Heitmeyer, John C

Smith, Donald W

English

NASA Technical Reports Server

I 
I 
RM A50K20 
NACA 
RESEARCH MEMORANDUM 
LIFT, DRAG, AND PITCHlliG MOMENT OF LOW-ASPECT-RA TIO WlliGS 
AT SUBSONIC AND SUPERSONIC SPEEDS - PLANE TRIANGULAR 
WlliG OF ASPECT RATIO 2 WITH NACA 0008-63 SECTION 
,By Donald W. Smith and John C. Heitmeyer 
Ames Aeronautical Laboratory 
Moffett Field, Calif. 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
February 1, 1951 
https://ntrs.nasa.gov/search.jsp?R=19930086427 2020-06-17T12:29:30+00:00Z

1 NACA RM A50K20 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
LIFl', DRAG, AND PITCHING MOMENT OF LOW-ASPECT-RATIO WINGS 
AT SUBSONIC AND SUPERSONIC SPEEDS - PLANE TRIANGULAR 
WING OF ASPECT RATIO 2 WITH NACA 0008-63 SECTION 
By Donald W. Smith and John C. Heitmeyer 
SUMMARY 
A wing-body combination having a plane triangular wing of aspect 
ratio 2 and NACA 0008-63 sections in streamwise planes has been inves-
tigated at both subsonic and supersonic Mach numbers. The lift, drag, 
and pitching moment of the model are presented for Mach numbers from 
0.24 to 0.95 and fram 1.30 to 1.70 at a Reynolds number of 3.0 million. 
The variations of the characteristics with Reynolds number are also 
shown for several Mach numbers. 
INTRODUCTION 
A research program is in progress at the Ames Aeronautical Labora-
tory to ascertain experimentally at subsonic and supersonic Mach numbers 
the characteristics of wings of interest in the design of high-speed 
fighter airplanes. Variations in plan form, twist, camber, and thick-
ness are being investigated. This report is the first of a series 
pertaining to this program and presents results of tests of a wing-body 
combination having a plane triangular wing of aspect ratio 2 and . 
NACA 0008-63 sections in streamwise planes. To expedite publication, 
these data are presented herein without analysis. 
NOTATION 
b wing span, feet 
mean aerodynamic chord (~~: c2 ~ , reet 
o c dY) 
2 
c 
L 
D 
M 
q 
R 
r 
s 
x 
y 
dCL 
do. 
dCm 
dCL 
RACA RM A50Km 
local ving chord, feet 
length of body including portion removed to accommodate sting, 
inches 
lift--drag ratio 
maximum lift--drag ratio 
Mach number 
free-stream dynamic pressure, pounds per square foot 
Reynolds number based on mean aerodynamic chord 
radius of body, inches 
maximum body radius, inches 
total ring area including the area formed by extending the 
leading and trailing edges to the plane of symmetry, square 
feet 
longitudinal distance from nose of body, inches 
distance perpendicular to plane of symmetry, feet 
angle of attack of the body axis, degrees 
drag coefficient ( drqSa
g) 
( lift) lift coefficient -qs-
pitching-moment coefficient about the 25-percent point of the 
( pitching moment) ring mean aerodynamic chord qSC 
slope of the lift curve measured at zero lift, per degree 
slope of the pitching-moment curve measured at ?,p.ro lift 
NACA RM A50K20 3 
APPARATUS 
Wind Tunnel and Equipment 
The experimental investigation was conducted in the Ames l2-foot 
pressure wind tunnel and in the Ames 6- by 6-foot supersonic wind 
tunnel. In each wind tunnel the Mach number can be varied continuously 
and the stagnation pressure can be regulated to maintain a given test 
Reynolds number. The air in these tunnels is dried to prevent formation 
of condensation shocks. Further information on these wind tunnels is 
presented in references 1 and 2. 
The model was sting mounted in each tunnel, the diameter of the 
sting being about 85 percent of the diameter of the body base in the 
12-foot wind tunnel and 73 percent of the diameter of the body base in 
the 6- by 6-foot wind tunnel. The pitch plane of the model support was 
vertical in the 12-foot wind tunnel and horizontal in the 6- by 6-foot 
wind tunnel. A balance mounted on the sting support and enclosed 
wi thin the body of the model was used to measure the aerodynamic forces 
and moments on the model. The balance was the 4-inch~iameter, four-
component strain-gage balance described in reference 3. 
Model 
A photograph of the model mounted in the Ames l2-foot pressure Wind 
tunnel is shown in figure 1. A plan view of the model and certain model 
dimensions are given in figure 2. Other important geometric characte:t-
istics of the model are as follows: 
Wing 
Aspect ratio ••• • • • • • 
Taper ratio • • • • • • • • • • 
Airfoil section (streamwise) 
Total area, S, square feet 
Mean aerodynamic chord, c, feet • 
. . 2 
• • • •• 0 
• • • • NACA 0008-63 
• • • • • 4.014 
Dihedral, degrees • • • • • • • 
1.889 
o 
Camber •••••• • • • • 
Twist, degrees •••• 
Incidence, degrees 
Distance, Wing-chord plane to body axis, feet • 
•• None 
o 
o 
o 
l 
4 NACA RM A50K20 
Body 
Fineness ratio (based upon length I; fig. 2) • • •• 12.5 
Cross-section shape • • • • • • • • • • • • • • • • Circular 
Maximum cross-sectional area, square feet • 0.204 
Ratio of maximum cross-sectional area to 
wing area . . . . . . . . . . • • 0.0509 
The wing was constructed by covering a steel spar with a tin-bismuth 
alloy. The body spar vas also steel but was covered with aluminum. The 
surfaces of the wing and body were polished smooth. 
TESTS AND PROCEDURE 
Range of Test Variables 
The characteristics of the model (as a function of angle of attack) 
vere investigated for a range of Mach numbers from 0.24 to 0.95 in the 
Ames 12-foot pressure wind tunnel and from 1.30 to 1.70 in the Ames 
6- by 6-foot supersonic wind tunnel. The Imjor portion of the data was 
obt ained at a Reynolds number of 3.0 million. Data were also obtained 
for Reynolds numbers up to 15.0 million at low subsonic »:Lch numbers and 
up t o 6.0 million at supersonic Mach numbers. 
Reduction of Data 
The test data have been reduced to standard NACA coefficient form. 
Factors which could affect the accuracy of these results and the correc-
tions applied are discussed in the following paragraphs. 
Tunnel-wall interference.- Corrections to the subsonic results for 
the induced effects of the tunnel walls resulting from lift on the model 
were made according to the methods of reference 4. The numerical value 
of these correctiollB (which were added to the uncorrected data) was: 
M 0.265 CL 
.6CD = 0.0046 CL2 
No corrections were made to the pitching-moment coefficients. 
The effects at subsonic speeds of constriction of the flow by the 
tunnel walls were taken into account by the method of reference 5. 
r 
NACA RM A50K20 5 
The correction was calculated for conditions at zero angle of attack and 
was applied throughout the angle-of-attack range. At a Mach number of 
0.95 in the l2-foot wind tunnel this correction amounted to a 2-percent 
increase in the Mach number over that determined from a calibration of 
the wind tunnel without a model in place. 
For the tests at supersonic speeds, the reflection from the t1.IDD.el 
wall of the Mach wave originating at the nose of the body did not cross 
the model. No corrections were required, therefore, for tunnel-wall 
effects. 
Stream variations.- Calibration of the l2-foot wind tunnel has 
shown that in the test region, the stream inclination determined from 
tests of a wing spanning the t1.IDD.el, with the support system at 00 angle 
of attack, is less than 0.080 • The variation of static pressure is less 
than 0.2 percent of the dynamic pressure. No correction for the effect 
of these stream variations was made. 
A survey of the air stream in the e:r by e:rfoot wind tunnel at 
supersonic speeds (reference 2) has shown a stream curvature only in the 
yaw plane of the model. The effects of this curvature on the measured 
characteristics of the present model are not known, but are believed to 
be small as judged by the results of reference 6. The survey also indi-
cated that there is a static-pressure variation in the test section of 
sufficient magnitude to affect the drag results. A correction was added 
to the measured drag coefficient, therefore, to account for the longi-
tudinal buoyancy caused by this static-pressure variation. This correc-
tion varied from as much as -0.0008 at a Mach number of 1.30 to +0.0009 
at a Mach number of 1.70. 
Support interference.- At subsonic speeds the effects of support 
interference on the aerodynamic characteristics of the model are not 
known. For the present tailless model, it is believed that such effects 
consisted primarily of a change in the pressure at the base of the 
model. In an effort to correct at least partially for this support 
interference, the base pressure was measured and the drag data were 
adjusted to correspond to a base pressure equal to the static pressure 
of the free stream. 
At supersonic speeds, the interference of the sting on the body, 
for a body-sting configuration similar to that of the present model, is 
shown by reference 7 to be confined to a change in base pressure. The 
previously mentioned adjustment of the drag for base pressure, therefore, 
was also applied at supersonic speeds. 
------ -- -- --~ -
6 NACA RM A50K20 
RESULTS 
The results are presented in this report without analysis in order 
to expedite publication. Figure 3 shows the variation of lift coeffi-
cient with angle of attack and the variation of drag coefficient, 
pitching-moment coefficient, and lift-drag ratio with lift coefficient 
at a Reynolds number of 3.0 million and at Mach numbers from 0.24 to 
1.70. The effect of Reynolds number on the aerodynamic characteristics 
at Mach numbers of 0.24, 0.60, 1.30, and 1.70 is shown in figure 4. The 
results presented in figure 3 have been summarized in figure 5 to show 
several important parameters as fUnctions of Mach number. The slope 
parameters in this figure have been measured at zero lift. 
Ames Aeronautical Laboratory, 
1. 
2. 
3. 
National Advisory Committee for Aeronautics, 
Moffett Field, Calif. 
REFERENCES 
Edwards, George G., and Stephenson, Jack D.: Tests of a Triangular 
Wing of Aspect Ratio 2 in the Ames 12-Foot Pressure Wind Tunnel. 
I - The Effect of Reynolds Number and Mach Number on the Aero-
dynamic Characteristics of the Wing with the Flap Undeflected. 
NACA RM A7K05, 1948. 
Frick, Charles W., and Olson, Robert N.: Flow Studies in the 
Asymmetric Adjustable Nozzle of the Ames 6- by 6-Foot Supersonic 
Wind Tunnel. NACA RM A9E24, 1949. 
Olson, Robert N., and Mead, Merrill H.: Aerodynamic Study of a 
Wing-Fuselage Combination Employing a Wing Swept Back 630 .-
Effectiveness of an Elevon as a Longitudinal Control and the 
Effects of Camber and Twist on the Maximum Lift-Drag Ratio at 
Supersonic Speeds. NACA RM A50A3la, 1950. 
4. Glauert, H.: The Elements of Aerofoil and Airscrew Theory. The 
University Press, Cambridge, England, 1926, ch. XIV. 
5. Herriot, John G.: Blockage Correction for Three-Dimensional-Floy 
Closed-Throat Wind Tunnels, with Consideration of the Effect of 
Compressibility. NACA RM A7B28, 1947. 
NACA RM A50K20 7 
6. Lessing, Henry C.: Aerodynamic Study of a Wing-Fuselage Combination 
Employing a Wing Swept Back 630 .- Effect of Sideslip on Aero-
dynamic Characteristics at a Mach Number of 1.4 with the Wing 
Twisted and Cambered. NACA RM A50F09, 1950. 
7. Perkins, Edward W.: Experimental Investigation of the Effects of 
Support Interference on the Drag of Bodies of Revolution at a Mach 
Number of 1.5. NACA RM A8Bo5, 1948. 

NACA RM A50K20 9 
Figure 1.- The model in the Ames l2-foot pressure wind tunnel. 
--- -------
Equation of fuselage ordinates: 
r (I 2x n~4 
- , 1-(1--) J r.. I 
All dimensions shown in inches 
unless otherwise noted 
? 
." ..., ----------------
-', .J _______ ----------
~ I 
~ 
Figure 2. - Plan view of the model. 
!:2: 
f) 
:x:-
§! 
(n 
~ 
o 
f-J 
f-J 
~ 
/,2 
/.0 4 L P LP P tL ~ V p L V P ../ P' P / 
.8 V ::l V V :/ P" .Ll ~ NIJC.4 0008-6J V 'f V ~ ~ ;r ./ V ~ f/"' PIa". wing L 
~ 
.6 
~, 
~ 
:<3 ~ .4 
Cb 
a 
~ .2 
..... 
...,j 
1ft' p:r V y / P V V V / I(' :> JJ' 
~ .L! v= ~ r-L -L y L P" ~ P L 
V fP lP ,.r:{ 1rr'P' V / ~ V ;r )f / .-( 
!L: V V /~ /';1 / V .J ;:V V" l? p 
v y if '1 R ;:f ;J ,L ../ P' ..L. P" ~ V ...rC. J:1' 
l? .... r ~ JI 'f"" p' J/ ., ;:r' p ;:f d If >f d rr' d 
0 ./ / ./ ~ 19' ~ ..P ~ IE ~ / / lP' ~ /f ? ~ /)I LP I~ if' ~ !if fa'" 
-.2 
u 0 if o 0' r;{ r:f cr [J u U v 
'r-I.? M= .24 - .40 -.60 !- 80 :-.86 
- .90 -.95 r- /.3 !- /.4 1-/.53 1- /.6 o Resu/ls from Ames 12'W. r. I 1 1 I I I J. 
-.4 I I I I I I o Results from Ames 6'x6'W:r. . 1 I I 
-4 0 4 8 12 16 20 24 28 for M= .24 
Angle of attack, a, deg ~ 
(a) CL vs a . 
Figure 3.- The variation of the aerodynamic characteristics with lift coeffir.ient at various Mach numbers. Reynolds number, ~~ ~ 
3.0 million. :r> 
~ 
Gi 
o 
~ 
o 
1 
~ 
r 
I 
I 
I 
r 
I 
1.2 
1.0 
0 p p :> 
J. 
V J p j ? p 
.8 
V J ! / V V 0 .P r 
~ oP k< (v !,J / -;; If' p c 
~ 
........ 6 
c:: 
.~ 
.~ 
:::: .4 
III 
a 
$: .2 
-..J 
0 
-.2 
~ ~ jJ lP l? j P p ~ ~ 
Jf / c1 } ~ l,P rV fl lei l.cf 6f1 ~ 
~ p I? ..,IV v lP V- ~ V rl c ./ 
I ? J v V' V fl ~ ~ V v li r: 0 
V Lr' 
.f rP rf cf ~ )C- D l;{ ../ Id /' tr ../ :r .f- A ~ i>"' ....... U ;:r J i ;::~ IP f jI .if ;:f d rf' if d ~ lP l;1 l"fl ~ LP ~ ~ lr ~ V ld 
ii IJ3 j1 1,1 ;f f!' ~ ./ V 17 ~ NACA 000.8-63 ~ rr ~~ 10 P 0 
( C d'}J? a
V 
I- r:-f-I ~3 I- ~ ~ 1- / ? I M=.24 -.40 t- .60 -
.80 - .85 - .90 -.95 1.3 - 1.4 . I' I ' 0 Results from Ames 12 w.r. 1 I 1 1 
I I I I I I I T. 
1 I 1 I 1 1 1 1 1 _ 1 1 1 0 Results from Ames 6 x 6 W . 
-·4.04 0 -.04 -.08 -.12 -.16 fOf M= .2 
Pitching-moment coefficient, Cm ~ 
':' 
(b) CL vs Cm · 
Figure 3. - Continued. 
~ 
:x> 
~ 
Gi 
o § 
I-' 
W 
1.2 
1.0 
.8 
~ 
~ .6 
.~ 
.~ ~ 4 II) • 
a 
~ .2 
-....J 
0 
-.2 
-.4 
r- f-~ -- .0 -:-- -0 f- f- ~ 
-
.-0 
-
.-Q' 
...-< 
..0-
- -
~ 
-- r-
:>--
--
-0 
- - -
;....-
.....c4 0008- 1>:1 
-' 
---
1">- ;....- ..D 
-
....-- ~ 
- f- -..., f...o' ....0-
-
,....cV 
---
-
~ ~ 
---
...cV ty' 
f.cY V (Y 1.t'Y' 
-
f..cY -' L,..o v v :r- .IJ 
./' 
ff' V ID'" /' ~ ~ ./" /" ~ 0"'" V'" ,/0"'" 
./ ~ ~/ ~ ~ v p- ,.........v k..-' k" ,...... A V ..a,.;CI 
r! d foV ~ )J ? V 0('" /' / ~ 
rf If ~ ~ hP 10 IP lo' rf' V 
55 8 P ~ P l t '- I cf ~ 
I~ Is I ~ S <> ~ ~ ~ 
~ ~ I ~ ( ( ~ ~ 
u u 
" 
'b 1: -I.~ b b 0 10 Mj.2( .40 - .6p-80 r:85 .90 - .95 r---1.4 r-1.53- '.6 - 1.7, 1 I I I I 
I I I I I I I I I I I I I I 
o .04 .08 .12 .16 .20 .24 .28 .32 .36 for M = .24 
Drag coefficient, CD 
(c) CL VS CD ' 
Figure 3. -Continued. 
"rJ' 
-
fo 
~ fr 
-
p-
,.......... fO'" 
..cr' 
kQ 
..l'] 
.-0 
-
...0- f--U 
,......... ko-' 1--0-
-IJ~ ...0- .-c v -"" 
...-.t. v-
o Results from Ames 12' W. r. 
o Results from Ames 6'x6'W.r. 
---
~ 
f-' 
~ 
~ 
:t> 
~ 
Gi 
~ 
14 
12 
~Q 10 
.tl'" 
.0::: 
~ 8 \,. 
~ 
~ ~ 6 
I 
;:: 
..... 
-..J 4 
2 
0 
o Results from Ames 12' W. T. 
o Results from Ames 6'x 6' W T. 
90"' ~ J:r ( 
\~ '\ ;S~ I ~h [0 ,<-.. ~ y 1\ ~ '\ "\ /" I~ r c rh T \ \ I~ ~ \ ~ 
I \r K \ \ \ 
" 
q NA<;;A 0008 - 63 
q \ ~ \
"' 
PIOM wIng 
\ 1\ \ 1\ N 1\ '\, 
\~ 't:l ~ \-, !,,, 'q ! ~ ~ ,....a-. ~ ro.... ~ F--r r ~ , 
1'0... '0. ........ ~" ~ '"\ 'n IJl'O- P"-~ ~ '0., D~ ~ pLJ..., P--n ~ '-. Fi Dc 
:l....'-'" '0. .......... 
'" 
R.... D../ p.......... n 1 "'0.... :..,.,/ / 1- / u.., ~ ~ L ~ '0... b.. 
v 
0 "Cl '9' r-a. 
-J r "0 p r 'U u D.o ou...c [no {] 
/ II I II I 
II ~ 
M=.24 .40 .60 .80 .85 .90 .95 1.3 1.4 1.53 1.6 1.7 I I I 
o .2 .4 .6 .8 1.0 1.2 for M= .24 
Lift coefficien~ CL 
(d) L/o vs CL' 
Figure 3. - Concluded. 
~ 
f) 
» 
~ 
(;; 
~ 
o 
I-' 
\Jl 
1.2 
/.0 
.8 
I--~ I--- ~ ~ ~ 
--
r- V / I--- -~ j ~ v V V 
........-I-- NACA 0008-63 
r""'" V f:I Plane wIng ~ 
c.S 
....... . 6 
c:: 
./ fY' f ~ 
~ / ~ Jf 
.Cb 
..... 
. ~ 
s::: .4 
-..;: 
Cb 
./rr /'l' )" 
1,.1 :/ r 
I:) 
~ 
~ .2 
..... 
-..J 
0 
~ ,,)V JI , 
tf' ~ J'r'" 
4~ ~ ~ 
!e~ / I 
ct;> a IQ 
-.2 
t:. R = 3.0 mlW,on 
o R = 15.0 mIllIon 
-.4 
o .04 .08 ,/2 ./6 .20 .24 .28 .32 .36 .04 0 ~04 -.08 ~/2 -.16 
Drag coefficient, Co Pitching-moment coefficient, Cm 
-4 o 4 8 12 16 20 24 
Angle of attock, a, deg ~ 
(0) M= .24. 
Figure 4. - The variation of the aerodynamic characteristics with lift coefficient at various Reynolds numbers. 
f-' 
0' 
~ 
> 
~ 
§ 
R3 
o 
1.2 
/.0 
.8 
f-- 4 -I-- f-6 /~ f f-- f-LY Ii! f" 
J... 
----
./ I r-- v ~ ,/ f-- NAC4 0008- 63 ~ ~ 
PlUM wIng 
l:i ,...-
V V tl 
~ 
.6 
........ 
~ 
.Cb 
.~ 
:f: .4 
Cb 
l~ / / 
" y 
V / II )V V ~V 
P / ~ i 
a 
~ .2 
... 
-..J 
0 
tl.rt u' ~ 
" .~ ;:I~ l~ ~ )I ,r i~ 
t .~ t 
~ ~ ~ 
-.2 
A R:I 3.0 million 
t::.. R:I 5.0 million 
-- -
-.4 
o .04 .08 .12 ./6 .20 .24 .28 .32 .36 .04 0 -.04 -.08 -.12 -.16 
Drag coefficient, Co 
-4 
Pitcl1ing- moment coefficient, em 
o 4 8 12 16 20 24 
Angle of attocK, a, deg 
(b) M· .60. 
Figure 4.- Continued . 
~ 
w 
~ 
!X> 
~ 
Gi 
o 
~ 
r' 
-.J 
.8 
- ~ 
.6 f-
NACA 0008-63 
Pion. wing 
f.S .4 
....... 
~ 
V ;oJ 
fl V 
IrJ( 
J ~ .2 
a Ir 
~ 
..J 0 
-.2 
-.4 
o 
<tf 
~)1i' 
~ 
~ f/ 
.04 .08 
./ 
(J1 
'iT 
/'V 
1/ 
.12 
/ ~ ~ 
AIl lP'" l/ 
P V )V 
,./ 1 
../ W' 
yo l>" 
D ~ 
V 
., V 
tI' 
V 
¢ R= 1.5 mil/ion 
t:. R= 3.0 mil/ion 
~ R=6.0 m/~/ion 
.16 .04 o -.04 -.08 -./2 
Drog coefficien" Co Pitching-moment coefficient, Cm 
-4 0 4 8 /2 /6 
-./6 
Angle of attack I al deg ~ 
(c) M·1.30. 
Figure 4. - Continued. 
f-' 
OJ 
~ 
:x> 
~ 
Gi 
o 
~ 
~ 
~ ... 
c::: 
.~ 
.~ ~ 
't () 
~ 
~ 
.... 
...J 
1.2 
/.0 
I--~ r--
r--
.8 r-- NACA 0008 - 63 ---o.~ ~ D Plane wing 
W --
~ ~ ~ ~ ~ 
.6 ~ V 
-"of ~ .;"r /' 
~ 
-----
~ la' ""'" d"' 
.4 ./-'1f" 
...... 
of rf J" 
;Y i->" ft ~ ~~ 
.2 ~lf V ~ P" l' d IQ ~ '1 
0 
,. / ~ 
':r ~ V 
-.2 " 
cr {'1 o R:I I. 5 mIllion 
A R = 3.0 mIllion 
t7 R = 6.0 mIllion 
-4 
. 0 .04 .08 .12 .16 .20 .24 .28 .04 0 -.04 -.08 -.12 -.16 -.20 
Orag coefficient, Co 
-4 0 
Pitching-moment coefflclen" Cm 
4 8 12 16 20 24 
Angle of attacK, a, deg 
(d) M= 1.70. 
~ 
Figure 4 . - Concluded. 
~ 
:> 
~ 
Gi 
~ 
f-' 
\0 
2 0 
.06 
-.3 
-.2 
JI~ 
"t:i'l 
-. 1 
o 
o 
.2 .4 
I 
.2 .4 
NAC4 0008 - 63 
Plane wing 
I--V 
-
Results 
---Results 
.6 .8 1.0 
Mach number, M 
dCL (a) da vs M . 
V 
-
+--
- -Results 
---Results 
.6 .8 1.0 
Mach number, M 
dCm (b) deL vs M. 
NACA RM A50K20 
--t- - i'--c 
from Ames 12'W. r 
from Ames 6 'x 6'W. r 
1.2 1.4 1.6 1.8 
from Ames 12'W. r 
from Ames 6'x6'Wr 
1.2 1.4 1.6 1.8 
Figure 5. - Summary of aerodynamic characteristics as a function of Mach number. 
Reynolds number, 3.0 mIllion. 
NACA RM A50K20 
~ 
---~ 
:::t.t::::I 
'-
"-~ 
~-.J 
16 
12 
4 
o 
o 
. 3 
.2 
.I 
o 
o 
-
-
.2 .4 
-
~ 
.2 .4 
.6 
NAC4 0008-63 
PIa". wing 
r\ 
\ 
Results from Ames 
- --Results from Ames 
.8 1.0 1.2 1. 4 
Mach number, M 
(c) (LrO)mox vs M . 
-
--
-
6 'x6' W. T. 
12'W. T. 
1. 6 
--
-----.., 
----
- - Results from Ames 12'W. r 
--- Results from Ames 6'x6' w.r 
.6 .8 1.0 1.2 1.4 1. 6 
Mach number, M 
(d) Ci. for (L1J}",oX vs M . 
Figure 5. - Continued. 
21 
1. 8 
1. 8 
22 
l_ 
.20 
. /8 
I--
. /6 I--
I--
.14 
~ ~ .12 
.06 
.04 
.02 
o 
o 
~ CL I I 
NAC4 0008 - 63 .n .6 
Plane wing 
,-
, , 
, , 
, 
, 
, . 
. , 
, 
" , 
. 
. , 
. 
. " 
.4 
/ 
--
.-' --f- ' - --
-
.2 
/ I 
- ~ - - - - t-- {; 
.Z .4 .6 .8 1. 0 
Mach number. M 
(e) Co vs M. 
Figure 5. - Concluded. 
NACA RM A50K20 
, 
/ 
I 
I 
I 
I 
I 
I 
/ 
/ 
/ , 
.-~ 
,-
--
!----- f-- -
1--
~ 
1. 2 1.4 1. 6 1.8 
NACA • Langley Field , V~ . 


Lift, Drag, and Pitching Moment of Low-aspect-ratio Wings at Subsonic and Supersonic Speeds : Plane Triangular Wing of Aspect Ratio 2 with NACA 0008-63 Section

Lift, Drag, and Pitching Moment of Low-aspect-ratio Wings at Subsonic and Supersonic Speeds : Plane  Triangular Wing of Aspect Ratio 2 with NACA 0005-63 Section

Lift, Drag, and Pitching Moment of Low-aspect-ratio Wings at Subsonic and Supersonic Speeds : Plane Triangular Wing of Aspect Ratio 2 with NACA 0005-63 Section

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