Summary Report of
FORCE
and pressure distribution were measured over an 18-in chord NACA 0015 airfoil in the Georgia Tech 21/2 by 9 ft, two-dimen-sional wind tunnel through 1800 angle of attack. An IAS = 80 mph was used, yielding an RN, = 1,230,000** under tunnel conditions. In general, the re-sults are presented as plots of the pres-sure distributions normal to the chord. plots of the force data compared with he pressure data, and tables of force and pressure data.
Of particular interest are the follow-ing:
76
(I) The drag coefficient of the airfoil
changes from 0.0092 at 0° to 1.770 at 90° and 0.0348 at 180°. (2) The slope of the lift curve from
0 to 6° is 0.096 per degree. and
from 180° to 186° is 0.090.
The complete report (100 Dam) is arailable at reproducton cost from the Director. Daniel Guggenheim School of Aeronautics, Georgia Institute of Technology. Atlanta. Ga. Reynold's Numbers quoted in this report are based on the airfoil chord at = 0.
1.6 14 510 La 6 -
04
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THE FORCES AND MOMENTS
OVER AN NACA 0015 AIRFOIL
Through 1800 Angle of Attack*
ALAN POPE
Associate Professor, Georgia Institute of Technology
12 i.0 0 8 (.3 2
The blocking due to the tunnel
walls appears to be quite negligi-ble, even for an 18-in model at cc.= 90°.
The downstream surface with the
airfoil at cc.= 90° is at approxi-mately 1.0 q over the entire area.
20 30 40 50 60 70 80 90 100 110 :20 50 140 150 160 :70 so NC-LE OF ATTACK , cc
Fig. 2: Comparison of force and pressure data for drag through the complete range of angle of attack
Fig. I: Comparison of force and pressure data for lift through the complete range of angle of attack
1 ncrociuction.On a helicopter in
forward flight, the inboard blade ele-ments that pass through the reverse-flow region experience a change in angle of attack of 360' once each rev-olution. In other flight conditions, the angles of attack of fairly large areas of the blades are beyond the usual values for which data are customarily pre-sented. This program was carried out to make available data for a represen-tative airfoil in the unfamiliar range, at a Reynolds Number close to flight
conditions.
Apparatus, Models, and Tests.The Georgia-Tech tunnel has a turbulence factor of L20. Using the model chord of 18 in and average temperature and pressure conditions during the pro-gram, the effective Reynolds Number was L230,000. Some force tests at 100
mph were also made, RN, being
1,530,-000 for this condition.
The model was made of wood with
flush pressure orifices and was checked
out with feeler gauges as being within
-',- 0. 0.005 in in contour. The
sm-rnrc was.. ae rorivnamically smooth.
F. Ft n tOo. ( LC) Lit I , I 1 . ',FORGE TESTS 1 TEsTs 1 ,-pscssunc
Fr
I\I
I ?. 1 i 10 20 30 40 50 60 70 80 90 ANGLE OF ATTACK ec i 20 50 :43 150 60 170 * 4, .; h : i NACA 0015 AIRFOIL I IAS 80 MPH, RN, 1,230,000 . I ' ; ' i zi 1 1 1 1 i I 1 I I I 1 1 I 1 I 1 1 al,t... ReEr4.4' I8 CI Cdo Cm1/4 LIST OF SYMBOLS Lift
section lift coefficient= RS tunnel dynamic pressure cor-rected for blocking, psi wing area, sq ft
section drag coefficient =
Drag
qS
section moment coefficient
about the airfoil quarter chord
Moment V.
Pressure orifices were located at the following chord stations on upper and lower surfaces: 0, 1.25, 2.50, 5.00, 7.50, 10.00, 15.00, 20.00, 25.00, 30.00,
35.00, 40.00, 45.00, 50.00, 55.00, 60.00, 65.00, 70.00, 75.00, 80.00,
85.00, 90.00, 95.00, 97.50, 100.00. The pressures were led out through plastic tubes to a 50-tube, multiple manometer, which was photographed.
The pressures were then read on a
ground glass with scales specially
pre-pared to subtract the tunnel static
pressures aid convert the remaining head into pressure coefficient.
04
Fig. 3: Effect of the angle ol attack on the normal force coefficient (pressure data)
36 -4P 510
1 '
NACA 0015 AIRFOIL
IAS. 80 MPH, N. 1,230,000
Resuits.lt is not the Purpose of
this report to present a lengthy aero-dynamic analysis of this well-known airfoil, but it appears in order to note some of the more interesting points of correlation with previous tests as well as the new data for the higher angles. The lift plot (fig. 1) shows several interesting features not previously re-ported. First, C(,31 was 1.20 with the airfoil leading edge forward, and 0.70 with the airfOil reversed (trail-ing edge forward). Unexpectedly, sec-ond lift peaks occurred at cc.= 50°
(ci = 0.97) and
cco =135° (c=
76 do 90 100 110 20 L30 40 LSO [60
' OF ATTACK, o.'
'
Fig. 4: Effectof the angle ofattack on
the chord coefficient (pressure data)
Fig. 5: Effect of the angle of attack on the moment coefficient (force data)
0.85). The lift-curve slope for the
airfoil reversed was .090 as compared to .096 with the leading edge forward. The lift data did not go through 0 at
= 90° and 180, missing those
values 50 and 1Y-.°'
respectively. The pressure lift cip, calculated from
0/9 = c
cod cc. cc
cco , and alsoplotted in fig. /, showed more
reason-able results.
The moment coefficient, fig. 5, be-came negative at 15' and remained so until 180° was reached. At CC 0 = 90°, .ct01/4 = 0.4,5 and C.% = cco3/4
3/4, ed. = 0.45 + 1.77 X .25
0.007, or the center of pressure was very nearly at the half chord. With airfoil leading edge forward, the cmac
= 0.006 and the a.
c. is at the22.5% c; with the airfoil reversed, the Cm a = 0.00 and the a. c. is at the 21.7% c from the (then) sharp lead-ing edge. No attempt was made to cal-culate the pressure-moment data.
The drag curve rose steadily beyond the stall and finally reached a maxi-mum of cdo =1.77 at cc. = 90°. This
is in good agreement with the value of 1.8 given for flat plates. The drag at CC = 0 was 0.0092 and at ace =--1800 was 0.0342, or about 3.8 times the minimum value.
A close correlation between force and pressure data was observed in the lift plot and in the drag plot, the latter
to a somewhat lesser degree. The
pres-(Continued on page 100)
r
I I I I I I 10 ' 30 40 50 1 60 70 110 90 100 ANGLE OF ATTACK a. I I I ! NAGA 110 120 30 140 I I [50 I SO 170 I 1 1 I I 1 I I AIRFOIL 005 1AS MPH ,RN. 0 to1,230,000 -I I I: N .
i I I I 1 1 1 1: NACA 00I5 AIRFOJ
Jj 1A5 80 MPH, RN.. 1,230,000 NIPP-! I . I I 10 20 , 30 40 50 60 70 BO 90 100 110 120 130 140 . 150 160 170 ANGLE OF ATTACK a 1 I I 1 c-v q5c airfoil chord, ft tunnel height, ft
cc airfoil angle of attack
correct-ed for tunnel wall effect,
de-grees
cc chord force coefficient form
integration of chord force pressure distribution
cs normal force coefficient from integration of normal force
pressure distribution
pressure lift coefficient = cs
cos cc,
C sin
cc0pressure drag coefficient .= cc cos cc + cN sin cc 0
.!.*
FORCES AND MOMENTS
(Continued from page 78) sure data should of course be increased
by some skin-friction value (say about
= 0.0080) to bring
the agree-ment a little closer.aP A very few incorrect values of
>
LO were noted, the greatest be-ing + L07. These points are shown t although the curves are properlyfaired through +LO) to illustrate the possible range of error. No correlation between CC and the overpressure
could be found.
Near CC 0 = 0, the pressure inte-grations showed a negative drag force of such proportions that even the ad-dition of skin friction would not con-vert them to reasonable overall values. The only explanation that seems pos-sible is to point out once again that chordwise pressure integrations
rep-resent the small difference between lame areas, and the resulting accuracy is hence not too good. Above cc, = 13°, the pressure and force measure-ments are much better.
REFERENCES
I. ALAN POPE. Wind Twine/ Testing. John Wiley & Sons, New York, 1947 p. 250. J. C. HUNSAXES and B. G. RIGHTS:MS,
En-gineering Applications of Fluid Mechanics. MeGraw-Hill Book Co-, New York. 1948 p.
199.
Pope. Letter to the Editor. Journal of the Aeronautical Sciences. November.
1947.
EASTMAN N. JACOBS, The Variation of
A.ir-foa Section Characteristics With Reynolds
Number. TR 566. 1937. i loUPPER SURFACE LOWER SURFACE ' ThL1 40 60 80 % FHORD 0-. .10°, z.. 0.942 I 1 1
lallailiMitr01111113ir
". . - '''
''''''' -.
zlcr o 0 ao i % CHORD a... 90°, z.. 1.658 I ; I -! I % CHORD cte 120°, z..L490 i 1 I ' % CHORD 20 4 80 1 1 CHORD cc.i70°, z.= 0.774 1 aSpan qf Flight
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Air Parcel
aI..70° z.= 1.613
I 1'
Picked From
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o Zlicr 0 Figure 6 cto .180°, z.. 0.0026 1 i
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