt leading edg~s.

The major contribution to the integrated local leading-edge thrust

is developed by those orifices near the wing leading edge (see DATA

REDUCTION). This suggests the possibility that a si·ngle leading-edge

pressure orifice m~ght suffice for a chordwise series of orifices around

the leading edge for leading-edge thrust determination on wind-tunnel

models. If so, it would result in significant savings in model con-

struction costs and wind-tunnel test time.' Figure 16 presents staggered

plots of CPLE versus eTloc at the six spanwise ~ressure stat~ons of

the research model. The linearity of the curves is interpreted as

attached flow and the, departure from linearity as the local onset of

separation. This indicates that CTloc can be calculated from CpLE

only in the attached flow regime. No systematic relationship between

CpLE and CTlocis apparent following the onset of local separation

0.32

0.28

0; 24

0.20

Co. I

" 0.16

0.12

0.08

0.04

o o

Co.= Co - CD" I 0

/ /

/ /

/

/

/'

y""",

/'

I /

/ '

I

" I

I /

I /

!

/ /

/ /'

0% suction I

I

I I I

I " I

/ /

I I I I

/ /

/ /

/ /

j":

/

100% suction /

I /

I /,

/

~ Experiment - - - - - VLM- SA theory

0.2 0.4 0.6 0.8 1.0 .

CL.

Figure 15.- Comparison of basic wing experimental data with VLM-SA theory.

57

, c

.0

\.

STiA 1

.0 • .02

2 , ;3

4

. .0

·1

5 6

0

-..,....'.,-j .. 0 STA 1(7]='.0.2.0) ~~-4(J----2:( ; '0.33) -.' . -~"-. , 3,( ='.0.45)

, ·...I~, "4( ,=;0.571 - ..... ....,. ~,.....,... -5 c( ="O~ 70)' - .. ' -~--.-, ;6'( =.0.82)

.0 .0 .0.02

Cr loe

0.04

Figure 16.- Basic wing le~ding-edge static pressure-thrust relationship.

'.::

, "

U1 00

59

due to the significantly altered Cp distribution around the leading

edge: Further discus~ion on this type of plotting will be ptesented in

conjunction with each leading-edge device.

/

;,/' /'

60

Chordwise Slots

Single and multiple chordwise slots cut into swept leading edges

produce performance improvements as reported iii reference 5. For refer-

ence purpo~es, basic configurations of cine and three chordwise slots per

leading edge were retested in the presentirwestigation.

The results with a sing'le slot were used to illustrate the effects

of the devi ce on tne wing leading-edge flow pattern and overall perfor

mance. T6 investigate the possibility of further performance improve-

mehts j the slot was modified with an internal contour (5C-l), which will

be discussed later. Figure 17 presents results of force and moment

measurements on a configuration utilizing a single slot at n = 0.625.

the CA curve shows the expected 10w-a drag penalty discussed previously

(BACKGROUND section). However, the slot acts to delay the onset of

separation and vortex lift, as indicated by the slight loss of normal

f6tc~ in the mid~a ra~ge. The result is ~~~~~-~~ l~~d·~n·g· ~~"~ C'lllfOfl\....CU J co -I I -CU':jC suction

and, thus, CA im'provement beyond 11 0 ct. However, a sudden loss of

slot effectiveriess at high a is ihdicated by the convergence of the

slot and basic wing curveSj beginning at approximately a = 17°.

Longitudinal stab,i1ity e·ffects of a single slot (Cm curve in fig. 17)

can be noted as a slight moderation of the initial pitch-down,' as well

as a delay of pitch-up to approximately 120 a. From a performance

standpoint; a penalty in max;mumL/D value occurs, but a substantial

improvement is obtained at mid a as a result of , the drag reduction.

Static pressure distributions around the leading edge at adjacent

st~tions on either side of the slot, in figure 18 (a = 16°), illustrate

CA

em

.02

-.02

-.04

-.06

.02

.015

.010

.005

-- ---- Basicwing

---{)-- 1 x SC-l @ 7J = 0.625 ----{}---- 3x SC-l @ 7f"'0.25, 0.50, 0.75

I I__J o 5 10

a, deg

15 20

Xcm = 0.589 cr -,

I

/ I

/ I

I I

25

o I I I '~,/-,-I ~---' o 5 10 15 20 25

a, deg

Figure 17.- Force and moment characteristics of single and multiple chordwise slot configurations.

61

62

1.0

.8

.6

eN

- - - - - - Basic wing

---()--:-- 1 x SC-l @ 7] = O. 625 ---0---. 3x SC-l@7]=0.25,O.50,O.75

20 25

a, deg

10

8

6

LID 4

2

o ~I ______ -L ______ -L ______ -L __ ~ __ _L ______ ~

o 5 10 15 20 25

a, oeg

Fi gure 17.- Conel uded.

63

-4 - -- - - - - Basic wing

---0--- 1 x SC-l @ 17 = 0.625 -3

7]=0.57

-2 C

P4 ~-~-O---O=-=-=-=-.::..=-=-=o

o ~--V----=o

1 I J_J ___ L __ ~ o 2 3 4 5

X, em

-4 ---'-

--3 1] = 0.70

CP5

------~ -1

1 _~ ____ L--,--I-,------, o 2 3 4

X, em

Figure 18.- Loca1 static pressure effects of the chordwise slot at a = 16°.

5

64

the effects in the vici~ity of t~e device. While little influenc~ is

detected 'inboa~d,the slotco~~artnientati~n"~ffect (see BACKGROUND

section) acts to retain attached flow at the leading edge on the out-

board side. \tJhere the basic wing is stalled. The resulting eff.e,cts on

leading-edge suction are depicted in figure 19, where sizable increases

in negative leading-edge pressure (leading-edge suction) are indicated

on the outboard side beyond 130 Ct. The spariwise variation of CPLE at

Ct = 160, in figure 20, reflects these bas,ic trends on either side of

the slot.

While the local leading-edge flow field seems to have been

favorably modified by the slot, an overall performance improvement

depends on the device's ability to increase the total leading-edge

thrust.' Figure 21 presents local thrust variations with angle of attack.

The local onset of separation at the two pressure stations outboard of

the single slot (STA's 5 and 6) has been delayed by approximately 1° Ct,

resulting in large thrust improvements beY9~d Ct ~ 11°. The loss of

axial force at high Ct, discussed previousiy, is shown to start at the

outboard stations as the local thrust and basic wingd~ta converge.

This is believed to result from a breakdown of attached flow on the

upper wing surface, which begins at the tip in view of the higher pre-

vailingupwash. The compartmentation effect of the slot is, thus, lost

and the primary vortex allowed to spread inboard. The inboard effect

of the slot results in a minor penalty in local thrust between 140 and

20° Ct. However, this unfavorable effect is highly localized, as

c .. PLE

C PLE

-3

-2

-1

o ! cr o

-3 I

-2 I

-1 ~

5

r;

;/

Orifice 210

2ll

10

a, deg

,~

;1

..ff

- - - - - - Basicwing ----10 1 x SC-l@"'i=O.625

Orifice 210 (7]= 0.602)

15 20 25

~ Orifice 211 (7]= 0.663)

-... ....... ......... -- -

o I if o 5 10 15 20 25

a, deg

Figure 19.- Local suction effects of the chordwise slot.

65

C PLE

c PLE

-5

,...4

-3

~2

-1

00

-31-

I -2

-1

",.

q fu .~ gb~\~

",'1 '~.

,2 .4

b

~\, .......

6

- - - - - -Basic wing

, ~ lxSC-l@ 7]=0.625

-':'~--o---- 3 x SC-l @ 7]= 0.25,0.50, 0.75

a = 16°

.6, .8 1.0 I

'1]1 I I

a,: 23°

o ·I<--_--"--__ -L-__ .L..-_---L __ .....J

o .2 .4 .6 .8 1.0

'T}

Figure 20 • .:.. Spanwis'e leadirig.:..edge static pressure distributions for single and multiple chordwise s10t configurations.

66

-----.--. Basic wing STA 1 ----{)-- 1 x SC-l @ 11 := 0.625

2 ----0------ 3 x SC-l @ n '" 0.25. 0.50. 0.75

.10

.08~-

3 4

5 6

CT .D6 loel

·.b4~ = 0.20

.02

01 fJ [ o 5 10 15 20 25

a, deg

.10,

.08 n = 0.45

CT 3 .06 loe

--D 0--'

.' /

/ ~- ...

. 04

.02

o o

.10

.08

CTloes .06

.04

.02

~ I 5 10 15 20 25

ix, deg

n = 0.70

o I ..... I I J o 5 10 15 20 25

a, deg

.I0r

.081--j

~ ... }.

? /

CTIOC2 .06

n,o"tv .04

.02

o I ifl ~-',_--' o 5 10 15 20 25

a, deg

.10 r-I

.08 n = 0.57

CTIOC4 .06

R p-IT ''':0

.04

.02

OL-HI o 5 10 15 2J 25

a, deg

.10r-

I .08 L_ n = 0.82

.06 ~- .. "0

+-fo ,o:r- ~ I I I j

o '-" 5 1 0 1 5 20 25

CTIOC6

ex, deg

Figure 21.- Effects of single and multiple chordwise slots on leading-edge thrust.

67

68

stati~~s inboard of - n -~ 0.57 rem~in unaffected~ The favorable out-

board effect is more widespread .'" , ;~.

A multiple slot configuration was tested ih an attempt to spread

the slot benefit over a larger portion of the span and to delay, or

eliminate, the loss of effectiveness at high a. The s16ts were located

at the 25, 50, andl'S percent semispan positions and Were again inter

nally contoured with 5C-l. The result was further delays in local onset

of separ~tion and, significantly higher local thrust values in the mi,d

and high-a range at all but the most inboard pressure station (fig. 21).

Apparently~ the favorable iriflUence of the ",upstream" slot acts to

nullify the adverse inboard effect of a single slot, resulting in thrust,

enhancement along a large portion of the span. the lack of a slot

inboard of stAl leads to early separation and l6ss cif thrust near the

apex. Compari son of spanwi se 1 eading-edge pressure variati ons at

a = 1 &:0 IV , in figure 20, also reflects suttion improvements outboard of

n = 0.25 withrilultiple slots. Based on these results, it is believed

that several slots located along the leading edge have the ability to

redirect the spanwise boundary layer flow in the chordwise direction

before it builds up in, magnitude. Again; however, loss of effectiveness

begins near the tip, as the c6rresp6~dihg CpLE data at a = 22~

(fig~ 20) converge to the basic wing curve. For a ready assessment of

the leading-edge effects- of multiple slots, the following table presents

a comparison of total leading-'edge thrust values for single and multiple

slot configurations:

r-

I CT

to

Basic Wing (BW) 0.04 1 x SC-l .04 3 x SC-l .05

C); == 14°

Percent t over Bt~

39 -86 10.7 84 33.0

69

a = 180 a = 220 I -.

Percent I CTtot

Percent CT over BW tot over BW .

0.0466 - 0.0534 -.0512 9.9 .0555 3.9 .0656 40.8 .0645 20.8

These data indicate that high-a thrust improvements of over 40 percent

(~elative to the basic wing) are obtained with a ttiple slot

configuration.

Figure 17 compares force da.ta for single anq multiple chordwise

slot configurations. The CA curve indicates that multiple slots pro

vide a substantial drag advantage beyond 120 a with a more gradual CA

reversal which is delayed from 120 to 190 a. Multiple slots also pro

duce a smoothet Cm curve, delaying pitch-up to 190 a. The additional

loss of lift beyond 120 a with the use of multiple slots is attributed

to the retention of attached flow (and, therefore, loss of vortex

induced lift) over a greater portion of the leading edge. Maximum

lift-to-drag is unchanged from the single slot case, but additional

improvements are indicated beyond 100 a before the final convergence to

the basic wing data at higher a.

The ability of chordwise slots to delay the local onset of separa-

tion is reflected by the extension of the linear portions of the CPLE versus CTl curves in figure 22. Delays of up to 20 a are indicated

oc just outboard of each slot in the triple slot configuration. The

extension of the curves along their initial slope suggests that attached

-5

-4

-1

STA 1

2 3

4 5

6

- -'- - -'- - - - Basic wing (all STA's)

--0--- STA 1 (7J = O~ 20) --[}-- 2 ( = 0.33) ~-o--' - 3 ( = 0.45) ~--ic:.' ' 4 ( = 0.57) ---Itr-----' 5 ( = 0.70)

3 x SC-1 @

7J = 0.25, 0.50, Q.75

--D--- 6 ( = 0.82)

o ~~ __ ~~ __ ~~~_~~ ___ L-~_~_~ __ -L __ ~~ ___ ~

o o o o o 0.02 0.04 0.06

I.. 0.02 J I

Figure 22.- Effect of multiple chordwise slots on leading-edge static pressure-thrust relationship. Selected high.;.a data points have been omitted for clarity. Angles in parentheses refer to basic wing data (see fig. 16).

....... o

.'

71

flow has been extended to higher a without modifying the local effec--

tive upwash. This suggests that the slot action is more due to compart-

mentation rather than to any vortex mechanism.

As mentioned in the BACKGROUND section, the object of internal con-

touring ofthechordwise slots was to reduce the low-a drag penalty

while improving the high-a performance. The low-a drag effects will be

discussed first. The following table compares zero-a drag measurements

on a triple slot configuration utilizing the various slot contour shapes

tested (fig. 3):

I CD Percent Device 0 ·over BW

Basi c wing 3 x open slots· 3 x SC-l 3 x SC-2 3 x SC-3 3 x SC-4

0.01261 .0l318 .01321 .01293 .01296 .01307

4.52 4.76 2.54 2.78.

. 3 h~ l __ ~~ ____ _

Comparison of open (uncontoured) slots with those making use of SC-1

shows relatively insignificant effects of contouring on low-a drag.

Therefore, stagnation pressure on the back face of the slot appears not

to be the primary cause of the low-a drag penalty. This is not sur-

prisi~g considering the relatively small area of this vertical face.

Two additional contour shapes tested, SC-2 and SC-3, effectively

reduced the depth of the slots. Zero-a drag measurements for configura

tions utilizing these contour shapes indicate a reduction of almost

50 percent of the drag penalty associated with the slots, with a

72

relatiy·eiYmi;;or:d·i·h:erenc~ betwe'"en the two. Since these contours

differ only insha:pe·~ ··and nbtlkng'th, these data further support the

conclusion that pressure on the vertical face at the end of the slot is : .,.

not the primary sour.ce of low-a. drag penalty. Compari son of zero-a drag

for SC-l and SC-2, which differ in length but have basically the same

contour shape, suggests that sTot depth is the determining factor for

low-a drag. The reducti on in drag penalty with SCM2 and SC-3~ there.;.

fore; results from redl,lced fticti-ondrag acting on the slot side

surfaces.

Comparisons of overall performance for the slot configurations just

described appear in figure 23, .Axial force shows a slight suction

ad~antage beyond 120 a and delayed loss of effectiveness with the u~e of

SC-1 relative to the shorter slots uti1 izing SC-2 and SC.;.3. Minor dif-

ferences ex; st between SC-2.· and SC-3. Since a sl ight high-a suction

advantage is realized with the use of 5C-1 relative to the uncolitoured

slots, high-a performance may be somewhat dependent on the shape of the

back face of the slots. However, slot depth is again the primary factor

determining high-a performance. Improved performance with increasing

slot depth is explained as follolt's: Since the postulated slot flow

mechani sm is paftia lly that of a "fl ui d fence," increasing slot depth

is analogous to increase in the chordwise length of a fence. Results of

previous research (ref. 5) indicate that increasing fence length results

in improved high-a drag performanCe and delayed loss .of effectiveness,

the same trends observed here.

CA

Cm

.02

o

-.02

-.04

-.06

o

.015

.010

.005

1.1 5 10

---0-----0-------<)----------6-------.-.-~-.--

ex, deg

15 20

x em = 0.589

c r

o ~I -----.~----~----~ o

ex, deg

73

Basic wing

3 x Open slots @ 'TJ = 0.25, 0.50, 0.75 3 x sC-ll 3 x sC-21 _ 3 x SC-3J @ TJ - 0.25,. 0.50, 0.75

3 x SC-4 . . .

CJE2/1 ~

25 ~.

~w SC-4 ~

25

Figure 23.- Effects of internal contouring on chordwise slot performance.

~ ~

74

S}otdepth is also of.prime i~portance to the longitudinal st~-.

bil ity of .the configu~~ti ~n.The most nbtabl e effect is accelerated

high;..a pitch-up with reduction in slot depth (fig. 23). Effects of slot

depth and contouring on normal; f6tce. and rolling moment are iJegligib le.

An additional slot contour; SC-4,'was fested in an attempt to

reduce thE: 10w-a drag penalty by contourtrig the forward po'H, on of the

slot, as shown in figure 3. However; sta.tic pr~ssure datCiindica1:e that

the contour may have obstructed th~ flb~ through the s16t~ resulting in

practically a nonexistent effect on drag perforinance (fig. 23). In

addition, 10rigitudina1 stability is severely degraded, with accentua.ted

pitch-up and pi tch-down throughout the ang1 e-of-a.tiaCk range. parain;..

eters such as n()rma1 force ahd lift-t6-drag ratio; not presented, also

agree closely with basic wing data, indicating that & slot starting

behind the leading edge is ineffective.

Slot width was investigated as a110ther parameter through which

further perforinahce improvements migHt be Teai ized. Compari son of

slots with widths of 0.25 and 0.50 percent of the semispanshm~ed,

however, that this is not a significant sizing cOhsideration within the

range tested.

In summary, the chordwi se slot device possesses dra'g-reduction

potential in addition to its ability to improve the 10hgltLidinal sta

bil ity characteri sti cs of the del ta wi ng. Multipl eslots spaced along

the span produce further performance improvements by bringing more of

the leading edge under attached flow a.nd by eliminating the adverse

effects inboard of a single slot. The slot effectiveness is primarily

75

throughcompartmentation of the leading edge to remove the 3~D effect

which causes earlier separation on swept leading edges, rather than ftom

any vortex action. The low-a drag penalty associated with this device

results from friction acting on the side surfaces inside the slot; the

pressure drag on the vertical face at the end of the slot is of secon

dary importance. L i kewi se, hi gh-a performance is primar'ily dependent

on slot depth rather than the shape (contour) of the back face.

Increasing slot depth produces progressively higher low-a drag penalty

but improved high-a performance.

76

Fences

As noted in the .BAC~GROUN.D~ection, fences have previously been

used on highly swept leading edges primarily for improving longitudinal

stability characteristics. The em ,curve in figure 24 shows that addi

tion of a single F-3,fence atT] = 0.625 improves the linearity of the

pitch curve, with reduced pitch:-downat a = 80 and generally a more

stable Gonfiguration throughOut theangle-of~attack range. The CA

curve in figure 24 shows that,fences also have drag-reduction capab,ility

at mid and high a, although some drag penalty at low a is inturred

due to friction drag on the fence. The advantage of the device is first

seen as a slight delay in the onset of separation to a ~ 10°. Beyond

approximately 11 0 a. a significant suction advantage is indicated with

the fence, with no sign of loss of effectiveness to the highest angle~

tested. The reduction in eN relative to the basic wing at angles of

attack greater than 'at separation onset also indicates delayed separa..:

tion and vortex formation with the fence attached. This is -further

supported by static pressure distributions around the leading edge on

either side of the fence (fig. 25). Inboard of the device, the flow

remains stalled at a = 16°. However, outboard, where leading-edge

separation has clearly occurred on the basic wing (as evidenced by

nearly uniform upper surface pressures), attached" flow fs maintained by

the fence. As shown by the C1 curve in figure 24, by delaying

separation. the fence also reduces the severity of the wing rock char-

acteristic of this delta planform.

77

.02 - - - -,- - - Basic wing

--0-- F-3 @ 7] = 0.625 --~-- F-3/SC-l combination @ 7]= 0.625

o "

<-.02

CA

-.04

-.06

I o 5 10 15 20 25

a, deg

Figure 24.- Force and moment characteristics of single fence and slot-fence configurations.

.".

'c .N -";,

cz

1.0

.8

.6

.4

.2

1 "n

- - -::- ~ -B~'sic~ing

. ~F-3 @7J=O.6~ ---:.o---F-3/SC-i cOmbination @ 7]=0.625

Ol,¢: J.

a, deg

.006 I

A II I I

.005 f-- ' \ I \ 1 I" I

.004~ . \ \ \ 1

:003, \ I

01 I

o 5 10 15 20

a, deg

Fi gllre" 24. - Corkl uded. "

78

.~

79

--4

- .- -- - - - Basic wing

-3 TJ = 0.57

.---0-- F-3 @ 1] = 0.625

-2 C

P4 - ..... ~--..,..;----

-1

~[ ""

I I I I I 0 1 2 3 4 5

X, em

f\\ / , 7J. 0.70

1\ \ -2

Cps

'-

o '-

~

1 I I I o 2 3 4 5

X, em

Figure 25.- Local static pressure effects of the fence at a= 16°.

80

The sepatatiorib6undary in figure c26 shows the effect of the F-3

fence on theinboard'movement'of leading.,.edge separation. The compart

mentation effect of the device (see BACKGROUND section) is seen to delay

the onset'of ~e~~ratton onit~outboatd side by approximately 30 Q.

This reducti on: in the"rat:e; Elf inboard movement of sepa rati on accounts

for the el im;:h'ati on-of' severe pitch-up,' as the forward movement of

the center;o{'vortex Tift is also delayed. This effect also diminishes

the asymmetry 'in the- spanwise lift distribution, responsible for the

erratic rolling moment behavior of the basic wing.

Local leading-edge thrust variations with a single F-3 fence at

n ::: 0.625 are presented in figure 27. At low Q, where the basic wing

flow remaihs attached, the fence has no influence on the leading-edge

flow conditions. However, beyond the lncal onset of separation for

the basic wing, substantial thrust improvements are evident outboard

of the device due to the delay in local stall and subsequent retention

of leading-edge suction. This effect can be attributed to the unsweep~

ing of the upper surface isobars in the vicinity of the fence, as

sketched below:

TJ

81

1.0

I I

- - - ....: - - Basic wing 0.9

--0- F-3 @ 7]: 0.625 t I I I

---0--- F- 31 SC-l combi nation 0.8

0.7

0.6

0.5

0.4

0.3

0.2

0.1

L.E. Attached

t I I \

@ 7]~ 0.625

, Location of device

l L.E.

Separated

-"

o LI ________ L-______ -L ________ ~ ______ ~ ______ __J

o 5 10 15 20 25

asep , deg

Figure 26.- Effects of the fence and slot-fence combination on leading-edge separation (C derived).

PLE

82

- - - - - - .Bas i c wi ng --0-- .F-3 @ n = 0.625

'T ~, I :+ .08 11 = 0.33

C :06r-- I ,( C T .06 -A ;;

Tloe1 loe2

.04 f-- 0 n = 0.20 .04

.0:1 A .02

0 0 5 10 15 20 25 0 5 10 15 20 25

lX, deg lX, deg

.1°i

n = 0.45 .1°i

n = 0.57 .08 ;08 r--

CTloe3 .06 CTloc4 .06

.04 .04

:02 .02

01 . I I

I I 0 0 5 10 15 20 25 0 5 10 15 20 25

lX,·.de.g ix, deg

.10 .10

.08 n =0.70 .08~ n = 0.82

CTloes .06 CTloe6 .06

.04 .04

,o:r I

.02

0 0 5 10 15 20 25 0 5 10 15 20 25

lX, deg a, deg

Figure 27.- Effects of the fence on leading-edge thrust.

basic delta wing with fence

-Cp

1

Fence on

-"~ Fence off \

"-"- ......

" .....

83

Inboard

~ outboard

Sketch I •

On the outboard side of the fence, suction peaks are reduced and occur

further aft of the leading edge. This reduces the tendency toward

separation by providing a more gradual pressure recovery on the upper

surface~ The opposite effects inboard facilitate earlier separation

and loss of thrust. However, this adverse effect is only localized, as

little influence of the fence is evident further inboard.

The effect of a single fence on the induced drag of the basic wing

is' shown in figure 28. At low a, before leading-edge separation has

spread to the fence location, the indicated increase in K results from

the friction drag on the fence. Beyond 100 a, however, the fence pro

vides an induced drag reduction by delaying leading-edge separation and

2.2

2.0

L8 ~

L6

K

1.4 I

L2 I--

LO

0.8

o

:''-.,

~CD . _ J.. K = -. 2~ (7T AR) - e

CL

~CD = CD - CD . min

~ I

tf / /

/ ./

II

84

... ....

- ~ - - - ~ Basic wing

~ F-3@'T]= 0.625 ---G--- F-3/SC-l combination

@ 7]= 0.625

1.. Fully attached flow

I .11 I .

5 10 15 20 25

a, deg

Figure 28.- Induced drag characteristics of single fence and slot-fence configurations. .

loss of suction. This benefit is visible up to the highest angles

tested, although the maximum relative improvement occu~s between 110 o and 13 a.

85

Figure 29 presents graphs of Cp versus LE

CTloc

at the six span-

wise pressure stations of the research model. Local separation is shown

to have been delayed by 5°, to a = 16°, at the station just outboard of

the fence, with a delay of approximately 3°, to a = 14°, at the station

nearest the tip. Inboard effects are minor. The extension of the

outboard curves along their initial slope again indicates flow modifi-

cation without alteration of the local angle of attack. As discussed

previously, the fence flow mechanism is rather one of isobar unsweeping

and leading-edge compartmentation.

Reca~l that thrust data in figure 27 indicated an early loss of

effectiveness just inboard of the fence. This was believed to be

partially the result of an accumulation of viscous fluid on the inboard

side of the device from the spanwise boundary layer flow. It was,

consequently, thought that by opening an adjacent slot inboard of the

fence, this viscous accumulation could be "blown off" by the slot jet

stream (see sketch in BACKGROUND section). The concept was tested on

the F-3 fence with the inboard slot contoured with 5C-l. Results of

force and moment measurements are shown in figure 24. Although there

is an additional low-a drag penalty as expected, definite axial force

improvements are indicated beyond 13° a. For instance, at a = 14.3°,

an 8.6 percent reduction in axial force is realized with addition of

the slot. The general nature of the pitching moment curve is

-5

-4

-3

-2

-1

5TA 1 2 .J

4 5 6'

- -' -~' - -' Basic wihg(aIISTA's)

---'-----10..-.> -~ 5T A "1 ( 7'J;' 0: 20) --'--:10': 2 ( = 0; 33)' ~~<> 3 ( = 0;45) -'-~' A' 4( =0;51) F-3@7'J=0.625 --'---iCS ," 5( = 0; 7d) ---'-ID ' 6( =0,82),

o ~~ __ ~~ __ -U¢-___ ~~~ __ ~=-__ ~=-__ -L __ ~~~ __ ~

0, o o o o o , 0.06

Figure 29.- Effect of the fence on leading-edge static pressure-thrust relationship. Selected high-a data paints have been omitted for clarity. Angles in parentheses refer to basic wing data (see fig. 16).

unaffected, but the configuration appears to be slightly more stable

beyond a ~ 110. No significant effects are evident in normal force

87

and rolling moment. Figure 28 shows that a slight improvement in

induced drag also results from addition of the adjacent inboard slot.

This advantage, however, dissipates by 190 a as the fence and slot-fence

data converge.

The reduction in drag with an adjacent slot inboard of the fence

is shown in figure 30 to result from thrust improvements outboard of

the device, rather than inboard as was expected. In fact, no influence

is evident inboard. The spanwise leading-edge pressure variation at

a = 160, in figure 31, shows the same trend. This leads to the con-

clusion that the inboard loss of fence effectiveness at high a is not

hastened by a viscous fluid accumulation, but rather is totally a result

of inviscid effects. As noted earlier, the unsweeping of the upper

surface isobars leads to increased suction peaks occurring closer to the

leading edge on the inboard side of the device. This results in a

steeper upper surface pressure recovery behind the "suctioh peak and.

consequently, earlier separation. Comparison with a single slot at

n = 0.625, in figure 31, indicates that the additional improvement out-

board is primarily a fence effect. It is believed that the spanwise

flow being lifted over the fence by the slot airstream results in the

formation of a vortex off the upper edge of the fence. This vortex lies

on the outboard side of the device and rotates in a sense so as to

induce a downwash velocity outboard. The resulting outboard reduction

88

STA ------ Basic \~;ng 1

-0---- F-3 @ T) = 0.625 2 3 ---o---~ F-3/SC-l combi nat; on}

4 -----<>----- F-4/SC-l combination @ n = 0.625

5 --------6------- F-5/SC-l combination

_10r- ,\/6 /l _10

°l I ! 0+ n = 0.33

CTloc'- _06 C Tloc2

_06

0'[ / II = 0.20 _04

_02 _02

01 o::J I b 0 5 10 15 20 25. 0 5 10 15 20 25

a., deg a., deg

~ .1Or- _10 r-

.08f- n = 0.45

°T n = 0.57 ~

C"O~0't CT .06 ~ loc4

~ .04 .04

_02 .02

01 VI I I 01 &f

0 5 10 15 20 25 0 5 10 15 20 25

a., deg a, deg

.10,.---'- .10

.08 n = U. /U 0+ n = 0.82

CT .06 C .06 loc5 Tloc6

.04 .04

.02 _02

0 0 0 .5 10 15 20 25 0 5' 10 15 20 25

a., deg a, deg

Figure 30.- Effects of slot-fence combinations on leading-edge thrust.

C PLE

~ @0J

-4.5 ~ -4.0

-3.5

-3.0

-2.5

-2.0

-1.5

-1.0

-.5

- - - - - - Basicwing

--0-- F-3 @ 7] = D. 625

----D-'--- F-3ISC-l COmbinatiOn} -----<>---- F-4ISC-l combination @ 7] = D_ 625 ---- -----&----- F-5ISC-l combination --.-~-.-- SC-l @ 7]- 0.625

~

" '-"-

Device

'. ,

"'-

4,~ \

.~

, \

89

(j ~~--~--~--~--~--~--~--~--~.~ 0 .1 .2 . .3 .4 .5 .6 .7 .8 1.0

T)

Figure 31.- Effects of slot-fence combinations on spanwise leading-edge static pressure distribution at a = 16°. -

in local effective angle of attack delays separation (see fig. 26),

with accompanying improvements in leading-edge suction and thrust.

90

In order to determine the role that the upper and lower edges of

the fence play in the flow"mechanism of the slot-fence combination,

tests were run with each edge alternately removed. Results of pressure

measurements and thr~st integrations appear in figures 31 and lO,

respectively. Removal of the lower edge of th~ fence (F-4) results in

little effect inboard but substantial high-a losses of leading-edge

suction outboard. The lower edge of the fence, therefore, may have

been helpful in guiding the lower surface flow through the slot (see

BACKGROUND section). Removal of the upper portion of the fence (F-5)

also results in outboard losses of leading-edge suction and thrust.

This is most likely attributed to elimination of the upper surface

vortex tripped by the upper edge of the fence, along with its acco~pany

ingdownwash velocity outboard. The upper surface flow characteristics,

therefore, revert back to those of a single slot, with the character

istic high-a loss of leading-edge suction and thrust.

In summary, the fence possesses significant drag-reduction poten

tial~ in addition to its ability to improve the longitudinal stability

characteristics of the delta planform. The high-a loss of leading-edge

thrust occurring inboard of the device was found to be an inviscid

effect (viz., isobar unsweeping), rather than due to viscous accumula

tion inboard as originally thought. Finally, addition of an adjacent

inboard slot resulted in a slight performance improvement, possibly

through the formation of a favorably rotating vortex outboard of the

fence.

91

Pylon Vortex Generators

The function of the pylon vortex generator relies on the formation

of a streamwise vortex at the sharp upper edge of a vertical blade

extending ahead of and below the wing leading edge (see BACKGROUND

section). Figure 32 summarizes the capabilities of this device (VG-l)

in single and multiple configurations. The VG-l had a leading-edge

sweepback angle of 300 and toe-in angle of 100 (see fig. 5). These

values were found in previous research (ref. 5} to produce the best

compromise between low- and high-a performance. The axial force and

percent drag reducti on (PD) "curves show the expected 1 ow-a drag pena 1ty

with a single VG-l at n = 0.50. This penalty continues to a = 80

(which is the angl~ for separation onset on the basic wing), at which

point the device begins to show a favorable effect. Beyond 100 a, a

single VG;l provides substantial drag reduction, with a maximum improve

ment of approximately 15 percent (PO curve) at a = 14°. In addition,

the mid-a CA reversal characteristic of the basic wing is eliminated.

A constant CA increment at higher a accounts for the tapering off

of the relative drag improvement seen in the PO curve.

The longitudinal stability effects of a single VG-J located at

n = 0.50 are reflected by improved linearity of the pitching moment

curve in figure 32. Alleviation of the initial pitch-down and a 20

delay in pitch-up, to a = 13°, are indicated. The pitch-up with the

VG is also very mild in comparison with the basic wing. The normal

force curve shows a loss of lift beyond 10° a, implying a reduction

of vortex lift contribution due to delayed separation. In addition,

CA

PD,

%

·02

o

-.02

-.04t -.06

o

mi 20 ~

10

o

-'10

-20

o

5

Drag reduction

if

I 5

92


--0-- lxVG-l @7]=0.50 ---G---- 2.x VG-l @7]=0.25, 0.50

----<>---- 3 x VG-l @7]= 0.25, 0.50, 0.75

. I 10 15 20 25

ex, deg

10 15 20 25

ex, deg

Figure 32.- Force and moment characteristics of single and multiple pylon vortex generator configurations.

93

.015

.010

Cm

.005 ''c(::~

OLI ______ ~ ________ L_ ______ ~ ______ _L ______ ~

o 5 .10 15 20 25

a, deg

Basic wing

--0-- Ix VG-I@7]=0.50 --~-- 2 x VG-l @ 7]= 0.25,0.50 -----<>---- 3 xVG-l @ 7] = 0.25, 0.50, 0.75

LO[ .8 ,/ .6

CN

.4

.2

I.. ~-----' 10 15 20 25

a, deg

Figure 32.- Continued.

f'

'-"l

.006

.005

.004

.003


----<0 'lx VG:-l@'TJ=0.50 ---0--- 2 x VG,.l@7J= 0.25, 0.50 ----<)---- 3 x VG::'l@ 'TJ= 0.25" 0.50, 0.75

A

1\ I \ I \ I \ I \ I \

\ \ \ I 1 \ \

~

.00?4-~

.001

o 1 ,I

o 5 10 15 20 25

ex, deg


94

95

the magnitude of the fluctuations in the C1 curve of the basic wing

are significantly reduced, indicating a more gradual and symmetric

growth of separation.

The performance of the vortex generator can be further enhanced

throu~h its use in multiple arrangements. Results of balant~ measure

ments On double (n = 0.25, 0.50) and triple (n = 0.25, 0.50, 0.75) VG-l

configurations are presented in figure 32. Generally, an increa~ing

number of VGls results in progressively higher low-a drag penalty (PO

and CA curves and Table III) but improved high-a performance. Drag

reductions of 17 percent with two and 24 percent with three VG-l IS

(relative to the basic wing) are obtained at a ~ 140. However, fUrther

losses of normal force are found,as expected, due to improved suppres

sion of flow separation and vortex lift. The longitudinal stability is

similarly improved progressively with increasing number of VGls, as

shown by the improved linearity of the Cm curve. Two VG-l IS have no

further i nfl uence on the pitch-up' at a = 130 but produce a more s'tab 1 e

configuration at higher a than a single VG-l. Three VG-lls result in

a more gradual initial pitch-down together with complete elimination of

instability., This is significant since the point of maximum percent

drag reduction now lies in a longitudinally stable a range. In

general, all VG arrangements eliminate the wing rock of the basic wing

in the intermediate a range.

As briefly noted in the BACKGROUND section, the formation of a

vortex at the swept-forward leading edge of the vortex generator isa

result of the prevailing sidewash velocity ahead of the wing leading

96

TABLE Ill. - VORTEX GENERATOR ZERO-a DRAGCHARACTERI STICS

Baseline

L.E. length reduction

Chord reduction

Diagonal cut

Wi th missil e

Without missile

{ { {

{

Device

Basi c wi ng

1 x VG-1

·2x VG-l

3 x VG-l

1 x VG-2

1 x VG-3

.1 x VG-4

1 x VG-5

. 1 x VG"6

1 x VG-7

1 x VG~8

1 x VG-.9

Percent increase in wetted area

over BW

-o .s

.9

1.4

.3

. 1

•. 2

.2

.3

. 1

·2 .~6

1.3

CD Peroentin creas e

0 in CD over BW 0

.'

0.01261 -

.01344 6.58

.01391 10.31

.01448 14~83

.01323 4.92

·."01312 4.04

.01315 4.28

.01327 5.23

. 01338 6.11

.01314 4.20

.01368 8.48

.01330 5.47

97

edge. At low a, this sideVlash velocity is not of sufficient magnitude

to cause vortex formation due to the toe-in angle of the device (see

fig. 5). When the sidewash is less than the toe-in angle of the VG, a

pressure drag is produced in addition to the skin friction drag of the

device. Thls accounts for the low-a drag penalty noted in the PO and

CA curves. Increasing a, accompanied by increasing sidewash velocity,

eventually leads to vortex formation at the VG leading edge. This

vortex travels over the wing upper surface and rotates in a sense so as

to induce a downwash velocity on its outboard side. Due to .the pre

vailing sidewash velocity, the upper surface path followed by this

vortex is angled toward the wing tip. Therefore, it should be noted

that any reference to inboard or outboard effects of the VG applies to

either side of this vortex rather than the device itself.

As shown by the separation boundary in figure 33, the vortex action

of a single VG-l at n = 0.50· results in substantial delays in local

separation on the outboard side. The extensive delay adjacent to the

VG is attributed to the exponential variation of the vortex tangential

velocity and, thus, the induced downwash. Slight delays of separation

inboard are attributed to the compartmentation effect of the device.

The sense of rotation of the VG vortex is opposite that of the primary

vortex, thereby reducing its strength and obstructing its inboard move

ment. Static pressure variations around the wing leading edge, in

figure 34, indicate that once inboa~d separation does occur (between 140

and 16° a at STA 3), the upper surface flow eventually stagnates at a

'T)

1.0

0.9

0.8

D.T.

0.6

0.5 L .

0.4

0.3

0.2

0; 1

oUtboard

98

- - - - -~ Basic wing

~ VG"'l @ 7]= 0.50 --(:: ..... VG-l @ 7,/= 0.50, flow·

stHlatta.ched. (!t highest test eX

~() n c; .' . _. ~ - ..... , .. 1 .. location of V{;

LE. . 'Separated

LE. Attached

--

o ! '"

o 5 10 15 20 25

Qsep,deg

Figure 33 • .;. Effect of the pylon vortex generatorbnleading-edge separation (C derived).

PLE

99

-4 -4 ------ Basic wing

ex = 140 --'0--- 1 x VG-l @ n = 0.50

Cp,3 CP4

STA 11 11 Y3('7}=0.45)

0 2 3 4 5 0 2 3 4 5 ...---4 ('7}:0.57)

X, em X, ·en:

-4,-- -4

ex = 16° -3"

-2

C, ~ >-,.:-0 C

P4 .3 , . , , -1 ,

o ~-

1 0 2 3 4 5 0 2 3 4 5

X, Cm X. em

-4r- -4

-3' ex = 18°

-2 C .

P.3 ~ CP4

-1

1 0 2 3 4 5 0 2 ~ 4

X, em X, em -5,--

~ -4,

ex = 23° -3

-2

C~ CP4 P3 \ "'"

-1 ' ,

0

1 0 2 3 4 5 0 2 3 4 5

X, cm X. em

Figure 34.- Local static pressure effects of the pylon vortex generator. Shaded regions indicate lower vortex contribution.

100

higher constant pressure than on the basic wing (see d = 180 and 230

plots) .. This results in a smaller contribution to leading~edge suction.

An aspect of the VG flow mechanism not yet discussed is the forma

tion of a vortex at the lbwer edge of the device, as shown in sketch D

of the BACKGROUND section. At the lowest angles of attack, this vortex .' ,

passes beneath the wing, undetected by the leading edge. At higher a,

howeVer, the vortex begins to impinge on the lower surface of the wing,

as indicated by the m.inor suction peak in the lower surface pressures

at STA 4 at a = 140 ,(fig. 34). This suction peak moves progressively

closer to the leading edge with increasing a, sug~esting that at angles

higher than those considered here .• this lower-edge vortex would pass,

completely over the wing. Once this occurs, its own contribution to the

wing leading-edge thrust (cross-hatched in fig. 34) is lost. In addi

tion, this vortex may act to degrade the upper VG vortex.

Local thrust variations for single and multipJe VG-l configurations

appear in figure 35. The high'-ci CA improvements noted in figure 32

are shown tc result from thrust enhancement outboard of each VG. The

figu:re also demonstrates the ability of m.ultiple VG"s to effectively

eliminate the adverse inboard effect of the single device. Note that

at low a (attached flow on the basic wing), local thrust values just

outboard of each VG (for all VG-lconfigurations) fall below those of

the basic wing. This is attributed to the induced downwash in these

reglons~ which effectively reduces the local angle of attack and, thus,

the sucti on forces around the leading edge. At higher a, however, thi s

~

.10

.08

.~---:--- Basic wing

----{)----- 1 x VG-l @ n = 0.50 _.-0--.- 2 x VG-l @ n = 0.25, 0.50 ..:... .. -V-.. - 3 x VG-l @ 11 = 0.25. 0.50, 0.75

.12

.10 l

.08

CT .06 locI

. CTIOC2 .06

.04 n = 0.20

.02

010'1' I I I---.J o 5 10 15 20 25

lX, deg

.10

.08 n = 0.45

CTIoc3 .06

.04

°tL o I· I o 5 10 15 20 25

lX, deg

.1°i .08~ n = 0.70 .il.

CTIOC5 '.06

.04

.02

OLI_~~_-L __ L-_-L_~ o 5 10 15 20 25

lX, deg

.04

.02

o I 0"1 o

.10

15 20 25

lX, deg

.08 n = 0.57 if; ,.

CTIOC4 .06

.04

o2~L 0~~1 I o 5 10 15 20 25

lX, deg

.1°i .08~ n = 0.82

CTloc6 .06

.04

.02

01 V/!

o 5 10 15' 20 25

lX, deg

101

Figure 35.- Effects of single and multiple pylon vortex generators on leadtng-edge thrust.

102

effect acts to delay separation and~thus, retain leading-edge suction

to the highest angles tested.

The spanwise leading-edge pressure distributions at Ci. = 140

(where the maximum percent drag reduction occurs) for single and

!l1ultiple VG-l configurations are presented in figure 36. The single

. VG:..: 1 curve s'hows only a local i zed ineffeCt i venes s on the i nboa rd side of

the device. In addition, the plots sho\'{how the mutual interaction of

multiple VG's acts to enhance the leading'-edge suction along the span.

The wavelike shape of the curves just outboard of each device is

attributed to the outboard drift of the VG vortex, discussed previously.

Compar'i son of the curves -indicates that the VG has a more pronounced

effect near the tip, where the flow has a tendency toward early separa

tion. This accounts for the substantial drag and longitudinal stability

improvernentswith the addition of a VG-l at n = 0.75 (triple VG-l

corifiguration). It is believed that two VG-l's located at n = 0.50

and 0.75 would provide almost the same leading-edge thrust enhancement

of the triple VG-lconfiguration tested nere.

Plots of CpLE versus CTloc at various spanwise positions for

the single VG-l configuration are presented in figure 37(a). The

linearity of the STA 4 curve at high Ci. is consistent with an attached

flow condition. The reduction in slope of the curve from its initial

value is att~ibuted to the suction peak produced by the lower VG vortex,

which additionally contributes to the leading-edge thrust but is not

detected in the CpLE measurements until it has passed over the leading

edge at higher Ci.. Without the suction contribution of this lower VG

...... -

C

103


-0-- lxVG-l @ 7]-0.50 ---0--- 2 x VG-l @ 7]= 0.25, 0_50 --.--0---- 3xVG-I@ 7]=0.25,0.50,0.75

-4.5

-4.0

-3.5

-3.0

-2.5~ NB

/ +------''-, PLE

-2.0

-1.5

-'1.0

.8 .9 1.0

7J

Figure 36.- Spanwise leading-edge static pressure distributions for single and multiple pylon vortex generator configurat~ons at a = 14°.

-5

-4

-3

-2

-1

0 0 0

I-

STA 1 2

3 4

5

0

0.02 -,

6

(a)

- - - - ..,... - Basic wing (all STNs)

0

---{o.---..,-----'l[)I----~O>---

a. ----1:1:33-----,

~---'lD~-

0 0

C Tloc

ST A 1 (7] = 0.20) 2 ( = 0.33) 3 ( = 0.45) 4 ( = 0.57) 5 ( = 0.70) 6 ( = 0.82)

0.02 0.04

Overall spanwise effect.

1 x VG-l @ 7] = O. 50

0.06

Figure 37.- Effect of the pylon vortE!X generator on leading-edge static pressure-thrust relationship. Selected high-a data points have been omitted for clarity. Angles in parentheses refer to basic wing data (see fig. 16).

~

C PLE

-5

-4

-3

-2

-1

o '/>(

o

.; ';

If

0.04

/

~ /.

Lower VG vortex th rust contribution

®/ /

/

~ Linear extension of / low-a portion of curve

--"-76 STA 4 (7]= 0.57)

® Data pOint recomputed based on elimination of lowe r VG vortex suction

0.06

C Tloe

0.08 0.10

(b) Local outboard effect of lower VG vortex.


105

0.12

-... ,. ..... '.

106

vortex, therefore, the CpLE versus CTloc curve would continue along

its initial slope. To check this hypothesis,. the influence of the lower

surface suction Was removed by reintegrating to obtain leading-edge

thrust using lower surface pressures measured on the basic wing at STA 4

for a = 180 and 230. The resulting data points are plotted in fig-

ure 37(b), along with a replotting of the STA 4 data from part (a).

Indeed, these recomputed pOints are found to shift much closer to the

linear extension of the low-a portion of the curve, reflecting the sig

nificant thrust contribution from the lower-edge vortex.

In an attempt to all evi ate the low-a drag penalty associ ated with

vortex generators, possible reduction of the VG size without adversely

affecting high-a performance was investigated. The size reductions were

in the form of progressive lower-edge cut-off, which effectively reduced

the leading-edge le~gth of the VG, and back-edge cut-off, which reduced

the chord of the device. In addition, a diagonal cutback (in effect

removing the chordwise tip) was included as another means of VG size

reduction. Geometric details of these size modifications are presented

in .figure 5.

Zero-a drag data for single VG configurations utilizing the various

geometries tested appear in Table III. As expected, lower-edge (VG-l

to VG-2, 3) and diagonal cutbacks (VG-l to VG-6, 7) resulted in sub

stantial zero-a drag reductions. The initial chord reduction (VG-2 to

VG-4) also produced a low-a drag improvement; however, further cutback

(to VG-5) resulted in a drag increase. This anomalous result may be due

to the construction method used. As previously noted, an additional

107

thickness was used on the outboard side of the VG for added stiffness to

avoid excessive deflection. As sketched below, cutback to VG-5, thus,

resulted in a thick base, with the additional pressure (base) drag

appearing in the overall zero-a drag of the configuration.

L. E.

VG-2 Added thickness

./

./ ./

I.E.

L.E •

VG-5

T.E. VG-4

VG-2

Sketch J

Figure 38 shows the overall drag and longitudinal stability effects

of variations in VG leading-edge length (lower-edge cutback). A 30 per

cent reduction from VG-l (to VG-2) results in the appearance. of a CA

reversal at a ~ 13°. In addition, substantial drag increase and severe

pitch-up are indicated at high a. Further reduction in leading-edge

length (to VG-3; 60 percent reduction from VG-l) results in little

further effect on drag up to 16° a but a loss of effectiveness at

higher a~· This is also reflected by increased severity of pitch-up at

a ~ 190.

CA

Cm

·02.l


~ ---0- I'VG-I} Of-------~-- Ix VG-2. @ 7]=0.50 ----<>---- I x VG-3

I b -.02

~""7::~ ---

~--=:B -.04 '

-.06

I o 5 ,10 15 20 25

lX, deg

.02~~

.015

.010

.005

Xcm = 0.589 cr

? f //-- .... tp

/ f

)

, I

\J

o LI ____ L-____ ~ ________ ~ ______ ~ ________ ~

o 5 10 15 20 25

lX, deg

IDS

Figure 3S.- Performance effects of pylon vortex generator leading-edge . 1 ength reducti on.

109

The primary function of the VG leading edge is to provide a sl'larp

edge along which the VG vortex may form and build in strength before

passing over the wing. A decrease in the length of this edge, there

fore, would be expected. to reduce the vortex strength. The-accompanying

reduction in downwash on the outboard side then results in earlier

separation. This effect is evident in the static pressure distribu

tions around the wing leading edge at a = 140 (STA's 5 and 6), in

figure 39(a). Figure 39(b) presents pressure distributions just out-

board of the VG (STA 4) at a = 16° and 190. With VG-3, there is no

evidence of the lower-edge vortex suction on the lower surface pressures

as was the case with VG-l. The lower vortex may, thus, be passing over

the wing leading edge at a lower a. The opposite sense of rotation of

this lower vortex would induce an upwash velocity outboard, leading to

earlier separation; in addition, the associated suction peak on the

lower surface and its contribution to the leading-edge thrust (shaded in

fig. 39(b)) is lost. A second advantage of a long VG leading-edge

length may, thus, be its ability to keep the lower VG vortex below the

wing and acting near the leading edge to higher a.

The effects of VG chord reduction on drag and longitudinal sta

bility are shown in figure 40. For structural reasons, VG-2 (30 percent

leading-edge length reduction from VG-l) was used as the baseline geome

try for analyzing this parameter. Axial force indicates that a25 per

cent reduction in VG-2 chord (to VG-4) produces a drag improvement

beyond 12° a, in addition to the low-a drag reduction noted earlier.

In addition, the severe pitch-up at a '" 19° of the VG-2 has been

Cps

~

-1

o

7] = 0.70

--D---rl LS-- ___

--0

- - - - - - Basic wi ng

---0-- lxVG-l @ '7=0.50

__ ~-- 1 x VG-3 @ '7= 0.50

VG-l

1 LI ______ ~ ____ ~ ______ _L ______ L_ ____ ~

o 2 3 4 5

X, em

-4, -3~ 7] = 0.82

Cp6

-1

o

1 o 2 3 4

X, em

(a) a = 14°.

110

5

Figure 39.- Outboard static pressure effects of pylon vortex generator leading-edge length reduction.

CP4

- - - - -- Basicwing

a z 160 . -0-- lxVG-1 @ 1]=0.50

---Q--- lxVG-3 @ 1]=0.50

VG-l

1 IL-__ -L ____ ~ __ ~ __ --~--~

o 234 5

X, em

\

~ a -190

~ C

P4

o .1 2 3

X, em

(b) a = 16° and 19°.


III

4

.02

~

o

-.02

CA

-.04

-.06

o

.02 ~v.,..;i(J

.015

.010

Cm

.005

00

5 10

Xcm =.0.589 Cr

5 10

- ~ - - - - Basicwing

-----0-- 1 x VG-2} ---D---" 1 x VG-4 @ '1]= 0.50 ----(>---- 1 x VG-5

'-

15 20 25

ex, deg

15 20 25

ex, deg

Fi gure 40. - Performance effects of pylon vortex generator chord reduction.

112

113

eliminated. Further chord reduction (to VG-5; 50 percent reduction from

VG-2), however, results in increased high-a drag and reappearance of

pitch-up at a ~ 190. Local thrust variations in figure 41 indicate

that these trends result primarily from f16w modifications outboard of

the device. Substantial high-a thrust enhancement is indicated at

STA's 5 and 6 with the initial chord reduction (VG-2 to VG-4), but this

improvement is lost with further cutback (to VG-5). Therefore, there is

a specific VG chord within the range tested which will produce the

optimum combination of low- and high-a performance.

The original design of the vortex generator presumed that the flow

mechanism depended totally on the vortex formed at its leading edge and.

passing over the wing upper surface. However, it now appears that the

vortices formed at the lower and back edges of the device may be of

importance. As the VG chord is reduced, the proximity of these edges to

the VGleading e~ge eventually becomes s~ch so as to cause interference

between corresponding vortices. In addition, at a high enough a, these

lower- and back-edge vortices may pass completely over the wing leading

edge and interfere with the upper surface flow. Specific effects on

performance would seemingly depend on' the. strength of these vortices,

which depends partially on the length of the corresponding edges. How

ever, at this stage there are insufficient data to provide definite con

clu~ions on the VG flow mechanisms producing the trends observed with

VG chord reduction. The geometries tested, however, did provide a

general idea of the VG proportions required for optimum performance.

·.;.,.

.10

.os

"'.' Crloe1 .06

C Tloe}

.04

.02

01.0'1 o 5

.10

. 0sL

.06

.04

.02

.;: .... ,

= 0.20

10 15 .20 ·:;25.

IX, deg

n = 0.45

I /i .

01 QI o 5 10 15 20 25

IX, deg

.10

.OS n = 0.70

CTloe5 .06

.04

.02

0 0 5 10 15 20 25

IX, deg

- - - - - - Bas i c wi ng

··~.lx VG,..2} '. -. .......u:-.- 1. x VG-4@Tl=0.50 -:-;-.. -:<>-.. - l' ,x VG-5

.10

.os n = 0.33

CTloe2 .06

.04

.02

0 0 '5 10 15 20 25

',12~ IX, deg

.10 .

T .OS

C . ,Tloe4 .06

,04

.02

0 0 5 10 15 20 25

:10

.os

Cr ' ' . . Ci6-loe6

,04

.02

01 v'l o 5

IX,deg

n = 0.82

10 15 20 25

IX, deg

,Figure 41.- Effects of pylon vortex generator chord reduction on leading-edge'tnrust.

114

115

Performance effects of simultaneous cutbacks of the lower and back

edges of the VG, in the form of diagonal cuts (see fig. 5), will now be

considered. Figure 42 presents force and moment data for single VG-l,

6,and 7 configurations. Axial force indicates a reduction in suction

effectiveness beyond 130 a with increasing cutbackr In addition, the

initial cutback (VG-l to VG-6) results in pitch-up at a ~ 190, which

becomes more severe with further reduction (to VG-7). Inspection of

local thrust variations, in figure 43, reveals that these effects on

high-a performance are again attributed to outboard flow modifications.

The losses resulting from cutback to VG-6 are most likely attrib

uted to an effective elimination of the lower edge of the VG. This

eliminates the lower surface suction peak produced by the lower-edge

vortex and, thus 1 its contribution to the leading-edge thrust. Another

possibility is that the proximity of the leading and back edges near the

tip of the VG-6 may cause a weakening of the VG leading-edge vortex as a

result of interference with the counter-rotating back-edge vortex. The

additional loss of effectiveness with cutback to VG-7 is believed to be

primarily the result of leading-edge length reduction. As previously

noted, this reduces the strength of the VG leading-edge vortex by

reducing the distance over which it forms.

The possible use of pylon vortex generators also as carriers of

slender external· stores (such as air-to-air missiles) will now be con

sidered. The VG tested was sim~lar to VG-l but had an extended chord

in order to provide a mounting position for the wooden dowel si~ulating

the external store (see fig. 6). This device was tested with (VG-8)

CA

Cm

.02

-0

-.02

-.04

-.06

~

Basic Wing

~ lXVG-l} ~~-o--- 1 x VG-6 @ 7]=0.50 _ .. --"<)-._- 1 x VG-7

~.~

o 5 10 15 20 25

.02~

.005

Xcm = 0.589 cr

Lx.; deg

o LI ___ ~ ______ ~L-______ L-____ ~~ ______ ~

o 5 10 15 20 25

a. deg

Fi gure 42. - Performance effects of pylon vortex generator di agona 1 cutback.

116

117

STA ------Basic wing 1

~1 x VG-11 2 3 ----0---1 x VG-6 @ 11 = 0.50

4 ----<>---- 1 x VG-7

.10..- '\ /5 .10 6

.08 LJ .08 L- n = 0.33

C cT T1oc1 .06

loc2 .06

.04 = 0.20 .04

.02 .02

0 0 0 5 10 15 20 25 0 5 10 15 20 25

a, deg .12,- a", deg

.~ .10..- .10 I-- /; /j

n = 0.45 .08~ >5;'

.081-- n = 0.57 b CTloc3 .06

C " Tloc4 .06

.04 .04

.02 .02

01 01' 0 0 5 10 15 20 25 0 5 10 15 20 25

a, deg a, deg

.10 .10

.08 n = 0.70 .08~ n = 0.82

CTIOC5 .06 C T loc6 .06

.04 .04

.02 .02

0 0 0 10 15 20 0 5 10 15 20 25

a, deg a, deg

Figure 43.- Effects of pylon vortex generator diagonal cutback on leading-edge thrust.

, '

'. i

118

and without (VG-9) the store in single VG configurations (at n = 0.50).

Table III indicates a low-a drag penalty due to the chord extension and

store addition; however, this is not of concern since a drag increase , .'

would in any case be obtained with a pylon. Selected results of farce

and, momen't measurements on VG-8 and VG-9 canfi guratians are presented

and campared with the baseline VG~l configuration (at n= 0.50) in

figure. 44. The CA curve indicates a loss of drag perfarmance beyand

120. C/. with VG-8 as campared wi th VG~ L Stati c pressure data~ nat pre-

sented, indicate that the external stare prevents the farmation of a

vortex at the lawer edge af the VG, thereby reducing the lawer surface

cantribution to' the wing leading.:.edge thrust. However, a, significant

high~C/. drag advantage aver the basic wing is s:till realized. In addi

tion, VG-8 delays the pitch-up af the basic wing by approximately 20.,

to c/., ~ 130 ~

Figure 44 shows that once the slender external store has been

re 1 eased, the, extended, chard VG:-9 conti nues to perfarm as a drag

reducer. Hawever, the drag reductian effectiveness af the device

appears to' have been hampered by the chard extensi on (from VG-l).

Sur,face pressure data indicate that this may be a result af the per-

sistence af the lawer-edge vortex along the extended edge, eliminating

its suctian effect an the wing leading~edge region. However, the

resulting drag reductian would, still be an improvement aver a conven-

~ianal pylon canfiguratian, which would cantinue to' praduce a drag

penalty even at higher C/.. Further research should be performed an

madificatians that will allow far, therealizatian of a greater portian

~

.02

o

-.02 --

CA

-.04

-.06

o

Cm

.005

0 0

---0-----0----_.-(>-._-

1 ·1

Basic wing

1 x VG-l} 1 x VG-8 @ 7]=0.50 1 x VG-9 .

.~

119

5 10 15 20 25

Xcm = 0.589 Cr

5 10

n, deg

15

n, deg

7""'-< (/ / ( ( / ( (,

,\. __ -- VG-8 w/mi ssile i

.... "'// VG-9 w/out missile

20 25

Figure 44.- Performance of an extended chord pylon vortex generator' utilized as a carrier of slender external stores.

120

of the VG drag-~eduction potential, both with and without the external . .,:

store. Variations in VG chord length and chordwise location of the

s~ore along the lower edge of the VG may be initial steps .

. In summary, the pylon vortex generator produces substantial

improyements in high-a drag and longitudinal stability when utilized on

highly swept leading edges. The lower VG vortex apparehtly plays an

important role through its own contribution to the leading-edge thrust

and, thus, should be considered in the design of the VG shape. The

mutual interaction of multiple VG's acts to further enhance the leading-

edge suction along the span. However, a low-a drag penalty is charac

teristic of the vortex generator. VG size reductions in the form of

leading-edge length and diagonal cutbacks reduce the low-a d~ag penalty

but also result in a loss of performance at high a. Reduction in VG

chord to a certain degree produces performance improvements at both low

and high a. The vortex generator has potential also as a carrier of

slender external stores. Although the performance of the extended chord

VG is not as efficient as that of the baseline VG-l, substantial drag

and longitudinal stability improvements over the basic wing are produced

both with and without the external store.

121


The common basis of the devices discussed thus far has been that of

maintaining attached flow at the wing leading edge in order to retain

leading-edge suction to higher a. An alternative to this conventional

approach of drag reduction is provided by the sharp leading-edge exten

sion (SLEE), or vortex plate, which manipulates the natural tendency

toward flow separation and vortex formation at a swept leading edge.

A sharp-edged plate lying in the plane of the wing lower surface and

projecting ahead of the wing leading edge forces separation along its

leading edge and subsequent vortex formation. Ideally, this vortex is

maintained just ahead of the wing leading edge along the entire span,

with its associated suction acting on the leading-edge thickness to

generate a thrust force. Flow reattachment just aft of the leading

edge curvature helps to maintain attached chordwise flow on the wing

upper surface (see sketch A in INTRODUCTION).

Balan~e data for a SLEE configuration utilizing an 0.71 cm exten

sion (see fig. 8) is presented in figure 45. This device had a spanwise

coverage of n = 0.25 to 0.93 and was tested with an F-2 fence at its

inboard edge (n = 0.25). The fence was intended to obstruct the span

wise boundary layer flow originating b~tweeri the wing apex and the

inboard edge of the SLEE, to allow the formation of a clean SLEE vortex.

This fence was shbwn in previous testing (ref. 5) to delay a mid~a

longitudinal instability produced by the SLEE and to maintain CA

improvement to significantly higher a. At low a, the vortex suction

is insufficient to offset the additional drag of the SLEE, accounting

--- ..... "-

o ~

" , ~

" -.02

CA

-.04

-.06

o 5

.04

o

-.04 em

-.08

-.12

o 5

,

-- - -- - Basic wing

---0-- 0.71 em ext. SLEE; wI F-2 @ 7]= 0.25

I· I· 10

a. deg

I

15 20

x . ~ = 0.474

cr

'-

"

25

'-

10 15 20 25

a, deg

122

Figure 45.- Force and moment characteristics of a sharp leading-edge extension configuration.

eN

LID

1.0

.8

.6

.4

.2

10 I

./ /

\ \

\ 8

6 11 4~ ,I I

2

a, deg

\

/

/ I

/ /

/

/ /

/

/ /

/

- - - - - - Basiewing --0- 0.71 em ext. SLEE;

wI F-2 @ 7]= 0.25

"-"-

"-"-

"-"-

OLI ________ L-______ ~ ________ L_ ______ _L ______ ~

o 5 10 15 20 25

a, deg

Figure 45.- Continued.

123

C'Z

.006

.005

.004

.003

.002 , .

. 001

- - - ....;; ~......;. ~asic wing ~,O.71cm ext. SLEE;

, I I

!

I

I

wI F2 @7J= 0.25

A

1\ 1 \ 1 \ I \ I 1

- \

\

\ .\

I I I

OLI ______ ~ ______ ~ ________ L_ ______ L_ ____ ~

o 5 10 15 20 25

a, deg-

Fi gure 45. - Cone 1 uded.

':.'

124

125

for the net CA increase. Therefore, to avoid a cruise drag penalty,

the SLEE should be designed as a retractable device. Beyond a ~ 12°,

the vortex strength and its optimum position produce substantial CA

improvements over the basic wing. The ability of the SLEE to maintain

its vortex ahead of the leading edge is also reflected in delayed onset

of vortex lift, reducing the CN beyond a ~ 10°. From a performance

standpoint, the low-a drag increase results in severe losses in LID.

Beyond 120 a, however, substantial improvements are noted, with the LID

increment tapering off gradually with increasing a. In addition, the

rolling moment instability of the basic wing between gO and 140 a is

eliminated. The longitudinal stability of the configuration is slightly

r~duced fro~ that of the basic wing, but the linearity of the Cm curve

is significantly improved. Here, the moment reference center has once

again been selected to better show up the effects of the device. The

new location (see DATA REDUCTION) is 3.82 cm (5 percent of root chord)

aft of the original moment reference center given in figure 2.

Static pressure distributions around the wing leading edge for the

0.71 cm ext. SLEE configuration at a = 120 are presented in fig-

ure 46. Attached leading-edge flow inboard of STA 6 on the basic wing

results in high suction peaks near the leading edge with subsequent

pressure recovery on the upper surface. However, the vortex-induced

reattachment due to the SLEE produces a stagnated flow condition, with

elimination of the peak negative pressures near the leading edge, but

with high negative pressures over a greater portion of the leading-edge

curvature. Note, especially, the high negative Cp on the lower

STA 1

2

- -'- - - -'Basi ewing

3 4

-0-----, Lo~er surface} 0.71 cm ext. SLEE; --'-'-0-----' Upper surface wI F-2 @ Tl = 0.25

- -4

-3

5 6

n =,0.20

CP1

-1

11 1 1 I I 0 2 3 4 5

-I x, em

-3 , n 0.4-5 \

\

-2~\ C I ",

P3 \ .--n..:--_

or "-1 _ I---~I----I----[

0 2 3 4 5

x, em

-I -3

, n = 0.70 \ \

-1: Cps \ ',---- ____

-1 \ \ ,

~r " ---------

0 2 3 4 5

X, em

Cp:z

CP4

:, CP6

-4

-3

1

, \

0

-4~

-3

k _ -2

-1

f" 0

-4

-3

I -2

0

1 0

n = 0.33

2 3 4 5

x, em

n = 0,.57

------------

2 3 4 5

X, em

n = 0.82

--------

-----------

2 3 4 5

X, em

126

Figure 46.- Effects of the ~harpleading7edge extension on static 0 pressure distributions around the leading edge at a = 12 .

127

surfa~e in contrast with the positive values measured on the basic wing.

With the 6asic wing at a lifting condition, the stagnation point occurs

just below the leading-edge curvature. The associated positive pressure

acting on the forward sloping surface (see sketch K, below) contributes

a positive axial force and, therefore, drag. On the other hand,-when

the stagnation point lies on the lower surface of the SLEE, the positive

pressure produces no axial force (sketch K). In addition, the high flow

velocity at the SLEE edge develops a high suction level to generate a

thrust force.

~

Basic wing wI SLEE

Sketch K

Leading-edge pressure variations, in figure 47, indicate that at

low a the leading-edge suction produced by the SLEE vortex is sub

stantially less than the potential flow suction of the basic wing.

However, once potential flow is lost, the SLEE eventually provides an

improvement over the basic wing (beyond a ~ 150 at n = 0.724,

a ~ 190 at n = 0.416). Static pressure distributions around the wing

C PLE

C PLE

-3 r-

-2 t -1

/

I ~~ 0

0 5

-3

-2

-1

o \ l o 5

128

Orifice 123 - - - - - - Basic wing

0 0.71 cm ext. SLEE; 221 wI F-2 @ 7] = 0.25

~ "'"' Orifice 123 (7]= 0.416)

/ ., / \

I '--

I /

/ /

/ /

/

10 15 20 25

a,deg

;---, I \

Orifice 221 (7] = 0.724)

/ \ / \

10 15 20 25

d, deg

Figure 47.- Leading-edge suction effects of the sharp leading-edge extension.

129

leading edge at ~ ~ 220 (fig. 48), indicate that the greatest contribu

tion to the leading-edge thrust improvement comes from the low pressures

induced on the lower surface. The accelerated loss of thrust at STA 1

may be attributed to the unsweeping of the upper surface isobars by the

fence (see Fence section). The spanwise CTl distributions, in. oc figure 49, are consistent with the CpLE data just presented. At

a = 120, the leading-edge thr0~t produced along the SLEE falls below

that of the basic wing; however, by 230 a, substantial improvements are

noted along the entire length of the device. Decreasing thrust incre

ment toward the tip suggests an expansion or diffusion of the SLEE

vortex core due to increasing upwash and entrainment effects. Means to

alleviate the loss of SLEE effectiveness due to vortex diffusion will

be discussed later.

Optimizing the SLEE extension \>li11 now be considered as a possfble

means of further performance improvements. The possibility of such

improvements is suggested by the fact that over-extension of the SLEE

would result in reattachment on the device itself, with local suction

peaks and the possibility of separation near the \!Jing leading edge. In

addition, the vortex suction would mainly act in the normal direction,

with little effect on axial force. This condition is depicted in

sketch L, below:

~

Cp1

CP3

Cps

STA 1 1 ------ Basic wing

3 4

5 6

-----cr----. _. Lower surface} 0.71 em ext. SLEE; ---~--- ~pper_surface wi F-2 @ n = 0.25

-5

-4

-3

-1

o

, ,

n = 0.20 --0---0-------0

1 LI ____ ~ ____ L_ __ -J ____ _L~~

o 2 3 5 4-

-4 X, em

-3 n

-1 ,

o ", [- "

1 I - I

-4

-3

-2

o

o

"' ,

2 3 4

X, em

n = 0.70

5

1 LI ____ -L ____ ~ ____ ~ ____ L_ __ ~

o 2 3 4 5

X, em

--'4 n = 0.33

-3

---------------'0::.:- --2 r- ~---o----_o c '-. P2 \

CP4

CP6

-1

o , ,

1 IL-____ L-____ L-____ L-____ L-__ ~

o 2 3 4 5

-4 X, em

-3 n = 0.57

---0-----

-1 , "

o

1 LI _-'--'-_--'-----'-_-'-__ '--_-'

o 2 3 4 5

X, em

-4

-3 n = 0.82

-2

-1

o

1 I'--_~ ____ _L ____ ~ ____ L-__ ~

o 2 3 4 5

X, em

l30

Figure 48.- Effects of the sharp leading-edge extension on static pressure distributions around the leading edge at a = 22°.

C Tloc

CT loc

.10 - - - - - - Basiewing

.08

--0--- 0.71 em ext. SlEE} I

---0--- 0.48 em ext. SlEE wi F-2 @ "l= O. 25 _ .. -(>-.. - 0.23 em ext. SlEE

- .. ---ts-.--- 0.00 em ext. SlEE

.06 a'12°

.04

.02

O~I ----~ ____ L-__ ~ ____ _L ____ ~

o

.12

.10

.08

.06

.04

.02

.2 I

\1 ,

~ SlEE apex;

F-2

.4

7J

" \ \

\ \

.6 .8

.. '~

" I ,

" "

,~

1.0

a = 23°

I I I OL- .1..6.8 1.0 o .2 _ I

7J

Figure 49.- Effects of SLEE extension on spanwise distribution of . leading-edge thrust.

131

( 132

Over-extension Under-extension

Sketch L

However, considering the range of extensions tested here, the over

extended condition is unlikely to be a factor except in a very narrow

a range. Nev~rtheless, this conditibn does provide a conceptual upper

bound for the SLEE extension. On the other hand, under-extension of the

SLEE may not allow the stagnation point to move to the SLEE even at

high a, leading to a large drag contribution, as discussed previously

(see sketch L). Therefore, the minimum reguirement for the SLEE exten

sion would appear to be that it project ahead of the stagnation point at

the highest a of interest.

Selected balance data for corifigurationsutilizing SLEE extensions

ranging from 0.71 to 0 cm (with an adjacent F.:..2 fence at the SLEE apex)

are presented in figure 50. Little effect from SLEE extension is evident

in CA up to a~· 160 . At higher a, how~ver, decreasing extension

results in progressively improve8 axial force characteristics, with the

CA

em

.0

o

-.02

-.04

-.06

o 5

.04

o

-.04

-.08 t·· -.12

. 0 5

10

---10)-------0---

----<>----_.---6---.. -

a, deg

"'

15 20

Xem = 0.474 c -r

"'-~

lO 15 20

a, deg

Basic wing

0.71 em ext. SLEE} 0.48 em ext. SLEE I

w F-2 0.23 em ext. SLEE

0.00 em ext. SLEE

25

I 25

Figure 50.- Effects of SLEE extension on performance.

133

@ 7]= 0.25

134

improvements continuing down to zero extension. In addition, the longi

tudinal stability of the configuration is slightly improved. Although

rtot presented here, effects on normal force and lift-to-drag are minor.

Comparison of static pressure distributions around the wing leading

edge for the 0.71 and 0 ·cm SLEE extensions (fig. 51) indicates a delay

in onset of vortex suction on the leading edge as the SLEE extension

is reduced. At. ~ = 120, the 0.71 cm ext. SLEE configuration is

.characterized by stagnated leading-edge flow; whereas with the 0 cm

extension, attached flow still persists on the upper surface. Appar

ently, the vortex size with zero extension is such so as to produce a

stagnated condition only on the lower portion of the leading-edge region

(see STA 5 inset sketches). This delay is shown in figure 52 to payoff

in improved leading-edge thrust at higher ~. At the highest angles

tested, a distinct advantage with reduced SLEE extension appears only

at STA's 5 and 6. However, at higher ~ngles, thrust improvements would

also be expected at the remaining spanwise stations, as the SLEE vortex

moves inboard. While the adverse inboard effect of the fence is again

seen at STA 1, the favorable outboard effect at STA 2 allows for con

tinued thrust improvement to the highest angles tested.

As already noted, the purpose of locating a fence inboard of the

SLEE was to aid in the formation of a clean SLEE vortex. The possi

bility of. substituting a chordwise slot (w/SC-l) for the fence was

investigated on the 0.48 cm ext. SLEE configuration. Selected

results appear in figure 53. Axial force is slightly improved in the

mid-~ range with the use of the slot; however, early loss of slot

CP1

c P3

Cps

STA 1

------ Basie wing

2 3 ~ 0.71 em ext. SLEE; wi F-2 @ n = 0.25 ---~--- 0.00 em ext. SLEE; wi F-2 @ n = 0.25

-4

-3

1 0

~:~ \ \

-2$:::

4 5

6

n = 0.20

-=---

2 3 4

X, em

n 0.45

Of-- '" ; I 1 -- - -,- - --,- - - -I

0 2 3 4

X, em

5

5

~ ~

~ ~\, \

' ''ri i '----_ n = 0.70 -1

~~~,~~--i-0 1 2 3 4 5

X, em

-4

-3 n = 0.33

CP2

1 0 2 3 4

-4~ X, em

-3 " r""7 , , n = U.Of

CP4

o ',_ 1'2- -o-'--D-:~:~-1· I ----1-----1-----1

o 2 3 4

X, em

~:I -2 b

. n = 0.82

C"_1~-~ \-y_---D---- ---o-~-

O~------------------1 1 1 1 1 o 2 3 4

X, em

135

5

5

5

Figure 51.- Effects of SLEE extension on static pressure distributions around the leading edge at a = 12°.

,

136

STA' ------Basie wing

1 ----cr- 0.71 em ext. SLEE; wI F-2 @ T) = 0.2.5 .·,2 ---G--- 0.00 em ext. SLEE; wI F-2 @ T) = 0.25

3

''[ ;8 4 5 ! .10r- ~../6 .10

.,

~ , n = 0.33 .os f---

, .os

CT .06 C .06~ )}'-/' loc1 Tloc2 I

.04 0.20

.04

.02 .02

0 0 0 .5 10 15 20 25 0 '='5 10 15 20 25

.12 r- a, deg p

a, deg

.10 //0 .10

.OS .OS

C .06 C .06 Tloc3 , / Tloe4

.04 .04

.02

/1 .02

01 0 0 5 10 15 20 25 0 5 10 15 20 25

a, deg a, deg

'T .10

.OS n = 0. 7Or:: .0sL n = 0.82 ---0

C .06 CT .06 Tloe5 loe6

.04 .04

.02 .02

0 0 0 5 10 15 20 25 0 5 10 15 20 25

a, deg a, deg

Figure 52.- Effects of SLEE extension on leading-edge thrust.

CA

LID

.0

o

-.02

-.04

-.06

o 5 10

':[ 6

4

2

137

- - - ---, - - Basic wing

-0-- 0.48 em ext. SLEE; w/F-2 @ 7]=0.25

---D--- 0.48 em ext. SLEE; w/SC-l @ 7]= 0.25

"

15

a, deg

_-0

20

'- ' - "

25

O~I ___ ~ __ -~---~---~---~ o 5 10 15 20 25

a, deg

Figure 53> Performance comparison of the fence and chordwise slot at theSLEE apex.

~:.

138

effectiveness (see Chordwise Slots section) results in a sudden loss of

SLEE performance beyond a ~ l~o. The same trends are evident in lift

to-drag. Although not presented, no significant effects are noted in

pitching moment. The chordwise slot, therefore, seems to be slightly

more effective than the fence in obstructing the spanwise flow near the

wing apex up to approximately 160 a. However, highly maneuverable air

craft encounter a wide range of angles of attack and, thus, the high-a

performance capability with the fence may be well worth the small

penalty at mid a.

As noted previously, the tendency of the SLEE vortex to expand and

migrate onto the wing surface will ultimately 1 imit the drag-reduction.

effectiveness of the device. The previous success of fences and chord

wise slots in compartmentation of the wing leading edge suggested that

their use along the SLEE would limit the growth of the SLEE vortex and,

thus, improve its outboard effectiveness. Each device (F-2 and slot

w/SC-l) was, thus, alternately tested at n = 0.625 (midpoint of SLEE)

on a 0.48 cm ext. SLEE configuration also utilizing an F-2 fence

at n = 0.25. Selected force and moment data appear in figure 54.

While the additional F-2 fence at n = 0.625 produces the expected

low-a drag (or CAl penalty, substantial improvement in SLEE suction

effectiveness is indicated beyond a ~ 16°, with no adverse effects on

longitudinal stability. Tuft photographs in figure 55 support the view

that this improvement is due to a compartmentation of the SLEE. In

interpretation of these photographs, a herringbone tuft pattern repre-

- sents a vortex core (shown as a dashed line), whereas a divergence of

c -A

Cm

D

-.02

-.04

-.06

-.08

o

.04 r-::

o

-.04

-.08

-.12

o

5 10

I 5 10

139


--0- 0.48 cm ext. SLEE; w/F-2 @ 7]=0.25 ---0---- 0.48 em ext. SLEE; w/F-2 @ 7]= 0.25, 0.625 - _.-<)----- 0.48 cm ext. SLEE; wI F-2 @ 7] = 0.25 and

"-

I J

IX, deg

15 20

Xcm = 0.424 c r

15 20

a, deg

25

25

- 5C-1@ 7]= 0.625

Figure 54.- Performance effects of the fence and chordwise slot in combination with the SLEE.

Figure 55.- Upper surface tuft photographs of 0.48 cm ext. SLEE with F-2 at n = 0.25 configuration with and without an F-2 fence at n = 0.625 (a = 140).

~

141

adjacent tufts indicates a flow reattachment line (solid line). In the

absence of the ~utboard fence, the SLEE vortex is free to expand, having

migrated completely onto the wing surface at a = 14°. A fence at

n = 0.625, however, splits the SLEE vortex into two smaller vortices;

one emanating just outboard of the fence and the other near the wing

apex. Thi s 1 imits the growth of the outboard voritex and keeps the

leading-edge region near the tip under the influence of vortex suction

to higher a. The inboard vortex is, subsequently~ forced onto the wing

surface by the fence and acts as a barrier to the streamwise migration

of the outboard vortex.

The effect of this additional F-2 fence at n = 0.625 on the

leading-edge suction is depicted in figure 56. At a = 160, a slight

suction improvement outboard of the fence roughly balances the charac-

teristic loss of suction on the inboard side. However, at a 23 0, a

substantial increase in the outboard thrust increment accounts for the

net improvement noted in the CA curve.

The chordwise slot, on the other hand, is shown in figure 54 to be

ineffective in the role of SLEE compartmentation. The spanwise iocal

thrust distribution at a = 160, in figure 56, shows signs of compart

mentation through a slight thrust increment just outboard of the slot.

However, by 230 a, this improvement has vanished due to loss of slot

effectiveness .. This slot at n = 0.625 was also tested in combination

with a slot (as opposed to F-2) at n = 0.25 . on the 0.48 em ext.

SLEE configuration, but with the same results (not presented).

- - -' - -- Basiewing

.10,-~ 0.48 em ext. SlEE; wi F-2 @ 7]' 0.25 _.--0-.- 0.48 em ext. SLEE; wi F-2 @ 7]'0.25,0.625 _._--<>-_.- 0.48 em ext. SLEE; wI F-2 @ 7]' 0.25 and

SC -1 @ 7]. O. 625

.08

I ~/rM'o 06t ' t ' C

\"'L, Tloc

.04 a -160

"-"- ,

I .02

0 0 .2 .4 .6 .8 1.0

I I 7}

I .12

I 110

.10f--- I~~, I I

I .08

I L~, / "- IB -

a' 230 06[ /

, 0

\'

CT \

\

Icc , .04 "- ,

F-2

g~>' F-2 02r <> F-2 SC-1

0 0 .2 .4 .6 _8 1.0

7}

Figure 56.- Effects of the fence· and chordwise slot on leading-edge ·thrust distribution along the SLEE.

142

· 143·

In summary, the sharp leading-edge ~xtension is effective in drag

reduction and rolling moment stabilization at high a, with no adverse

effects on longitudinal stability. The drag reduction is insensitive

to SLEE extension at low and mid a; however, there is a distinct

advantage to having a shorter extension at high a, assuming the exten

sion is beyond the stagnation point. When utilized at the apex of the

SLEE, the chordwise slot is slightly more effective than the fence in

obstructing the spanwise flow near the wing apex at mid a. However,

far better performance at high a makes the fence more desirable for

use on highly maneuverable airciaft. Unlike the chordwise slot, the

fence also possesses the ability to compartment the SLEE when used along

its length, thereby retaining leading-edge suction to higher Ci,.

_" ~ • , •• ~~~. 110_

144

.. Lea:di ng- Edge Vortex Flaps

As an ext"erision of thE{ SLEE concept; 1 eading-edge vortex fl aps were

tested as a possible means'6f obiaining furtnerimprovements in leading-

edge thru~t at high a. This device is similar in appearance to a con

ventional leading-edge flap which is deflected to align it w5th the

local upwash in order to provide a smooth onflow at the wing leading

edge and, thus, av6id5eparation. The vortex flap, however, relies upon

a separation vortex forced~by underdeflection relative to the oncoming

flow to generate a suction force on the flap and, thus, provide a thrust

component. Ideally, this vortex induces reattachment just aft of the

leading-edge curvature, thereby allowing attached flow on the upper

surface. The thrust component generated on the vortex flap, thus, leads . to" alleviation of the lift-dependen't drag (see sketch A in •

INTRODUCTION) .

The flow mechanism of the vortex flap will now be discussed in

detail with respect to the 300 downward-deflected (+) VF-3. This was a

constant chord, flap (flap chord 7.3 percent of mean geometric chord)

beginning at n = 0.2~ and extending out to the wing tip (see fig. 7).

Elimination of the first ~~ percent of the flap near the wing apex

(n = 0 to 0.25) was based on its relative ineffectiveness due to low

prevailing upwash in that region (ref. 3). Figure 57 presents wing

leading-edge static pressure variations at selected spanwise positions

for a configuration utilizing a 300 + VF-3. As an aid in the discussion,

a qualitative summary of the various stages of leading-edge flow

development with iricreasing ~,at constant deflection angle, is also

C PLE

Orifice 117

_ 205

227 ©zpm

-2.0

-1. 5

-1. 0

.

~

~\ ."

:9 .. \\i

\ I

-u ,>I I

-0.5 /1D1 .... q? I c( \0 I

, /

~~-~ ~\ ''0; q

o J \ I

o 5 10 15

a, deg

p--

117 ("1= 0.33)

--0- _ P 227 ("1 = 0.82)

@

20 25

145

Figure 57.- Leading-edge static pressure variation and flow develop~ent on a 30° + VF-3 configuration. Stages of flow development are arbitrarily indicated on the orifice 117 curve.

146

included. Due to entrainment effects and the spanwise upwash distribu

tion ahead of the leading edge, these successive stages of flow develop

ment will also be found with increasing outboard distance for constant

flap deflection and angle of attack. At low a (stage A),the vortex

flap is ~ffectively overdeflected relative to the small upwash and,

thus, induce~ ~6rtex formation on its lower side. The resulting orien

tation of the flap normal force vector produces a drag increase and loss

of lift. Eventually, a smooth onflow at the flap edge (stage B) occurs,

as the flap angle matches the upwash. Note that to this stage, CpLE

has become more negative with increasing a due to the rapid develop

ment of potential flow suction around the leading edge. Therefore,

separation at the knee is likely due to the high degree of leading-edge

curvature. At stage C, the upwash angle nominally exceeds the flap

angle, resulting in vorte~ formation and reattachment on the flap upper

surface. Tuft photographs (fig. 58) indicate that the transition of

the flow pattern from over- to under~deflection occurs between a = 60

and 100 for the 300 + VF-3. At a = 60 , tufts near the flap leading

edge are hidden, indicating separation on the lower side. At a = 100,

howeVer, these tufts lie on the face of the flap, indicating favorable

separation. With further increases in a, or outboard distance,

reattachment moves toward the leading edge, eventually coinciding with

the leading edge at stage D. The relatively high pressure at the

reattachment position corresponds to the low point on the -CpLE dip

in figure 57. This may be regarded as the "design" point since the full

chord of the flap is under vortex suction. The locus of the minimum

1"-, o:::j-.--~

+-' rO

c: 0

-r-+-' rO $-:::s en

-r-4-c: 0 u

(Y)

I I.l.. > -7-

0 0 (V)

4-0

U1 .c: 0. CO >-en 0 +-'0 0 0

.c: rl 0-

-0 +-' c: 4- CO :::s +-'0

\.0 OJ

-0 II -r-V) (j

ro I.!)

QJ >-:::s 01

-r-I.l..

148

-CpLE points along the span may, thus, be used as a guide to design an

optimum twisted flap for a given average deflection angle (see next " ,

paragraph). As the flow reattachment position and, subsequently, the

flap vortex, move onto the wing upper surface (stage E), the flap thrust

effectiveness diminishes.

The locus,of CpLE derived ,deSign points along the 300 + VF-3 is

plotted in figure 59. As previously noted, the ideal flow condition is

first met near t,he wing tip and moves, inboard with increasing a. This

curve also provides a general idea of ,the relative flap twist necessary

to produce the ideal flow condition simultaneously along the entire

flap. However~ this twisted flap, apart from being impractical, may not

be desirable since loss ~f effectiveness would also occur at once along

the entire span.

Figure 59 also presents the locus of the initial relative maximum

-C PLE points (see fig. 57) along the 30° + VF-3. This curve may be

thought of as a boundary between over- and under-deflection of the flap

relative to the oncoming flow. Again, due to the spanwise upwash dis

trib~tion, vortex formation initially occurs near the tip and spreads

inboard with increasing a. This is also evident in the tuft photo

graphs in figure 60. Although these were taken of the 600 + VF-3~ the

same basic trends would be expected with a 300 + deflection but at

lower angles of attack. The chordwise orientation of the tufts on the

flap at a = 140 is indicative of attached flow. By 160 a, however,

the flap vortex has formed and advanced to approximately the midpoint

of the flap (position 'b'). Due to expansion of the vortex core

0,

deg

20

18 .

16

14

12

10

8

6

o

-Cp LE

r

a

----(0)----'-----iDI---

149

Low point on -CPLE dip Initial -CPLE maximum

J,9 Attachment / - \'

on flap t--'

Attachment on wing

0.2 0.4 0.6

7]

Q. . P ___ Smooth -'Dr::;r:Y .. ~ on flow

.. / ~

0.8 1.0

Figure 59.- Local design point distribution (Cp derived) along 30° + VF-3 (see fig. 57). LE

Figure 60.- Side tuft photographs of 60° + VF-3 configuration at a = 14°, 160, 180 , and 23°.

150

151

toward the tip, reattachment between positions 'b' and 'c l occurs on the

flap face (indicated by the diverging tufts on the flap), with nearly

. an ideal flow condition aft of position 'cr. Further increases in a

are accompanied by continued inboard movement of the vortex, v/ith its

apex lying near the flap apex at a= 230. At this stage, reattachment

aft of position 'c' occurs on the wing upper surface, as evidenced by

the realignment of the tufts along the length of the flap.

Leading-edge thrust variations for the 300 ~ VF-3 configuration are.

presented in figure 61. No effect is evident on the leading-edge thrust

at low a since separation occurs on the lower side of the flap. At

mid a, however, the flap induces a thrust loss outboard of STA 1, cor

responding to the leading-edge suction loss noted in figure 57. However,

as the ideal flow condition is reached near the tip, the flap effective

ness begins to appear as a substantial thrust improvement. At angles of

attack higher than those considered here, further inboard movement of

the ideal flow condition would result in thrust improvements also at

the inboard stations.

The effects of th~ 300 ~ VF-3 on the overall performance character

istics of the configuration are given in figure 62. A large axial force

improvement over the basic wing is found at high a and is sustained to

the. highest angles tested. The severe low-a drag penalty is not of con

cern since at cruise the flap would be undeflected or retracted; as is

the loss of normal force at low a, which is attributed to the effective

negative camber induced by the downward flap deflection. The migration

of the primary vortex onto the wing upper surface leads to onset of

152

STA B' . . ~-----aSlC wlng 1 -0-- 30° ... VF 3 2 -

.10 r- ~ _3 .10 4

5 / - . . n - 0 33 .08 f--\,\ 6' .08· - •..

. 02 • . .02

o - 1 0 o 5 1 0 1 5 20 25 0 5 1 0 1 5 20 25

ex, deg IX, deg

.10

t .10

.08 n = 0.45 .o8L n = 0.57

cT .06·,' CT .06

'0'' :04 [ ,/ V-- - '00< ' .. o.;r /,/\ .02 .02

o ,'1 ·1 0 o .5 10 15 20 25 0 5 10 15 20 25

ex, deg IX, deg

.10r- .10

1 ,08/:::- n= 0.70. .08f- n = 0.82

CT

[OC5 .06t /"'. ~ c"o,' .06L .04 / Y .04

.02~r - .02

o 1 ,'1 1 1 0 o 5 1 0 1 5 20 25 0 5 1 0 15 20 25

ex, deg ex, deg

Fi~ure 61.- Effects of 300 + VF~3 on wing leading-edge thrust.

C A

CN

.08

o

-.04

"'-.08

-.12

o

1.0

.8

. 6

.4

.2

- - - - - - Basic wing.

~ 300 .VF-3 ----0---- 45°. VF-3

-----<>---- 60° t VF-3

I . I 5 10 15

lX, deg

20

I~ i)/~ / '

/¥ " .

25

C );;r4/ ~ 25

lX, deg

153

Figure 62.- Effects of vortex flap deflection angle on force and moment characteristics,

'.

154

10 I

/ Basic wing /

--- -I

8f- I ----0- 30° t VF-3

I J

---0--- 45° t VF-3 J _c_-(>-___ (I:P t VF-3

I 6

LID

4~ <;> "-

"-"-

"-"-

---I 2

0 0 _2 A _6 _8 LO

CL

-006r ~ 1\ 1\ r _005 f- I \ I \ f

I \ / J \

.004 r- \ \

\ \

I _003

Cz

_002

_001

5 10 15 20 25

0:, deg


155

vortex lift at approximately a = 140 , accounting-for the convergence of

the CN curve toward the basic wing data at hi~her a. The resulting

effect on LID is a loss of performance at low a but substantial

improvement beyond CL ~ 0.3. The C1 data indicate that the vortex

flap also eliminates the severe mid-a rolling moment instability of the

basic wing.

Longitudinal stability effects of the 300 -t VF-3 are deP5cted in

figure 63. The moment reference cent~~ has once again been moved to the .,". ~.

same position used with the SLEE data (3.S2 em aft of the original ref

erencecenter indicated in fig. 2) to better show up the effects of the

device. The 300 -t VF-3 produces: a slight reduttion in longitudinal

stability at mid and ~igh ~ but eliminates the nonlinearities -of the

basic wing at a ~ So. The chordwise location of the center of preSSll;re

for thi~ same configuration (derived from Cm and eN data) is plotted

as a fUnction of a in figure 63. The forward-movement of xcp rela

tive to the basic wing beyond a ~ 60 (the angle at which vortex forma

tion is initiated on the flap face) is attributed to a forward movement

of the center of pressure along the flap. Whereas attached flow on the

basic wing positions the suction peak near the leading edge, the vortex

suction effect on the flap moves this suction peak forward onto the flap

surface (sketch M). Since pitch-up is the result of a forward movement

of the center of pressure, those flap configurations producing the most

aft center of pressure will be considered the most desirable as far as

longitudinal stability is concerned.

em

x ~

c r

0.04

o

-0.04

-0.08

-0.12 .

I o 5

x ~ = 0.474

cr

10

-- - - - - Basicwing

~--{Ol--- 30° t VF-3

° ---0--- 45 t VF-3 ----<>---- (fJ0 t VF-3

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

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

L J ___ I 15 20 25

a, deg

a, deg

o 5 10 15 20 25 o I I . . I

0.1

0.2

0.3

0.4

0.5

Figure 63.- Effects of vortex flap deflection angle on longitudinal stabil ity character; s ti cs.

156

-c p

Basic wing

Sketch M

157

wi Vortex flap

Comparison of experimental results with theory (ref. 15) for the

300 + VF-3 configuration appears in figure 64. The downward shift in

the experimental lift data as compared VJith theory is attributed to the

effective neg.ative camber induced by the asymmetric trailing-edge region

(see fig. 2). Otherwise, the experimental lift and drag data show

relatively good agreement with theory throughout the angle-of-attack

range.

For a more basic assessment of the aerodynamic effectivene~s of the

various flap configurations tested, an analysis on a per unit flap area

basis will also be used. This will provide an indication of the effi

ciency with which the surface area of a particular flap configuration

is being utilized for drag reduction. Figure 65 presents balance and

leading-edge pressure-integrated thrust data for the 300 + VF-3 configu

ration. The axial force curve has been shifted downward to zero to

CL

Co

1.0

0.8

0.6

0.4

0~2

10

a, deg

0;3 r-

0.2

0.1

/

/ /

/

/

-'-0------ Expe ri me nt

- - - - - VLM-SA theory

15 20

/,.

'l' I'

25

o I -,~ .. 1.-1 II o 5 10 15 20 25

a, deg

Figure 64.- Comparison of 300 + VF-3 experimental data with VLM-SA theory. . .

158

-c Ttot

and

flCA

Cx,

1m2

a, deg

o 5 10 o ITP;;: I

...... 'D... t... t..v

...... ,Q

'Q Wing leading-f-" 'Du" ,dgo th,," -0.01 . ~~ -0.04

-0.06

-0.08

-0.10

-0.12

2.0 I

1.6

1.2

0.8

0.4

_ (-CTtot) - (flCA) Cx - A

VFtot

"-h.

...... -- -0 -Crtot Vortex flap

contribution

f Total thrust

Onset of vortex lift

,6CA = CA - CAo

Oe1------L-----~-------L------~----~ o 5 10 15 20 25

a, deg

Figure 65.- Thrust contribution from 30° + VF-3 flap surface.

159

160

eliminate the profile drag component .. CTtot represents the contribu

tion of the wing leading-edge suction to the axial force obtained by a

spanwise summation of CTloc

The difference in the two curves, there-

fore, represents directly the thrust contribution of the flap. This

difference is divided by the total flap surface area and plotted in fig-

ure 65. The curve indicates an improvement in flap efficiency with

increasing a. The sudden reduction in slope at a ~ 160 indicates

the onset of loss of flap thrust effectiveness due to vortex migration

onto the wing upper surface. This is consistent with the onset of vor-

tex lift noted in the CN curve (fig. 62) at approximately the same a.

Additional evidence is provided by the photographs in figure 66. At

a = 14°, the mini~tufts near the wing leading edge are generally chord-

wise, implying reattachment either on the flap or at the knee, with

attached flow prevailing on the wing upper surface. At a = 170,

however, these same tufts are angled toward the wing tip, suggesting

vortex action and reattachment on the upper surface.

The effects of flap deflection on overall performance are shown in

figures 62 and 63. Axial force shows an improvement at high a with an

increase in downward deflection of VF-3 from 30° to 45°, but a loss of

performance with further deflection to 60°. The effective increase in

negati ve ca·mber with increasing fl ap defl ecti on results in a reducti on

of normal force at low a. The same trends are evident at higher a

due to a delay in vortex migration onto the wing upper surface. Lift-

to-drag appears relatively insensitive to flap deflection beyond

CL ~ 0.4. This is attributed to the self-compensating for thrust effect

Figure 66.- Upper surface tuft photographs of 30° + VF-3 configuration at a = 14° and 17°. I-' m I-'

163

A major concern of the aircraft design engineer is that of produc

ing adequate lift to reduce the landing speed. In addition, decelera

tion of the aircraft from cruise to landing speed has been a problem

requiring the use of speed brakes and high approach angles of attack.

An up-deflected (t) vortex flap was tested in the present investigation

as a possible aid in alleviation of these problems (ref. 17). It was

believed that the suction effect of a vortex formed on the lee side of

the flap (see sketch F in BACKGROUND) would produce both lift and drag

increments ideal for the landing approach. The vortex flow induced on

the wing upper surface would provide additional lift.

Results of balance measurements on a 300 t VF-7 configuration (see

fig. 7) are presented in figure 67. Normal force indicates onset of

vortex lift almost simultaneously with departure from a = 00. The

strong spanwise inclination of the tufts near the wing leading edge at

a = 60, in figure 68, is indicative of this upper surface vortex flow.

Note the well-defined reattachment line with chordwise flow downstream.

The resulting improvement in CN implies that the required lift at

landing may be obtained at a lower flight speed. The rolling distance

after touch-down, which is proportional to the square of the landing

speed, can thereby be significantly reduced. At a 00, the 300 t VF-7

produces a relatively small increase in axial force since there is no

addition of frontal area (see fig. 7). However, at mid and high a,

large increases in drag are available for deceleration of the aircraft

during the landing approach. In addition, t1e linearity of the pitching

moment curve is significantly improved over that of the basic wing. The

.08

.04

CA

-.04

-.08

1.2

1.0

.8

.6

CN

.4

.2

o 5

"

/ /

/

10

- - - - Basic wing

---0- 30° t VF-7

15 20 25

lX, deg

I I

I I

I

I

/ /

I

I I

I /

/

/

I /

I I

I

I I

/

lX, deg

164

Figure 67.- Effects of upward vortex flap deflection on force and . moment characteristics.

em

C2

.04

o

-.04

-.08

-.12

o

.006 r--

,005

.004

.003

.001

'--,

Xcrn = 0.474 C

r

5 10

, '-

ct, deg

"" ,I

h f\ I' I \ I \ I \

\

\

\ I


--0---- 30° t VF-7

, , ,

15

, ,

I 20

, ,

25

0, ~ o ,~

cx. deg


.. '.

165

" -J.

to c ..., (0

Q)

00

OJ c: rl--O

U Q n)

-') II

Ul O)C 0-,

-n Q.I (') (0

rl'" C -n rl'"

-0 ::y 0 rl'" 0 to ..., Q.I

"0 ::y

0 -n v..) C)

0

...,..

< -n !

'-.J

(') 0 :::::; -I)

to C -'l OJ rl'" -J. ~

0 Q)

::l Q)

~

167

rolling mOment characteristics are excellent up to approximately

ex. = 16°, at which point a minor instability appears. However, the

severe wing rock of the basic wing between 80 and 14° ex. is el-lminated.

These characteristics are significant since any unstable behavior at

near-stall landing conditions may be critical.

An inverse tapered vortex flap was also tested in the present

investigation to take advantage of certain flow phenomena characteristic

of the device. It was reasoned that increasing the flap chord toward

the wing tip would better accommodate the expanding vortex core and,

accordingly, produce a more efficient drag reducer. This inverse

tapered flap (VF-2) was designed with approximately the same surface

area as the constant chord VF-3 (see fig. 7), allowing for direct per

formance comparison. Balance data for the 300 + VF-2 and VF-3 configu

rations are presented in figure 69. The axial force data indicate a

slight improvement in drag-reduction effectiveness beyond ex. ~ 16° with

the inverse tapered VF-2. However, there is a slight loss of normal

force within this same ex. range, The longitudinal stability character

istics (hbt pr~sented) are not significantly different from those of the

constant chord flap.

Static pressure distributions around the wing leading edge

(fig. 70) confirm that drag performance improvements with the inverse

tapered VF-2 are indeed attributable to its ability to better maintain

the vortex on the flap and delay its migration onto the upper surface

in the outboard region. Stagnated pressures at n = 0.70 and 0.82 wIth

the constant chord VF-3 at ex. = 120 suggest that reattachment has moved

.08

.0

o CA

-.04

-.08

-.12

o 5

'T .8

.6

CN

.4

.2

10

Basic wing

~ 300~VF-3 ---D--- 30°. VF-2 -----0---- 30°. VF-l

---

15 20 25

a.. deg

168

Figure 69.- Effects of inverse tapered flap chord distribution and apex shape on vortex flap performance.

STA 1

2 -4

-3

3 4

5 , 6

-- -- -- - - Basic wing -0-- 30° {- VF-3 ----0---- 30° {- VF-2

-4-

-3

(Constant chord) (Inverse tapered)

n = 0.33

Cp1 = 0.20 C

p2

-1 -1

1~~ 2 3 4 5 0 1 2 3 4 5

X, em X, em

-4 r- -'[ -3!- -3

\ n = 0.45

-2 "'" n = 0.57

-2 l \ l \ , l

CP3

, , CP4

\

\

-1 -1

---------- -------

1 I 1 0 2 3 4 :5 0 2 3 4 5

X, em X, em

-4

-3 ,

1 = 0.70 \

~. -2 \

Cp " \'------ CP6 5, _

1 1 I I 11 1 1 I.~ o 2 3 4 5 02345

X, em X, em

169

Figure 70.- Effects of inverse tapered vortex flap chord distribution . on static pressure distributions around the leading edge

at a = 12°.

170

to the wing surface. The inverse tapered VF-2, however, has retained

attached flow at the leading edg~, with reattachment still on the flap.

Tuft photographs taken at a = 160 (fig. 71) indicate the same trends.

The chordwise orientation of the tufts near the wing leading edge with

the VF-2 indicates attached upper surface flow, whereas the spanwise

inclined tufts with the VF-3 configuration are the result of vortex

spillover. The delay in onset of vortex lift with the VF-2 accounts

for the reduced normal force noted in figQre 69.

A comparison of the CpLE derived locus of design points along the

.300 + VF-2 and VF-3 appears in figure 72. The relatively smaller flap

chord near the apex of the VF-2, as compared with the VF-3, reduces

the a at which the design point is met. The opposite effect is evi

dent outboard of n ~ 0.60. Delays in vortex spillover of up to 30 a

at these outboard stations with the VF-2 account for the high-a improve

'ments noted in the CA characteristics.

Comparison on a pe~ unit flap area basis, in figure 73, indicates

that the inverse tapered flap area distribution is somewhat more effi-

cient than the constant chord ~ariation up to o .

a ~ 16 . -More pronounced

improvements are indicated at higher a, as the VF-3 begins to lose out

board suction effectiveness. An implication of this result is that

savings in system ~'ieight are possible through appropriate geometric

design of the vortex flap.

An additional inverse tapered vortex flap (VF-l), characterized by

a chordwise-cut apex, as opposed to the sweptback apex of the VF-2, was

tested to determine the importance of flap apex shape. It was thought

Figure 71.- Upper surface tuft photographs of 30° + VF~2 and VF-3 configurations at a = 16°. I-' "'-J ......

20

i

18

.-

14

ades ,

deg 12

10

8

6

(} " 0.2

-.~ - - -:--

----IO~-

.~ ,.

\ \ \

300 t Vf-3

300 t VF-2

172

I nverse tapered

\. .

" / Constant ~hord " / - ./

TJ at which the chords of (VF-2 and VF~3 are iden.tical

0.4 0.6 0.8

'7]

l.0

Figure 72.- Effects of inverse tapered flap chord distribution on local design point distribution (C derived) along th~vbrtex flap~ _. PLE

; . .;.

r \...IX,

1m2

2.0

1.6

(-C l C = Ttot - (!',CAl x

AVFtot

1.2

0.8

0.4

/'

./' --./

-- ---- - - 30° t VF- 3

--0- 30° t VF-2

-----G---- 30° t VF-l

./

/ /

./

173

o ~ry/~/ ____________ ~ ______________ L-____________ ~L-____________ ~ ______________ ~

5 10 15 20 25

a, deg

Figur~ 73.- Effects aT lnverse tapered vortex flap chord distribution and apex shape on flap thrust parameter.

174

that a counter-rotati~g, vortex formed at this chordwise edge would

interfere with the formation of the primary fl ap vortex. Balance data

in figure 59, however, indicate th~t this particular apex modification

has insignificant effects on overall performance. The close agreement

of the flap thrust coefficient data in figure 73 provides added support.

As evidenced by the vortex flap data presented thus far, streamwise

migration of the flap vo~tex with increasing outboard distance is the

primary 1 imi tati on to effici ent fl ap performance at hi gh Ci.. Segmenta

tion of the flap was suggested ~s one method of alleviating this problem

through the formation of an independent vortex on the outboard segment,

thereby delaying thrust loss near the tip. The geometries of the seg-

mented flaps tested appear in figure 7. VF-4 was derived from VF-3 ,

simply by segmenting the flap at its midpoint and, thus, afforded the

opportunity to isolate the performance effects attributed solely to seg-

mentation. VF-5 and VF-6were included to investigate the possibility

of further performance improvements through variations in flap chord

distribution.

Result~ of balance mea~urements on configurations utilizing the

segmented vortex flaps at 300 downward deflections are presented in fig

ure 74. Comparison of VF-3and VF-4 axial force data indicates little

overall effect from segmentation. However, slight reductions in normal

force and longitudinal stabillty result at high Ci.. Comparison with

VF-5 and VF-6 data suggests that a substantial amount of flap area may

be eliminated with little sacrifice of CA performance up to Ci. ~ lSo.

Beyond Ci. ~ 180, theinvers~ tapered V~-5, with 42 percent less surface

CA

CN

.08..-

o

-.04

-.08

-.12

o

1.0

1 .8

.6

.4

.2

0 1,,/'

0/:;

175

Basic wing

--0-:- 30° t VF-3

----0--- 30° t VF-4 ------'0---- 30° t VF-5 - ... --6--... - 30° t VF-6.

5 10 15 20 25

0:, deg

0:, deg

Figure 74.- Effects of segmentation and flap geometry on vortex flap performance (8 = 30°).

o

-.04 em

-.08

-.12

o

--'·~"O.-"

- - - - - --'- Basic wing

--'----(0)-----.,. 30° t VF- 3

° ----0---. 30 t VF-4

----<)---- 30° t VF-5 " ---- -6------ 30v t VF-6

" ~.~ .. ::~:-\ .. ~'" '--- --........ ....... ....

" ", '- .. '-

'-'-

5

Xcm = 0.474 ~r

10 15

<1, deg


..... '-

'- .......

20

176

25

177

area than the constant chord VF-4, loses effectiveness, while the para

bolic VF-5(13 percent reduction in area) continues to perform on par

with the VF-4. Reduction in flap area actually results in an improve

ment in longitudinal stability beyond a '" So, v.Jith little effect on

normal force. Balance data for 450 downward deflections of these seg

mented flaps (fig. 75) indicates the same trends. However, the 300

deflections will be used for description of the underlying flow

mechani sms.

The performance of the various segmented vortex flaps on a per unit

flap area basis is summarized in figure 76. As previously noted, the

flap thrust coefficient, Cx, provides an indication of CA effects

attributed solely to vortex suction on the flap surface. Segmentation

of the full length VF-3 into VF-4 results in substantial improvement in

flap area efficiency beyond. a:'" 160, as loss of flap suction has

apparently been delayed. The parabolic (VF-5) and inverse tapered

(VF-6) segmented flaps provide further improvements at lower a, but

with earlier loss of effectiveness with decreasing flap area. Each

particular segmented flap geometry seems to be the most efficient of

those considered within a specific a range, implying that the actual

geometry selected would depend on the design angle of attack.

Tuft photographs in figure 77 confirm that Cx improvements

resulting from segmentation are indeed attributed to retention of vortex

suction on the outboard flap segment. The mini-tuft pattern and static

pressure distribution at STA 6 with the full length VF-3 suggest flow

reattachment on the upper surface. However, the aft portion of the

.08

.0

o CA

-.04

-.08

-.12

.04

o

em -.04

-.08

-.12

u Basic wing

-0---:--- 45° t VF-3

° ----fr---- 45 t VF-4 -----<)---- 45° t VF-5 - ... ---L:r--.-.- 45° t VF-6

-~ - --

.--~

~~

'-------"----'---------'------'----,. o

o

5 10

5

Xcm = 0.474 Cr .

10

15

a, deg

15

a, deg

20 25

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

"

20 25

178

Fi gure 75. - Effects of segm.enta~i on and fl ap geometry on vortex fl ap performance (8 = 45 ).

ex, 1m2

179

2.0

1.6

1.2

0.8

0.4

:--_-0

V' / /<"'0-- / /

d~:/X/ /~~ ~/ '

" //--- . '",~

:;:>p' / '->

C = (-CTtot1 - (.6.CAl x

AVFtot

1d #

/:! !~ <1/7

,'10(//

A,)-l---Y

~

/;~ /~/

- - - - - - 30° t VF-3

--0-- 30° t VF-4

---0--.- 30° t VH

----<)---- 30° t VF-6

O~~S/~ ____ ~---------~------------~---------~------_-J 5 10 15 20

a, deg

Figure 76.- Effects of segmentation and vortex flap geometry on flap thrust parameter.

25

~:L -2~

I

-i~--~------~

ob--1 I ~----I----I----I- ) o 3 4

180

-4F -3 ------ Basicwing

-2k

Figure 77.- Side tuft photographs of 300 + VF-3, VF-4, VF-5, and VF-6 configurations at a = 14°.

-4, -:It -2

.~ i~

-4,---I I

o , 2 :I

-3~

~~~I, __ ~ __ - ,, __________ __

(-" ,,:::. 0

1 '-----

,1 I -;----1-- --;-023 ~


181

_4

f -3 - - - - - - Basic wing

-2 _ ~----------------,

, -, -~ ,1 1--'-1----; ---I

o

182

segmented VF-4 appears to still be fully under the influence of vortex

suction (as shown by the mini-tufts on the flap slanting away from the

wing leading edge) with reattachment in the vicinity of the knee

(attached flow indicated on the wing leading edge). The flow through

the break has apparently forced the vortex on the forward flap segment

to peel off and, most likely, travel over the upper wing surface. The

parabolic VF-5 and inverse tapered VF-6 show the same baSic trends but

with accelerated forward movement of the outboard vorte~. Compar~son

with static pressure distributions at STA 3, where segmentation would

appear to have little influence, suggests that this is purely a geo-

-metric effect. As a result, the tuft photographs in figure 78 indicate

a reduction in upper surface vortex flow in the outboard region upon

segmentation of VF-3 (to VF-4), but with the opposite effect with cut-

back from VF-4 to VF-5 and, subsequently, VF-6. As noted in prev; ous

sections, an increase in upper surface vortex flow tends to increase the

overall lift of the wing but also the severity cif the lift-dependent

drag.

Figure 79 presents -the CPLE ~erived locus of- design points alnng

each of the 300 down-deflected segmented flaps. Although -the data

points are scattered, certain trends are consistent with those observed

in the tuft photographs. Sinc~ effects inboard of n 0.625 are-

pu~ely geometric, the VF-3 and VF-4 data points approximately coincide. - . Reduction in flap chord with VF~5 and VF-6 results in satisfaction of

the design condition at progressively lower a alorig'this forward

segment. The convergence of the data in the vicinity of the break

183

Fi gure 78. - Upper surface tuf'tphotographs of 30° -} VF -3. VP.;;'4'; VF-5. .. and VF-6 configurations at a = 18°.

Odes,

deg

- - - - - 30° t VF-3 --~o 30° t VF-4

---0--- 30° t VF-5

20;....,......- ----<>--.-. 30° t VF-6

18

16

14

12

10

8

Segment

'- / ....... _/

6 ., I_ I I . . . . o 0.2 0.4 0.6 0.8

'TJ

184

1.0

Figure 79.- Effects of segmentation and flap geometry on local design point distribution (C derived) along the vortex flap. . ~E

185

(n = 0.625) i~ attributed to the equality of the respective flap chords

(see fig. 7). The effect of segmentation (VF-3to VF-4) is evident

along the outboard segment in the form of a delay of several degrees in

vortex migration onto the upper surface. The effects of chord reduction

noted on the inboard flap segment are also evident here.

The utilization of vortex flaps for roll augmentation was briefly

investigated using the inverse tapered segmented VF-6. It was antici

pated that asymmetric flap deflections would alter the spanwise lift

distribution so as to produce adequate rolling moments at high lift,

when conventional control surfaces, such as ailerons, are degraded by

flow separation. Rolling moment data for a configuration utilizing a

300 i VF-6 on the left-hand and 450 + VF-6 on the right-hand wing panel

are presented in figure 80. The positive shift in the ~t curve cis com

pared with the basic wing is purely a planform effect. The imbalance of

planform area addition produces an additional lift increment on the

left, with a resulting right-wing-down rolling moment. Beyond ex '" 160 ,

the flap deflection effect becomes evident. The strong positive rolling

moment initiated at this ex is due to the earli~r migration Qf the

300 i flap vortex onto the left wing panel as compared with the vortex

emanating from the 450 i VF-6 on the right. The"additional vortex lift

on the left produces a maximum positive rolling moment of 0.0041 at

ex '" 180. The subsequent migration of the 450 i VF-6 vortex onto the

right wing panel, with its lift acting at an n value greater than that

of the left-hand vortex, produces a reversal in the rolling moment at

higher ex. The fact that strong positive and negative rolling moments

C'l

- - :- - ~ . ..,- Basicwing'

--'-'-_. """""""0 30° t VF-6~n left; 45° ~ VF-6 on right

.006 ~ '\ 1\

.0051--- I \

.003

.002

"'--- .... , f '" ,

.001

o

-.001

o 5

I . I . I

I

I I

I

I

10

\ \ \ \

15 20 25

lx, deg

186

Figure 80.- Effect of differential vortex flap deflection on rolling moment characteristics.

r,'.

..

187

are separated by only 40 or 50 a may produce undesirable handling

qualities at high lift. In addition, the magnitudes of the rolling

moments produced by the configuration tested here are inadequate to be

: of practical value on an actual aircraft. However, a greater differen

tial in flap deflections may provide the required rolling moments.

Based on the previous performance of fences and chordwise slots,

individually and in combination with the SlEE, their potential for

improving the performance of the vortex flaps was also investigated.

The 300 + VF-3 was tested with a single slot (w/SC-l) and then with a

single F-3 fence at its apex (n = 0.25), again in an attempt to aid in

the formation of an undisturbed flap vortex by blocking the leading-edge

cross-flow near the wing apex. An additional test run was made with a

single chordwise slot cut into the flap itself (SI) ~t n ~ 0.625, with

the slot in the wing leading edge remaining sealed (see fig. 7). As

with segmentation, the flow through the slot was expected to aerody

namically compartment the flap and, thus, induce the formation of an

independent primary vortex on its outboard side. Ideally, the entire

. flap surface would thereby remain under the influence of vortex suction

to higher u.

Axial and normal force data for the vortex flap-slot/fence con

figurations just described are presented in figure 81. Both graphs

indicate the overall ineffectiveness of each concept .. Effects on pitch

ing and rolling moment, not presented, are also negligible. Static

pressure data for the configurations with either a slot or fence at

n = D.25 show no evidence of a loc~l effect from the additional device.

.08 - - - - - - Basic wing

---0-- 3fP t VF-3 . . ° ---D---o~-. 30 t VF-3; wI SC-l@7]=0.25 ----0---- 30° t VF-3; wI F-l @ 7]= 0.25 -_._-£r_._- 300 t VF-3;w/S' @ 7]=0.625

o CA ----

-.04

-.08

-.12

o 5 10 15 20 25

ex. deg

'I I .6

CN

A

_2

a,

Figure 81.- Performance effects of the fence and chordwise slot in combination with the vortex flap.

188

"

--

189

The fact that these devices did not extend out ahead of the wing leading

edge, where the formation of the flap vortex occurs, may be responsible.

In that case, the chordwise slot cut into the flap at n = 0.625 would

be expected to perform effectively, assuming the flow through the slot

is sufficient for compartmentation. Indeed, the tuft photographs in

figure 82 show evidence of a second flap vortex outboard of the slot.

As depicted by the static pressure distributions in the same figure, the

wing leading-edge flow at STA 5 (n = 0.70) in the case of the VF-3 with

out a slot has already reached a stagnated condition, implying reattach

ment on the upper wing surface. The additional vortex formed as a

result of the slot at n = 0.625, however, retains attached flow at the.

wing leading edge, with the ideal flow conditi~n to be met at higher a,

as indicated in figure 83. Apparently, however, the sum of the suction

effects from the two smaller flap vortices (with a slot at n = 0.625)

is approximately equal to the suction produced by 'the single expanded

. vortex (without a slot), since no CA effect is evident. Additional

support is provided in figure 84, which shows that on a per unit flap

area basis, the 5uCtioneffectiveness of the flap surfaces· are

equivalent.

Finally, since the SLEE and vortex flap make use of basically the

same flow mechanism, a brief comparison of their suction effectiveness

is appropriate. Comparison of axial force data for the 450 + VF-3 and

o cm ext. SLEE (each found as the most effective in its respective

class), in figure 85, shows a significant advantage beyond a ~ 100

with the vortex flap. However, on a 6CA (relative to the basic wing)

190

-3f- ------ Basic wing

-2

-......\-.

o r '------- ___ h ___ h_

1 I I i I o 2 3 4

-4

f --

-3

-2

-- - -- ----- ---- -- ---

-'~ , ,

o~ 1 o 3

X, em

Figure 82.- Side tuft photographs of 30° + VF-3 configuration with and without a slot in the flap at n = 0.625 (a = 140 ).

.,

a des,

deg

'_~,7"'~~ ,

20

18

16

141 12~

10

8

6

191

- - - - - 30° t VF-3

----0-- 30°f VF-3 wI slot in

\A~ ~ I b-o

\ i

"-"- /

....... ./

Slot in flap

, flap '@ 7]= 0.625 (see fig. 7)

I. L------L_-.l.----L--~---'

a 0.2 0.4 0.6 0.8 1.0

7]

Figure 83.- Effect of a chordwise slot in the vortex flap on spanw·ise local design point distribution (C derived).

PLE

ex, 1m2

2.0

1.6

1.2

0.8

0.4

~ #'

~

~ ~

(-CT ) - (t.CAJ C = tot

x AVFtot

/,

~ ~

/,

V I On . o 5 10

'; J

I

Q., deg

Y

/ "7

- - - - - 30° ~ VF-3

--0-- 30° t VF-3 w/slot in

15

flap @ 7]= 0.625 (see fig. 7)

20

Figure 84.- Effect of a chordwise slot in the vortex flap on flap thrust parameter.

192

25

C>

CA

Sref (,0..C A) Afront

.08

-.04

-.08

-.12

.4

-.4

-.8

-1. 2

o 5

o 5

10

10

193

- --- -- Basicwing

---0- 0.00 em ext. SLEE; wI F-2 @ 7"J = 0.25 o

---~-- 45 t VF-3

15

ex, deg

15

20 25

fleA = CA - CA BW

Sref = Basic wing pianforrr area

Afront = Total frontal area (excluding housing)

20 25

ex, deg

Figure 85.- Comparison of sharp leading-edge extension and vort2x flap performance.

194

and frontal are~ basis., the two. devices perform identically. Therefore,

although the high-a drag performanc~ of th~ vortex flap is significantly

better than that of the SLEE, the additional weight of the vortex flap : I.

and the fact that the. SLEE may be rapidly deployed must be considered.

In summary, the vortex flap.prov~des substantial leading-edge

thrust improvements, resulting i\l 5izable axial force and lift-to-drag

increments at high a and elimination of the severe mid-a rolling

moment instability of the basic wing. In order to eliminate the charac-

teristic low-a drag penalty, however, the vortex flap would have to be

retracted or undeflected at cruise. Forty-five degrees appears to be

the optimum flap deflection angle based strictly on drag-reduction per

formance; however, lift-to-drag is relatively insensitive to flap

deflection in the range CL = 0.4 to 0.9. An upward deflected vortex

flap, by producing aerodynamic drag and vortex lift increments, is also

effective as a decel~ration and high-lift device for approach and land-

ing. A large downward deflection of the vortex flap upon touch-down can

provide continued. aerodynamic braking, with the added advantage of

increased downpressure on the wheels. High-a performance may be

improved through the use of an inverse tapered, as opposed to a constant,

flap chord distribution, which is better able to accommodate the expand-

ing vortex core and, thus, delay vortex spillover in the outboard

region. A two-segment constant chord flap improves the flap thrust

efficiency at high a through the formation of a separate vortex on

the outboard segment, which delays the outboard loss of leading-edge

suction. Elimination of the relatively ineffective portions of the

195

constant chord flap, leading to the parabolic (13 percent area ~eduction)

and inverse tapered (42 percent reduction) segmented flap configurations,

results in little or no loss of drag-reduction performance except at the

highest test a. Therefore, on a per unit area basis. these segmented

flap geometries are more efficient than the constant chord flap. Dif

ferential deployment of the vortex flap appears to also have roll aug

mentation potential at high lift. The use of a chordwise slot cut into

a full length flap at its midpoint generates an aerodynamic segmentation

effect by splitting the primary vortex into two smaller vortices but

with no significant influence on the total leading-edge thrust.

196

CONCLUSIONS

This report has presented the results of a wind-tunnel investiga-

tion undertaken to examine the potential for further drag reduction

through refined versions of leading-edge devices such as chordwise

slots, fences, pylon vortex generators, sharp leading-edge extensions,

and leading-edge vortex flaps. The results were based on low-speed

balance and static pressure measurements taken on a 60-deg cropped delta

wing model in the NASA-Langley Research Center 7- by lO-foot high-speed

tunnel. The intended use of the devices tested would be to reduce the

severe lift-dependent drag penalties associated with highly swept wings

at high lift. This section highlights what are believed to be the most

significant findings of the investigation.

The chordwise slot, fence, and vortex generator devices produce

substantial high-a drag and longitudinal stability improvements when

utilized on highly swept leading edges. The use of these devices in

mUltiple arrangements further enhances the leiding-edge suction along

the span by alleviating the adverse inboard effect of each particular

device. However, each device is characterized by a low-a drag penalty.

In the case of the chordwise slot, the main contribution to this drag

increase is from friction acting on the internal side surfaces, with the

pressure drag acting on the vertical face at the end of the slot of

secondary importance. Sealing the slot at cruise vJOuld be one method of

avoiding this drag penalty. Likewise, high-a performance is improved

with increasing slot depth, but with no effect from the contour of the

back face.

c· I

197

,

J In the case of the fence, the characteristic high-a loss.of

leading-edge thrust on the inboard side of the device is attributed to

its unsweeping of the upper surface isobars, rather than to any viscous

accumulation on the inboard side.

Vortex generator size reduction in the form of leading-edge length

and diagonal cutbacks reduce the low-a drag of the device but also

adversely affect high-a performance. Reduction in VG chord to a certain

extent results in performance improvements throughout the a range;

however, further reduction actually produces an increase in low-a drag.

The utilization of variable VG toe-in along the span to match the pre-

vailing sidewash ahead of the leading edge may be one method of simul-

taneously reducing the low-a drag and improving high-a performance. The

lower VG vortex apparently plays an important role through its own con-

tribution to the leading-edge thrust and, thus, should also be con-

sidered in the design of the VG shape. An extended chord VG provides

the added advantage of possible use also as a carrier of slender exter-

nal stores, providing substantial drag and longitudinal stability

improvements both with and without the external- store.

The vortex suction effect of the sharp leading-edge extension and

leading-edge vortex flap devices produces substantial lift-to-drag

increments at high a, in addition to improving the linearity of the

pitching and rolling moment curves. The drag-reduction effectiveness

of the SLEE is insensitive to extension at low and mid a; however,

there is a distinct advantage to having a shorter extention at high a.

as long as the device extends at least beyond the stagnation point.

198

The performance of the SLEE is enhanced by the utilization of either a

fence or chordwise slot at its apex to obstruct the spanwise flow

originating near the wing apex and allow for the formation of an undis-

turbed primary vortex. The fence also possesses the ability to compart-

ment the SLEE when used along its length, delaying outboard loss of

effectiveness.

In the case of the vortex flap, the lift-to-drag is insensitive to

deflection angle in the range. CL ~ 0.4 to 0.9. High-a drag-reduction

effectiveness is improved through the utilization of an inverse tapered,

as opposed to a constant, flap chord distribution to better accommodate

.the expanding vortex core. In addition, segmentation of the vortex flap

improves the high-a performance of the device on a per unit flap area

basis through the formation of a separate vortex on the outboard segment.

A chordwise slot in the vortex flap has the same aerodynamic effect ! .

but has no influence on the total leading-edge thrust. Parabolic and i . inverse tapered segmented. flap chord distributions provide f~rther ! -

improvements in flap area efficiency. The performance of these flaps

~uggest that -similar variations in SLEE extension may ~nprove the

device's ability to hold the primary vortex ahead of the leading edge.

Twisting the SLEE and vortex flap in future tests may also aid in that

respect.

By means of an upward deflection, the vortex flap is also effective

as a deceleration and high-lift device for approach and landing. Large

downward deflection upon touch-down provides continued aerodynamic

braking, along with the added advant~ge of downpressure on the wheels.

199

In addition, differential deployment of the vortex flap appears to pro

vide roll augmentation at high a.

Finally, it should be noted that the devices tested here are still

in the early stages of development and, therefore, future tests must be

performed on scale ~odels of actual aircraft to determine the effects of

parameters such as camber, sweep, twist, and fuselage interference prior

to making any final decisions on their effectiveness. In addition, it

is recommended that future research include testing of the lateral/

directional characteristics of the devices to complete a data base to

be used in the final design.

200

, 'R~FERENCES

. "~' '. ,

1. Rao, D. rv'!.: Leading-Edge IlVortex Flapsll for Enhanced Subsonic

AerodYnami.cs of Slender Wings. ICAS-80-13.5, 12th Congress of ( ..... ;<PO ....

the International Council of the Aeronautical Sciences (Munich),

1980. ':,'

2. Rao, D. M. and Johnson,T .. ,D."Jr.: Subsonic Wind-Tunnel Investi-

gation of Leading-Edge Devi~es.0n Delta Wings (Data Report).

NASA CR 159120, ,1979.

3. Rao. D. M.: Leading Edge Vortex-Flap Experiments on a 74-Deg.

Delta Wing. NASA CR 159161,1979.

4. Bobbitt, P. J.: Modern Fluid Dynamics of Subsonic and Transonic

Flight. AIAA Paper 80-0861, 1980.

5. Johnson, T. D., Jr. and Rao. D. M.: Experimental Study of Delta

Wing Leading-Edge Devices for Drag Reduction at High Lift.

NASA CR 165846, 1982.

6. KUchemann, D., FRS: Types of Flow on Swept Wings. J. Royal Aero.

Soc., Vol. 57, pp. 683-699, November 1953.

7. Harris, C. D. and Bartlett, D. W.: Wind Tunnel Investigation of

Effects of Underwing Leading-Edge Vortex Generators on a Super-

critical-Wing Research Airplane Configuration. NASA TM X-2471,

1972.

y

8. Rao, D. M. and Johnson, T. D., Jr.: An Investigation of Delta

Wing Leading-Edge Devices. J. Aircraft, Vol. 18, No.3,

pp. 161-167,1981.

201

9. Fox, C. H., Jr. and Huffman, J. K.: Calibration and Test Capabili

ties of the Langley 7- by 10-Foot High Speed Tunnel. NASA

TM X-74027, 1977.

10. Fox, C. H., Jr.: Real Time Data Reduction Capabilities at the

Langley 7- ~y 10-Foot High Speed Tunnel. NASA TM 78801, 1980.

11. Rao, D. M. and Tingas. S. A.: Subsonic Balance and Pressure

Investigation of a 60-Deg. Delta Wing with Leading-Edge Devices

(Data Report). NASA CR 165806, 1981.

12. Gillis, C. -L., Polhamus, E. C., and Gray, J. L., Jr.: Charts for

Determining Jet-Boundary Corrections for Complete Models in 7- by

10-Foot Closed Rectangular Wind Tunnels. NACA NR L-123, 1945.

13. Herriot, J. G.: Blockage Corrections for Three-Dimensional-Flow

Closed-Throat Wind Tunnels, with Consideration of the Effect of

Compressibility. NACA Rep. 995, 1950.

14. Ridder, S.: On the Induced Drag of Thin Plane Delta Wings. An

Experimental Study of the Spanwise Distribution of the Leading

Edge Forces at LOlli Speeds. KTH AERO TN 57, Royal Institute of.

Technology (Stockholm), 1971.

202

15. Margason, R. J. and Lamar, J. E.: ~ Vortex-Lattice Fortran Program

for Estimating Subsonic Aerodynamic Characteristics of Complex

P1anforms. NASA TN 0-6142, 1971.

16. Stadler, E. L. and Trammler, S. J.: F-4C Performance Data and

Substantiation. McDannel Aircraft Corporation Report G 853,

Vol. I, 1968.

17. Marchman, J. F.: The Aerodynamics of Inverted Leading Edge Flaps

on Delta Wings. AIAA Paper 81-0356, 1981.

-'

..,.

1, Reoo;;:-No. -- 2. Governmen~Acces~ion No. ! 3. Recipient's Cataloo No.

NASA CR- 165923 ! . -4. Title and Subtitle f 5: RePort Date

Subsonic Balance and Pressure Investigation of a 60-Deg ~ . Ma~ 1982 Delta Wing with Leading-Edge Devices I o. Performing OrganIzatIOn Code , I 505-31-43-03

7. Author(s) I 8. Performing Organization Report No.

Stephen A.Tingas* and Dhanvada M. Rao** I------------~------------I--_____________________ -,-__ '--___________________ . ____ ,~ 10, Work Unit No.

9, Performing Organization Name and Address !

I North Carol ina State University Ii------------~--' I Mechanical and Aerospace Engineering Department ! 11. ~C~rI~406 Grant No,

I Raleigh, NC 27650.

I 13. Type of Report and Period Covered

12. Sponsoring Agency Name and Address Contractor Report I National Aeronautics and Space Administration

"

Washington, DC 20546 14. Sponsoring A,gency Code

I 505-31-43-03 15. Supplementary Notes

*North Carolina State University, Raleigh, NC 27607 **Vigyan Research Associates, Inc., 28 Research Drive, Hampton, VA 23666

116. Abstract

'I' Low supersonic wave drag makes the thin highly swept delta wing the logical choice for use on aircraft designed for supersonic cruise. However, the high-lift maneuver capability of the aircraft is limited by severe induced-drag penalties attributed

I I I

·1

I I i i I

I I

I I

i I

I \17.

I

to loss of potential flow leading-edge suction. Th'is drag increase may be alleviated through leading~edge flow control to recover lost aerodynamic thrust through eith~r retention of attached leading~edge flow to higher angles of attack or exploitati'on of the increased suction potential of separation-induced vortex flow. A low-speed wind-tunnel investigation was undertaken to examine the high-lift devices such as fences, chordwi'se slots, pylon vortex generators, leading-edge vortex, flaps, and sharp leading-edge extensions. The devices were tested individually and i~ combinations in an attempt to improve high-alpha drag performance with a minimum of ! low-alpha drag penalty. This report presents an analysis of the force, moment, and !

static pressure data obtained i'n angles of attack up to 230, at Mach and Reynolds

numbers of 0.16 and 3~85 x 106 per meter, respectively, The results indicate that all the devices produced drag and longitudinal/lateral stability improvements at high lift with, in most cases, minor drag penalties at low angles of attack.

Key Words (Suggested by Author(s)) 18, Distribution Statement

Leading-edge devices Delta Wing , Unclassified ~ Unlimited

,. I Subsonic wind-tunnel test Ba)ance and pressure data High-lift drag reduction

19. Security Oassif. (of this report)

Unclassified 20. Security Classif. (of this page)

Unclassified

Subject Category - 02

21. No. of Pa;;-es

216 1 22, Price

! A.10

N-305 For sale by the National Technical information Service, Sprinefieid, Virginia 22161

End of Document

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