Semispan-wing models were tested at angles of attack from 0 to 180 deg at low subsonic speeds. Eight plan forms were considered, both swept and unswept with aspect ratios ranging from 2 to 6. Except for a delta-wing model of aspect ratio 2. all models had a taper ratio of 0.5 and an NACA 64AO10 airfoil section. The delta-wing model had an NACA 0005 (modified) airfoil section. With two exceptions, the models were tested both with and without a full-span trailing-edge flap deflected 25 deg. The Reynolds numbers based on the mean aerodynamic chord were between 1.5 and 2.2 million. Lift, drag, and pitching-moment coefficients are presented as functions of angle of attack. Approximate corrections for the effects of blockage were applied to the data

Koenig, David G.

English

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< NASA MEMO 2-27-59A
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NASA
MEMORANDUM
LOW-SPEED TESTS OF SEMISPAN-WING MODELS AT
ANGLES OF ATTACK FROM 0 ° TO 180 °
By David G. Koenicj
Ames Research Center
Moffett Field, Calif.
NATIONAL
SPACE
AERONAUTICS AND
ADMINISTRATION
WASHI NGTON
April 1959
https://ntrs.nasa.gov/search.jsp?R=19980228368 2020-06-15T23:32:45+00:00Z

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
MEMORANDUM 2-27-59A
LOW-SPEED TESTS OF SEMISPAN-WING MODELS AT
ANGLES OF ATTACK FROM 0° TO 180 °
By David G. Koenig
SUMMARY
Semispan-wing models were tested at angles of attack from 0° to 180 °
at low subsonic speeds. Eight plan forms were considered_ both swept and
unswept with aspect ratios ranging from 2 to 6. Except for a delta-wing
model of aspect ratio 2, all models had a taper ratio of 0.5 and an NACA
64A010 airfoil section. The delta-wing model had an NACA 0005 (modified)
airfoil section. With two exceptions, the models were tested both with
and without a full-span trailing-edge flap deflected 25 °. The Reynolds
numbers based on the mean aerodynamic chord were between 1.5 and
2.2 million.
Lift, drag_ and pitching-moment coefficients are presented as
functions of angle of attack. Approximate corrections for the effects
of blockage were applied to the data.
INTRODUCTION
Recent development of some vertical take-off and landing aircraft
has indicated the need for information concerning the aerodynamic charac-
teristics of wings for angles of attack well above the point of wing
stall. An ab_uudance of results exists for two-dimensional airfoils
through large angles of attack and some of these results are either pre-
sented or referred to in references i and 2. Some information on three-
dimensional models are presented in references 3 through 8.
The present test program was undertaken in order to augment existing
results to the extent of considering larger ranges in aspect ratios and
angles of sweep. Wing plan forms were chosen for the tests for which
data at low angles of attack were already available. Except for an
aspect ratio 2 delta wing, all wings had a taper ratio of 0.5 and aspect
ratios ranging from 2 to 6. Data at low angles of attack for these wings
may be found in references 9 through 14. The models were tested through
an angle-of-attack range of from 0° to 180 °. Results with a full-span
trailing-edge flap deflected 25 ° were also obtained_ for six of the
eight models investigated.
NOTATION
A
c
C
cL1CD
Cm
eL t
CDt
Cmt
K
Kref 3
qt
S
(L
A
aspect ratio
mean aerodynamic chord, ft
cross-sectional area of the wind-tunnel test section, sq ft
(67.8 sq ft for the present case)
lift, drag, and pitching-moment coefficients, respectively,
corrected for blockage by multiplying CLt , CDt , and Cmt ,
respectively, by K
lift
uncorrected lift coefficient, qt S
uncorrected drag coefficient, dra_____g
qt S
uncorrected moment coefficient_ pitching moment
_qt S
blockage correction factor
blockage factor presented in reference 3
dyna_nic pressure measured at a point slightly ahead of the model
and corrected for the normal pressure gradient existing in the
test section without the model installed between the point of
measurement and the _/4 point on the model, ib/sq ft
wing semispan area, sq ft
angle of attack, deg
flap deflection, deg
sweep of the quarter-chord line, deg
MODELS
Geometric data for the eight semispan models tested are presented
in figure i. The models used for the tests were identical to those for
which results for the unstalled angle-of-attack range are presented in
references 9 through 14.
Aspect ratios of the models ranged from 2 to 6. Except for the
aspect-ratio-2 delta wing, all the wings had a taper ratio of 0.5 and
the NACA64A010airfoil section. The delta wing had an NACA0005 (modi-
fied) airfoil section. With the exception of the delta wing, and the
unswept aspect ratio 4.5 wing models, a full-span 30-percent-chord
trailing-edge flap was installed on the models. A small gap existing
between the wing and the trailing-edge flap was sealed for all the tests.
TESTS
All tests were madein one of the Ames7- by lO-foot wind tunnels.
Lift, drag, and pitching-moment data were obtained through an angle-of-
attack range between 0° and 180° . For those models provided with trailing-
edge flaps, the tests included measurementswith the flaps deflected
about 25° .
The data were obtained with a dynamic pressure, qt, of approximately
25 pounds per square foot. The Reynolds numbers corresponding to this
dynamic pressure ranged between 1.5 and 2.2 million.
CORRECTIONS
Corrections which were essentially blockage corrections were made
to all data presented herein as follows:
cD --
where K was obtained from the solid faired curve presented in figure 2.
The values of K presented in figure 2 were obtained as follows:
Two models identical in plan form and wing section to the aspect-ratio-3
unswept wing and the aspect-ratio-2 swept-wing models, but having one-half
their wing area, were constructed and tested. The resulting values of
CL+ and CD+ as obtained from these smaller models were reduced by the
bl$ckage cSrrection factor presented in reference 3 (and shown in fig. 2
as Kre f 3) for corresponding values of (S sin _)/C. These modified
values of CD and CL were then divided by the corresponding values of
CDt and CLt for the two larger models (for a given _ and model plan
form) and plotted against (S sin _)/C, as shown in figure 2. The curve
faired through the mean of the resulting values shown in figure 2 was
employed, as indicated above, to correct the force and moment data for
the effects of blockage.
The corrections madecan only be held approximate since the wind-
tunnel cross section was circular for the tests of reference 3 as con-
trasted to the rectangular cross section of the wind tunnel of the present
tests. It should also be emphasizedthat possible wind-tunnel boundary
effects were not considered in correcting the data.
RESULTS
Drag, lift, and pitching-moment data are presented in figures 3, 4,
and 5, respectively. As was described in the previous section, approxi-
mate corrections for wind-tunnel model blockage have been applied to all
force and momentdata.
There are somedifferences between the present data and those
presented in references 9 through 14. If the wind-tunnel wall correction
factors are taken into account, the present data are practically identical
to those of the reference reports for angles of attack below wing stall.
However, for somemodels, differences exist after stall takes place.
These differences seemexplainable from the standpoint of slight variations
in test conditions and the blockage corrections madeto the data for the
present tests.
AmesResearch Center
National Aeronautics and SpaceAdministration
Moffett Field, Calif., Nov. 26, 1958
REFERENCES
i.
t
,
,
Wick, Bradford H.: Study of the Subsonic Forces and Moments on an
Inclined Plate of Infinite Span. NACA TN 3221, 1954.
Critzos, Chris C., Heyson, Harry H., and Boswinkle, Robert W., Jr.:
Aerodynamic Characteristics of NACA 0012 Airfoil Section at Angles
of Attack From 0° to 180 °. NACA TN 3361, 1955.
Knight, Montgomery, and Wenzinger, Carl J.: Wind-Tunnel Tests on a
Series of Wing Models Through a Large Angle-of-Attack Range.
Part I - Force Tests. NACA Rep. 317, 1929. (Supersedes NACA
T_ 294)
Boisseau, Peter C.: Low-Subsonic Static Stability and Damping
Derivatives at Angles of attack From 0° to 90 ° for a Model With
a Low-Aspect-Ratio Unswept Wing and Two Different Fuselage Fore-
bodies. NASA MEMO 1-22-59L, 1959.
5. Johnson, Joseph L., Jr.: Low-Speed Measurements of Static Stability,
Damping in Yaw, and Damping in Roll of a Delta, a Swept, and an
Unswept Wing for Angles of Attack From 0o to 90o. NACA RM L56BOI,
1956.
t Hewes, Donald E.: Low-Speed Measurement of Static Stability and
Damping Derivatives of a 60 ° Delta-Wing Model for Angles of Attack
of 0° to 90° . NACA RM L54G22a, 1954.
. Johnson, Joseph L., Jr.: Measurement of the Static Stability and
Control and Damping Derivatives of a O.13-Scale Model of the
Convair XFY-1Airplane. NACA RM SL54J04, 1954.
. Kuhn, Richard E., and Draper, John W.: Investigation of the Aero-
dynamic Characteristics of a Model Wing-Propeller Combination and
of the Wing and Propeller Separately at Angles of Attack Up to
90 ° . NACA Rep. 1263, 1956.
. Dods, Jules B., Jr.: Wind-Tunnel Investigation of Horizontal Tails.
I - Unswept and 35 ° Swept-Back Plan Forms of Aspect Ratio 3- NACA
RMA7K24, 1948.
i0. Dods, Jules B., Jr.: Wind-Tunnel Investigation of Horizontal Tails.
II - Unswept and 35° Swept-Back Plan Forms of Aspect Ratio 4.5.
NACARMA8BI1, 1948.
ii. Dods, Jules B., Jr.: Wind-Tunnel Investigation of Horizontal Tails.
III- Unswept and 35° Swept-Back Plan Forms of Aspect Ratio 6.
NACARMA8H30, 1948.
12. Dods, Jules B., Jr.: Wind-Tunnel Investigation of Horizontal Tails.
IV - Unswept Plan Form of Aspect Ratio 2 and a Two-Dimensional Model.
NACA RMA8J21, 1948.
13. Dods, Jules B., Jr.: Wind-Tunnel Investigation of Horizontal Tails.
V - 45° Swept-Back Plan Form of Aspect Ratio 2. NACA RM A9D05, 1949.
14. Dods, Jules B., Jr.: Control Effectiveness and Hinge-Moment
Characteristics at Low Speed of Large-Chord, Low-Balanced, Flap-Type
Controls on a Triangular Wing of Aspect Ratio 2. NACA RM A52F13,
1952.

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66. oo
A
/
i //
I
29.33
(b) Aspect ratio 4, 5, and 6 models; S = 10.08 sq ft.
Figure I.- Concluded.
9K
1.00
.96
.92
.88
.8O
•?6
.72
.68
.64
.6O
%
\
K
ref 3
C
K, derived from
@ Brag data
[] Lift data
\ O
O o
[]
[]
D
O
]
O
o
0 .02 .04 .06 .08 .10 .12 .14 .16
S sins
C
Figure 2.- Blockage correction used to correct force and moment data.
i0
CD
1.6
1.2
.4
o 20 40 60 80 I00 120 140 160 180
(I
(a) A = 2; A = 16.7 °
CD
1.6
1.2
.8
.0
0 20 40
6, deg
o 0
60 80 loo 120 3_)40 160 3.80
(b) A = 2; delta wing
Figure 3.- Drag characteristics of the models.
ii
CD
1.6
1.2
.8
.4
0
(c) A = 2; A = 45°
CD
1.6
1.2
.8
0
0
[] 28 ° _:_
20 _0 6O 80 100 :]20 1_0 !60 !80
(a) A = 3; A = ]_]_.3°
Figure 3.- Continued.
12
CD
1.6
1.2
.8
.4
0
f
/
/
5, deg
Q 0
(e) A = 4.5; A = 7.6 °
CD
1.6
1.2
.8
.2
0
5
20 40 60 80 lO0 120 l_O 160 180
(f) A = 4.5; A = 35 °
Figure 3.- Continued.
13
CD
1.6
1.2
.8
.4
0
J
i
(g) A = 6; A = 5.7°
CD
1.6
1.2
.8
.4
0
0
6, deg __
0
25
20 4o 6o i2o lho i6o80 I00 180
(h) A = 6; A = 35°
Figure 3.- Concluded.
14
CL
1.6
1.2
.8
.4 /
I
6j deg
rn 25
-1.2
1.2
.8
.4
CL
0
,
#
I
I
I
(a) A = 2; A = 16.7 °
6)
20 40 60 80 lO0 120 140 160
(b) A = 2; delta wing
Figure 4.- Lift characteristics of the models.
i8o
15
CL
1.2
.8
.2
-1.2
(c) A = 2; i = 45°
%
.8
/
I
!
0
-.I
6j degO 0
[] 28 _ _
-1.2
0 20 20 6o 80 I00 120
(d) A = 3; A = ll.3 °
l_O 160 180
Figure 4.- Continued.
16
CL
1.2
.8 I'_
!
I
• g I
I
!
0
-1.2
(e) A = 4.5; A = 7.6 °
CL
.8 "1 /_'_ O[] O, deg_2_0 _
I
-._ __o /
-1.2
0 20 hO _ 80 100 120 _0 l_ 180
(f) A = 4.5; A = 35 °
Figure 4.- Continued.
17
%
%
1.2
.8
.4
-.4
-.8
-I.2
!
6_ deg /_
(g) A = 6; A = 5.7 °
!
/
/
/
/
-1.2
0 20 40 60 80 100 120 140 160 180
(h) A = 6; A = 35 °
Figure 4.- Concluded.
18
%
%
.I
o __
-.2_A__
-°3
-.2
-.F
.1
0
-.1
-.2
-.3
-.2
I
6, deg II
II
oo jtrn 2F
\
\
%
(a) A = 2; A = 16.7 °
5, deg
-'Fo 2o 2o 6o 8o i0o 12o lho 16o
(b) A = 2; delta wing
Figure 5.- Pitching-moment characteristics of the wings.
180
19
%
-.I
-.2
-.3
-.4
-.5
.1
O
[]
6, deg
0
25
(c) A = 2; A = 45°
%
2 °
--,2
--.3
-.4
-.5
o 20 4o 60 80 I00 120
(Z
(d) A : 3; A =ll.3 °
!
,q,
].40 16o 18o
Figure 5.- Continued.
2O
%
%
O •
-.1
-.2
-.3
-.4
-.5
.1
\
• v
(e) A = 4.5; A = 7.6 °
/
8, deg I
-.I o o I1l'o 25 _ I
-.2 !
-o_
20 ho 60 80 loo 12o _o 16o 18o
Q;
(f) A = 4.5; A = 35 °
Figure 5.- Continued.
21
=f
-4
!
%
C m
-.I
-.2
-.3
-.4
5egIo 0E] 25
(g) A = 6; A = 5.7 °
.1
o_
/
-o2
-.3
-.4
0 20 &o 60 80
-.5
8, deg
o 0
o 25
II
I I
_j
t;
I00 120 I_0 160
/
II"
Z80
(h) A = 6; A = 35°
Figure 5.- Concluded.

_j



Low-Speed Tests of Semispan-Wing Models at Angles of Attack from 0 to 180 degrees

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