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NASA Contractor Report 2948

A

Pressure Distributions for. the GA( W)-2 Airfoil With 20% Aileron, 25% Slotted Flap and 30% Fowler Flap

W. H. Wentz, Jr., and K. A. Fiscko

GRANT NSG-1165 FEBRUARY 1978

FOR EARLY DOMESTIC DISSEMINATION

Because of its significant early commercial potential, this information, which has been developed under a U.S. Government program, is being disseminated within the United States in advance of general publication. This information may be duplicated and used by the recipient with the express limitation that it not be published. Release of this information to other domestic parties by the recipient shall be made subject to these limitations.

Foreign release may be made only with prior NASA approval and appropriate export licenses. This legend shall be marked on any reproduction of this information in whole or in part.

February 1980 Date for general release .-.__-.

https://ntrs.nasa.gov/search.jsp?R=19810010497 2018-04-26T16:39:01+00:00Z

TECH LIBRARY KAFB, NM

NASA Contractor Report 2948

Pressure Distributions for the GA( W)-2 Airfoil With 20% Aileron, 25% Slotted Flap and 30% Fowler Flap

W. H. Wentz, Jr., and K. A. Fiscko

Wichita State University Wichita, Kansas

Prepared for Langley Research Center under Grant NSG-1165

National Aeronautics and Space Administration

Scientific and Technical Information Office

1978

FOR EARLY DOMESTIC DISSEMINATION

Because of its significant early commercial potential, thi:s information, which has been developed under a U.S. Government program I is being disseminated within the United States in advance of general publicat.ion. This information may be duplicated and and used by the recipient with the express 1imitatio:n that it not be published. Release of th%s information to other domestic parties by the recipient shall be made subject to these limitations. Foreign release may be made only with prior NASA approval and appropriate exp0r.t licenses. This legend shall be marked on any reproduction of this information in whole or in part. Date forgeneral release: --.___ February 1980.

iii

SUMMARY

Surface pressure distributions have been measured for the 13% thick GA(W)-2 airfoil section fitted with 20% aileron, 25% slotted flap and 30% Fowler flap. All tests were conducted at a Reynolds number of 2.2 x lo6 and a Mach number of 0.13. Pressure distribution and force and moment coefficient measurements are compared with theoretical results for a number of cases. Agreement between theory and experiment is generally good for low angles of attack and small flap deflections. For high angles and large flap deflections where regions of separation are present, the theory is inadequate. Theoretical drag predictions are poor for all flap-extended cases.

INTRODUCTION

This report documents experimental surface pressure distributions for the GA(W)-2 airfoil section fitted with 20% aileron, 25% slotted flap, and 30% Fowler flap. Pressure distributions and aerodynamic characteristics of the basic GA(W)-2 airfoil section have been reported earlier (ref. 1). Wind tunnel force measurements of the airfoil with high-lift and control devices including optimizations of flap settings have been conducted at WSU, and the results of that research have been reported in ref. 2.

Theoretical computer calculations of pressure distributions using the methods of refs. 3 and 4 are presented for a number of cases. The purpose of the present research is to determine actual pressure distributions for the new airfoil with high-lift and control devices, and to compare both experimental pressure distributions and overall aerodynamic force and moment results with theoretical predictions.

SYMBOLS

Dimensional quantities are given in both International (SI) Units and U.S. Customary Units. Measurements were made in U.S. Customary Units. Conversion factors between the various units may be found in ref. 5. The symbols used in the present report are defined as follows:

C Airfoil reference chord (flap nested)

cd Coefficient of drag, section drag/(c x dynamic pressure)

CL? Coefficient of lift, section lift/(c x dynamic pressure)

cm Pitchin 9

moment coefficient, section moment about . 25c/(c x dynamic pressure)

cma Airfoil forward section moment coefficient, moment about leading edge/(c2 x dynamic pressure)

Cmf Flap moment coefficient, moment about leading edge/

(c2 x dynamic pressure) cna Airfoil forward section normal force coefficient,

normal force/ (C x dyn-amic pressure)

Cnai Aileron normal force coefficient, normal force/ (c x dynamic pressure)

Cnf Flap normal force coefficient, normal force/ (c x dynamic pressure)

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

%! Flap cove length

P Pressure

X Coordinate along airfoil chord

Z Coordinate normal to airfoil chord

a Angle of attack, degrees

6 Xotation of control surface from nested position, degrees. (Trailing edge down is positive.)

Subscripts

a Airfoil forward element

ai Aileron

f Flap

P Pivot point for flap

co Free-stream conditions

APPARATUS AND TEST METHODS

Model Description

The GA(W)-2 airfoil section is a 13% maximum thickness airfoil section derived from the 17% GA(W)-1 section (ref. 6). For tests in the WSU two-dimensional facility, models are sized with 91.4 cm (36 inch) span and 61.0 cm (24 inch) chord. All models were equipped with 1.07 mm (0.042 inch) diameter pressure taps.

2

Model geometric details and flap pivot locations are given in figure 1.

Instrumentation

Pressure measurements were made using as many as 96 pressure channels multiplexed into 4 pressure transducers through a series of pressure switches. The unbonded-strain gage type transducers are connected to precision digital strain indicators for conversion from anal.og to digital data. The digital data are recorded

on punch cards for off-line processing through the WSU Digital Computing Center. System resolution is +2.4 newtons/meter2 (0.05

psf) which corresponds to +0.2% of dynamic pressure for the pre-

sent tests. Figure 2 shows a pressure measurement schematic.

Test Procedure

All tests were conducted at a Reynolds number of 2.2 x 10 6

and Mach number of 0.13. Transition strips consisting of 2.5 mm (.lO inch) wide strips of #80 Carborundum grit were applied to the upper surface at 5% chord and to the lower surface at 10% chord. All pressure data have been converted to coefficient form. Tunnel dynamic pressure has been corrected for solid and wake blockage, and model angle of attack has been corrected for induced effects, using the linear correction methods of ref. 7. Surface pressure measurements were integrated numerically to calculate component normal force coefficients, and moment coefficients about the component leading edge or hingeline.

Wind Tunnel

The WSU Walter Beech Wind Tunnel is a closed return tunnel with atmospheric test section static pressure. With two-dimensional inserts installed the test section is 0.91 m x 2.13 m (3 ft x 7 ft). Complete description of the insert and calibration details are given in ref. 8.

3

THEORETICAL METHODS

For selected cases, studies have been conducted to determine theoretical pressure distributions and overall force coefficients for comparison with the experimental measurements. These theoretical studies were conducted utilizing sophisti- cated computing routines which include boundary layer effects. For the flap-nested configuration, the computations utilized the programs of refs. 3 and 4. The principal difference between these programs is that ref. 4 includes a drag computation by the Squire-Young method (see ref. 91, but is restricted to single-element analysis. The flap-extended configurations were analyzed by the program of ref. 3, which is capable of analyzing multi-element configurations, but does not utilize the Squire-Young drag computation.

As discussed in ref.10, for improved simulation of pressures near the airfoil trailing edge of flap-extended configurations, the lower surface flow was assumed to separate at the entrance to the flap cove and reattach ahead of the slot lip. This technique for modeling involves using an effective cove shape derived by assuming a straight line from cove entrance to the 75% cove location, as shown in the sketch which follows:

Assumed Effective Shape

Sketch A - Flap Cove Theoretical Modeling

4

EXPERIMENTAL AND THEORETICAL RESULTS

Flap Nested

For the flap-nested configuration, data from the present tests are compared with NASA data from ref. 1 and theoretical results in figure 3. The two sets of experimental data show good agreement at all angles up to 14O. At 18", the NASA data show more scatter than the WSU data, indicating a somewhat less stable separation. Agreement between experiment and theory is good for cases with little or no separation (CI c 14O). For cases with separation ahead of .9 x/c, the theory is sub- stantially in error.

20% Aileron

Pressure distributions with 20% aileron with 0.5% gap are shown in figure 4 for aileron deflections from -60" to +60" and a nominal angle of attack range of -8O to 16". These data show trends very similar to the pressure distribution results reported earlier (ref. 10) for an aileron applied to the GA(W)-1 airfoil section. For large aileron deflections (&a 2 20°) separated flow on the suction side of the aileron is indicated, as evidenced by a region of constant pressure. Because separation is ordi- narily present for moderate and high aileron deflection, no theoretical studies were conducted with aileron.

25% Slotted Flap

Experimental force characteristics for optimum flap settings as reported in ref. 2 are shown in figure 5, along with theoretical force characteristics for selected cases. The experimental

cI1 vs- a. curve for 35" flap shows an increase in slope just prior to stalling, an indication of flow improvement just prior to mas- sive separation. During optimization force studies with 35' and 40" flap deflections, many non-linearities in force characteristics

5

were observed. This non-linear behavior is interpreted as evi-

dence of separation over some portion of the airfoil or flap at virtually every angle of attack for these flap deflections. It is seen that the improvement in cQ values between 30" and

35" flap is quite small, and the performance with 40" flap is essentially the same as 35" flap.

For the flap-nested case, the methods of refs. 3 and 4 are compared in figure 5. These data show that the Squire-Young drag computation routine of ref. 4 provides considerable improvement in drag prediction.

For angles of attack less than 10" and flap deflections less than 30", theoretical lift and pitching moment agree reasonably well with experiment. For higher angles of attack and flap deflections the agreement becomes progressively poorer. These trends are attributed to inadequate theoretical modeling for situations with nearly-separated or partially-separated boundary layers. The theoretical drag predictions are poor for all flap-extended cases. The theory is very inconsistent, even predicting a reduction in drag as flap deflection is in- creased in some instances.

Theoretical pressure distributions are presented in figures 6 through 9. The results of these theoretical studies compare quite favorably with experiment for angles of attack below stall, and lower flap deflections. For higher angles of attack or flap deflection angles, the agreement becomes progressively poorer.

Detailed experimental pressure distributions for this configuration for optimum flap settings are presented in figure 10. For flap deflections up to 20", the pressure data indicate attached flow for angles of attack up to 12'. For 30" flap deflection, flow separation on the flap is indicated at 16.2" angle of attack, and a large step in pressure is indicated on the airfoil upper surface. This condition is beyond cIlmax as indicated by the force data of figure 5. The apparent jump in pressure at about 20% chord is attributed to an unsteady flow situation with

6

intermittent separation, and is believed to be associated with the slow-scan method of pressure recording utilized for these tests.

For 35O flap, the flap flow tends to separate at the trailing edge at the lower angles of attack. At higher angles, the flap flow improves.

The situation with 40° flap is very similar to 35" flap. In this case, the flap is evidently fully attached only at 12.2O, and re-separates at 14.4". These observations for 35' and 40" flap correlate very well with trends observed in the force measurements.

30% Fowler Flap

Results of optimum force tests for this flap from ref. 2 are shown in figure 11, along with theoretical results for selected cases. As with 25% flap, the theory significantly over-predicts lift at high angles of attack and high flap deflections.

Close comparison of the flap-nested data from the 25% and 30% flap models (figures 5 and 11) shows that the 30% flap model provides slightly more lift at low angles than the 25% flap model, even though cQmax and stalling- angle are unchanged. The 30% flap model had a clean, continuous upper surface while the 25% model had a slight step at the spoiler trailing edge, and four spoiler hinges which created small .protuberances at four span-wise stations. The differences in aerodynamic data are attributed to these geometric variations.

Experimental pressure distributions are compared with theory in figures 12 through 15. Again, the principal dispari- ties occur at high angles and large flap deflections.

Detailed experimental pressure distributions for various flap settings are shown in figure 16. For flap deflections up to 30", the distributions indicate attached flow on both airfoil and flap for all but the highest angles of attack. For 35" and 40° flap deflections the distributions indicate separation

7

on the flap at most angles of attack. At the post-stall (u=12.3') condition, separation is indicated over the aft

half of the flap, and the airfoil forward section pressure shows a jump indicating an unstable pressure distribution as discussed with the 25% flap.

CONCLUSIONS

1. Pressure distributions for the GA(W)-2 airfoil with 20% aileron show trends similar to the GA(W)-1 airfoil with aileron.

2. Lift and pitching moment predictions from theory agree reasonably well with experimental measurements for cx less than loo and flap deflections less than 35'. For cases with nearly separated or partially separated boundary layers, present theories are inadequate.

3. Drag prediction using the Squire-Young formulation is adequate for single-element airfoils without separation. The multi-element drag computation is poor for all cases.

4. Pressure distributions for the GA(W)-2 airfoil with 25% and 30% flaps show good agreement with theory at low angles and small flap deflections, but poor agreement for high angle or large flap deflection cases which involve flow separation.

REFERENCES

1. McGhee, R.J.; Beasley, W.D. and Somers, D.M.: Low-Speed Aerodynamic Characteristics of a 13-Percent-Thick Airfoil Section Designed for General Aviation Appli- cations. NASA TMX-72697.

2. 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.

3. Stevens, W-A.; Goradia, S.H. and Braden, J.A.: Mathematical Model for Two-Dimensional Multi-Component Airfoils in Viscous Flow. NASA CR-1843, 1971.

4. NASA Langley Research Staff: Viscous Flow Single-Element Airfoil Analysis Computing Routine (Program AIRFOIL), 1976.

5. Mechtly, E.A.: The International System of Units - Physical Constants and Conversion Factors (Revised). NASA SP- 7012, 1969.

6. McGhee, R.J. and Beasley, W.D.: Low-Speed Aerodynamic Char- acteristics of a 17-Percent-Thick Airfoil Section De- signed for General Aviation Applications. NASA TND- 7428, 1973.

7. Pope, A. and Harper, J.J.: Low-Speed Wind Tunnel Testing, John Wiley and Sons, 1966.

8. Siew, R.J.: Calibration of a Two-Dimensional Insert for the WSU 7' x 10' Wind Tunnel, AR 73-2, Wichita State University, 1973.

9. Schlichting, H.: Boundary Layer Theory. McGraw-Hill Book co., Sixth edition, 1968.

10. Wentz, W.H., Jr.; Seetharam, H.C. and Fiscko, K-A.: Force and Pressure Tests of the GA(W)-1 Airfoil with a 20% Aileron and Pressure Tests with a 30% Fowler Flap. NASA CR-2833, 1977.

9

UPPER SURFACE LOWER SURFACE

x/c 0.0000

0020 :0050

0125 :0250 -0375 . 0500

0750 :1000 .1250 .1500 .1750 .2000 .2500 .3000 .3500 .4000 .4500 .5000 .5500 .5750 .6000 .6250 .6500 .6750 .7000 .7250 .7500 .7750 -8000

8250 :8500

8750 : 9000 .9250 .9500 .9750

1.0000

Z/C x/c 0.0000

.0103 0163

:0246 0336

:0400 .0451 .0528

0588 :0637 .0677

0712 :0742 .0788 .0820 .0840

0849 :0846

0833 :0807

0789 :0767

0739 :0708

0672 :0633

0591 :0545 .0497 .0447 .0395

034i :0285 .0228 .0170 .OllO .0049

-.0015

0.0000 .0020 .0050 .0125 .0250

0375 :0500 .0750 .lOOO .1250 .1500 .1750 .2000 .2500

3000 13500 .4000 .4500 .5000 .5500 .5750 .6000 .6250 .6500 .6750 .7000 .7250 .7500 .7750

8000 :8250 .8500 .8750 .9000 .9250 .9500 .9750

1.0000

Z/C

0.0000 -.0066 -. 0097 -.0144 -.0188 -.0223 -.0250 -.0294 -.0328 -.0357 -.0380 -. 0398 -.0415 -;0438 -. 0449 -.0452 -.0449 -.043? -.0417 -.0386 -.0362 -.0337 -.0307 -.0276 -.0243 -.0210 -.0175 -.0143 -.OllO -.0078 -.0051 -.0028 -.0012

0000 :0001

-.0007 -.0028 -.0071

(a) Basic GA(W)-2 Airfoil.

Figure 1 - Geometry.

10

5 = 0.0263~

r2 = 0.0313c

(b) 20% Aileron.

Figure 1 - Continued.

Pivot(0.8125c,O.Olc)

Skirt #2 .7625c

Flap Upper Surface

x/c Z/C

0.7500 .7531

7562 :7593 . 7625

7750 17875 . 8000

8125 :8250 . 8375 . 8500 . 8625 . 8742 . 8875 . 8992

-0.0010 0.0072

. 0109

. 0138

. 0164

. 0234 0276

: 0298 . 0307

0308 10306 . 0302 . 0288 . 0271 . 0250 . 0229

Nose Radius = 0.012~

Nose Radius Location (x/c,z/c) = (0.762c,-0.00087~)

Note: Remainder of flap contour matches basic airfoil.

Skirt #l

Radius = 0.012~

Location (x/c,z/c) = (0.738c,-0.00087~)

(c) 25% Slotted Flap. Figure 1 - Continued.

12

Pivot(0.775c,O.Oc)

Skirt

x/c z/c . 675 -.0231

:685 680 -.0215 -.0204

:700 690 -.0158 -.0105 . 705 . 0030

Flap Upper Surface

x/c z/c

:705 700 -. 0069 0030 :720 710 .0119 :0075

. 740 .0171

:775 760 -0194 .0190

:825 800 . .0184 0172 -850 0156 . 875 :0137 :925 900 -0086 .0114

. 950 0051 1:ooo 975 -.0044 :0014

Nose Radius = 0.0117~

Nose Radius Location (x/c,z/c) = (0.7119c,-0.0071c)

Note: Remainder of Flap contour matches basic airfoil.

(d) 30% Fowler Flap. Figure 1 - Continued.

13

Airfoil with Pressure Tubes Model

96 Pressure Tubes

24 Pressure Tubes-

4 Pressure Switches

Pressure Tube Pressure Data System

4 Pressure Transducers }NMHT"l

Electrical Signal

4-Channel Digital Strain Indicate #l #2 #3

Analog to Digital and Recording

Punched Cards

To Computer for Processing and Plotting

Figure 2 - Pressure Measurement System Schematic.

14

-2

cP

-1

0

Present Tests NASA Tests(Ref. 11

Theory predicts no separation.

0.2 0.4 0.6 0.8 1.0

Figure 3 - Comparisons of Pressure Measurements with Theory and NASA experiments.

15

(b) a=8'

x/c Figure 3 - Continued.

16

-1

0


17

I-.

x/c Figure 3 - Continued. 18

-1c

-9

-a

-7

-$

cP t -E

-4

-3

-2

-1

0 ,

4

Present Tests

Note: Theory predicts separation at x/c = . 67 (upper surface).

WC

Figure 3 - Concluded. 19

(a) AILERON DEFLECTION = 0.0 DEGREES

-8*oc

-7-a

-6-a

-5a

$ -4-a

-3*oc

-2coc

-1.00

Or00

I-00

MACH NO. = 0.13 REYNOLDS NO. = 2.2 E 06

C SrtvEuL PLpHh n a

+ -7.9" -.55 x 0.2O 27 D 8.30 1:08 7 12.4O 1.43 4 16.4O 1.64

Flagged Symbols Denote Lower Surface Pressures

C n ai 04

:09 10

:11 . 13

I - Figure 4 - Pressure Distributions with 20% Aileron.

20

(b)- AILERON DEFLECTION = 5.0 DEGREES MACH NO. = 0.13

REYNOLDS NO. = 2.2 E 06

-7-a

-6-a

-5-a

$ -4-a

-3*m

-2-m

-14Xl

O-00

l*W

+ -7.90 x 0.20 D 8.3" w 12.4O Q 16.4='

-16-00

C n a -.38

45 1124 1.61 1.80

C n ai 09

:12 13

:14 . 15

Figure 4 - Continued. 21

(c) AILERON DEFLECTION = 10.0 DEGREES

-7-a

-6*oc

-s*oc

6 -4-00

-3-00

-2*oo

-1-w

O-00

1-W


C SYMBX ALFHA

n a + -7.9" -.25 x 0.20 59 D 8.3' 1:39 T 12.4' 1.77 4 16.4' 1.95

46-00 T t

-14-00

c;, -12-00

t

-10-00

-8-00 IL- o*o x/t 030


C n ai 11

114 . 16

16 :18

(d

-8-w

-7-w

-8.00

-5vw

5 -4*cm

-3-00

-2.00

-l*W

0-w

1.00

AILERON DEFLECTION = 20.0 DEGREES MACH NO. = 0.13


C

SykBu-hlpHh n a

+ -7.90 -.12 x 0.20 74 0 8.3O 1:55 v 12.4O 1.91 0 16.4' 2.09

-10-W


C n ai 16

:20 20

:20 .22

.

(e) AILERON DEFLECTION = 40.0 DEGREES

-7.oc

%-cm

-5-W

$ -4-m

-3-m

-2-w

-1-w

o-w

1-W


C n SYhHlL ALPHA a

+ -7.90 27 x 0.20 1:12

.24

.27 8 8.30 1.93 .26 w 12.4O 2.27 .25 4 16.4O 2.45 .23

C n ai

: -


(f) AILERON DEFLECTION = 60.0 DEGREES MACH NO. = 0.13


%*W

-7-m

-6-w

-5aJ

$ -4*m

-3-m

-2-w

-1-w

O-00

I-00

SYma-ALFlin C n a

+ -7.9O .57 x 0.20 1.39 D 8.3' 2.16 T 12.4O 2.49

-16-W t t

-14-00

5 -12-m I

-lO*W

-8~00 L- O-0 x/c o-20

C n ai 30

132 .32 .32

c - Figure 4 - Continued.

25

(g

-8vCC

-7wocl

-6-w

-5-W

C P

-4.00

-3-m

-2*w

-1.00

I

O-00

AILERON DEFLECTION = -5.0 DEGREES MACH NO. = 0.13


C SYMBOL ALRIA n a

+ -7.9O -.62 x 0.20 09 D 8.3O :92 w 12.4O 1.27 4 16.4O 1.50

O-80

C n ai 00

:04 .08 . 09 . 11


(h). AILERON DEFLECTION = -10.0 DEGREES

-8-E

-7.oc

-6-W

-5-w

5 -4-00

-3-m

-2-00

-1-00


C n C

~~ ALf?-lA a n ai .A -7.90 -.74 -.04 x 0.2O -.12 -.02 D 8.3O .73 .02 v 12.3O 1.09 .05 4 16.4O 1.33 . 08


(i) AILERON DEFLECTION = -20.0 DEGREES MACH NO. = 0.13


%=CE

-7-m

-6-W

-5-w

$ -4.00

-3-w

-2-00

-1-W

O=OO

C

!aMBL1- ALmA n a

b -7.9" -.91 * 0.20 -.34 D 8.3O .40 w 12.3O -75 4 16.4O 1.06

C n ai -.lO -.08 -.09 -.06 -.Ol

- -


(j) AILERON DEFLECTION = -40.0 DEGREES MACH NO. = 0.13


-8-W

-7-m

-6-00

-5-w

%

-4-m

-3-w

-2-w

-1-W

O-W

1-W

C C n n SYMBOL- a ai

+ -7.9O -1.09 -.16 x O.1° -.75 -.16 D 8.3O -.Ol -.16 w- 12.3' .40 -.14 4 16.4O .83 -.09


(k) AILERON DEFLECTION = -60.0 DEGREES MACH NO. = 0.13


C n C

SybrBa-m a n ai

+ -7.8" -1.18 -.23 x O.1° -1.13 -.24 D 8.2" -.40 -.23 w 12.3O -.Ol -.22 . 16.3=' .41 -.18


. Theo

'Notes: (1) Shaded symbols denote theoretical

values using the method of Ref. 3. (2) Flagged symbols from method of Ref. (3) See pressure distributions for

computer predicted separation.

4.

(a) Lift. sretical and Experimental Force Characteris

with 25% Slotted Flap.

\

tics

31

Notes: (1) Shaded symbols denote theoretical values using the method of Ref. 3.

(2) Flagged symbols from method of Ref. 4. (3) See pressure distributions for

Symbol 6, x- (%I z,(%) computer predicted separation. -L L

O0 0

10" 9.0

2o" 13.0

30" 12.5 . 35O 13.0

40" 13.0

LJ

0

4.0

4.0

3.5

2.5

2.0

1.0 2.0 3.0 4.0

CR

(b) Drag. Figure 5 - Continued.

Notes: (1)

(2) (3)

Shaded symbols denote theoretical values using the method of Ref. 3. Flagged symbols from method of Ref. 4. See pressure digtributions for computer predicted separation.

I~-- / ;f7fr~ I:;: .:I: :ti;J;~~~~:~ If:,Itii~ili’t-:iit:::.lir.rlll::ti:i-l:.::}:1::l:‘: 76: ::I: 1. .I.:: :, :,: : I:::!.::. .:i

---~Z :.:. . ..-~~~~..::I :f: _ , ,. j..

..___ -- ._,. ..---.-. .~_.. ..-. . . . (a) ~1 = 0.2' ;t-............ ,, ' :L:::.ii .1 _ ._

_ -. - _... L.-l. . (1) Theory predicts no separation. - -

(2) Confluent boundary layer error encountered.

--_ Theo, ‘- - -’

x/c Figure 6 - Pressure Distributions with

25% Slotted Flap, 10" Flap Deflection. 34

-


35

(c) a = 16.4'

-1 Note: (1) Theory predicts separation at x/c = . 87 (upper surface).

(2) No confluent boundary layer error encountered.

-1

--- Theory(Ref. 3)

x/c Figure 6 - Concluded.

36


38

I - ! .j : I

-; - i

-8

(1) Theory predicts no separatio (2) No confluent boundary layer

error encountered.


39

::

:

:. -

-.

1:

--1

-2

cP

-2

. . _ - . _ _. . .~ ..__ .-, -_. . .._ 0 Experiment

--- Theory(Ref. 3)

(1) Theory predicts no separation.


: I. ’ I,. :t I i. -i---i: :/. .i:::.i. -i-- .j ..:: 1. ./. :i:.y:i Li ..,_ \::.:i / .... / . ..i... j ‘3& ‘1 / ; / : I- A :

-1. ./ :I /.I /.I/ -2 I- .i.\ / ; / : /

0.2 0.4 0.6 0.8 1.0

X/C Figure 8 - Pressure Distributions with

25% Slotted Flap, 30° Flap Deflection. 40

x/c FS.gure 8 - Continued.

41

Note : (1) Theory predicts no separation.


” Lo.jyCL 1111C;1‘L --- Theory(Ref.

-6

- 0.2 0.4 0.6 0.8 1.0


42

. - - Theory (Ref. 3 ) YI :::.'l~:l'::::'....'::. '. :., :-. :1:: :::: I,

..I. .L...... ,-..I. ..I. I I_ / (1) Theory predicts no separatio (2) No confluent boundary layer

error encountered. i:i:,l:iiiliiiilii:ii~~::l:j: I;:iil:;\:l:;.;jy qTqTqT-qY=

0.2 0.4 0.6 0.8 1.0 10.75 1.0



0.6 0.8 1.0



--- Theory(Ref.


45

-8

-6

, _. ! ! ! ! , , , , , , ,

t---- .” ‘... .‘. ... :: :.: ._..’ :.. .:

-

.~..

a = 14. 4 0 ==lf===y iii: :i:~:::: .1~~~I, ::1: :: : :1-. -i:;,-

.; ) 1



.i.:.i 11.. ,.:/

-- 0 -0.2 0.4 0.6 0.8 1.0

WC Figure 9 - Concluded.

46

-8-M:

-7-m

-6-oc

-5-m

C P

-4-00

-3-00

-2.00

-1-00

O-00

(a) FLAP DEFLECTION = 0.0 DEGREES, LOW a's MACH NO. = 0.13


sY?vBlALPHA a f + -7. go -.53 04 x 0.2" 32 :OS D 8.3" 1:14 09 v 12.4O 1.42 :11

Flagged Symbols Denote Lower Surface Pressures

Figure lo- Experimental Pressure Distributions with 25% Slotted Flap. 47

(b) FLAP DEFLECTION = 0.0 DEGREES, HIGH cl's MACH NO. = 0.13


-9-m

-7-m

-6-00

-S-o0

$

-4-m

-3-m

-2-m

-1.oc

0-w

C n C SYNEQ_ ALFl-w a nf

+ 8.3O 1.14 09 x 12.4O 1.42 :11 0 16.4O 1.55 .15 v 18.3' 1.21 .18

Figure lo- Continued. 48

(c) FLAP DEFLECTION = 10.0 DEGREES , LOW a's MACH NO. = 0.13


-7*!m

-6~00

-5-00

C P

-4.00

-3*oo

-2*oo

-l*OO

0-W

l-00

SYtvlmL ALPHA

+ -8.0° x 0.2O D 8.3" 7 12.40

-16-w

-14-m

5 -E!.W

-10.00

-El*00 L O-0 x/c Oe20

C C n m a a

-.19 -;045

1:78 79 -.31 -.57 2.22 -.68

Figure lo- Continued. 49

C nf

21 :27 -30 . 31

C mf

-.34 -.42 -.47 -.48

-7*oc

-6.01:

(d) FLAP DEFLECTION = lo.0 DEGREES, HIGH a'S MACH NO. = 0.13


C C n m SYhecL ALFl-iA a a

+ 8.3O 1.78 -.57 x 12.4O 2.22 -.68 P 16.4O 2.49 -.73 T 18.2O 1.45 -.50

-16-W t

-14=w t

C C nf "f

. 30 -.47

.31 -.48

.31 -.49

.41 -.68

Figure10 - Continued.

50

(c)FLAP DEFLECTION = 20.0 DEGREES, LOW cl's MACH NO. = 0.13

REYNOLDS NO. =

+ -8.0” t 0.10 D 8.3’= T 12.3”

-14.00 -14.00

cb cb

\

-12-00 -12-00

-IO=00 -IO=00

:\ a

-8*oo -8*oo

i\

I; I; 0 0.0 -0 x/c 0.20 x/c 0.20

2.2 E 06

C C n a ma

13 1115

-.16 -.44

2.12 -.70 2.58 -. 81


51

C nf

C mf

. 37 -.54

:44 42 -.59 -.61 . 44 -.61

.-...- -....

(f) FLAP DEFLECTION = 20.0 DEGREES, HIGH cl's

MACH NO. = 0.13

-a-a

-7.01

-6-U

-S*U

C P

-4.M:

-3*Dc

-2-00

-1.00

O-00

REYNOLDS NO. =

SYtmu ALPHA + 8.3" I 12.3O D 14.40 v-

16.4'

-16~00

2.2 E 06

C C n a ma

2.12 -.70 2.58 -.81 2.76 -.86 2187 -.87


52

C C nf mf

44 :44

-.61 -.61

.43 -.61 -42 -.60

-8-00

-7-00

-6-00

C P

-4=oo

-3.00

-2.00

-1-m

O-00

(g)FLAP DEFLECTION = 30.0 DEGREES, LOW a's MACH NO. = 0.13


5YhEKL AlPHA + -8.0° I 0.2O D 8.3O r 12.4O

C n C m C a a nf

C mf

1158 60 -.62 -.36 :46 46 -.62 -.62

2.56 -.88 47 -.64 2.97 -. 98 146 -.64

-16 * 00

-14-00

5 -12voo

-lO*oo

-8-00 i 0.0 - 0.x3

Figure10 - Continued. 53

(h) FLAP DEFLECTION = 30.0 DEGREES, HIGH cl's MACH NO. = 0.13


slM8n ALPHA

+ 8.3O x 12.4O 0 14.40 v 16.2O

C n C a m C C

a "f mf

2.56 -.88 .47 2.97

-.64 -.98 46 3.04 -.64 -.97 143

1.85 -.63 -.67 .31 -.58


(i) F'LAP DEFLECTION = 35.0 DEGREES, LOW U'S MACH NO. = 0.13


-8-00

-7.00

-6-W

-El*00

C P

-4.00

-3*oo

-2.00

-1.00

0.00

SYMBU- N-R-IA

+ -8.0° L 0.20 D 8.3O v 12.4='

-16-00

-14.00

C n C a m a

73 1172

-.41 -.67

2.72 -.94 3.08 -1.02

C nf

C mf

47 146

-.62 -.63

46 147

-.64 -.66


(j) FLAP DEFLECTION = 35.0 DEGREES, HIGH a's MACH NO. = 0.13


-8*O

-7.0

-6-U

-5-a

=P

-4-a

C n C a m C

a nf C

mf 2.72 -. 94 46 3.08

-.64 -1.02 :47

3.12 -.66

-1.01 .43 -.63

l 8.3O t 12.4O P 14.40


56

(k) FLAP DEFLECTION = 40.0 DEGREES, LOW a'S MACH NO. = 0.13


-7-m

-6-00

-5-i-w

$ -4*uc

-3-00

-2.00

-1-w

O-IX

1-W

+ -8.0° 1 0.2O P 8.3O 3- 12.4O

-16-W t

C n C

a m a

68 1171

-.39 -.68

2.70 -.93 3.15 -1.05

C

nf C

mf

143 46 -.64 -.61

-44 -.62 . 47 -.65


57

(1) FLAP DEFLECTION = 40.0 DEGREES, HIGH a's MACH NO. = 0.13


-8*a

-7-a

-6-a

-5.a

C P

-4-a

-3 UC

-2*rx

-1-W

0-w

1-w

SYMBOL bL.PHA

+ 8.3O x 12.4O P 14.40

C C n m a a

2.70 -. 93 3.15 -1.05 3.18 -1.04

C c .’ nf mf

44 :47

-.62 -.65

-47 -.75

Figure10 - Concluded. 58

symbol

0 5O 18.0 7.50 a loo 23.0 7.50

II, 2o" 25.0 3o" 30.0

theoretical values

(2) See pressure distributions

separation.

Figure11 - Theoretical and Experimental Force Characteristics. with 30% Fowler Flap.

0-l 0

Symbol

0"

5O

10°

20"

3o"

35"

400

xp

0

18.0

23.0

25.0

30.0

30.0

10

05

-

n

Notes: (1) Shaded symbols denote theoretical values using the method of Ref. 3.

(2) See pressure distributions fo& zp(%)

0

7.50

7.50

7.50

6.75

6.75

nmmlt-e!r Dredicted separati .on‘:

B

2.0 3.0 4.0 c!L

(b) Drag..

Figure ll-*Continued.

Symbol 6f x (%) zp(%l p

0 0 18.0 7.50

n 10" 23.0 7.50 cl 2o" 25.0 7.50 0 3o" 30.0 6.75 0 35O 30.0 6.75 n 40" 30.0 6.00 0 5o" 27.0 8.00 0 60° 28.0 8.75

(1) Shaded symbols denote theoretical values using the method of Ref. 3.

(2) See pressure distributions for computer predicted separation.

x/c Figure 12 - Preskre Distributions with 30% Fowler Flap,

1O.O Flap Deflection. 62

x/c Figure 12- Continued.

63

’ 0.2 0.4 0.6 0.8 1.0


64

(d) cx = 16.1"

Note:' (1) Theory predicts separation at x/c = . 92 (upper surface).


x/c Figure 12- Concluded .

65

I-


(2) No confluent boundary layer

-_- Theor '- - -' y (Ker . 31

or encountered.

x/c Fig'ure 13 - Pressure Distributions with

30% Fowler Flap, 20' Flap Deflection. 66

x/c Figure lx- Continued.

67

lo-- 0.2 0.4 0.6 0.8 1.0


68

-8

Note:

(d) IT = 16.1°

(1) Theory predicts separation at x/c = .95 (upper surface).

(2) NO confluent boundary layer error encountered.


69


30% Fowler Flap, 30° Flap Deflection. 70

-8

(

-7

-fl

cP -5

-4

-3

-2

( -1

1

1 .- 0.2 0.4 0.6 0.8 1.0

X/G Figure 14- Continued.

71

-9

-8


72

0.7 1.0

-10 C p = -26.0 (d) a = 16.0° -n-in

Note: (1) Theory predicts sepdration at x/c = . 94 (upper surface).


--- Theory(Ref. 3)

A. 0 -- 0.2 0.4 0.6 0.8 1.0


73

1 I ,n L-. .-

lHj&/Tq I J j 0.7"

,! ; 'iij i!r: ,'- 1.. 0

A- 0 0.2 0.4 0.6 0.8 1.0

x/c Figure 15- pressure Distributions with

30% Fowler Flap, 40' Flap Deflection. 74


75

I - -

0.2 0.4 0.6 0.8 1.0

x/c Figure 15- Concluded.

76

(a) FLAP DEFLECTION - 5.0 DEGREES MACH NO. = 0.13

REYNOLDS NO. '= 2.2 E 06

-7-a

-64x

-s*oc

6 -4-m

-3-m

-2-m

-l-al

O-00

l-00

C C n C C m SYMBDL ALPHA

a a nf mf

+ 0.2O 1:42 87 -.39 -13 -.24 x 4.2O -.54 .16 -.25

8.3" 1.93 -.68 12.3O 2.35 -.77

. 16.2O 1.71 -.64

-l6*00

-14*00

5 -l2*00 I

Flagged Symbdls Denote Lower Surface Pressures

. 20 -.33 24

:35 -.39 -.58

Figure 16 - Experimental Pressure Distributions with 30% Fowler Flap.

77

(b) FLAP DEFLECTION = 10.0 DEGREES

-8.00

-7-00

-6.00

-5-00

5 -4-m

-3-m

-2-00

-1-00

O-00

l-00


C C n a m !aMBoL ALPHA a

+ O.1° 1.17 -.52 . 4.2' 1.74 -.68 b 8.3' 2.28 -.82 t 12.3" 2.71 -.93

16.1" 1.81 -.i3

C nf

C mf .

22 :25

-.34 -.37

28 131

-.41 -.44

-42 -.66


I I.

(C) FLAP DEFLECTION = 20 .O DEGREES

-8-a

-7-a

-6-OC

-5-CC

$ -4-m

-3-m

-2-00

-1-m

O-00

1-m


C n C SYMBOL ALPHA

a sma

+ O.1° 1.70 -.73 * 4.2O 2.28 -.90 D 8.3“ 2.80 -1.04 7 12.3 3.15 -1.10 4 16.1° 1.38 -.79

-l6-00

C "f

C mf

141 39 -.50 -.52

141 41 -.53 -.54

. 53 -.77


FLAP DEFLECTION = 30.0 DEGREES MACH NO. = 0.13


-84X

-7-a

-6-a

-5CIC

5 -4-00

-3-m

-2-00

-1-00

O-00

l-00

C n C m SYhEQ ALFi-h a a

+ O.1° 2.27 -1.00 x 4.1" 2.83 -1.16 b 8.2' 3.31 -1.28 7 12.3O 3.51 -1.28 4 16.0" 2.06 -. 93

C C

nf mf

:53 53 -.61 -.61 :45 51 -.58 -.60

. 60 -.83


-8-00

-7-m

-6-00

-5*cx3

C P

-4-m

-3-00

-2*oo

-1-00

o=oo

1-m

(e) FLAP DEFLECTION = 40.0 DEGREES MACH NO. = 0.13


SYMBU. ALPHA a a

i

+ O.1° 2.48 -1.10 x 4.1° 2.97 -1.22 b 8.2" 3.51 -1.38 w 12.3" 2.88 -1.02 4 16.0" 2.49 -1.21


C C

nf mf

154 57 -.66 -.66

. 55 -.63

:51 39 -.60 -.79

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

I 4. Title and Subtitle

Pressure Distributions for the GA(W)-2 Airfoil With 20% Aileron, 25% Slotted Flap and 30% Fowler Flap

7. Author(s)

W. H. Wentz, Jr. and K. A. Fiscko

9. Performing Organization Name and Address

5. Report Date

February 1978 6. Performing Organization C&e

8. Performing Orgamzation Report No.

AR 76-3 10. Work Unit No.

505-06-33-10 Wichita.State University Wichita, Kansas 67208

1 11. Contract or Grant No

NSG 1165 1 13. Tvpe of Report and Period Covered

2. Sponsoring Agency Name and Address

NASA Langley Research Center Hampton, Virginia 23665

5. Supplemenrary Notes

Langley technical monitor: Robert J. McGhee Topical Report

6. Abstract

Surface pressure distributions have been measured for the 13%-thick GA(W)-2

airfoil section fitted with 20% aileron, 25% slotted flap and 30% Fowler flap.

All tests were conducted at a Reynolds number of 2.2 x lo6 and a Mach number of

0.13. Pressure distribution and force and moment coefficient measurements are

compared with theoretical results for a number of cases. Agreement between

theory and experiment is generally good for low angles of attack and small flap

deflections. For high angles and large flap deflections where regions of

separation are present, the theory is inadequate. Theoretical drag predictions

are poor for all flap-extended cases.

- ‘. Key Words (Suggesred by Author(s)) 18. Distribution Statement

Low-speed Experiment Airfoil Theory FEDD Distribution Aileron High-lift Pressure distributions

Subject Category 02 I. Security Clanif. iof this report) 20. Security Classif. (of this page1 21. No. of Pages 22. Rice’

. . lflcwl Unclassified 85 $6.00

*Available from NASA's Industrial Application Centers NASA-Langley, 1978

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