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P Ta 5aa DAVID W TAYLOR NAVAL SHIP REARC" AND' D6WLOPMMW CC-gic p/ao AQADA XMRININTAL INVESIATION OF A CIRCULATION CONTROL AILEON3(UIJUL 79 5 0 PRINCKUN1AhhhgoOTNMhhE"/ N

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EXPERIMENTAL INVESTIGATION OF A CIRCULATION

CONTROL AILERON

by

Steven W. Prince

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SX1VIATION AND SURFACE EFFECTS DEPARTMENT

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CID July 1979

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S XPERIMNTAL.INVESTIGATION OF A V7 ial~ptCIRCULATION CONTROL AILERON . Oct i77 Sep 78

6. PERFO AMIN ORG. REPORT NUMBER

7. AUTHOR(s) 0. CONTRACT OR GRANT NUMBERe)

/ , : Steven W.jPrince9. PERFORMING ORGANIZATION NAME AND ADDRESS 10. PROGRAM ELEMENT. PROJECT, TASK

AREA & WORK UNIT NUMBERS

David W. Taylor Naval Ship R&D Center Program ElementAviation and Surface Effects Department- Task AreaBethesda, Maryland 20084 Work Unit

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1. SUPPLEMENTARY NOTES

I19. KEY WORDS (Continue on reverse side It neceeery and Identify by block number)

I- AileronCirculation Control

20 A kRACT (Continue or% reveras el$ If necoesry and Identify by block number)

--,A Circulation Control (CC) aileron was tested on, Asemispan wing- A

fuselage model at a dypamic pressure equal to 20 lb/t L'(957 N/mg) and aReynolds number of 0.3 x 106 /ft (2.62 x 106 /m)). Three different trailingedge geometries were used on CC ailerons of 10 and 20 percent of the halfspan. Blowing was controlled to produce jet momentum coefficients fromr 0.0017 to 0.0124. Rolling moment coefficients as high as 0.035 were recorded

for the 20-percent CC aileron for angles of attack between 0 and 12 deg. The 10f-

M (Continued on revprne aide)

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-ICaileron was at least three times as effective as a pure reaction jetfor the same amount of bleed air. Adverse yaw was large, on the order ofone-half of the rolling moment.,

.Ac,r

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TABLE OF CONTENTS

Page

I LIST OF FIGURES ..................................................... iii

I NOTATION ............................................................ iv

ABSTRACT ............................................................ 1

ADMINISTRATIVE INFORMATION ........................................... I

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

MODEL ............................................................... 2

- WIND TUNNEL EXPERIMENT ................................................... 3

RESULTS AND DISCUSSION .............................................. 4

SUMMARY AND CONCLUSIONS .............................................. 6

j LIST OF FIGURES

1 - Semispan Wing-Fuselage Model and Groundboard .................... 9

2 - Fourteen-Percent Supercritical Airfoil with CirculationControl Trailing Edge ........................................... 10

3 - Geometry of Circulation Control Aileron Trailing EdgeConfigurations .................................................. 11

I 4- Wing Fence Inboard of Circulation Control Aileron Section ....... 12

5 - Smispan Wing-Fuselage Model with Double Slotted Flaps .......... 13

6 - Rolling Moment Coefficient Versus Angle of Attack ............... 15

7 - Rolling Moment Enhancement ...................................... 21

I 8 - Adverse Yawing Moment of the Twenty-Percent CirculationControl Ailerons ................................................ 24

i,tit

I I

NOTATION jA Area of slot

AR Aspect ratio

b Wing span

C Rolling moment coefficient

C Yawing moment coefficientn

C Momentum coefficient based on wing area

C Momentum coefficient based on aileron section area

jiA

m Mass flow IPd Wing plenum total pressure, psig

P Free-stream static pressure, psig 1q Free-stream dynamic pressure, lb/ft

2

R Universal gas constant IS Wing reference area, ft I

s 1715 ft 2/sec 2 R

Td Wing plenum total temperature IV Jet velocity

S Angle of attack (deg) Iy Ratio of specific heats I

iI

ivi - I |i

ABSTRACT

A Circulation Control (CC) aileron was tested on ajsemispan wing-fuselage model at a dynamic pressure equal

to 20 lb/ft 2 (957 N/m2 ) and a Reynolds number of 0.8 x 106 /ft

(2.62 x 106 /m). Three different trailing edge geometrieswere used on CC ailerons of 10 and 20 percent of the halfspan. Blowing was controlled to produce jet momentumcoefficients from 0.0017 to 0.0124. Rolling moment coeffi-cients as high as 0.035 were recorded for the 20-percentCC aileron for angles of attack between 0 and 12 deg. TheCC aileron was at least three times as effective as a purereaction jet for the same amount of bleed air. Adverseyaw was large, on the order of one-half of the rollingmoment.

ADMINISTRATIVE INFORMATION

This study was authorized and funded by the Naval Air Systems Command

(NAVAIR) 320D under Program Element 62241N and Task Area WF 41 421 000.

The work was completed in FY 79 at th6 David W. Taylor Naval Ship Research

and Development Center (DTNSRDC) under Work Unit 1-1600-079.

INTRODUCTION

Insufficient roll control power is a problem with many vertical/short

takeoff and landing (V/STOL) aircraft at transition speeds. A large amount

of roll control power is required to trim out a lift-jet induced rolling

moment due to sideslip. Rolling moment coefficient can approach 0.3 in a

30-deg sideslip.' A Circulation Control (CC) aileron is a potential method

of providing adequate roll control power at transition speeds without exces-

sive bleed air requirements.

A CC aileron is a powered roll control device in the same location on

the wing as the aileron. The CC aileron uses tangential blowing over a

round trailing edge to produce increased section lift coefficients. The

jet blown over the round trailing edge stays attached by the Coanda

principle. It moves the trailing edge stagnation point to the underside,

increasing circulation.2 Roll control is achieved by blowing the CC

aileron on one wing to raise the wing. The trailing edge can be retracted

or faired in by various means for high-speed flight.

II

MODEL DESCRIPTION

The model is a semispan wing fuselage mounted on a circular ground-

board. The principal dimensions of the model are shown in Figure 1. The

Circulation Control airfoil is a 14-percent thick supercritical wing with

a circulation control round trailing edge (Figure 2). The wing had an

aspect ratio of 4.0. Blowing air was supplied through a wing plenum. The

CC aileron section was made by closing the slot over the inboard section

of the wing with gasket material so only the outboard section was blown.

The gasket material extended approximately 0.25 in. (0.64 cm) out of the

slot to insure separation of the flow over the rounded trailing edge.

Intermittent flow attachment over the round unblown section was suspected

to have caused the considerable scatter found in the previous data. Two

CC aileron spans were investigated; one of 20-percent half span with a

slot extending 4.94 in. (13.56 cm) from the tip and the other of 10-percent

half span with the slot extending 2.47 in. (6.27 cm). Three different

trailing edge geometries were investigated; see Figure 3. A fence was used

to separate the blown CC aileron section from the rest of the wing as shown

in Figure 4.

The wing-fuselage model was mounted in the wind tunnel test section

such that only the wing was attached to the balance frame. The wooden fuse-

lage was mounted to the groundboard and was independent of the balance

frame with a small gap existing between the wing root and the fuselage body.

The forces and moments measured by the balance frame were essentially wing-

alone data in the presence of a body. A large fence was installed around

the wing root area very close to the fuselage. The fence is shown in fFigure 2, and its position is noted in Figure 1.

The circular ground board is 8 ft (2.33 m) in diameter and serves as ja reflection plane for the half model. The groundboard is constructed

from 0.5 in. (1.2 cm) plywood and is mounted to the test section floor on

2.5 in. (6.35 cm) wooden spacers, leaving a gap between the groundboard and

the tunnel floor for boundary layer bleed. The groundboard is shown in

Figure I and additional details are shown in Figure 5.

2 I___ __ _ ,_

IIf

WIND TUNNEL EXPERIMENT

The investigation was conducted in the DTNSRDC 8- by 10-foot north

subsonic wind tunnel. The model was floor mounted in a vertical position

using Strut System 8. Blowing air was supplied through a pipe in the

center of this strut. The strut system is located beneath the tunnel floor

and transfers the aerodynamic loads to an external Toledo mechanical

balance system. The Toledo balance system records six-component force and

moment data on magnetic tape using a Beckman 210 high speed data acquisition

system.

Three total pressure transducers and one thermocouple temperature

probe were mounted in the wing plenum. Due to the difficulty of actually

f measuring the jet velocity, it was calculated assuming isentropic expansion

of the air from the total pressure in the wing plenum to free-stream static

pressure. The jet velocity can be determined from the expression

(Reference 2):

The mass flow of the jet m was measured by a venturimeter in the airsupply line. The blowing coefficient C is determined from the expression:

i qSIFor each configuration, the model was set at an angle of attack and

data was taken at different blowing coefficients by varying the plenum

pressure. The model was then set at the next angle of attack, and the

process was repeated.

All forces and moments were resolved about the mean aerodynamic

chord and reduced to standard coefficient form using the stability axis

[system. Coefficients were based on twice the reference area of the semi-span model to simulate the case of a full-span model. Net rolling and

3

"II

yawing moments (AC, and AC ) are the differences between the momentncoefficients with and without blowing.

Each data point was taken as the average of 10 data samplings over a

5-sec interval. Model weight and air pressure line tare corrections were

applied to the balance data. The only aerodynamic corrections applied to

the force and moment data consisted of the standard downwash corrections

as outlined in Reference 3; angle of attack and drag coefficient were the

two parameters affected.

All data were recorded at a dynamic pressure of 20 lb/ft 2 (957 N/m )

with a Reynolds number of 0.8 x 106 (2.62 x 10 6/m). The 20- and 10-percent

CC ailerons were each tested at five momentum coefficients. Angle of

attack was varied from 0 to 24 deg in 4-deg increments.

RESULTS AND DISCUSSION

The rolling moment coefficient about the centerline of the half model

for the configurations evaluated is shown in Figures 6a through 6f. The

application of roll control to lift one wing is modeled as blowing on one

CC aileron and no blowing on the other. The net rolling moment that would

be felt by a whole model with a certain amount of blowing on one CC aileron

is the difference between that blown curve and the unblown curve.

The four 20-percent CC ailerons show a net rolling moment that is

fairly even with angle of attack for each momentum coefficient. A much

smaller net rolling moment was generated by the two 10-percent CC ailerons

with blowing. Scatter in the data is probably responsible for most of the

unevenness of the increments.

Figures 7a through 7f show net rolling moment versus momentum coeffi-

cient for all six configurations evaluated. Lines fanning out from the

origin represent levels of constant rolling moment enhancement. Rolling

moment enhancement is the net rolling moment achieved with the device

divided by the net rolling moment that would be produced by using the air

for a tip jet (assuming no losses in the tip jet).

4

!)

i AC qSb

Enhancement = ZV.(b

AC~qSb

C qS(b/2)

2AC )

C

For the 20-percent CC aileron, net rolling moment increases nonlinearly

with momentum coefficient, with less incremental net rolling moment being

gained by additional blowing. At the maximum momentum coefficient tested,

the device was still showing increasing net rolling moment with increasing

blowing. A rolling moment enhancement greater than four was achieved by

the 20-percent CC ailerons in most conditions, and enhancement greater than

12 was achieved by the short edge configuration at low momentum coefficients.

The rounded edge and square edge configurations performed better than

the long edge configuration with a maximum net rolling moment of 0.037

compared to 0.030. Removing the fence from the round edge configuration

decreased maximum net rolling moment by 10 percent, from 0.037 to 0.033.

The fence increased the performance of the low aspect ratio blown section by -

making it more like a two-dimensional section. The 10-percent CC ailerons

showed similar nonlinear trends with momentum coefficient, but performance

jwas only one-third as good as the 20-percent CC ailerons. The comparatively

poor performance of the 10-percent CC aileron is probably due to the very

low aspect ratio of the blown portion of the wing. Again, the square edge

configuration performed better than the long edge configuration.

Adverse yaw plots for the 20-percent square edge and long edge

configurations are shown in Figures 8a and b. The adverse yawing moments

measured were large, on the order of one-half of the rolling moment. The

square edge configuration shows increasing adverse yaw with angle of attack

with the exception of zero angle of attack.

5

4,

SUMMARY AND CONCLUSIONS

A CC aileron was evaluated in a low-speed wind tunnel test for

potential use on V/STOL aircraft. A large amount of roll power for a small

expense of bleed air was desired. (The highest rolling moment achieved was

Cz = 0.037 for a momentum coefficient of C = 0.014, which represents a

rolling moment enhancement of 5.29. The adverse yawing moment measured was

approximately one-half of the rolling moment.) The 20-percent CC aileron

performed three times as well as the 10-percent CC aileron. A fence on the

inboard edge of the CC aileron increased performance. The rounded edge and

square edge configurations performed better than the long edge configuration.

Various improvements are recommended for further investigation. A

spoiler raised on the opposite wing would increase rolling monent and reduce

adverse yaw by increasing diag on that wing. Up blowing on the opposite

wing from a slot on the underside of the trailing edge would also increase

rolling moment, but its effect on adverse yaw is uncertain.

6

)

I

REFERENCES

1. Marguson, R.J. and C.L. Gentry, Jr., "Aerodynamic Characteristics

of a Five-Jet VTOL Configuration in the Transition Speed Range," NASA

Langley Research Center TN D-4812 (Oct 1968).

2. Englar, R.J., "Subsonic Wing Tunnel Investigation of the High Lift

Capability of a Circulation Control Wing on a 1/5-Scale T-2C Aircraft Model,"

David W. Taylor Naval Ship Research and Development Center Report ASED-299

(May 1973).

3. Pope, A. and J.J. Harper, "Low Speed Wind Tunnel Testing,"

John Wiley & Sons, Inc., New York (1966).

I.7II

NMNLCHORD SWEEP Ideg)

ASPECT RATIO TIP ROOT L.E. T.E. SEMISPAN

(21.34cm) (46 .58cm) (0.582m)

4 .40 in. 18.34 in. 25.8 0 .12 ft

(16.31cm) 146.58cm) 2. 0 (0.646m)

________________AR =_______ 4_______ ___

AR =43

I ROOT FENCE\

ITUNNEL FLOOR. SIDE VIEW 8 ft (2.4384m) DIA.SPLITTER PLATE

.1 TOP VIEW

Fiue1-Smsa ig-Fuselage Model and Ground Board

Al

LU

00

02Z

040cu

"-4

0 4-4

z -4w :13LA-J

'4-4

4-J

94S-A ,-00

P4.J

LUU

z to4Lu)

100

m _ _ _ um _ _ _J

3.5 deg-

I CIRCULATION CONTROL AILERON

1/16 in.

CIRCULATION CONTROL AILERON WITH SIMULATED SHORT FLAP

i J 90 deg

CIRCULATION CONTROL AILERON WITH SIMULATED EXTENDED FLAP

Figure 3 -Geometry of Circulation Control Aileron1Trailing Edge Configurations

1 11

If

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Figure 4 - Wing Fence Inboard of CirculationControl Aileron Section

12

Figure 5 - Semispan Circulation Control Wing-Fuselage Model

t1.3

1t

Figure 6 -Rolling Moment Coefficient versus Angle of Attack

0.22511

Pd (09i) A

0 0.00.200 -l 1.95 0.0030

L.8.50 0.0099

Q.7 12.90 0.0141

0.150

zU

LL 0.1250

zLU

0.100

0

0.075

0.050-

0.025

10 40 4 8 12 16 20 24ANGLE OF ATTACK, ot (deg)

I Figure 6a -Twenty-Percent CC Aileron, Simulated Short Flap

15

Figure 6 (Continued)

0.225 111

Pd Ipsag) C M

0.200 - 0 0.0 0

o 2.0 0.0030

A 3.95 0.0050

0.175 f 8.60 0.0099o 13.00 0.0141

S0.150

zLUULA.LA.LU 0.1250

0S0.100-

-j-j0

0.075

0.50-

0.024

ANGLE OF ATTACK, a (dog)

Figure 6b -Twenty-Percent CC Aileron, Simulated Extended Flap

16

Figure 6 (Continued) WT EC

0.225

0.200 - pd (09i) CY

o 0.0 0

o 13.0 0.0141

0.175

0.150

Ui 0.125U.U.0.0

UZ 0.100

0.05

0

-4 00751216202

ANGLE OF ATTACK, a (dog)

Figure 6c -Twenty-Percent CC Aileron, with Fence

17

Figure 6 (Continued)

0,225111

Pd (psiq) C IjNo FENCE

0.200 0 o.0 0Q 13.0 0.0141

0.175

0.150

2

U.. 0.1250

o 0.100

2j

0cc 0.075

0.050

ANGLE OF ATTACK, a~ Ideg)

Figure 6d -Twenty-Percent CC Aileron, without Vence

18

I Figure 6 (Continued)

I 0.175

Pd (psig) CUI 0.0 00.150 0 2.0 0.0017

I A 3.95 0.0032L 8.65 0.0060

0.125 Q 13.1 0.0084

I LUj U 0.100-

U-

0

2 0.075-0

0 0.050

II ANGLE OF ATTACK, a (dog)

Figure 6e -Ten-Percent CC Aileron, Simulated Short Flap

1 19

Figure 6 (Continued)

0.200

Pd (pseg) CM

0 0.0 00.175 - C 2.05 0.0017

4.0 0.0032

L 8.7 0.0060

0.150 - 13.1 0.0084

LU 0.125

Uj0.U

0.0Ui

0

0.0750

0.050

0.025-

00 4 8 12 16 20 24

ANGLE OF ATTACK, a (deg)

Figure 6f -Ten-Percent CC Aileron, Simulated Extended Flap

20

IFigure 7-Rolling Moment EnhancementENHANCEMENT =24 16 12

0.3I PARAMETE R//

0.020

.0

0I Figure 7 a -Twenty-Percent CC Aileron, Simulated Short Flap

12z ENHANCEMENT

PARAMETER8611W 0.03

4-

1 0.02

~10 0.005 CM 0.010 0.015

10 0.05 C JAA 0.10MOMENTUM COEFFICIENT

I Figure 7b -Twenty-Percent CC Aileron, Simulated Extended Flap

21_ ___1

Figure 7 (Continued)

0.04

6ENHANCEMENT-

PARAMETER ool-

0.03 - loe--

0.02 --o l o o

Il Ce - (deg)

0.01 4---

z- 8

o 0I16

0

20.04z-j _j ENHANCEMENTo PARAMETER8

'U~ 0.03

2 - a-o

0.02 -'0, o,.1 o

0.001 C A3 s 0.1 0.015

0 0.06 0.10

MUAMOMENTUM COEFFICIENT

Figure 7d -Twenty-Percent CC Aileron, wi~thout Fence

22

Figure 7 (Continued)

0.04

ENHANCEMENTPARAMETER

12 8 60.03/

U 2

0.021

z

Zj Figure 7E,- Ten-Percent CC Aileron, Simulated Short Flay!

0

~0.04o (deg)-j ENHANCEMENT

0PARAMETER Q

I- 04Z 0.03 <

ts,16

0.212 8

0.02 2

0.005 C 0.010 0.015

01.1 0.20

MOMENTUM COEFFICIENT

Figure 7f -Ten-Percent CC Aileron--Simulated Extended Flap

i2

-0.03a(dg

0 05 4

-0.02 L16f24

U0 -0.01

zLU

U-U-LU0 0

z Figure 8a -Simulated Short Flap

o 0.03

z

-0.02

-01

0 0.01 0.02 0.03 0.04

NET ROLLING MOMENT COEFFICIENT, ACC

Figure 8b - Simulated Extended Flap

Figure 8 -Adverse Yawing Moment of the Twenty PercentCirculation Control Ailerons

24

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