NASA Contractor Report 3439 Contractor Report 3439 ... control surface normal force and hinge...

NASA Contractor Report 3439

Wind Tunnel Force and Pressure Tests

of a 13% Thick Medium Speed Airfoil

With 20% Aileron, 25% Slotted Flap

and 10% Slot-Lip Spoiler

W. H. Wentz, Jr.

Wichita Stale University

Wichita, Kansas

Prepared for

Langley Research Centerunder Grant NSG-1165

N/LSANational Aeronautics

and Space Administration

Scientific and Technical

Information Branch

1981

https://ntrs.nasa.gov/search.jsp?R=19830008016 2018-05-21T16:08:32+00:00Z

SUMMARY

Force and surface pressure distributions have been measured

for a 13% medium speed (NASA MS(I)-0313) airfoil fitted with

20% aileron, 25% slotted flap and 10% slot-lip spoiler. All

tests were conducted in the Walter Beech Memorial Wind Tunnel at

Wichita State University at a Reynolds number of 2.2 x 106 and

a Mach number of 0.13. Results include lift, drag, pitching mo-

ments, control surface normal force and hinge moments, and surface

pressure distributions. The basic airfoil exhibits low speed

characteristics similar to the GA(W)-2 airfoil. Incremental

aileron and spoiler performance are quite comparable to that

obtained on the GA(W)-2 airfoil. Slotted flap performance on

this section is reduced compared to the GA(W)-2, resulting in

a highest C£max of 3.00 compared to 3.35 for the GA(W)-2.

iii

INTRODUCTION

As part of NASA's recent program for developing new airfoil

sections (Ref. i), Wichita State University is conducting flap

and control surface research for the new airfoils. One of the

new airfoils designed for medium (subsonic) Mach number cruise

conditions is the NASA MS(I)-0313 airfoil. The present report

documents two-dimensional wind tunnel tests of this airfoil with

20% aileron, 25% slotted flap and 10% slot-lip spoiler.

All experimental tests reported herein were conducted in

the Walter Beech Memorial Wind Tunnel at Wichita, at a Reynolds

number of 2.2 x 106 and a Mach number of 0.13. NASA tests of

this airfoil at higher Reynolds number and Mach number have

been reported in reference 2.

SYMBOLS

The force and moment data have been referred to the .25c

location on the flap-nested airfoil. Dimensional quantities are

given in International (SI) Units. Measurements were made in

U.S. Customary Units. Conversion factors between the various

units may be found in reference 3. The symbols used in the

present report are defined as follows:

c Airfoil reference chord (flap-nested)

c d Airfoil section drag coefficient, section drag/(dynamic pressure x c)

cf

ch

c£

cm

Cma

Cmf

Cn

Cna

Cnai

Cnf

CpAh

p

RN

x

z

6a

_f

_s

Flap chord

Control surface hinge moment coefficient, section

moment about hingeline/(dynamic pressure x controlsurface reference chord 2)

Airfoil section lift coefficient, section lift/

(c x dynamic pressure)

Airfoil section pitching moment coefficient with

respect to the .25c location, section moment/

(c 2 x dynamic pressure)

Airfoil forward section moment coefficient, moment

about leading edge/(c 2 x dynamic pressure)

Flap moment coefficient, section moment about leading

edge/(c 2 x dynamic pressure)

Airfoil normal force coefficient, section normal

force/(c xdynamic pressure)

Airfoil forward section normal force coefficient,

section normal force/(c x dynamic pressure)

Aileron normal force coefficient, section normal

force/(c x dynamic pressure)

Flap normal force coefficient, section normal force/

(c x dynamic pressure)

Coefficient of pressure, (p- p )/dynamic pressure

Spoiler projection height normal to local airfoil

surface

Static pressure

Reynolds number

Coordinate parallel to airfoil chord

Coordinate normal to airfoil chord

Angle of attack, degrees

Increment

Rotation of aileron from nested position, degrees

Rotation of flap from nested position, degrees

Rotation of spoiler from nested position, degrees

Subscripts:

a Aileron

f Flap

p Pivot

Remotefree-stream value

TESTMETHODS

Instrumentation, test procedure, tests facility and data

correction methods have been described in reference 4.

Resolution values for the various instrumentation systems

are given in Table i.

Table 1 - Instrumentation Resolution

Measurement

lift (force balance)

drag (wake survey)

drag (force balance)

pitching moment

(force balance)

hinge moment

pressure transducers

dynamic pressure

angle of attack

flap and aileron angles

spoiler angle

flap longitudinal and

vertical settings

Resolution

Dimensional Form

±0.9N

±0.06N

±0.2N

±0.1N-m

±0.02N-m

±4.SN/m 2

±4.SN/m 2

±0.05 °

±0.5 °

±0.25 °

±0.6mm

Coefficient Form

±0.001 (Ac£)

±0.00009(Ac d)

±0.0003 (Ac d)

±0.0003 (Ac m)

±0.006 (Ac h )

±0.004 (ACp)

±O.O01c

MODELDESCRIPTION

The MS(I)-0313 airfoil section is a 13%maximumthickness

section designed for high cruising efficiency at medium (=.72)

Machnumber. Model geometric details are given in figure i.

For tests in the WSUtwo-dimensional facility, models were sized

with 91.4 cm span and 61.0 cm chord. The 20%chord aileron was

designed with a 0.5% chord leading edge gap. The 25%slotted

flap was designed with an airfoil forward section terminating at

87.5% chord. The 10%spoiler was arranged in a slot-lip config-

uration with the 25%slotted flap. The model was fitted with

2.5 mmwide transition strips of #80 carborundum grit located

at 5%chord on the upper surface and at 10%chord on the lower

surface. The more aft grit location on the lower surface was

selected to place the transition strip aft of the stagnation

point for high flap deflection conditions.

RESULTS AND DISCUSSION

Presentation of Results

Test results and comparisons with theory and other experi-

mental results are shown in the figures as listed in Table 2.

Table 2 . List of Figures

Configuration

airfoil, aileron,

flap and spoiler

basic section

basic section

basic section

20% aileron

20% aileron

20% aileron

25% flap

25% flap

25% flap

25% flap

25% flap

25% flap

25% flap

10% spoiler

10% spoiler

Type Data

model geometry

c£,cd,Cm

pressures

tufts

c£,cd,Cm

Ac£,ACd,ACm,C h

pressures

optimum flap settings

CZmax contours

c£,cd,c m

flap effectiveness

experimental pressures

pressures

tufts

effect of spoilers on

lift for various flap

settings

incremental spoilereffectiveness and

hinge moments

Comparisons

data of Ref. 2

theory

theory

GA (W)-2

theory

Figure

1

2

3

4

5

6

7

8

9

I0

ii

12

13-16

17-20

21

22

Discussion

Flap Nested: (figures 2 through 4). Comparisons of WSU

data with NASA data show that the lift and pitching moment data

agree quite well, even including stalling effects. The drag

data do not compare as well. The agreement is good at low lift

coefficients, but the WSU tests indicate somewhat higher drag

levels at moderate lift coefficients. At near-stalling lift

coefficients, the data again show reasonable agreement.

The pressure distributions show good agreement with the

theoretical methods of reference 5 at angles of attack below

separation. Separation predictions agree rather well with ex-

periment, with separation appearing first at the trailing edge,

and gradually progressing forward. At high angles of attack

with massive separation, the discrepancies between experimental

and theoretical pressure distributions are large.

Table 3 shows a comparison of this airfoil with the GA(W)-2

airfoil of reference 6.

Table 3 - Comparison of Section Properties

(RN = 2.2x 106 , Mach = 0.13)

thickness/chord

c£ @ _ = 0 °

cm @ _ = 0 °

c d @ _ = 0 °

C£max

GA(W)-2 (Ref. 6)

0.13

0.43

-0.107

0.0109

1.67

MS(I)-0313

(Present Tests)

0.13

0.31

-0.075

0.0100

1.66

These data show a reduction in c£ @_= 0° and corresponding re-

duction in c m @ a = 0 ° for the MS(I)-0313 airfoil, as expected

for the lower design lift coefficient. The C£max for the MS(1)-

0313 is essentially the same as for the GA(W)-2, in spite of the

reduction in design lift coefficient. The drag level for the

MS(I)-0313 airfoil is essentially the same as the GA(W)-2 at

the Reynolds number and Mach number of the present tests.

20% Aileron: (figures 5 through 7). Aileron characteristics

for this airfoil are quite similar to the LS(I)-0421, GA(W)-I,

and GA(W)-2 airfoils (given in refs. 4 and 6-9). Control effective-

ness is somewhat non-linear but positive for all angles below

stall. Integrations of pressure distributions are tabulated to

provide individual component normal force coefficients for struc-

tural design purposes.

25% Flap: (figures 8 through 20). Optimum flap settings

are quite similar to other airfoils (such as refs..4 and 6-9).

The theory of reference 10 under-predicts the lift for i0 ° flap,

and over-predicts the lift for larger flap deflections and for

zero flap. As reported earlier (ref. 4), the reasons for these

trends are not understood. The highest CZmax obtained for this

airfoil-flap combination was 3.00, at a 30 ° flap deflection, com-

pared to the value of 3.35 obtained with the GA(W)-2 airfoil with

25% chord flap deflected either 35 ° or 40 ° (ref. 6).

The flap effectiveness data show that the increments in

C£max for high flap deflections are substantially lower than the

increments in c£ at _ = 0 ° . The increments in C£max for the 25%

flap at high flap deflections with this airfoil are consistently

lower than increments obtained with the GA(W)-2 airfoil. Incre-

ments in C£maxwith 20%plain flap are essentially the sameas

obtained with the GA(W)-2 airfoil. Limitations in C£maxcan only

be understood by study of separation patterns.

Separated regions are observed from tuft photos and from

interpretation of surface pressure distributions. Separation

is evidenced in surface pressure distributions by two charac-

teristics:

a) Trailing edge pressure changes from Cp _ 0 to Cp = -0.i

to -0.2 when separation occurs.

b) Pressure becomes essentially constant from the trailing

edge forward to the point of separation.

Integrations of pressure distributions are tabulated to

provide individual component normal force coefficients. Com-

parisons of theoretical pressure distributions with experiment

show good agreement except for cases where regions of separation

are present. No theoretical results are shown for the case of

0 ° angle of attack with 10 ° flap deflection. With this geo-

metry the computer program failed to run, in spite of repeated

attempts and numerous checks of input geometry, flap nose geo-

metry smoothing, etc.

Pressure distributions and tuft surveys indicate that for

25%flap deflections of i0 ° and 20° , initial separation takes

place at the airfoil trailing edge, and movesprogressively for-

ward, while the flap flow remains attached. With 30° and 35° flap

deflections, the flow over the flap was separated from about mid-

flap chord aft for low angles of attack. At angles of attack near

C£max the flap flow was attached, but flap separation reappeared

rather quickly at angles just beyond C£max, along with separation

at the trailing edge of the main airfoil.

Studies of tuft photos and pressure distributions from the

GA(W)-2 tests (refs. 6 and 9) show quite different separation

characteristics than the MS(I)-0313 airfoil. The GA(W)-2 tuft

and pressure studies show attached flow over the flap at all

angles of attack for all flap deflections up to 30 ° . These dif-

ferences in boundary layer separation and surface pressure dis-

tributions are entirely consistent with the lower C£max performance

from force measurements of the MS(I)-0313 airfoil-flap combination.

The reduced flaps-down performance of the MS(I)-0313 airfoil

is somewhat surprising, since tests of the GA(W)-2 and MS(I)-0313

airfoils show that both airfoils achieve the same C£max without

flaps. Reference 2 confirms that the unflapped airfoils have the

same C£max at a Reynolds number of 2 x 106 , but it should be noted

that at higher Reynolds numbers the GA(W)-2 has higher C£max than

the MS(I)-0313. The thickness distributions of these airfoils are

nearly identical, so the principal difference between the airfoils

is a reduction (average reduction : 25%) in camber of the MS(I)-0313.

This reduction in airfoil section camber reduces flow turning angles,

particularly in the region of 75%to 85%chord. This region forms

the flap leading edge camber and the camber of the important flap

slot lip. Comparable theoretical runs for the GA(W)-2 and

MS(I)-0313 airfoils with 30 ° flap deflection and experimental

optimum gap and overlap for each show that the GA(W)-2 should pro-

duce 0.14 higher c£, due to these added camber effects. The fact

that the experimental increment in C£max is 0.35 is evidently a

consequence of non-linear boundary layer behavior associated with

conditions near separation. Evidently the difficulties in attain-

ing attached flap flow for high deflection angles on the MS(I)-0313

airfoil are a consequence of this camber reduction.

10% Slot-Lip Spoiler: (figures 21 and 22). Spoiler con-

trol effectiveness and hinge moment characteristics are quite

similar to those observed for slot-lip spoilers on similar air-

foils in earlier research (refs. 4 and 6). Control effective-

ness with flap nested is positive and nearly linear for normal

angles of attack. At -8 ° angle of attack a lack of response

(deadband) appears for small deflections, but this negative

lift condition does not represent a realistic flight situation

for normal operations. Control effectiveness increases as the

flap is deflected, showing a strongly non-linear characteristic,

but without reversal or deadband tendency.

i0

CONCLUSIONS

i. The MS(1)-0313 basic section exhibits lift and drag

characteristics similar to the GA(W)-2 section at RN= 2.2 x 106

and Mach = 0.13. Pitching moments are reduced somewhat due to

the reduced camber of the MS(I)-0313 section.

2. Aileron control effectiveness and hinge moments for

the MS(I)-0313 are similar to comparable parameters for the

GA(W)-2 section.

3. Incremental C£max performance of a 25% slotted flap

on the MS(I)-0313 section is somewhat lower than a similar

flap applied to the GA(W)-2 section. Incremental performance

of a 20% plain flap on this section is similar to a 20% plain

flap applied to the GA(W)-2 section.

4. The highest C£max for this airfoil flap combination

is 3.00 compared to 3.35 for the GA(W)-2 airfoil with a simi-

lar flap.

5. Slot-lip spoiler control effectiveness on the MS(I)-0313

section is non-linear but positive for normal angles of attack

and spoiler deflection angles. Spoiler incremental effective-

ness and hinge moment values are similar to comparable values

for the GA(W)-2 section.

ll

REFERENCES

1.

2.

Pierpont, P.K.: Bringing Wings of Change. Astronautics and

Aeronautics Magazine, October 1975.

McGhee, R.J., and Beasley, W.D.: Low-Speed Aerodynamic

Characteristics of a 13-Percent-Thick Medium-Speed Air-

foil Designed for General Aviation Applications. NASA

Technical Paper 1498, Aug. 1979,

3o

4°

Mechtly, E.A.: The International System of Units--Physical

Constants and Conversion Factors (Revised). NASA SP-7012,1969.

Wentz, W.H. Jr., and Fiscko, K.A.: Wind Tunnel Force and

Pressure Tests of a 21% Thick General Aviation Airfoil with

20% Aileron, 25% Slotted Flap and 10% Slot-Lip Spoiler.

NASA CR-3081, 1979.

5.

6.

Smetana, Frederick 0., Summey, Delbert C., Smith, Neill S., and

Carden, Ronald K.: Light Aircraft Lift, Drag, and Moment

Prediction - A Review and Analysis. NASA CR-2523, 1975.

Wentz, W.H. Jr.: Wind Tunnel Tests of the GA(W)-2 Airfoil

with 20% Aileron, 25% Slotted Flap, 30% Fowler Flap and 10%

Slot-Lip Spoiler, NASA CR-145139, 1977.

7.

8,

9°

Wentz, W.H. Jr., and Seetharam, H.C.: Development of a

Fowler Flap System for a High Performance General Aviation

Airfoil. NASA CR-2443, 1974.

Wentz, W.H. Jr., Seetharam, H.C., and Fiscko, K.A.: Force

and Pressure Tests of the GA(W)-I Airfoil with a 20% Aileron

and Pressure Tests with a 30% Fowler Flap. NASA CR-2833, 1977.

Wentz, W.H. Jr., and Fiscko, K.A.: Pressure Distributions

for the GA(W)-2 Airfoil with 20% Aileron, 25% Slotted Flap,

and 30% Fowler Flap. NASA CR-2948, 1978.

10. Stevens, W.A., Goradia, S.H., and Braden, J.A.: Mathematical

Model for Two-Dimensional Multi-Component Airfoils in Viscous

Flow. NASA CR-1843, July 1971.

12

UPPERSURFACEx/c

0.0000

0020

0050

0125

0250

0375

0500

0750

i000

1250

1500

1750

2000

2250

2500

2750

3000

3250

3500

3750

4000

4250

4500

4750

5000

5250

5500

5750

6000

.6250

6500

6750

7000

7250

7500

7750

8000

8250

8500

8750

.9000

.9250

.9500

.9750

1.0000

z/c

0.0010

0095

0151

0243

0345

0418

0474

0552

0606

0648

0682

0709

0733

0752

0767

.0780

.0789

.0796

.0801

0803

0803

0800

0795

0787

07770765

0748

0729

0706

0679

0649

0615

0577

0537

0494

0449

0402

0353

0303

0252

0201

0149

0098

0047

- 0005

LOWER

x/c

0.0000

.0020

.0050

0125

0250

0375

0500

0750

1000

1250

1500

1750

2000

2250

2500

2750

3000

3250

3500

3750

4000

4250

4500

4750

5000

5250

5500

5750

6000

6250

6500

6750

7000

7250

7500

7750

8000

8250

8500

8750

.9000

.9250

.9500

.9750

1.0000

SURFACE

z/c

0.0010

-.0063

- 0099

- 0153

- 0206

- 0244

- 0275

- 0323

- 0361

- 0392

- 0418

- 0440

- 0458

- 0473

- 0485

- 0494

- 0501

- 0506

- 0509

- 0511

- 0509

- 0505

- 0498

- 0488

- 0475

- 0459

- 0440

- 0418

- 0393

- 0364

- 0333

- 0300

- 0266

- 0231

- 0196

- 0160

- 0128

- 0098

- 0073

- 0051

- 0035

- 0026

- 0025

- 0035

- 0061

(a) Basic MS(I)-0313 Airfoil

Figure 1 - Geometry.

i3

OLq00

0 -,--I

,-4 0-,-I,<

Id_0 ,--I

bm.,-I

].4

.875c

_ .00125c

.... .7 _ _0c

.758c

Flap Upper Surface

x/c z/c

0.7500 -.0053

.7531 .0030

.7562 .0068

.7594 .0097

.7625 .0122

.7750 .0192

.7875 .0231

.8000 .0251

.8125 .0259

.8250 .0262

.8375 .0260

.8500 .0254

.8625 .0248

.8750 .0239

Nose Radius = .012c

Nose Radius Location

(x/c,z/c) = (0.7620,-0.0054)

Note: Remainder of flap contour

matches basic airfoil.

(c) 25% Flap Geometry.

Figure 1 - Continued.

15

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19

¢1 .,4

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C

I 0i Z

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O Experiment

Theory (Ref. 5 )

Theory predicts no separation.

i _ x/c

Figure 3 - Pressure Distribution for the Basic Section.

20

-I@

9

-8

-7

G

5

cP

(b) e = 8.4 °

O Experiment

--Theory (Ref.5)

Note: Theory predicts no separation.

-3

2

-i

@

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21

-i@

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4

3

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Note :

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O Experiment

--Theory (Ref. 5 )

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x/c = 0.89 (upper surface)

x/c


22

-i@

-9

-8

Note :

(d) e = 16.7 °

O Experiment

_Theory (Ref.5)



-7

8

-5

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-3

-2

-I

@


23

Note:

(e) a = 19.0 °

O Experiment

Theory (Ref. 5 )



@

®OQO00

000 °

x/c

Figure 3 - Concluded.

24

c_ = 0 ° _ = 4 °

c_ = 8 ° c_ = 10 °

(a) Low angles of attack

Figure 4 - Tuft Patterns with 25% Slotted Flap, 0 ° Flap Deflection.

25

= 12° = 14°

= 16° = 18°

(b) High angles of attack


26

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i!M_!!

iilgli!!HiiNii

!I!!il!l_hit_

itl] ....

tiitlittitl%tl-

iiii!ii!1i!I11II7

t _i JtHH!Iih_hi!ilt_ti__ii1iii!ttt_

iiiJtiiilit_ !iiti,+ii,, iiii!14,,"!!!iHil_NN._:,!H!iI%%t[lit ' " +<'_'i_ilihif::iiTi_

,_,,li!iii_i!7i!7

!ilhiilHi_

!iil

l_':llH!!l_;_-iiiii!iih'1_

(d) Hinge Moment.


_33'I

s

-8-

i

_7 _

(a) AILERON DEFLECTION = 0.0 DEGREES

MACH NO. = 0.13

6

REYNOLDS NO. = 2.2 x i0

-6-

C

P

SYMBOL ALPHA c cn . nal a

_' -8.0 0.03 -0.66

-3.9 0.05 -0.24

0.2 0.06 0.21

+ 4.2 0.08 0.64

X 8.3 0.08 1.02

-4

-i

1

"T[

!

t

I I

i 0 .8 t.0

C XI

Figure 7 - Pressure Distribution with 20% Aileron.

=

-s TJ

J(a) AILERON DEFLECTION = 0.0 DEGREES

MACH NO. = 0.13

REYNOLDS NO. = 2.2 x 106

SYMBOL ALPHA c cnai n a

12.4 0.09 1.38

14.4 0.i0 1.51

_ 16.4 0.14 1.55

+ 18.3 0.17 1.24

_" 20.3 0.17 1.06


35

-8_

-7

-6

Cp

-5

-4

(b) AILERON DEFLECTION = 5.0 DEGREES

MACH NO. = 0.13

REYNOLDS NO. = 2.2 x i0 _

SYMBOL ALPHA c cn • nal a

[_ -8.0 0.07 -0.48

0.2 0.i0 0.39

_ 8.3 0.ii 1.17

0 _0 x/c 1 . 0 •8 1 ,0


0

-8 TI

++i

(c) AILERON DEFLECTION = 10.0 DEGREES

MACH NO. = 0.13


cP


[] -8.0 0. I0 -0.34

_3 0.2 0.13 0.51

8.3 0.14 1.31

0 0 x/c 1 -0

I.......................... _.-------

f

(.


38 !I

-BT{

t

0

(b) AILERON DEFLECTION = 5.0 DEGREES

MACH NO. = 0.13


SYMBOL ALPHA c c

nai n a

Ca 12.4 0.12 i. 52

C9 16.3 0.20 1.44

18.3 0.21 i. 36

C

x/c


-_'37

I

C

P

(c) AILERON DEFLECTION = i0.0 DEGREES

MACH NO. = 0.13


SYMBOL ALPHA c

nai

[] i2.4 o.ls16.3 0.23

_ 18.3 0.23

Cn

a

1.64

1.53

1.33

F

I

•8 1 .0


C

P

-7 !I

I!

-61I

(d) AILERON DEFLECTION = 20.0 DEGREES

MACH NO. = 0.13



C_" -8.0 0.18 -0.14

0.2 0.20 0.73

8.3 0,.19 1.52

-4 T

-3

L

0 •0 x/c I ,0 .8 I .0


40 ¸

J

!

o

f

(d) AILERON DEFLECTION = 20.0 DEGREES

MACH NO. = 0.13

REYNOLDS NO. = 2.2 x 10 _

SYMBOL ALPHA c c

: nai n a

[] 12.4 0.20 1.84

16.3 0.27 1.61

18.3 0.27 1.39

x/c

i

i

J

I,8 I 0

Figure 7 - Continued. i

41

-8-

_TJ-

(e) AILERON DEFLECTION = 40.0 DEGREES

MACH NO. = 0.13

REYNOLDS NO. = 2.2 x I0 _

cP

i

SYMBOL ALPHA c c

nai n a

-7.9 0.24 0.24

0.2 0.26 1.09

•'_ 8.3 0.24 i. 83

T


42

(e) AILERON DEFLECTION = 40.0 DEGREES

MACH NO. = 0.13


SYMBOL ALPHA c c

nai .na

"[_ 12.4 0.24 1.95

16.3 0.29 1.67

•_ 18.2 0.29 i. 53

-2

-i

x/c

k

I

i .0 ._ i .0


i43

-7

-5

cP

(f) AILERON DEFLECTION = 60.0 DEGREES

MACH NO. = 0.13


SYMBOL ALPHA C c

nai n a

-7.9 0.28 0.59

0.24 0.29 i. 38

_'_ 8.4 0.28 2.05

0

,_3 L .0


44

-7_

(f) AILERON DEFLECTION = 60.0 DEGREES

MACH NO. = 0.13


SYMBOL ALPHA c cn . nel a

12.3 0.30 2.03

16.2 0.33 1.70

18.2 0.34 1.65"

T

I I, , l_

t .O ,_ I .0


45

-8-

Cp

_4

2

0

(g) AILERON DEFLECTION = -5.0 DEGREES

MACH NO. = 0.13


SYMBOL ALPHA c cn . nal a

-7. 8 -0.02 -0.84._ 0.2 0.01 -0.00_'_ 8. 3 0.05 0.84

0.0x/c

TI

I

[ .0 .8 [ ,0


=

_ 46

-6

cp

-5

-4

-3

-2

-[

0

(g) AILERON DEFLECTION = -5.0 DEGREES

MACH NO. = 0.13


SYMBOL ALPHA c c

nai n a

_ 12.4 0.06 1.20

16.3 0.13 1.31

_ 18.3 0.16 1.22

#------_----4

0 •0 xlc t .0 .B t .0


147

cP

-5

0

I

0.0

C

(h) AILERON DEFLECTION = -i0.0 DEGREES

MACH NO. = 0.13


SYMBOL ALPHA c n cai na

-7.9 -0.06 -0.99

tL' 0.2 -0.04 -0.21

_'_ 8.3 -0.01 0.65

x/c


48_

-ST

--7

(h) AILERON DEFLECTION = -I0.0 DEGREES

MACH NO. = 0.13


cP


[] " 12.4 0.02 i. 04

rD 16.4 0.07 1.26

,'_ 18.2 0.12 i. 29

0

x/c

,r

"7


149

cP

-[!

cP

-t0 I

-9 i

0 0

(i) AILERON DEFLECTION = -20.0 DEGREES

MACH NO. = 0.13


0.0

SYMBOL ALPHA c c

nai n a

[_ -8.0 -0.ii -1.25

0.2 -0.10 -0.43•'_ 8.3 -0.08 0.39

2o [X/C

x/c [ .0 ,8 t .0


_3

50_

-S

-7

-5

cp

(i) AILERON DEFLECTION = -20.0 DEGREES

MACH NO. = 0.13


SYMBOL ALPHA c cn , nal a

12.3 -0.07 0.76

16.4 -0.04 1.02

,_ 18.4 0.01 1.08

0

C

x/c


51

(j) AILERON DEFLECTION = -40.0 DEGREES

MACH NO. = 0.13


I I

Z.0X/C

T

I

Figure 7 - Continued. I

52]

-S+

-7

(j) AILERON DEFLECTION = -40.0 DEGREES

MACH NO. = 0.13



12.3 -0.17 0.36

16.4 -0.i0 0.81

_ 18.4 -0.07 1.05

0

0.0 x/c t +0 .8 t .0

Figure• +

7 - Continued.

i

153I

cP

4

c

(k) AILERON DEFLECTION = -60.0 DEGREES

MACH NO. = 0.13

REYNOLDS NO. = 2.2 x 10 _

-I I I

0 .0 x/c


54_

0

L0.0

<

(k) AILERON DEFLECTION = -60.0 DEGREES

MACH NO. = 0.13


SYMBOL ALPHA c c

nai n a

12.3 -0.22 -0.06

_" 16.3 -0.18 0.39

,_ 18.4 -0.14 0.62

T!l

ii

i

I I

x/c .8


_55,

U

U

0

0

0

f /

I

0

0 0

56

.4-1

.._

4-1

0

I

=,-t

N

U

Airfoilc£ max = 2.60

z/c

.01

.02

O3

04

05

06

07

O8

/I"

f

/J J

/ ¢ //_/_ 2_'-5-_ ,. Flap

_/ 2.58

09

.04 .03 .02 .01 0

x/c

-.01 -.02 -.03 -.04 -.05

Note: Contours are for locus of flap nose point, (see p.56).

(a) i0 ° Flap Deflection

Figure 9 - C£max Contours.

I--

_57

z/c

01

02

03

O4

O5

06 --

07

.08

.09•04 .03 .02 .01 0

x/c


(b) 20 ° Flap Deflection

Figure '9 - c£ max Contours•

58i

Airfoil

.0

.0_

.0._

.04

z/c.0_

.06

I

./ill,2.80

2.90

2.95

2.90_ 2.0

2.80 /2.40

max=3.0

"- Flap

.07

• 08!

.0S

.04 .03 .02 .01 0

x/c

-.01 -.02 -.03 -.04 -.05


(c) 30 ° Flap Deflection

Figure 9 - c£ max Contours•

59

Airfoil

01 _

02

03

O4

z/c

05

06

07

max= 2.97

_2.95

/,-

/ /".... 2.3 / _.

/2.1

Flap

%

08

09

.04 .03 .02 .01 0

x/c

-.01 -.02 -.03 -.04 -.05


(d) 35 ° Flap Deflection

Figure 9 - c£ max Contours.

60 _

%,g

............... I:i :::: ...... :_._::-::L t..... :.: ' L_

i::Tl Note: 1. Flagged symbols denote.... !_ theoretical values using

i_!!:!i_i2:±I . :-i: i:F!_-:-_ the method of Ref. 5.i;i_:_) _ " 2. Dashed symbols denote

........,/-i -.I: -_ [ _:::l_--Ii_:_:_-17-27:/I !_?I-_ _- -_ " thetheoreticalmethodofvaluesRef. 1o.USing

(a) Lift. -- ............_-_

Figure 10- 25% Slotted Flap Performance, .: . .......... _!_:!!!

i:22;!:!:__iii_iii_T61I

i

Note: i. Flagged symbols denote theoretical

values using the method of Ref.5.

I:2. Dashed symbols denote theoretical

values using the method of Ref,lO. '_

.__ ! i _ _ i 'I]' i -: li ' i ......... i,F :- Symbol 6f Xp/c Zp/c

-IT-- [_ 0 ° 0.0 " 0.0

i • , _ .j • Ii0 ° 0.126 0.052

20 ° 0.136 0.039

30 ° 0.135 0.027

!

621

!

-+-

- - i ........ "_:: ': ._ • _1 " : I: : :: li!i i i I!_

::.... _u_ .,_ _:- : :t_!: _ :_:-

:LE :: .t :. 4a _ _: ::......... '* ...

:::t ..........

:!'i :i!! I:I _) 0 • ili : i :- ' ,:il:: --,.

_:-_ ' : _ ,--I :: l:::: I_

...... i ........ i--1 :::1 _ > ........: : ..... .... • .:t ' :';: --_--4

! -,-,.

'i',,:tF: ::-: -"....

::.t" _'1 i .ta _ C

i . I _.):_ ::_:: z ..... i_!il : _- i| o

• J_ :..I ::_ :::.I . , . .

::: it" : t,ILL _:_ .... ::_LI:L_Lf::: i:: =_..... ' [_-_::::1 : I::.: " "

LL_: _

_'_ '_ _ _'_r '_" ''

iC :i_i_ :ii ....i'G:i ,,

_2_12:., i 12ti 2:2i11_1:2 !iii ,,_

i

64

(a) FLAPDEFLECTION= 0.0 DEGREES

-7

I

-6J-

C

P

-4

MACH NO. = 0.13

REYNOLDS NO.= 2.2 x 106

SYMBOL

+

X

-3 1_'__-Z

0

i0.0

ALPHA c cn ma a

-8.0 -0.65 0.12

-4.0 -0.21 0.01

0.0 0.26 -0.ii

4.0 0.69 -0.21

8.1 1.09 -0.30

c cnf mf

0.03 -0.05

0.05 -0.10

0.07 -0.13

0.08 -0.16

0.08 -0.17

i!

i

Figure 12 - Pressure Distribution with 25% Slotted Flap.

i65

(a) FLAPDEFLECTION= 0.0 DEGREES

MACH NO. = 0.13

REYNOLDS NO.= 2.2 x 106

SYMBOL ALPHA c c cn a m a nf

12.1 1.42 -0.37 o. I0

14.1 1.50 -0.37 0.12

16.0 1.51 -0.37 0.14

+ 17.8 1.21 -0.37 0.20

_ 19.8 1.07 -0.33 0.20

(.

x/c


66_

(b) FLAPDEFLECTION= 10.0 DEGREES

-S T MACH NO. = 0.13

REYNOLDS NO. = 2.2 x l0 s

-6

C

P

i

SYMBOL ALPHA c c c c

n a m a nf mf

-8.0 -0.38 0.04 0.24 -0.36

0.2 0.58 -0.21 0.30 -0.44

8.3 1.55 -0.45 0.34 -0.49

-4

-3

-I

0

i0 •0 x/c i .0 .8 I .0

Figure 12- Continued........ -i

i67

cP

(b) FLAP DEFLECTION = i0.0 DEGREES

-3-

-6

MACH NO. = 0.13


%

SYMBOL ALPHA c c

n a m a

12.4 1.98 -0.56

_ 16.2 1.68 -0.54

18.2 1.38 -0.48

c

nf

0.35

0.48

0.52

L __ x/c L 0


681

c

mf

-0.50

-0.70

-0.77

(C) FLAP DEFLECTION = 20.0 DEGREES

Ji

-7 1

I

P

-4 1

i 'i

MACH NO. = 0.13


SYMBOL ALPHA

[] -8.0

0.2

L%8.3

c cn ma a

0.01 -0.11

1.02 -0.37

1.96 -0.62

1O,O' . x/c. i ,0


c cnf mf

0.38 -0.53

0.42 -0.57

0.42 -0.57

]

• 8 i. 0

'69

-6L!

tI

-[I._

I

P

-9

(c) FLAP DEFLECTION = 20.0 DEGREES

MACH NO. = 0.13


7

SYMBOL ALPHA c cn m

. a a

[] 12.4 2.37 -0.72

16.2 1.59 -.59

18.2 1.46 -0.53

c cnf mf

0.42 -0.57

0.48 -0.69

0.55 -0.83

f

--.._


70_

i

°icP

-4 ¸

0

(d) FLAP DEFLECTION = 30.0 DEGREES

MACH NO. = 0.13


SYMBOL ALPHA

-8.01-'4

m_. 0.2

± 8.3

c C c

n a m a nf

0.48 -0.30 0.46

1.42 -0.54 0.43

2.29 -0.76 0.40

x/c

c

mf

-0.62

-0.62

-0.58


_i71

(d) FLAP DEFLECTION = 30.0 DEGREES

MACH NO. = 0.13


ALPHA cna

12.4 2.67

16.3 1.75

18.2 1.73

c m Cnf Cmfa

-0.85 0.43 -0.60

-0.66 0.33 -0.60

-0.66 0.35 -0.62

f_

x/c

Figure 12 - C oDtinued.

72!

-s T

(e) FLAP DEFLECTION = 35.0 DEGREES

MACH NO. = 0.13


SYMBOL ALPHA cna

[_ -7.9 0.59

0.2 1.52

8.3 2.41

Cma Cn f Cmf

-0.35 0.46 -0.61

-0.58 0.42 -0.61

-0.81 0.42 -0.59

\

x/c

Figure 12- Continued.

_73

0

(e) FLAP DEFLECTION = 35.0 DEGREES

MACH NO. = 0.13


SYMBOL

[]

C 0 Z -0

x/c

ALPHA c c cn a m a nf

12.4 2.74 -0.88 0.43

16.2 1.80 -0.68 0.36

18.2 1.76 -0.68 0.38

Ti

I

D

.0

cmf

-0.60

-0.65

-0.68

I


-10

-9

-8

-7

-6

-5

Cp

-q

-3

-2

-]

@

Note:

(a) s = 8.4 °

O Experiment

Theory (Ref.10)

(i) Theory predicts separation at

x/c = 0.81 (lower surface)

(2) No confluent boundary layererror encountered.

OQQQ

-3

-2

-1

' :.'oo0

i

13°

Figure 13 - Pressure Distributions with 25% Slotted Flap,

10 ° Flap Deflection.

(b) _ = 12.7 °

O Experiment

Theory (Ref.lO)

Note: (i) Theory predicts separation atx/c - 0.81 (lower surface)


(2) No confluent boundary layererrror encountered.

C

P

2

-i0

-9

8

-7

-G

-5

Cp

-3

-2

-I

@

@

0

0000

(C) e = 17.2 °

O Experiment

Theory (Ref.10)

Note: (i) Theory predicts separation at -

x/c = O. 79 (upper surface)


' 0.'20

x/c

00000

x/c


77

-I0

9

-8

Cpmin

= -18.1

(d) _ = 20.i °

O Experiment

Theory (Ref.10)

Note: (I) Theory predicts separation at



0.20 (_)

x/c O

781

-I0

-9

-8

-7

-G

-5C

P

-3

-2

-i

@

(a) e = 0.2 °

O Experiment

Theory (Ref.10)

Note: (i) Theory predicts no separation.


-3

Q

OOOo

1 x/c 1

-2

@

Figure 14 - Pressure Distributions with 25% Slotted Flap,20° Flap Deflection.

-I@

9

-8

(b) u = 8.7 °

O Experiment

-- Theory (Ref.10 _)

Note: (i) Theory predicts separation at



7

G

-5

Cpl

-y

-3

2

-i

-3

-2

1


03O

-I0

-9

8

7

-@

S

cP

-4

-3

2

i

0

15

14

cP

-13

(c) _ = 13.3 °

O Experiment

Theory (Ref.10)



x/c

I

0.20

-3

2

1

1.'oo 0 3o

1Figure 14 - Continued.

81

I0

-9

-8

-7

-@

-5

C

P

3

-2

-i

@

-16

O

c = -19.7pmin

(d) e = 18.3 °

O Experiment

Theory (Ref.10)





0,20

-3

-2

-I

x/c

Figure 14- Concluded.

@

O

%

I

,30

82

-I0

9

8

7

-6

5

C

P

3

2

I

@

(a) e = 0.1 °

O Experiment

_Theory (Ref_10)



-3

% 2O o

I

q-:.'oo0

x/c

Q

30



83

-I0

9

-8

7

@

S

c

-4

-3

2

-i

0

-14

Cp !

-13

12

(b) u = 9.0 °

O Experiment

_Theory (Ref.10)




i

0.50

xlc


2

84

-]0

-9

-8

-7

-5

-S

cP

-3

-2

-i

@

cP

(C) _ =13.4 °

O Experiment

---- Theory (Ref.10)


x/c = 0.13*(lower surface)

(2) No confluent boundary layer

error encountered.

* Near transition point

x/c 1Figure 15- Continued.

13 0

85

cpm.in

(d) e = 18.3 °

O Experiment

Theory (Ref, 10)



(2) No confluent boundary layer

error encountered.

-3

-2

-i


0I. O0 .30

86

-I0

-9

-8

-7

5

-S

cP

(a) e = 0.6 °

O Experiment

-------Theory (Ref.10)

Note: (i) Theory predicts separation atx/c = 0.83 (lower surface)

(2_ Confluent boundary layererror encountered.

-3

-2

1

@

0

x/c



87

-I0

8

-8

7

-G

S

cP

-q

3

-2

1

0

Note:

0.50

x/c

(i)

(2)

(b) u = 9.2

O Experiment

-- Theory (Ref.10)

Theory predicts no separation.

Confluent boundary layererror encountered.

(

.30


88

-I0

9

-8

-7

G

-S

cP

4

3

-2

-I

0

cpmin

(c) u = 13.6 °

O Experiment

-- Theory (Ref.10)



(2) Confluent boundary layererror encountered.

x/c


89

-3

-2

-I

0I

I 00 3O

-5

cP

-3

Cpmin-26.5

Note: (I)

(2)

(d) _ = 18.6 °

O Experiment

Theory (Ref.10)




NO confluent boundary layer

error encountered.

-3

-2

-i

0

0

q_)OOOOOOOOOOq O0

I i l I I I t I

0 O0 x/c

I

1,00


-2

i

00.30

90

o, = 0 o c_ = 4 °

= 8 ° _ = i0 °

(a) Low Engles of attack

Figure 17- Tuft Patterns with 25% Slotted Flap, I0 ° Flap Deflection.

91

c_= 12° = 14°

o f>

= 16 ° c_ = 18 °

(b) High angles of" attack


92

c_ = 0 ° c_ = 4 °

...q

C_ = 8 ° C_ = 12 °


Figure 18- Tuft Patterns with 25% Slotted Flap, 20 ° Flap Deflection.

93

o_

= 14 °= 16 °

a = 17 °

= 18 °



94

c_ = 0 o c_ = 4 °

o :c>

CL = 8 °= I0 o



95

= 12° _ = 14 °

= 15 ° = 16 °



96

c_ = 0 ° a = 4°

= 8 ° = i0 o



97

e_= 12° _ = 14°

= 15° e = 16°



98

Figure 21- Effects of Spoiler Deflection on Lift for 25%Flap.99

100

Symbol

[]

+

X

II_LI

i:ilf_

t-

;'4.

hl-

2d+J..

!'G,

'¢_-

L_:_

!hE!!qt_-_

UP

_h

iii!H!!

ill+.T

-Lii/

"2[:

2.5 ° I F-:- _ T:::-;.! [ .7- :[:_T: ::_]T i'T_q;r:T: :-]T_:_I-: :FF::-_--_----I :: -_ -;_: T/[T_- ---_ : : :- " _ l-": " :: :_ _I: _- :.... " ............ _,:_'_:--

10 ° _: .... i ..... _ :1

20o : ..... ':_:i,, ......:. " .::=:: ::'::' _ ....... ............ :_ ' :::_,_i::_

, _. _, _ q ......::i !::?T r :!:!]:: :i :: _!:;TF-{il

60° : ':: :_-t ii [[:: ]]'L!;!LLI! 4: .: :: f ,_

x_: r.. : :: _:1 _ " "...... :"': .....: :t!i 1_I_: :: :_,,_i,_ ::: '_ ::!: :: -::

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1, Report NO.

NASA CR-3439

2, Government Accession No.

4. Title and Subtitle

WIND TUNNEL FORCE AND PRESSURE TESTS OF A 13%

THICK MEDIUM SPEED AIRFOIL WITH 20% AILERON,25% SLOTTED FLAP AND 10% SLOT-LIP SPOILER

7. Author(s}

W. H. Wentz, Jr.

9, Performing Organization Name and Addre_

Wichita State University

Wichita, Kansas 67208

12. S_nsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, D.C. 20546

15 Supplementary Notes

Langley Technical Monitor:

Topical Report

Robert J. McGhee

3. Recipient's Catalog No,

5. Report Date

JUNE 1981

6. Performing Organization Code

8. Performing Orgamzation Report No,

WSU AR 78-4

10. Work Unit No.

1. Contract or Grant No.

NSG-II65

13. Type of Report and Period Covered

Contractor Report

14 Sponsoring Agency Code

505-31-33-05

16. Abstract

Force and surface pressure distributions have been measured for a 13%

medium speed (NASA MS(1)-0313) airfoil fitted with 20% aileron, 25%

slotted flap and 10% slot-lip spoiler. All tests were conducted in the

Walter Beech Memorial Wind Tunnel at Wichita State University at aReynolds number of 2.2 x 106 and a Mach number of 0.13. Results

include lift, drag, pitching moments, control surface normal force and

hinge moments, and surface pressure distributions. The basic airfoil

exhibits low speed characteristics similar to the GA(W)-2 airfoil.

Incremental aileron and spoiler performance are quite comparable to

that obtained on the GA(W)-2 airfoil. Slotted flap performance on

this section is reduced compared to the GA(W)-2, resulting in a

highest CZmax of 3.00 compared to 3.35 for the GA(W)-2.

17. Key Words (Suggested by Author(s))

Airfoil designControl surface

Pressure distributions

Aerodynamic forces

Flap

19. S_urity Cla_if. (of this report] 20. Security Classif. (of this _ge)

Unclassified Unclassified

18, Oistribution Statement

,- , __ LNn- n - ,J v

21. No. of Pages 22. Price

116

.... ' - " -.... : ..... '"--- li_..... 1111-I

NASA-Langley, 1981

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