Fournier, Paul G

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IDENTIAL Copy 364 RM L53G14a 
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NACA 
RESEARCH MEMORANDUM 
WIND-TUNNEL INVESTIGATION OF THE

AERODYNAMIC CHARACTERISTICS IN PITCH AND SIDESLIP AT HIGH

SUBOC SPEEDS OF A WING-FUSELAGE COMBINATION IIAA

TRIANGULAR WING OF ASPECT RATIO 4

By Paul G. Fournier

Langley Aeronautical Laboratory rQ 
Langley Field, Va.
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CLASSIED DOCUMENT 
This material contains Information affecting the National Defense of the United States within the meaning 
of the espionage laws, Title 18, U.S.C., Secs. 79 and 794, the transmission or revelation of which in any
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manner to an unauthorized person Is prohibited by law.
	
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NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
August 28, 1953 
CONFIDENTIAL 
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https://ntrs.nasa.gov/search.jsp?R=19930087666 2020-06-17T12:11:47+00:00Z
NACA HM L53G14a	 CONFIDENTIAL 
NATIONAL ADVISORY COMNITTEE FOR AERONAUTICS 
RESEARCH MENORA1JDUM 
WIND-TUNNEL INVESTIGATION OF THE 
AERODYNAMIC CHARACTERISTICS IN PITCH AND SIDESLIP AT HIGH 
SUBSONIC SPEEDS OF A WING-FUSELAGE COMBINATION RAVING A 
TRIANGULAR WING OF ASPECT RATIO 4 
By Paul G. Fournier 
The results presented in the present paper are part of a program 
conducted to investigate the effect of wing plan form on the aerodynamic 
characteristics in pitch, in sideslip, and during steady roll. This 
paper presents the aerodynamic characteristics in pitch and sideslip at 
high subsonic speeds of a wing-fuselage combination having a triangular 
wing of aspect ratio 4, a leading-edge sweep angle of 45 0 , and with an 
NACA65AOO6 airfoil section parallel to the plane of symmetry. The range 
of Mach number was from 0. 1 0 to 0.95 at a range of Reynolds number from 
2.4 x 1o6
 to 3.9 x 106 
The results indicate that the effective-dihedral parameter C1
'-'CL 
determined at low lift coefficients for the wing-fuselage combination is 
considerably higher than would be expected from available methods - 
particularly in the range of Mach number from 0.75 to 0.92. 
The wing-fuselage combination was longitudinally unstable at lift 
coefficients from approximately 0.50 to 0.70 and at Mach numbers from 
O. li-O to 0.80. At Mach numbers above 0.80, the high-lift stability 
appeared to improve.
INTRODUCTION 
A systematic research program is being conducted in the Langley 
high-speed 7- by 10-foot wind tunnel to determine the aerodynamic char-
acteristics of various model configurations in pitch, in sideslip, and 
during steady roll up to a Mach number of about 0.95. 
The data presented in this paper were obtained from tests of a 
wing-fuselage combination having a triangular wing of aspect ratio Ii- with 
leading-edge sweep angle of 45° and an NACA 65A006 airfoil section par-
allel to the plane of symmetry. The pitch and sideslip data for the 
CONFIDENTIAL
2	 CONFIDENTIAL	 NACA RM L53G14a 
fuselage alone are presented 
range of Reynolds number for 
from 2. 4 x 106 to 3.9 x 106. 
results, only a limited anal; 
acteristics is presented.
in references 1 and 2, respectively. The 
the sting-supported model tested varied 
In order to expedite the issuance of the 
Tsis of some of the more significant char-
COEFFICIENTS AND SYMBOLS 
The stability system of axes used for the presentation of the data 
together with an indication of the positive forces, moments, and angles 
are presented in figure 1. All moments are referred to the quarter-
chord point of the wing mean aerodynamic chord. 
CL	 lift coefficient, Lift/q.S 
CD	 drag coefficient, Drag/cis 
Cm	 pitching-moment coefficient, Pitching moinent/qS 
C1	 rolling-moment coefficient, Rolling moment/q.Sb 
Cn	 yawing-moment coefficient, Yawing moment/qSb 
Cy	 lateral-force coefficient, Lateral force/cis 
drag due to lift, CD - CDC 
Dbp	
base pressure-drag coefficient 
CDCo	 drag at zero lift 
L/D	 lift-drag ratio, CL/CD 
Cl	 dynamic pressure, pv2/2, lb/sq ft 
P	 mass density of air, sings/cu ft 
V	 free-stream velocity, fps 
M	 Mach number
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NACA RM L53G14a	 CONPIUENTIAL 
R	 Reynolds number, pV/p. 
absolute viscosity of air, slugs/ft-sec 
A	 aspect ratio, b2/S 
S	 wing area, sq ft 
b	 span, ft 
c	 local chord, ft
b/2 
mean aerodynamic chord,
	
	 c2dy, ft 
SO 
y	 spanwise station, ft 
a.	 angle of attack, deg 
13	 angle of sideslip, deg 
C	 lift-curve slope per deg, 6CL/& 
C1  
6C z
per deg60
Cn 
Cn13 = s.— per deg 
Cy	
Cy
13 = s— per deg 
C113c = CL per deg
MODEL AND APPARATUS 
The wing-fuselage combination tested is shown in figure 2. The wing, 
made of aluminum alloy, had an NACA 65A006 airfoil section parallel to the 
plane of symmetry and was attached in a mlthiing position to the aluminum 
fuselage. The ordinates of the fuselage are presented in reference 1. 
The wing aspect ratio is -I and the leading-edge sweep angle is 
CONFIDENTIAL
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	 MACA RN L53G14a 
The model was tested. on the sting-type support system shown in fig-
ures 3 and 4. With this support system the model can be remotely oper-
ated through a 280 angle range in the plane of the vertical strut. By 
means of couplings in the sting, the model can be rolled through 900 so 
that either , angle of attack (fig. 3) or angle of sideslip (fig. ii. ) can 
be the remotely controlled variable. With the model in a horizontal 
position (fig. 3), couplings can be used to support the model at angles 
of sideslip of approximately _40 and lia while the model is tested through 
the range of angle of attack. The data were obtained by use of an inter-
nally mounted electrical strain-gage balance. 
TEST AND CORRECTIONS 
The tests were conducted in the Langley high-speed 7- by 10-foot 
wind tunnel through a range of Mach number from 0.40 to 0.95. The size 
of the model caused the tunnel to choke at a corrected Mach number of 
about 0.96. The blocking corrections which were applied to the data were 
determined by the velocity-ratio method of reference 3. 
Jet-boundary corrections, which were applied to angle of attack, 
lift, and drag, were calculated by the method of reference I-i. . The correc-
tions to pitching moment, lateral force, yawing moment, and rolling moment 
were considered negligible. 
Sting-support tares were determined, but were found to be negligible 
for a wing-fuselage combination except for drag. The drag tare results 
from the influence of the sting on the external model pressure - particu-
larly near the rearward end. This tare value amounted to a drag-coefficient 
increment of about 0.002 throughout the test ranges of angle of attack 
and Mach number and should be added to the drag data presented. The 
correction has not been applied as some uncertainty exists regarding its 
exact magnitude and also because the correction was obtained after the 
results of some investigations (for example, ref. i) which are related 
to the results of the present investigation had already been published. 
The angle of attack and angle of sideslip have been corrected for the 
deflection of the sting-support system and balance under load. 
The drag data have been corrected to correspond to a pressure at 
the base of the fuselage equal to free-stream static pressure. For this 
correction, the base pressure was determined by measuring the pressure 
CONFIDENTIAL
NACA EM L53G14a	 CONFIDENTIAL	 5 
inside the fuselage at a point about 9 inches forward of the base. The 
following corrections were added to the measured drag coefficients: 
M CDb 
O.40 0.0022 
.60 .0022 
.80 .0027 
. 90 .0032 
.95 .0033
No corrections were made for the effects of aeroelastic distortion. 
Experience with other related plan forms indicate that these effects 
would be small. The variation with Mach number of the mean test Reynolds 
number based on the wing mean aerodynamic chord is presented in figure 5. 
RESULTS AND DISCUSSION 
The results of the present investigation are presented in the fol-
lowing figures: 
Basic data:
	
Figures 
Longitudinal ...........................6 
Lateral 
Summary plots: 
Effects of Mach number ...................... 8
 Drag due to lift ......................... 
Effective dihedral parameter ...................10 
The basic longitudinal and lateral data for the fuselage alone were 
previously presented in references 1 and 2, respectively. 
The basic pitching-moment characteristics (fig. 6(b)) indicate 
longitudinal instability of the wing-fuselage combination at a range of 
Mach number from 0. 1 0 to 0.80 and at a range of lift coefficient from 
approximately 0.50 to 0.70. The rearward shift in low-lift aerodynamic-
center position which begins at about 0.80 Mach number seems to be accom-
panied by a somewhat improved high-lift stability. 
The basic data of figure 7(a) also show that at high lift coeffi-
cients the values of Cj are considerably larger at high subsonic Mach 
numbers than at low speeds. The effective dihedral parameter
CZPCL 
CONFIDENTIAL 
M.
	 CONFIDENTIAL
	
NACA RIVI L53G14a 
determined at low lift coefficients through the range of Mach number is 
presented in figure 10. The experimental values are compared with the 
theoretical values computed by the method of reference 5 and corrected 
for the effects of Mach number by the method of reference 6. The experi-
mental values are considerably higher than the estimated values, partic-
ularly in the range of Mach number from 0. 75 to 0.92. For the wing tested, 
CIPCL 
exhibited a rapid reduction in magnitude with Mach number above 
the force-break Mach number. 
Some of the more important stability derivatives are presented in 
figure 8. The experimental values of C:r< show fair agreement with the 
wing-alone values calculated by the method of reference 7. The experi-
mental maximum lift-drag ratio is compared with two theoretical curves. 
The upper curve (ref. 8), 1 , was obtained from the experimental 2 fDC
 
drag at zero lift CDC0 and the induced drag factor for potential 
flow for an elliptical loading l/tA and represents the condition of 
the resultant force normal to the local relative wind (full leading-edge 
F0^D_suction). The lower curve, , was obtained from the experi-  
meital drag at zero lift CDC0 and the experimental lift-curve slope 
CL, and represents the condition of the resultant force due to angle of 
attack normal to the wing chord. 
For the present wing, the experimental values of (L/D)max lie 
between the two theoretical curves up to a Mach number of about 0.85, 
then approach the curve which is based on the assumption that the 
resultant force is normal to the wing chord. The decrease in experi-
mental values of (L/D) 11,	 above a Mach number of 0.85 is due to a 
higher drag due to lift at the higher Mach numbers (fig. 9). 
CONCLUSIONS 
The results of the present investigation of the aerodynamic char-
acteristics in pitch and sideslip at high subsonic speeds of a wing-
fuselage combination having a triangular wing of aspect ratio ii-, a 
leading-edge sweep angle of 45 0 , and with an NACA 65A006 airfoil section 
parallel to the plane of symmetry indicate the following conclusions: 
CONFIDENTIAL
NACA RM L53G1I a	 CONFIDENTIAL	 1 
1. The effective-dihedral parameter Cj,, CL determined at low lift 
'  
coefficients was found to be considerably higher than would be expected 
from available methods - particularly in the range of Mach number from 
0.15 to 0.92. 
2. The wing-fuselage combination was longitudinally unstable at 
lift coefficients from approximately 0.50 to 0.70 and at Mach numbers 
from 0.40 to 0.80. At higher Mach numbers the high-lift stability 
appeared to improve. 
Langley Aeronautical Laboratory, 
National Advisory Committee for Aeronautics,

Langley Field, Va., July 1, 1953. 
CONFIDENTIAL
8	 CONFIDENTIAL	 NACA BM L53G14a
REFERENCES 
1. Kuhn, Richard E., and Wiggins, James W.: Wind-Tunnel Investigation 
of the Aerodynamic Characteristics in Pitch of Wing-Fuselage Combi-
nations at High Subsonic Speeds. Aspect-Ratio Series. NACA 
RM L52A29, 1952. 
2. Kuhn, Richard E., and Fournier, Paul G.: Wind-Tunnel Investigation 
of the Static Lateral Stability Characteristics of Wing-Fuselage 
Combinations at High Subsonic Speeds - Sweep Series. NACA EM L52Glla, 
1952. 
3. Hensel, Rudolph W.: Rectangular-Wind-Tunnel Blocking Corrections 
Using the Velocity-Ratio Method. NACA TN 2312, 1951. 
4. Gillis, Clarence L., Poihamus, Edward C., and Gray, Joseph L., Jr.: 
Charts for Determining Jet-Boundary Corrections for Complete Models 
in 7- by 10-Foot Closed Rectangular Wind Tunnels. NACA WE L-123, 
194 5. (Formerly NACA ARE L5G31.) 
Toll, Thomas A., and Queijo, M. J.: Approximate Relations and Charts 
for Low-Speed Stability Derivatives of Swept Wings. NACA TN 1581, 
1948. 
6. Fisher, Lewis R.: Approximate Corrections for the Effects of Com-
pressibility on the Subsonic Stability Derivatives of Swept Wings. 
NACA TN 1854, 19I9. 
7. DeYoung, John: Theoretical Additional Span Loading Characteristics 
of Wings With Arbitrary Sweep, Aspect Ratio, and Taper Ratio. NACA 
TN 1491, 1917. 
8. Jones, Robert T.: Estimated Lift-Drag Ratios at Supersonic Speed. 
NACA TN 1350, 1947. 
5 .
 
CONFIDENTIAL
NACA RM L53G14a	 CONFIDENTIAL 
L ateral force 
Relative wind
Yawing moment
Drag 
moment
Lift I 
Rolling moment
Pitching moment 
-	 >Drdg 
Relative wind 
Figure 1.- Stability system of axes shoving positive direction of forces,

moments, and angles. 
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Figure 6.- Longitudinal stability characteristics of a wing-fuselage 
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CONFIDENTIAL
NACA RN L5G14a	 CONFIDENTIAL	 15 
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CONFIDENTIAL 
16	 CONFIDENTIAL	 NACA RN L53G14a 
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CONFIDENTIAL
NACA RN L5G14a	 CONFIDENTIAL	 17 
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CONFIDENTIAL
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Figure 7 . - Lateral stability characteristics of a wing-fuselage combination 
having a triangular wing of aspect ratio ii. Data are not corrected for 
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of sideslip was varied.
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NACA-Langley - 8-28-53 - 325


Wind-tunnel investigation of the aerodynamic characteristics in pitch and sideslip at high subsonic speeds of a wing fuselage combination having a triangular wing of aspect ratio 4

Wind-tunnel investigation of the aerodynamic characteristics in pitch and sideslip at high subsonic speeds of a wing fuselage combination having a triangular wing of aspect ratio 4

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