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Wind Energy Reports, no.1 Center for Energy Studies

Two Dimensional Tests of GA(W)-1 and GA(W)-2 Airfoils at Angles- of-Attack from 0 to 360 Degrees

Dale Satran and M.H. Snyder Wichita State University

Recommended citation Dale Satran and M.H. Snyder. Two Dimensional Tests of GA(W)-1 and GA(W)-2 Airfoils at Angles- of-Attack from 0 to 360 Degrees. Wichita, Kan: Wichita State University, 1977. -- 37 p. Digitized by University Libraries and posted in Shocker Open Access Repository

Citable Link: http://soar.wichita.edu/dspace/handle/10057/5692

Terms of use: in the Public Domain

WER-l

P6-IZ-06-0(l

Wind Energy Report No. 1

TWO-DIMENSIONAL TESTS OF GA(W) -l

AND GA(W) -2 AIRFOILS AT ANGLES - OF-ATTACK

FROM 0 TO 360 DEGREES

by

Dale Satran

and

Melvin H. Snyder

Wind Energy Laboratory

Wichita State University

Wichita, Kansas

January, 1977

FOR EARLY DOMESTIC DISSEMINATION

Because of its s'ignificant early commercial potential, this information, which has been developed under a State of Kansas program using NASA infor­mation, 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 This information may be duplicated and used by the recipient with the express limitation that it not be published.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. Date for general release: January. 1979

ABSTRACT

Tests of the GA(W) - l and GA(W)-2 airfoil s were conducted in the 2- D test section o f the Wichita State University 7' x 10' low speed wind tunnel to de te rmine the lift , drag, and pitching moment coefficients of the airfoils. The chords of the aluminum models were six inches. Thg airfoils ~ere tested a6 three Reynolds numbers: .37 x 10 I .51 x 1 0 , and .67 x 10 , and at angles of attack from a to 360 degrees . Hysteresis studies were also performed from negative stall to posi tive stall .

INTRODUCTION

As a part of the continuing wind energy re s earch program at Wichita State University, a two- bladed wind turbine was de­signed and built (re ference 1). It became evident, during the design process and during the calculation of off-design perfor­ma nce, that there is a serious lack of airfoil data for low Reynolds numbers a nd for large ranges of angle of attack .

From the limited data availabl e , it appeared that the GA(W)-l airfoil would be a better section for a windmil l blade (i.e., shorter chord and lighter weight) than the o l der sec­tions. For this reason, the wind turbine , described in refer­ence I, was constructed using a GA(W)-l airfoil.

In order to provide desi g n data for future wind turbine designs, tests have been conducted in the Walte r H. Beech Wind Tunnel at Wichita State University . These tests, re­ported in this report, were two-dimensional te s ts of the GA(W)-l and GA(W)-2 airfoils .

MODELS AND TEST PARAMETERS

The models tested were the GA(\'l)-l and GA(\v) - 2 airfoils. The contours and surface coordinates of the GA(W)-l airfoil are shown in figure 1. Figure 2 illustrates the GA(W) - 2 air­foil , but the coordinates have not been released for general publication. Persons desiring these coordinates must apply to NASA. The models were constructed of aluminum with six­inch chords.

The standard 2-D tunnel walls are 3 feet aRart, and usually the models have- ·for:ty-two-inch diameter end plates_ (for models with c = 2 ft.) . . These 6-inch chord models were equipped with l2-inctldiameterend plates an~the angle of attack drive was modified to permit testing through an angle of attack range of 360 degrees. See figure 3.

The Reynolds numbers attained were fixed by t he limits to the dynamic pressure. Minimum dynamic pressure, fixed by repeatability of data, was 8 psf. and corresponded to Reynolds number of about 350,000. Maximum dynamic pressure was limited ~_ bending ~ of ..-the mQgel which produced deflec­tions of the end plates. For these small models, ~y'namic pressure was limited to 48 pst. corresponding to a Reynolds-' fiumber of approximately 7·0-0,000.

TESTING

The models were tested for a full 360 0 and were tested for hysteresis by rotating the airfoil from deep negative

1

stall t o deep positive stall and back to deep negative stall. The models were tested at three dynamic pressures, 8 , 18 and 4 8 psf . These data were taken o n t h e force b a lance. Wake s urveys were also performed using a five-tube probe to measu re the total pressure variation through the wake (see fiqure 4).

DATA REDUCTI ON

Force data were reduced by a comput er program whi ch corrected the data for f l ow angularity, solid blockage , wake blockage, horizontal buoyancy, and turbulence as shown in the Appe ndix. The' program (refe rence 2) computed lift , drag , pitching moment about the quar ter chord , normal and axial co­efficients for each angle of attack. Drag data from the force balance included interference and tare d r ag for the end plates . The dif ferences between the drag data from the force balance and from the wake survey were removed by the p r og r a m f or each model to yield the actual drag of the model.

Pressure data from the f i ve-tube wake survey p r o be were "reduced into ve l oci ties whi ch were integrated by another p ro­gram to yield drag of the model only . The wake survey was taken too far (six inches) behind the model. Velocity p ro­files through the wake were very flat wh ich hindered deter­mination of the wake boundaries. By changing the integra­tion limi ts by on ly one pe r cent , the drag could change as much as fifty to one h undred pe rcent . By determining the best range of d r ag f or the wake sur veys at each angle of at­tack and by plotting a curve through the average of these drag values, a table of the d rag coefficients was gene r a t ed for each air foil. Minimum drag coeffic i ents are compared , in figure 5 , wi th ski n f ricti on coefficients for all-laminar and all- turbule nt flow.

Wake da t a were used t o correct the drag when the air­foil was not stalled. When the airfoil was stalled , inter­ference and t are drag we r e considered c onstant and were subtracted from the force data.

RESULTS

Corrected aerodynamic characteristics of the are presented graphically in figures 6 throug h 8 . results are presented in tables I and II.

2

airfoils Principal

REFERENCES

1. Dunn, C.H., and Snyder, M.H.: "A Prototype wind Generator Sys tem Supplying Energy to the Electric Utility Grid," University of Missouri-Rolla Energy Conference , Rolla, MO, Octobe r 1975.

2. Rotramel, Breiden thal, and Wentz; Computing Routines for Airfoil Section Wind Tunnel Data Reduction , Aer onautical Engineering Department, Wichita State Univers ity, June 1973.

3 . Wentz, W.H.,Jr., a nd Seetharam, H.C.; A Fowler Flap System for a High-Performance General Aviation Airfoil, Paper no. 740365, S.A.E. Business Aircraft Meeting, Wichita, Kansas, April 1974.

3

SYMBOLS

b airfoil span, d i stance between end plates ft.

C airfoil chord ft.

Cd section drag coefficient, D/qbc

c l section l i ft coefficient, L/qbc

c rnc section pitching moment coefficient, Mi-/qbC 2

T

o drag lb.

L lift lb .

pitching moment about the quarter-chord ft.-lb.

q d · 1 v2 ynam~c pressure, 2P p . s. f.

RN Reynolds number, pVc/v

S wing area = be sq . ft.

t maximum thickness of airfoil ft.

v wind tunne l velocity ft . /sec.

a angle of a t tack. deg.-

coefficient of viscosity slugs/ft. sec .

air density s l ugs/cu. ft.

4

TABLE I

Pr incipa l Parameters fo r GA(W) - l Airfoil

Reynolds Number . 37 x 106 .51 x 1 0 6 .67 x 1 0 6

c l = 0.0 at a = -4.2 ° - 4.3° -4.4°

a = 0.0 at c l = .47 .50 .52

clmax = 1. 35 1. 39 1. 43

.0/10

Cdmin = ~ge_ .0 102 . 0077

L/Dmax = 66.69 76 . 16 86 . 27

at

a = 6.20 6 .20 2.20

Stall + a 16.2° 16. 2° 16.2 °

Range - a - 11.8° - 11.8 0 - 11.8°

5

TABLE II

Principal Parameters for GA(W)- 2 Airfoil

Reynolds Number .37 x 10 6 . . 51 x 10 6 .67 x 106

c 1 ~ 0.0 at a ~ - 4.6° _4.7° _ 4.8°

a ~ 0.0 at c 1 ~ .50 .52 . 54

Clmax ~ 1. 37 1. 42 1. 45

Cdmin ~ . 0122 . 0091 .0067

L/Dmax ~ 66.65 84.44 87.64

at

a ~ 4.20 4.2° 4.2 0

Stall + a 12 . 2° 14.2° 14 . 2°

Range - a - 11 . 8' - 9.8° -9.8°

6

chord line

tic ~ 0 .17

GA(WH Airfoil Coo!dinales

Uppe! SUlface Lower Surface:

XI' ~ XI, ~

0.00000 0.00000 0.00000 0.00000 0.00200 0.01300 0.00200 -0.00930 0.00500 0.02040 0.00500 - 0.01380 0.012,50 0.03070 0.01 250 -0.02050 0.02500 0.041 70 0 .02500 • -0. 02690 0.03750 0 .04965 0.03750 -0.03 190 0.05000 0.05589 0.05000 - 0.03580 0.07500 O.06SSI 0.07500 - 0.()4210 0.1 0000 0.07300 o.looon · 0.04700 0.12500 0 .07900 0 .12500 - 0.05 100 0.15000 0.08400 0.15000 -0.054)0 0.17500 0 .08840 0.17500 -0.05700 0 .20000 0 .09 20!'! 0 .20000 - 0.05930 0.25000 0 .09770 0 .25000 - 0.06270 0.30000 0.10160 0 .30000 -0.06450 0.35000 0 .10400 0 .35000 - 0. 06520 0.40000 0.10491 0.40000 - 0.06490 0.45000 0.10445 0 .45 000 -0. 06350 0.50000 0.10258 0 .50000 -0.06100 0.55000 0.099 10 0.55000 -0.05700 0 .57500 0.09668 0.51500 - 0.05400 0.60000 0.09371 0.60000 -0.05080 0.62500 0.09006 0 .62500 -0.04690 0.65000 0.08599 0.65000 -0.04280 0.67500 0 .08136 0 .67 500 -0.03840 0.10000 0.01634 0.10000 -0.03400 0.72500 0 .07092 0.72500 - 0.02940 0.75000 0 .065 13 0.75000 -0.02490 0.17500 0.05901 0.17500 -0.02040 0.80000 0.05286 0.80000 -0.01 600 0.82500 0.04646 0 .82500 -0.01200 0.85000 0.03988 0.&5000 - 0.00860 0.87500 0.03315 0 .87500 -0.00580 0.!l/0000 0.026)9 0.90000 -0.00360 0.92500 0.0 1961 0.92500 -0.00250 0.95000 0.01287 0.95000 - 0.00260 0.91500 0.00609 0.91500 -0.00400 1.00000 -0.00070 1.00000 -0.00800

Fig ure 1. GA{W) - l Airfoil.

B

chord line

Figure 2. GA(W)-2 Ai rfo il.

9

Figure 3. View Of 2-D Test Section From Downstream Of Model .

Figure 4. Wake survey Probe In Position,

Viewed From Downstream Of The Model.

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Figure 5. Airfoil Minimum Drag Coefficients.

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