Watson, James M

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RM L50H28a J FILE Copy	 CONFIDENTIA :
NACA 
RESEARCH MEMORANDUM 
EFFECT OF AN END PLATE ON THE AERODYNAMIC CHARACTERISTICS 
OF A 20.55 0 SWEPTBACK WING WITH AN ASPECT RATIO OF 2.67 
AND A TAPER RATIO OF 0.5 
	
TRANSONIC -BUMP METIOD	 4 
By James M. Watson
N"L viP)	 TTff 
Langley Aeronautical Labora
	
- 
Langley Air Force Base, Va. 
	
c:r	 • - 
UNCLASSIFIED DATE: 8-23-54. 
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 the revelation of Its contents In any 
Information so classified may be Imparted only to persons in the military and naval services of the United 
States, appropriate civilian officers and employees of the Federal Government who have a legitimate Interest 
thereto, and to United States citizens of mown loyalty and discretion who of necessity must be Informed thereof. 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
December 21, 1 
ON Fl DENTIAL
Restriction/Classification Cancelled
https://ntrs.nasa.gov/search.jsp?R=19930086374 2020-06-17T23:31:56+00:00Z
NACA 1*4 L5OH28a	
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
EFFECT OF AN END PLATE ON THE AERODYNAMIC CHARACTERISTICS. 
OF A 20.550
 SWEPTBACK WING WITH AN ASPECT RATIO OF 2.67 
AND A TAPER RATIO OF 0.5 
TRANSONIC-BUMP METHOD 
By James M. Watson 
SUMMARY 
An' investigation by the transonic-bump method conducted in the 
Langley high-speed 7- by 10-foot tunnel to determine the effect of an 
end plate on the aerodynamic characteristics of a 20.550 sweptback wing 
with an aspect ratio of .2.67 and a taper ratio of 0.5 indicated that 
the end plate affected the aerodynamic characteristics of the wing at 
transonic Mach numbers in the same manner as has been observed for swept 
and unswept wings at low speeds. 
INTRODUCTION 
The feasibility of using end plates on wings is being investigated 
as part of the National Advisory Committee for Aeronautics program to 
improve the aerodynamic characteristics of wings at transonic speeds. 
Previous investigations at low subsonic speeds (references 1 and 2) have 
generally shown that the addition of end plates to wings resulted in 
increased lift-curve slopes with only small decreases in maximum lift-
drag ±atios. In addition, improvements in the lateral stability charac-
teristics and increased control effectiveness have resulted from the 
addition of end plates to swept wings (reference 1). 
The present investigation was made by the transonic-bump 'method 
(reference 3) in the Langley high-speed 7- by 10-foot tunnel . to determine 
the aerodynamic characteristics in pitch at transonic speeds of two semi-
span wings having an aspect ratio of 2.67, a taper ratio of 0.5, and 
sweepback of 20.55°,. one wing with and one wing without an end plate. 
The data were obtained through an angle-of-attack range of -li.° to 120 
and a Mach number range of 0.6 to 1.18. 
Restriction/Classification Cancelled
Restriction/Classification Cancelled
2	 CONFIDENTIAL	 NACA EM L501128a 
SYMBOLS 
CL	 lift coefficient (ce pane lift) 
qS 
CD
	
drag,-coefficient (wice panel drag) 
qS 
Cm	 pitching-moment coefficient referred to 0.25 
(Twice panel pitching moment" 
qSE	 I 
CB	 bending-moment coefficient about root chord line (atplane of 
fRoot bending moment\ 
symmetry)	
b	 ) q22 
q	 effective dynamic pressure over span of model, pounds per 
square foot (pV2/2) 
S	 twice wing area of semispan model, 0.167 square foot 
(L/D)max maximum lift-drag ratio 
CL/a, per degree of angle of attack 
mean aerodynamic chord of wing, 0.25 foot (^2 b1'2 1 c2) 
local wing chord, feet 
twice span of semispan model, 0.67 foot 
spanvise distance from wing root, feet 
air density, slugs per cubic foot 
free-stream velocity, feet per second 
effective Mach number over span of model 
local Mach number 
CLa 
c 
c 
b 
y 
P 
V 
M 
M1
I 
CONFIDENTIAL
NACA RN L5OH28a	 CONFIDENTIAL	 3 
Ma	 average chordwise local Mach number 
B	 Reynolds number of wing based on C 
a	 angle of attack, degrees 
Subscripts: 
M	 at constant Mach' number 
CL=O	 at zero lift
MODEL AND APPARATUS 
The two semispan wings used in the investigation had 20.550
 
of
sweepback referred to the quarter-chord line, an aspect ratio 2.67, and 
a taper ratio 0.5. These wings were constructed of i -inch sheet steel 
with the leading and trailing edges beveled (fig. 1). A - - inch sheet-
steel end plate with beveled leading and trailing edges was brazed to 
the tip of one of the wings. The wings were mounted on an electrical 
strain-gage balance enclosed within the bump. The pitching moments were 
measured about the quarter chord of the mean aerodynamic chord and the 
bending moments were measured about the root chord. The air flow into 
and out of the balance chamber around the root of the wing was kept at 
a minimum by using a very finely knit, sponge-like material glued beneath 
the bump surface and free to slide on the surface of the models as a seal. 
(See fig. 2.)	 -
TESTS 
The tests were conducted in the Langley high-speed 7- by 10-foot 
tunnel and used the transonic-bump method. This method is an adaptation 
of the NACA wing-flow technique for obtaining transonic speeds and co1a-
sists in the mounting of a model in the high-velocity flow field generated 
over the curved surface of a bump located on the tunnel floor. (See 
reference' 3.)	 -	 - 
Typical contours of local Mach numbers in the region of the model 
location on the bump, obtained from suveys with the clear-bump, are 
shown In figure 3. A Mach number gradient exists over the span of the 
model; this gradient results in a difference of about 0.05 at the lowest 
Mach number, 0.06 at the highest Mach number, and a maximum difference 
of about 0.08 at a Mach number of about 1.0. The chordwise Mach number 
CONFIDENTIAL
CONFIDENTIAL
	 NACA EM L50H28a 
variation is generally, less than 0.01 except at the higher Mach numbers 
where the variation becomes as great as 0.03. No attempt has been made 
to determine the effects of these spanwise and chordwise variations in 
Mach number. The long-dashed lines shown near the wing root indicate 
the extent of the bump boundary layer and represent a local Mach number 
5 percent below the maximum value at each chordwise station. The effec-
tive test Mach number was obtained from contour charts similar to those 
of figure 3 and from the relationship 
b/2 
M=f 
The variation of average test Reynolds number with Mach number is 
presented in figure Ii. 
Force and moment data were obtained for the wings through a Mach 
number range of 0.60 to 1.18 and an angle-of-attack range of - l IP to 120. 
Jet-boundary corrections have not been applied inasmuch as the 
boundary conditions to be satisfied have not been accurately defined. 
These corrections are believed to be small, however, inasmuch as the 
effective flow field is large compared to the small size of the wings. 
- RESULTS AND DISCUSSION 
The aerodynamic characteristics in pitch of the wings, one wing 
with and one wing without an end plate, are presented in figure 5. A 
comparison of the various parameters for the two configurations is given 
in figure 6. The slopes summarized in figure 6 were taken over a lift-
coefficient range of about ±0.1. 
Lift and Drag Characteristics 
The lift-curve slopes (CL)M (fig. 6) of the wing without an end 
plate increased from a value of about 0.056 at a Mach number of 0.6 to 
0.016 at a Mach number of 0.94. Mach number increases above 0.94 
decreased the value of (CL ) to about 0.062 at M = 1.18. The addition 
of the end plate caused an increase in (CL"\ (at the low -lift coefficients) 
of 0.015 at M = 0.6, 0.024 at M= 0.94, and 0.015 at M 1.18. The 
higher lift-curve slope caused by the wing with the end plate indicates 
CONFIDENTIAL
NACA RM L5OH28a	 CONFIDENTIAL	 5 
an increase in the effective aspect ratio. A calculated effective aspect 
ratio of 3.94 was obtained from reference 4 for the wing with end plate. 
The estimated lift-curve slopes shown in figure 6 were obtained 
from references 4 and 5 (by the use of the Mach number expansion presented 
therein) and indicate excellent agreement at the lower test Mach numbers. 
The drag coefficient (fig. 6) of the wing with an end plate, at a 
lift coefficient of zero, was slightly higher than that of the wing with-
out an end plate. The drag rise occurred at a slightly lower Mach number 
for the wing with an end plate. At higher lift coefficients (approx. 0.5) 
the drag of the wing with an end plate was approximately the same as that 
of the wing without an end plate (fig. 5); this fact indicated that the 
reduction in induced drag of the wing with an end plate, due to the higher 
effective aspect ratio, approximately counterbalanced the drag of the end 
plate. The value of (L/D)max of the wing with an end plate was about the 
same as the value for the wing without an end plate (fig. 6) throughout 
the Mach number range investigated. 
Root-Bending-Moment Characteristics 
The variation of root-bending-moment coefficient with lift coeffi-
cient (fig. 5 or 6) of the wing with an end plate was about 10 percent 
larger over most of the Mach number range than that of the wing without 
an end plate and indicated that the end plate caused an outward movement 
of the lateral center of pressure (fig. 6). 
Pitching-Moment Characteristics 
The wing with an end plate (fig. 5) exhibited an increase in longi-
tudinal stability at the .10w lift coefficients over that of the wing 
without an end plate. The aerodynamic center, near zero lift coefficient, 
of the wing without an end plate was located at about 0.14Eat M = 0.6 
and moved rearward to about 0.43 at M = 1.18. The aerodynamic center 
of the wing with an end plate showed the same general trends with Mach 
number but was from 4 to 6 percent of the chord behind that of the wing 
without an end plate (see fig. 6). 	 - 
CONCLUSIONS 
An investigation by the transonic-bump method was conducted in the 
Langley high-speed 7- by 10-foot tunnel to determine the effect of an 
end plate on the aerodynamic characteristics of a 20.55 0 sweptback wing 
with an aspect ratio of 2.67 and a taper ratio of 0.5. The addition of 
CONFIDENTIAL
6	 CONFIDENTIAL	 NACA EM L5OH28a 
the end plate increased the lift-curve slope, had a negligible effect 
on the maximum lift-drag ratio, and increased the longitudinal, stability 
at all Mach numbers investigated. Similar results have been previously 
observed from the addition of end plates to swept and unswept wings at 
low speeds. 
Langley Aeronautical Laboratory 
National Advisory Committee for Aeronautics 
Langley Air Force Base, Va. 
REFERENCES 
1. Riebe, John M., and Watson, James M.: The Effect ,
 of End Plates on 
Swept Wings at Low Speed. NACA TN
	 , 1950. 
2 Hemke, Paul E.: Drag of Wings with End Plates. NACA Rep. 267, 1927. 
3. Sclmeiter, Leslie E., and Ziff, Howard L.: Preliminary Investigation 
of Spoiler Lateral Control on a 42 0
 Sweptback Wing at Transonic Speeds. 
NACA RN L7F19, '191-7. 
1. Von Krmn, Th., and Burgers, J. M.: General Aerodynamic Theory - 
Perfect Fluids. Airfoils and Airfoil Systems of Finite Span. Vol. II 
of Aerodynamic Theory, div. E., oh. IV, sec. 19, W. F. Durand, ed.., 
Julius Springer (Berlin), 1935, pp. 211-212. 
5. DeYoung, John: Theoretical Additional Span Loading Characteristics of 
Wings with Arbitrary Sweep, Aspect Ratio, and Taper Ratio. NACA 
TN 1491, 1947.
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NACA RN L50H28a	 CONFIDENTIAL	 9 
Figure 2.- Typical model installation with sponge seal on the transonic
bump in the Langley high-speed 7- by 10-foot tunnel. 
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of 0.5, one wing with and one wing without an end p la te .  


Effect of an End Plate on the Aerodynamic Characteristics of a 20.55 Degree Sweptback Wing with an Aspect Ratio of 2.67 and a Taper Ratio of 0.5. : Transonic-Bump Method

Effect of an End Plate on the Aerodynamic Characteristics of a 20.55 Degree Sweptback Wing with an Aspect Ratio of 2.67 and a Taper Ratio of 0.5. : Transonic-Bump Method

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