NASA Contractor Report 178046 

SUBSONIC FLOW INVESTIGATIONS ON A CRANKED WING 

DESIGNED FOR HIGH MANEUVERABILITY /I 

Dhanvada M. Rao 

(NASA-Cl3-178046) SUBSCNIC PLOW M86-20361 INVESTIGATIONS CN A CRANKED WING DESIGNED 

FOB HIGH fiANEUVEBABILITP final Beport 

€iC AO3/tlF A01 CSCL OlA G3/02 05686 

I 

I (Vigyan Research Associates, Ioc.) 45 p Unclas 

VTGYAN RESEARCH ASSOCIATES, INC. 

Hampton, Virginia 23666 

Contract NAS1-17420 

February 1986 

, 

National Aeronautics and 

Space Administration 

bngley Research Center 

Hampton, Virginia 23665 
TABLE OF CONTENTS 

Page 

INTRODUCTION ........................... 1 

LIST OF SYMBOLS ......................... 1 

EXPERIMENTAL DETAILS ....................... 2 

BASIC WING CHARACTERISTICS .................... 2 

TIP PANEL FLOW CONTROL CONCEPTS ................. 5 

RESULTS OF FLOW CONTROL CONCEPTS ................. 6 

N.C.S.U. SEMI-SPAN MODEL: ................... 6 

Fence ............................ 6 

Pylon Vortex Generator ................... 7 

Mid-Span Strake ....................... 7 

Cavity Flap ......................... 8 

LRC WING/BODY MODEL: ..................... 8 

THEORETICAL PREDICTION OF BASIC WING FORCE CHARACTERISTICS .... 9 

CONCLUSIONS AND RECOMMENDATIONS ................. 10 

ACKNOWLEDGEMENT ......................... 11 

REFERENCES ............................ 11 

i 
INTRODUCTION 

This report summarizes the main results of exploratory experiments on 

flow control devices applied to a cranked-wing planform, with the object of 

increasing the subsonic maneuverability of supersonic fighter designs under 

evaluation in the NASA Langley Advanced Concepts Branch, HSAD. The first part 

of the study was devoted to understanding the basic flow phenomena responsible 

for usable lift limitation on an uncambered 70 deg./50 deg. cranked wing, due 

to onset of pitching moment non-linearity and longitudinal instability. With 

the insight gained into flow separation phenomenon and the formation and growth 

of vortices with increasing angle of attack, several flow control device concepts 

were proposed for improving the cranked wing pitch characteristics in order to 

increase the usable CL. 

explored, and recommendations made for further research. 

The prospects and limitations of these concepts were 

C 

cD 

‘Do 

cL 

‘m 

cN 

cP 

cP,u 

H.L. 

PVG 

X 

LIST OF SYMBOLS 

Local chord 

Drag coeiii ci ent 

Zero-1 ift drag coefficient 

Lift coefficient 

Pitching moment coefficient 

Normal force coefficient 

Pressure coefficient 

Upper-surface pressure coefficient 

Hinge line station 

Pylon vortex generator 

Chordwise distance from leading edge (positive 

upper surface, negative lower surface) 

1 
Angle of attack 

Upwash angle 

Spanwise distance as fraction of wing semi-span 

Sidewash angle 

EXPERIMENTAL DETAILS 

Two separate wind tunnel investigations were conducted. A semi-span wing 

model of composite construction, incorporating upper-surface static pressure 

taps along two chordwise rows was tested in the North Carolina State University 

32 inch x 45 inch subsonic tunnel. This model was also used for extensive flow 

visualizations. The major dimensions of this model and its vertical mounting 

on a reflection plane are shown in fig. 1. The pressure and oil-flow tests 

were conducted at a nominal free-stream velocity of 90 ft/sec., corresponding 

to a mean-chord Reynolds number of 0.9 x 10'. 

the test speed was approximately 20 ft/sec. 

For helium-bubble visualizations 

The balance model in the form of a wing/body configuration, tested in the 

NASA Langley 12-ft Low Speed Tunnel, is shown in fig. 2. The tests were run 

at a free-stream velocity of 70 ft/sec. giving a mean-chord Reynolds number of 

0.92 x lo6. 

The geometry and major dimensions of the flow control devices as tested 

on the NCSU semi-span model are shown in fig. 3. The fence and pylon vortex 

generators on the Langley force model were scaled with respect to the wing 

chord at the crank. 

BASIC WING CHARACTERISTICS 

To help understand its low-speed aerodynamics characteristics, the cranked 

wing may be viewed as a conventionally swept wing having a relatively large apex 

2 
extension (or strake). 

where a combination of high aerodynamic loading and boundary layer build-up 

leads to early separation. 

1 eading-edge flow separation covering the full chord 1 ength, producing marked 

lift loss and drag rise. 

The conventionally-swept wing stalls in the tip region, 

The thin wing sections near the tip rapidly develop 

The addition of a highly swept apex region generates a pair of counterrotating leading edge vortices, which persist downstream to the wing trailing 

edge. 

surface. 

generates additional lift, which rapidly increases with angle of attack. 

Furthermore, the rotational field of the vortex induces a sidewash component 

directed towards the tips, which increases the effective sweep angle of outboard 

leading edges. Thus, an outboard leading edge vortex originates from the crank, 

which serves to stabilize the flow separation and generates vortex lift on the 

outer panel. 

These apex vortices are stabilized by flow reattachment to the wing 

A high suction level induced on the wing underneath each vortex core 

The apex vortex also induces a vertical velocity on the crank vortex, 

tending to raise it away from the outer panel surface. When at some angle 3f 

attack the apex vortex has gained considerable strength, it rapidly displaces 

the crank vortex upwards as depicted in fig. 4, resulting in a prounounced loss 

of outer-panel lift. 

of the apex vortex in the wake towards the trailing edge. Progression of the 

vortex breakdown forward over the wing surface reduces the vortex lift on the 

inboard regions of the wing. 

Increasing angle of attack also moves the breakdown position 

The upper-surface flow visualizations and pressure measurements on the 

basic cranked wing broadly confirm the flow development pictured above. 

sequence of oil-flow photographs at increasing angle of attack is presented 

in fig. 5. The upper-surface flow at a = 5 deg. is fully attached. At a = 10 

A 

3 
L 

and 15 deg. the characteristic tracks produced by the apex and crank 

vortices on the wing surface are separate and traceable back to the trailing 

edge. At a = 17 deg. however, the crank-vortex track abruptly disappears well 

ahead of the trailing edge, and the spanwise boundary layer flow penetrates 

farther into the outer panel. 

in fig. 6 shows the crank-vortex track continuing to shrink with further 

increase in angle of attack. 

A second sequence of oil-flow photographs shown 

A series of helium bubble visualizations is presented in fig. 7. Each 

picture is a composite of two separate photographs, obtained with the bubblegenerator moved from the apex to the crank position for better definition of 

the respective vortices. 

At higher angles of attack, massive flow separation on the outer panel defeated 

attempts to get the bubbles closer to the wing surface, and the crank vortex 

could only be glimpsed occasionally. The apex vortex core however remained 

easily identifiable; the onset and progress of its breakdown is seen in the 

photographs for c1 = 25 deg. and 30 deg. (A video-tape record of the he1 ium 

bubble visualization tests can be obtained from ViRA, Inc.) 

At a = 15 deg. the two vortex cores remain distinct. 

Chordwise pressure distributions at the inboard and outboard stations with 

increasing angles of attack are presented in fig. 8. 

remains attached at a = 5 deg., but ci = 10 deg. local suction peaks reveal the 

presence of apex and crank vortices at the respective stations. The inboard 

suction peak continues to develop with increasing angle of attack, but the 

outboard pressure distributions rapidly flatten out indicating full-chord 

separation starting ata= 15 deg. 

The upper-surface flow 

The local normal force coefficients, obtained by integrating the above 

pressure distributions, are plotted versus angle of attack in fig. 9. This 

comparison highlights the load trends above c1 = 10 deg., at which point the 

4 

~ ~ 
outboard panel stalls while the inboard section continues to generate 

vortex 1 ift. Consequently a progressive longitudinal instability wou 

cipated above ci = 10 deg. 

The lift and pitching moment measurements (on the wing-body mode 

strong 

d be antiof fig. 

2) are shown in fig. 10. 

instability onset at ci = 10 deg.; however the lift-curve slope remains continuous 

across a = 10 deg. implying that the growth of vortex lift over the inboard region 

is compensating for the decreasing lift on the outboard panel. 

to emphasize the large disparity between usable and available maximum lift 

coefficients typical of a crank. wing, and indicate considerable potential for 

improving its maneuver and low-speed capability through effective control of the 

outer panel flow at high angles of attack. 

The moment data confirm the expected longitudinal 

These results serve 

TIP PANEL FLOW CONTROL CONCEPTS 

The effectiveness of conventional leading-edge deflection to prevent outer 

panel stall on this wing would appear to be constrained by excessive upwash 

induced lcca!!y,e.g. the chord-plane flow angle exceeding 50 deg. at N, = 20 deg. 

as indicated by potential flow calculations (using PAN AIR code) presented in 

fig. 11,A. 

flow separation may be expected which largely cancels the advantage of achieving 

a smooth on-flow. 

however the typically large tip chord of such flaps poses structural and 

actuation problems. 

reduction,other concepts than those relying on camber variation for outer-panel 

flow control appeared worth considering. 

At such high deflections of a leading edge flap, severe hinge-line 

This problem may be avoided by the use of vortex flaps; 

While a leading edge flap offers the advantage of drag 

In the present study, three alternative schemes were conceived as depicted 

in fig. 12 and explained below: 

5 
A). Generate a control vortex at the crank, opposite in rotation 

to the leading-edge vortices. 

would be to induce a downward velocity on the crank vortex 

resisting its lift-off tendency. The fence and pylon vortex 

generator are two different means of creating this control 

vortex, and may be employed in combination. 

The purpose of this vortex 

9). Generate a control vortex near the crank having the same 

rotational sense as the leading edge vortex. This new vortex 

would serve to energize the stalling flow on the tip panel, 

much as a wing-root strake. A relatively small folding strake 

that can be deployed at a specified angle of attack is 

env i saged. 

A lower-surface hinged flap to trap a vortex just below the 

leading edge. This spanwise vortex serves the dual purpose 

of (a) assisting the incident flow at high upwash angle to 

turn towards the outer wing panel surface, and (b) moderating 

the intensity of vortex lift generated on the inner wing and 

thus reducing the unstable pitching moment. 

induced in the 'cavity' produces a vortex-flap type of thrust 

component for the additional benefit of drag reduction. 

The vortex suction 

RESULTS OF FLOW CONTROL CONCEPTS 

N. C. S. U. SEMI-SPAN MODEL: 

Fence - The effect of a chordwise fence on the outboard upper-surface 

pressure distribution at several angles of attack is shown in fig. 13. In 

comparison with the basic wing, the fence generates higher levels of suction 

6 
over the outboard panel. 

with increasing angle of attack, an opposite effect occurs between a = 25 deg. 

and 30 deg. 

a rapid growth of the fence vortex can be inferred especially at ct = 20 deg. and 

a = 30 deg., when 

While this suction increment gradually diminishes 

In the corresponding oil flow visualizations presented in fig. 14, 

25 deg. Apparently the fence vortex is further intensified at 

its spanwise influence extends to the whole of the tip region. 

Pylon Vortex Generator - The effects produced by a crank- ocated PVG are 

shown in fig. 15. A distinct suction peak at a = 15 deg. suggests the persistence 

of a crank vortex, in contrast to the nearly uniform pressure on basic wing data. 

At higher angles of attack, the overall suction level increases relative to the 

basic wing. 

the basic wing in fig. 16. 

fence-PVG combination. 

increment in the angle of attack before reaching CNmax; the rate of decline of 

the local normal force after CNmax is also moderated. 

improvement is even greater. These trends are distinctly favorable to pi tch-up 

a1 1 evi at i on. 

The pressure-integrated local normal force with PVG is compared with 

Included are results with the fence, and also a 

Both these devices individually provide a 2 to 3 deg. 

In combination the 

Mid-Span Strake - The outboard pressure distributions on the wing with an 

in-plane strake are presented in fig. 17. At all the test angles of attack a 

well-established vortex originating from the outboard junction of the strake 

and wing leading edge is indicated. A typical oil flow pattern at a = 20 deg. 

presented in fig. 18 reveals the track of this vortex, and demonstrates the 

beneficial effect of the strake on the tip flow in comparison with the basic 

wing. Note also the considerable vortex action on the strake itself. 

direct vortex lift on the strake being close to the wing centroid should not 

greatly influence the center of pressure. 

have yet been obtained for this configuration. 

The 

No overall force and moment data 

7 
Cavity Flap - A preliminary test was performed with a 10- deg. deflected 

cavity flap attached under the outboard leading edge. 

pressure at one location underneath the leading edge (i.e. in the flap cavity 

region) is plotted versus angle of attack in fig. 19. 

negative pressure (relative to the basic wing) starts at a = 10 deg., implying 

the onset of a cavity vortex. 

with angle of attack. 

the cavity corner revealed the suction level to be relatively uniform (fig. 20). 

The chordwise upper surface pressure distributions with the cavity flap on the 

outer panel (fig. 21) are not particularly informative, although the presence of 

a crank-vortex suction peak at a = 15 deg. (in comparison with the flap-off case) 

suggests that a reduction of the local incidence might have occurred. 

The lower surface 

With the flap on, a 

Thereafter, the cavity suction steadily increases 

Spanwise exploration with a static pressure probe along 

Results with a 30 deg. deflected cavity flap attached to the inboard (70 

deg. swept) leading edge, are presented in fig. 22. In this case the uppersurface distributions clearly reveal a diminished primary vortex suction peak 

relative to the basic wing. 

this section will help to improve longititudinal stability of the cranked wing. 

A pronounced suction generated in the flap cavity also is noted.* On the basis 

of these pressure results, force tests to evaluate the potential of cavity flaps 

The associated reduction of the nose-up moment on 

for usable-lift enhancement as well as for drag reduction appear worthwhile. 

LRC WING/BODY MODEL: 

The main purpose of this test was to assess the fence and PVG devices in 

extending the longitudinally-stable CL range of the cranked wing. 

moment characteristics with and without the devices are compared in fig. 23. 

The pitching 

*Note: The capture of a stable cavity vortex was confirmed in water tunnel 

experiments performed cooperatively with Northrop Aircraft Corporation; a 

video-tape is available with ViRA, Inc. 

8 
An evaluation in terms of the stability derivative dCm/dCL of the devices is 

presented separately in fig. 24 and 25. 

eliminate the basic wing pitch-up at CL = 0.65 (CY = 15 deg.) The fence 

however is progressively destablizing above CL = 0.4, an effect not found 

with the PVG. 

from 0.65 to 0.8. 

could possibly be alleviated by the use of a second PVG located on the inboard 

leading edge, based on author's previous experience with multiple PVGs on delta 

wings (ref. 1). 

Both the fence and the PVG clearly 

In either case, the longitudinally stable limit of CL is raised 

A slrongpitch up noted just above CL = 0.8 with the PVG 

- THEORETICAL PREDICTION OF BASIC WING FORCE CHARACTERISTICS 

The vortex lattice method incorporating the Polhamus suction analogy, or 

VLM-SA (ref. 2) has been generally successful in predicting the lift and drag 

characteristics of highly swept delta and strake-wing planforms with fully 

developed leading-edge vortices. 

follows (see fig. 26) 

The available options in VLM-SA are as 

(L.E.S. = Leading Edge Suction) 

Option Inboard L.E. Outboard L.E. 

A Vortex Vortex 

B Vortex Attached (full L.E.S.) 

C Vortex "Separated" (zero L.E.S.)" 

D Attached (full L.E.S.) Attached (full L.E.S.) 

E 

* * 

"Separated" ( zero L . E. S . ) "Separated" (zero L . E. S . ) 

* 

This is a fictitious flow model, assuiiled to represent a fully-stalled 

condition. 

9 
In applying the VLM-SA method to the present cranked wing, a two-planform 

model was adopted: (I) the apex extension and, (11) the rest of the wing. (See 

sketch in fig. 27.) The 'augmented lift' concept was applied at the crank position 

to acccount for the induced lift of the shed apex vortex. 

The 1 ift characteristics predicted by the various VLM-SA options are 

compared with wind tunnel data in fig. 27. Even before vortex breakdown, the 

experimental lift capability of this wing is seen to fall considerably short of 

the full vortex lift prediction. 

breakdown) is obtained with the VLM-SA option C, i.e. assume full vortex lift on 

the inboard leading edge and zero vortex lift outboard. 

drag polar prediction, as shown in fig. 28. 

of the attainable performance if full vortex lift could be sustained on the outboard leading edge at higher angles of attack. 

Indeed, the best lift prediction (up to vortex 

This is also trueof the 

These calculations provide an indication 

CONCLUSIONS AND RECOMMENDATIONS 

Exploratory flow visualizations and pressure tests were performed on a 

70 deg./50 deg. cranked wing model to assess the potential of device concepts 

in controlling the tip-panel stall, in order to alleviate or postpone the 

onset of longitudinal instability to a higher lift coefficient. 

studied were (1) fence, (2) pylon vortex generator, (3) a folding strake and 

The concepts 

(4) cavity flap. The fence and PVG were also tested on a balance model. 

Both the fence and PVG, individually and in combination, produced distinct 

and favorable effects on the vortex flow development and local normal-force 

characteristics on the tip panel, resulting in a 'L increment of 0.15 before 

instability onset. Refinement of the design and placement of these devices is 

expected to yield additional improvements. 

10 
Although the scale of these experiments was relatively small, the results 

arebelievedto be basically valid for full-scale Reynolds number, because the 

crank wing pitch-up problem as well as the proposed concepts for its alleviation 

involve sharp edge separation vortices and therefore should be only weakly 

Reynolds number dependent. 

The limited but promising pressure and flow visualization results on the 

folding strake concept warrant balance testing in order to establish its liftimprovement potential. 

Similarly, pressure measurements and flow visualization studies of the 

cavity flap concept were encouraging and a force test program is recommended. 

ACKNOWLEDGEMENT 

The assistance of Robert Vess, North Carolina State University, and 

Gregory Riebe, Advanced Concepts Branch (HSAD) , NASA Langley, during the wind 

tunnel tests reported here is thankfully acknowledged. 

REFERENCES 

1. Rao, D. M. and Johnson, T. D., Jr.; "Alleviation of the Subsonic PitcnUp of Delta Wings", J. Aircraft, Vol. 20, No. 6, June 1983, pp. 530-535. 

2. Lamar, J. E.; "Summary of Some Recent Studies of Subsonic Vortex Lift 

and Parameters Affecting the Leading-Edge Vortex Stabil i ty", AIAA Paper 

76-414, July 1974. 

11 
I +-TUNNEL I 32 CROSS SECTION 

SHOWING WING 

ON REFLECTION PLANE 7 3,15 

OUTBOARD 

PRESSURE 

5% THICK CIRCULAR ARC . ). . ./ 1 

AIRFOIL 

50 ' 

... 

1 

15.75 

/ 4,5 I W77' I I 

I- 29,15 

Figure 1 - Semi-span Wing Model tested in N.C.S.U. subsonic 

wind tunnel (dimensions in i nches) 
- TYPICAL WING SECTION 

46 O mi1 

Figure 2 - Wing-body balance model tested in NASA Langley 

12-ft low speed tunnel (dimensions in inches) 

13 
'/ \\\\\\\\\\\\ \\\\\\\ \\\\\ \ \ -., , . \\\\\\\\\ \\\\\\\\\\\\\\\\\ ' ' ' . . 

9 

WING SECTION AT CRANK I I 

A> Fence and Pylon Vortex Generator 

Figure 3 - Flow control devices (dimensions in inches) 

14 
B) Folding Mid-Span Strake 

Figure 3 - Continued 

POSITION 

15 
C) Leading Edge Cavity Flap 

Figure 3 - Concluded 

16 
(VIEW FROM T,E, LOOKING UPSTREAM) 

LOW ALPHA 

I \\ APEX VORTEX 

CRANK VORTEX 

\\ CRANK I 

I 

Figure 4 - Hypothesized development of leading edge vortices 

on cranked wing. 
Figure 5 - Oil flow visualizations of basic cranked wing 

showing progression of tip panel stall (a = 5' to 17") 

18 
n 

0 

M 

0 

Ln 

N 

0 

0 

N 

0 

Ln 

1 

0 

0 

4 

r: 

0 

*r 

v) 

v) 

aJ 

L 

a, 

0 

L 

Q 

v) 

cco 

04 

v) I1 

S 

08 

.C w 

I 

19 
Figure 7 - Helium bubble visualizations of basic cranked wing 

20 
- INBOARD STN, ---- OUTBOARD STN , 

1, ----- 

'-cp I, ---.-- 

_--- 1 d= 5" 

0 y 

_-------- c c@ 

0 - -_---- \, d = 15" 

2 

----- ---- -- 

0 & = 25" 

0 0.25 0,s 0,75 1,O x/C 

Figure 8 - Pressure distributions on basic cranked wing 

.. 

21 
1.6 

1.4 

CN, LOCAL 

.1,2 

1,o 

0.E 

0,f 

0, 

0, 

\ INBOARD \f STN, 

/ 

/ 

0 5 10 15 20 

Figure 9 - Local normal force characteristics from integrated 

pressures on basic cranked wing. 

22 
1.2 

CL 

1 

1.0 

0.8 

0, E 

0,4 

0.2 

( 

-b 

I 

0 10 20 30 40 50 

PC (DEG,) 

0 

Crn 

-,02 

-,04 

-,06 

-,08 

-,lo 

Figure 10 - Lift and pitching moment characteristics of basic 

cranked wi ng/ body model 

23 
60 

o+s -UPWASH ANGLE 

50 

40 

30 

20 

10 

0 082 0,4 7 0.6 0,s 1,o 

ANGLE OF ATTACK 

A) UPwash 

Figure 11 - Potential flow (PAN AIR) calculation of basic cranked 

wing 

24 

- 
50 

+ 

(DEG. 1 

40 

30 

20 

10 

CRANK A 

0 0,2 0,6 0,8 1,o 

Ob4 71 

B> Sidewash 

Figure 11 - Concluded 

25 
FENCE 

A) Counter-Rotciting 

Crank Vortex 

PVG 

otating 

k Vortex 

C) Cavity Flap 

Figure 12 - Outboard panel flow control concepts 

26 
----- BASIC WING; - WITH FENCE 

1.0 

0 

2.0 

-cp, u 

le0 

- 

8 I I 1 

0 

1.0 

r 

I I I 

p< = 15" 

0' I I 1 I 

od = 20" 

bc= 25" d = 30' 

d = 25", 30" 

" 

0 0,25 0,5 0,75 x/c 

1,o 

Figure 13 - Effect of fence on outboard panel upper-surface 

pressures 

27 
0 

28 

0 

M 

E 

0 

0 

Ln 

cv 

0 

0 

cv 

0 

Ln 

d 

9- 

0 

cc 

cc 

aJ 

.. 

-I- 3 00 

I 

'aJ L 

0 

0 

ri 

It 

.I 
----- BASIC WING; - WITH PVG 

1,o 

0 

2.0 

-Cp, u 

d= 25" 

-- --- - ' d- - f 

I I ! 1 

1.0 

0 I 

0' I I I I 

Figure 15 - Effect of PVG on outboard panel upper-surface 

pressures 

OUTBOARD 

STN, 

29 
1,2 

CN, LOCAL 

1.0 

0,8 

0.6 

0,4 

0.2 

WITH FENCE 

ONLY 

WING SECTION 

0 5 10 15 20 25 

b( (DEG,) 

Figure 16 - Local normal force characteristics of outboard 

panel showing effect of devices 

30 

~ 
2.0 

-cp, u 

0 

0 

1.0 

0 

1.0 

----- BASIC WING; - WITH STRAKE 

I I I I 1 

I 

H = 25" 

0 

0 0.25 0.5 0.75 x/c 1.0 

Figure 17 - Effect of folding strake on outboard panel 

upper-surface pressures 

31 
I 

I3 

U 

W 

z 

.. 

v) 

4 

cn 

32 

0 

c 

C, 

-0 

E: 

CrJ 

ma, 

3Y v)m 

I 
LOWER SURFACE PRESSURE a X/C = ,075 

(OUTBOARD STAT I ON 1 

CP 

-1,o 

-0.5 

0 

0.5 -- - ---I----- 

1 I I I I I 

0 5 10 15 20 25 

& (DEG, 1 

Figure 19 - Variation of cavity pressure with angle of attack, 

outboard cavity flap. 

33 
. 

-2,o 

CP,CAVITY 

-1,o 

0 I I I 

0.6 0,7 0.8 27 0,9 1,o 

I 

CRANK TIP 

Figure 20 - Spanwise distribution of cavity pressure with 

outboard cavity flap at U= 25 deg. 

34 
----- BASIC WING; - WITH FLAP 

1,o 

o( = 25' 

- 

------ 

----------- 

0' I I I J 

0 0,25 0.5 0.75 x/c 1.0 

Figure 21 - Effect of outboard cavity flap on outboard panel 

upper-surface pressures 

35 
L.E. 

-2,o 

-1,o 

0 

T.E. 

+x 

-0,25 0 0,25 0,5 0,75 1,o 

x/c 

Figure 22 - Effect of inboard cavity flap on inboard panel 

upper-surface and cavity pressures 

36 
0 0.2 0,4 0.6 0,8 1,o 1.2 

cL 

WING PVG 

Figure 23 - Effect of fence and PVG on pitching moment 

characteristics of wing-body model. 

37 
0.3 

0.2- 

0.1’ 

0 

-.01 

BAS I C 

WING 

-.02 

0 0,2 0,4 0,6 0.8 1.0 1#2 

cL 

Figure 24 - Longitudinal stability derivative of wing-body 

model with and without fence 

/ /‘ FENCE 
0.3 

acm, 

acL 

0,2 

0,l 

0 

BAS I C 

-,02 0 0,2 0,4 0,6 0,8 1,o 1,2 

cL 

Figure 25 - Longitudinal stability derivative of wing-body 

model with and without PVG 

39 
I 

n 

POTENT I AL FLOW 

( FULL LoEo 

SUCTION 

LoEo VORTEX SEPARATED& FLOW 

(FULL VORTEX (L,E,SUCTION AND 

LIFT) VORTEX LIFT BOTH 

O) ----- 

Inboard 8 Out board 

ZERO 1 

Vortex 

Inboard Vortex; Outboard Attached 

Inboard Vortex; 

Outboard Separated 

a@ Inboard Attached 8 Outboard 

----- 

Inboard 8 Outboard 

Separated 

----- 

Figure 26 - Options in VLM-SA computation of basic wing 

40 
. 

\ PLANFORM 

VLM-SA OPT1 ONS 

(SEE FIG, 26) 

0 10 20 30 .40 

d. (DEG, 1 

c 

2,o 

Cl 

1.6 

1,2 

8!8 -- 

0.4 

Figure 27 - Comparison of basic wing-body model lift data 

with various VLM-SA options (body not represented 

in VLM computation) 

41 
VLM-SA OPTIONS 

(SEE FIG,25) 

1.6 

cL 

1,2 

0,6 

0,4 

0 

I- 

------ - e/ -- /-- 

/ 

7 

DATA 

f I 

I I I 1 

Figure 28 - Comparison of basic wing-body model drag polar 

with various VLM-SA options (body not represented 

in VLM computation) 

42 
1. Report No. 

NASA CR-178046 

2. Government Accession No. 3. Recipient’s Catalog No. 

- ~~ 

5. Report Date 

February 1986 

6. Performing Organization Code 

19. Security Classif.(of this report) 

Uncl assif ied 

8. Performing Organization Report No. 

20. Security Classif.(of this page) 21. No. of Pages 22. Price 

Unclassified 44 A03 

10. Work Unit No. 

11. Contract or Grant No. 

NAS1-17420 

13. Type of Report and Period Covered 

Contractor Report 

14. Sponsoring Agency Code 

505-61-71-05 

Langley Technical Monitor: William J. Small 

Final Report 

16. Abstract 

The characteristic pitching-moment nonlinearity of cranked wings limits 

their usable lift coefficient well below CLrnax. The potential of several 

aerodynamic devices, viz. fences, pylon vortex generators (PVG), mid-span 

strakes and cavity flaps, in delaying the pitch-up onset on a 70/50 deg. 

cranked wing was explored in low-speed tunnel tests. Upper-surface pressure 

measurements and flow visualizations were conducted on a semi-span wing model 

to observe the vortex flow development with increasing angle of attack, and then 

to assess the effectiveness of the devices in controlling the collapse of vortex 

lift over the wing panel outboard of the crank. Force tests on a full-span wing 

and body model were also conducted to assess the fence and PVG in improving the 

usable CL. 

17. Key Words (Suggedted by Authom(s)) I 18. Distribution Statement 

Pi tch-Up Control 

Aerodynamic Devices 

Vortex Management 

Cranked Wing 

Uncl assi f i ed-Unl imi ted 

Subject Category 02 

For sale by the National Technical Information Service, Springfield, Virginia 22161 

NASA Langley Form 63 (June 1985)