Two dimensional tests of GA(W)-1 and GA(W)-2 airfoils at angles

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Wichita State University Libraries

SOAR: Shocker Open Access Repository

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
P6-IZ-06-0(l
WER-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 infor
mation, 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with the express limitation that it not be published.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 in
formation to other domestic parties by the recipient shall be made subject
to these limitations. Foreign release may be made only with prior NASA ap
proval 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 airfoils were conducted
in the 2- D test section of the Wichita State University
7' x 10' low speed wind tunnel to determine 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 .51 x 10 , and
I
a
.67 x 10 , and at angles of attack from to 360 degrees .
Hysteresis studies were also performed from negative stall
to positive stall .
INTRODUCTION
As a part of the continuing wind energy research program
at Wichita State University, a two- bladed wind turbine was de
signed and built (reference 1). It became evident, during the
design process and during the calculation of off-design perfor
mance, that there is a serious lack of airfoil data for low
Reynolds numbers and for large ranges of angle of attack.
From the limited data available , it appeared that the
GA(W)-l airfoil would be a better section for a windmill blade
(i.e., shorter chord and lighter weight) than the ol der sec
tions. For this reason, the wind turbine , described in refer
ence I, was constructed using a GA(W)-l airfoil.
In order to provide design data for future wind turbine
designs, tests have been conducted in the Walte r H. Beech
Wind Tunnel at Wichita State University. These tests, re
ported in this report, were two-dimensional tes 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 air
foil , 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 six
inch 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 deflec
tions 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 3600 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
48 psf. These data were taken on t he force ba lance. Wake
surveys were also performed using a five-tube probe to measure
the total pressure variation through the wake (see fiqure 4).
DATA REDUCTION
Force data were reduced by a comput er program which
corrected the data for fl ow angularity, solid blockage, wake
blockage, horizontal buoyancy, and turbulence as shown in
the Appendix. The' program (reference 2) computed lift, drag ,
pitching moment about the quar ter chord , normal and axial co
efficients for each angle of attack. Drag data from the
force balance included interference and tare dr ag for the
end plates . The differences between the drag data from the
force balance and from the wake survey were removed by the
progr am for each model to yield the actual drag of the model.
Pressure data from the fi ve-tube wake survey probe were
"reduced into vel ocities whi ch were integrated by another pro
gram to yield drag of the model only. The wake survey was
taken too far (six inches) behind the model. Velocity pro
files through the wake were very flat which hindered deter
mination of the wake boundaries. By changing the integra
tion limits by only one percent , the drag could change as
much as fifty to one hundred percent . By determining the
best range of dr ag for the wake sur veys at each angle of at
tack and by plotting a curve through the average of these
drag values, a table of the drag coefficients was gener ated
for each airfoil. Minimum drag coeffici ents are compared,
in figure 5, wi th ski n fricti on coefficients for all-
laminar and all-turbulent flow.
Wake dat a were used t o correct the drag when the air
foil was not stalled. When the airfoil was stalled , inter
ference and t are drag were considered constant and were
subtracted from the force data.
RESULTS
Corrected aerodynamic characteristics of the airfoils
are presented graphically in figures 6 through 8. Principal
results are presented in tables I and II.
2
REFERENCES
1. Dunn, C.H., and Snyder, M.H.: "A Prototype wind Generator
System Supplying Energy to the Electric Utility Grid,"
University of Missouri-Rolla Energy Conference, Rolla,
MO, October 1975.
2. Rotramel, Breidenthal, and Wentz; Computing Routines for
Airfoil Section Wind Tunnel Data Reduction , Aeronautical
Engineering Department, Wichita State University, June
1973.
3. Wentz, W.H.,Jr., and 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, di stance between end plates ft.
C airfoil chord ft.
Cd section drag coefficient, D/qbc
cl section l i ft coefficient, L/qbc
2
c section pitching moment coefficient, qbC
rnc Mi-/
T
o drag lb.
L lift lb.
pitching moment about the quarter-chord ft.-lb.
q d · pressure, 21 Pv 2 p. s. f.
ynam~c
RN Reynolds number, pVc/v
S wing area = be sq. ft.
t maximum thickness of airfoil ft.
v wind tunnel velocity ft ./sec.
a angle of at tack. deg.-
coefficient of viscosity slugs/ft. sec.
air density sl ugs/cu. ft.
4
TABLE I
Principal Parameters for GA(W) - l Airfoil
6 6 6
Reynolds Number . 37 x 10 .51 x 10 .67 x 10
c = 0.0 at a = -4.2° - 4.3° -4.4°
l
a = 0.0 at c = .47 .50 .52
l
cl = 1. 35 1. 39 1. 43
max
.0/10
Cdmin = ~ge_ .0102 . 0077
L/D = 66.69 76 .16 86.27
max
at
a = 6.20 6.20 2.20
Stall + a 16.2° 16. 2° 16.2°
-
Range a - 11.8° - 11.80 - 11.8°
5
TABLE II
Principal Parameters for GA(W)-2 Airfoil
Reynolds Number .37 x 106 . . 51 x 106 .67 x 106
c 0.0 at a - 4.6° _4.7° _4.8°
1 ~ ~
a 0.0 at c .50 .52 . 54
~ 1 ~
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° 4.20
~
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 0.01250 -0.02050
0.02500 0.04170 0.02500 • -0.02690
0.03750 0.04965 0.03750 -0.03190
0.05000 0.05589 0.05000 -0.03580
0.07500 O.06SSI 0.07500 -0.()4210
0.10000 0.07300 o.looon ·0.04700
0.12500 0.07900 0.12500 -0.05100
0.15000 0.08400 0.15000 -0.054)0
0.17500 0.08840 0.17500 -0.05700
0.20000 0.0920!'! 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 0.45000 -0.06350
0.50000 0.10258 0.50000 -0.06100
0.55000 0.09910 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 0.67500 -0.03840
0.10000 0.01634 0.10000 -0.03400
0.72500 0.07092 0.72500 -0.02940
0.75000 0.06513 0.75000 -0.02490
0.17500 0.05901 0.17500 -0.02040
0.80000 0.05286 0.80000 -0.01600
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 0.01961 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
Figure 1. GA{W) - l Airfoil.
B