Naeseth, Rodger L

O'Hare, William M

English

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NACA 
RESEARCH MEMORANDUM 
THE EFFECT OF AILERON SPAN AND SPANWISE LOCATION ON THE LOW- I
SPEED LATERAL CONTROL CHARACTERISTICS OF AN UNTAPERED

WING OF ASPECT RATIO 2.09 AND 450 SWEEPEACK 
By Rodger L. Naeseth and William M. O'Hare 
Langley Aeronautical Laboratory

Langley Air Force Base, V, 
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NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON
February 10, 1950
https://ntrs.nasa.gov/search.jsp?R=19930086182 2020-06-17T13:23:35+00:00Z
NACA EM L9LO9a
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
THE EFFECT OF AILERON SPAN AND SPAEWISE LOCATION ON TEE LOW-

SPEED LATERAL CONTROL CHARACTERISTICS OF AN UNTAPERED

WING OF ASPECT RATIO 2.09 AND 450
 SWEEPBACK 
By Rodger L. Naeseth and William M. O'Hare 
SUMMARY 
A low-speed lateral-control investigation was made of an untapered 
complete wing of aspect ratio 2.09 and 450 sweepback with 0.25-chord 
flap-type ailerons having various spans and spanvise locations. The 
tests were made at a Reynolds number of 3,100,000. 
Aileron effectiveness increased with increase in aileron span. An 
outboard aileron was appreciably more effective than an inboard aileron 
of the same span. The variation of rolling-moment coefficient with 
aileron-deflection angle was linear over the range of angles of attack 
for an aileron-deflection range of ±200. Aileron effectiveness could be 
accurately predicted by the methods of NACA TN 1674 for low angles of 
attack.
INTRODUCTION 
An investigation was made in the Langley 300 MPH 7- by 10-foot 
tunnel to determine the effect of aileron span and spanwise location 
on the aileron effectiveness of a 25-percent-chord sealed aileron on an 
untapered complete wing of aspect ratio 2.09 and 14.5° sweepback. Lateral-
control data are presented for the wing equipped with several spans of 
outboard and inboard ailerons. The basic aerodynamic characteristics in 
pitch and the lateral-stability parameters of the wing are also presented 
herein. The results of the investigation are compared with the aileron-
effectiveness values calculated by the method presented in reference 1.
2
	 NACA RM L9LO9a 
COEFFICIENTS AND SYMBOLS 
The data are referred to the stability axes (fig. i), which area 
system of axes with the origin at the center of moments (0.25 mean aero-
dynamic chord (fig. 2)), and in which the Z-axis is in the plane of 
symmetry and perpendicular to the relative wind, the X-axis is in the 
plane of symmetry and perpendicular to the Z-axis, and the Y-axis is 
perpendicular to the plane of symmetry. 
The coefficients and symbols are defined as follows: 
CL	 lift coefficient (Lift/qS) 
CD	 drag coefficient (Drag/qS where Drag = -X when lit = 00) 
Cy	 lateral-force coefficient (Y/qS) 
Cm	 pitching-moment coefficient (M/qS) 
C 1	 rolling-moment coefficient (L/qSb) 
Cn	 yawing-moment coefficient (N/qSb) 
Y	 lateral force, pounds 
L	 rolling moment about X-axis, foot-pounds 
M	 pitching moment about Y-axis, foot-pounds 
N	 yawing moment about Z-axis, foot-pounds 
S	 wing area, 6.16 square feet 
q	 free-stream dynamic pressure, pounds per square foot (V2)2 
V	 free-stream velocity, feet per second 
p	 mass density of air, slugs per cubic foot 
C	 local wing chord, feet 
wing mean aerodic chord, feet
 (fb/2
* 	 c2) 
b	 wing span, 3.59 feet 
ba	 aileron span, feet
3 NACA BM L9LO9a 
Y	 lateral distance from plane of symmetry, measured parallel 
to Y-axis, feet 
Yo	 lateral distance from plane of symmetry to outboard end of 
aileron, measured parallel to Y-axis, feet 
yj	 lateral distance from plane of symmetry to inboard end of 
aileron, measured parallel to Y-axis, feet 
CG	 angle of attack of wing chord plane, degrees 
*	 angle of yaw, angle between relative wind and plane of 
symmetry measured in fl-plane 
ba	 aileron deflection relative to wing chord plane, measured in a plane perpendicular to aileron hinge axis and 
positive when trailing edge is down, degrees 
CL	 L =
Ncj 
Cla =
ICZ C, = - 
Li, 
Cfli,W 
Cy 
CY* =
Rolling-moment and yawing-moment coefficients presented represent 
the aerodynamic moments on a complete wing produced by the deflection 
of the aileron on only the right semispan of the wing. 
CORRECTIONS 
Jet-boundary (induced upwash) corrections based on unswept-wing 
theory (reference 2) were applied to the angle of attack and to the 
drag data. The data were also corrected for blockage effects by the 
methods of reference 3, and for model support strut tares.
14	 NACA RN L9LO9a

MODEL AND APPARATUS 
The complete-Wing model was mounted horizontally on a single strut 
support in the Langley 300 MPH 7- by 10-foot tunnel in such a manner as 
to permit measurement of all forces and moments acting on the model. 
The wing model of 1450
 sweepback and aspect ratio 2.09 had a taper ratio 
of 1.0, with neither twist nor dihedral, and had NACA 614A-010 airfoil 
sections normal to the leading edge (fig. 2). The model was fabricated 
by means of a sandwich-type construction consisting of a laminated 
mahogany core enclosed in a covering composed of i--inch sheet aluminum 
glued between sheets of i-inch fir. The right semispan of the wing was 
equipped with a 0.25-chord aluminum flap divided into four parts by cuts 
parallel to the plane of symmetry. The various flap segments were set 
at given deflections by means of hinge clamps on each, flap segment. All 
flap gaps, except the one between the deflected and undeflected flap 
segments, were sealed for all tests. Outboard ailerons (outboard end 
at 0. 983_) having spans of 0.983*
	
.- 
, 0.714l, 0.500*, and 0.258* as 
well as inboard ailerons (inboard end at plane of symmetry) having spans 
of O.483L and 0.242 were tested. 2	 2
TESTS 
All the tests were performed In the Langley 300 MPH 7- by 10-foot 
tunnel at an average dynamic pressure of approximately 99 pounds per 
square foot, which corresponds to a Mach number of 0.26 and a Reynolds 
number of 3.1 x 106, based on the wing mean aerodynamic chord of 1.72 feet. 
Data for each test were obtained through an angle-of-attack range 
from -60
 to beyond the wing stall. Lift, drag, and pitching-moment data 
were obtained for the wing at
	
= 00
 with 6a = 00, and lateral-
stability derivatives of the wing were obtained from tests at 
'V =±5. 
Lateral-control data were obtained with various spans of inboard and 
outboard ailerons at deflections ranging between ±200.
NACA RN L9LO9a	 5
DISCUSSION
Wing Aerodynamic Characteristics 
The lift, drag, and pitching-moment characteristics of the wing 
model are presented in figure 3. The value of CL, of 0.042 computed 
by the method presented in reference 4 agrees very well with the experi-
mental lift-curve slope (measured near CL = 00) of 0.041. The pitching-
moment-data indicate that the aerodynamic center was about 3 percent 
mean aerodynamic chord ahead of the /1 at low lift coefficients. The 
wing became stable for lift coefficients above approximately 0.5. 
The variation of the wing lateral-stability derivatives with lift 
coefficient is given in figure i. 
Lateral Control Characteristics 
Rolling-moment-coefficient and yawing-moment-coefficient data for 
various spans of inboard and outboard ailerons are given in figures 5 
and 6. The variation of rolling-moment coefficient with aileron 
deflection at a, = 00
 and 160 for various spans of aileron is shown in 
figure 7. The control effectiveness parameter C 1	 was obtained from 
cross plots of the data of figures 5 and 6 and is presented in figure 8 
as a function of angle of attack. Figure 9 shows the variation of 
aileron effectiveness with aileron span at a = 00
 for outboard ailerons. 
Rolling-moment characteristics. - The variation of C1 with ba 
was fairly linear over the range of angles of attack for the deflections 
tested, and aileron effectiveness increased with increase in aileron 
span (figs. 5 to 7). However, the effectiveness of inboard ailerons was 
appreciably less than for outboard ailerons of the same span (fig. ). 
The values of C 1	 for outboard ailerons decreased slightly with 
increase in angle of attack up to a Z. 150 (fig, 8). Above a 150, the 
effectiveness increased, and then decreased above the wing stall. The 
effectiveness of the inboard ailerons increased slightly with increase 
in angle of attack up to a 9. Above this angle, the variation of 
C 1	 with a was similar to the variation of the outboard aileron 
effectiveness. The values of C 1	 of outboard ailerons obtained from 
figure 8 at a = 0 are presented in figure 9 plotted against relative 
location of the inboard end of the aileron. This outboard spanwise 
effectiveness curve may be used to compute the effectiveness of inboard 
ailerons because the value of Cj
	 for an aileron spanning any portion
rA
	
NACA RM L9LO9a 
of the wing is the difference between the values of Cj
	
at the 
inboard end and Cj	 at the outboard end of the aileron span being 
considered (reference 5). Thus, values of C15 calculated by this 
method (from fig. 9) for the two inboard ailerons tested are found to 
agree with the experimental values shown in figure 8. 
Aileron effectiveness C 15 computed for the various aileron spans 
a 
on the wing of the present investigation by the theoretical and empirical 
relationships given in reference 1 show excellent agreement with the 
experimental values of figure 9 - indicating that C 1	 can be pre-
dicted satisfactorily by this method. 
Yawing-moment characteristics.- The sum of the yawing moments 
produced by equal up and down aileron deflections for each aileron 
configuration was generally negligible for small angles of attack 
(figs. 5 and 6). At moderate and large angles of attack, however, the 
yawing moments were adverse and generally increased with increase in a. 
and 5a except for the inboard aileron of	 = 0.242, which produced 
negligible yawing moments throughout the angle-of-attack and aileron-
deflection ranges (fig. 6(h)). At the higher angles of attack, 
the Cn/Ci ratio amounted to as much as -0.7, indicating that in some 
instances a sizeable rudder deflection may be required to perform a 
coordinated roll with an airplane using a. wing of this plan form 
(reference 6).
CONCLUSIONS 
A low-speed lateral-control investigation was made of an untapered 
complete wing of aspect ratio 2.09 and 11.50 sweepback with 0.25-chord 
ailerons having various spans and spanwise locations. The tests were 
made at a Reynolds number of 3,100,000. The results of the investigation 
indicate the following conclusions: 
1. Aileron effectiveness increased with increase in aileron span. 
An outboard aileron was appreciably more effective than an inboard 
aileron of the same span. 
2 The variation of rolling-moment coefficient with-aileron-
deflection angle was linear over the range of angles of attack for an 
aileron-deflection range of ±20.
NACA RN L9LO9a	 7 
3. Aileron effectiveness could be accurately predicted by the 
method of NACA TN 167 1 for low angles of attack. 
Langley Aeronautical Laboratory 
National Advisory Committee for Aeronautics 
Langley Air Force Base, Va. 
REFERENCES 
1. Lowry, John G., and Schneiter, Leslie E.: Estimation of Effectiveness 
of Flap-Type Controls on Sweptback Wings. NACA TN 1674, 1948. 
2. Gillis, Clarence L., Polhamus, 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 ARR L7G31, 
1945. 
3. Herriot, John G.: Blockage Corrections for Three-Dimensional-Flow 
Closed-Throat Wind Tunnels, with Consideration of the Effect of 
Compressibility. NACA RN A7B28, 1947. 
4. PoThamus, Edward C.: A Simple Method of Estimating the Subsonic Lift 
and Lumping in Roll of Sweptback Wings. NACA TN 1862, 1949. 
5. Weick, Fred E., and Jones, Robert T.: Re"sume and Analysis of N.A.C.A. 
Lateral Control Research. NACA Rep. 605, 1937. 
6. Fehiner, Leo F.: A Study of the Effect of Adverse Yawing Moment on 
Lateral Maneuverability at a High Lift Coefficient. NACA ARR, 
Sept. 1942.
Xb 
0 
Re/at
NACA RN L9LO9a 
Ii 
Lift 
t 
a 
Re/ati
Sect/on B-B 
z 
Figure 1.- System of stability axes. Positive values of forces, moments,

and angles are indicated by arrows.
NACA RN L9LO9a	 9 
oa) 
 
cd 
(j
0) 
..' a) 
1'..	 \D•H 
•0) 
'.0
0 
II N.
co 
0) 
Q)	 'dr-i 
q)	 tto
0 
a). 
Hil 
EZ 10 
0 •H 
04-' 
__	
N Cj +) cI—I 
0c 
o 
0)O'd 
,&	 •0) 
rcj
C'J0a) 
0)4) 0) 
l
10
	
NACA RM L9LO9a 
.1 
sI 
E
4 
.3 
28	 .2 
24
	
/ c) 
20
	
0 
/6 
/2 
8 
4 
-4 
-8 J	 I	 I	 I	 I	 I	 I	 I	 I	 I	 I	 I 
-4	 -.2	 0	 .2	 4	 .6	 .8	 tO 
Lift coefficient, CL 
Figure 3.- Aerodynamic characteristics in pitch of the wing. 5a = 00
CY31r 
Cal, 
Ci
NACA RM L9LO9a	 11 
rISM 
-g	 :2	 0	 .2	 4	 .6	 .8	 /0
Lift coefficient, CL 
Figure 4. - Variation of the parameters C, C, and C
	
with lift
coefficient. ba =
12
	
I'IACA RM L9LO9a 
02 
E
CZ-
 
-. 0/ 
-O2 
O3
	 & (deg) 
o-20 
Q-/0 
E-5 
CZ- 	 o2 
E5 
Q)	 20 
E 
E
O2 
-8 -4
	 0	 4	 8	 12	 16 20 24 28 
Angle of attack, cc, deg 
(a)	 = 0.983. 
Figure 5.- Variation of the lateral control characteristics with angle 
of attack. Outboard ailerons.
NACA BM L9LO9a	 13 
.03 
.02 
:
ICj 
rz 
Qt
02 
MI 
.02 
1°' 
E 
>	
-8 -4	 0	 4	 8	 12	 16 20 24 28 
Angle of attack, cc, deg 
(b)	 = 0.71.
Figure 5. -
 Continued.
114.
	
NACA RM L9LO9a 
.0/
6 (deg) 
o-20
D 4 
02 
05 
Io 
'20 
.
-02 
-S	 -4	 0	 4	 8	 12	 16 
Angle of ctfack,a,deg 
b (c) E7 = 0.500. 
Figure 5.- Continued.
20 24 28
NACA RM L9LO9a	 15 
•1.' 
rz 
c) 
Qo 
E 
E 
Ot
	
6 (deg) 
o-20 
°-1O 
•q)	 02 05 
20 
E 
IZI 
E
I	 I	 I	 1	 I	 I	 1	 I	 I	 I	 I• 
	
-4	 0	 4	 8	 12	 16 
Angle of attack, ct, deg 
(d) ba
 = 0.258.
Figure 5.- Concluded.
20 24 28
16	 NACA EM L9LO9a 
.02------- - -- - - - 
i,  
ll 0
•= •= __ - 
- 
-02------- - - - - 
50 (deg)- - - - 
Q-20 
0
-10 - - - - 
0/0 - - - 
rZ, 20
-
-40	 4	 8	 12	 16	 20	 24	 28 
Angle of attack, ', dog
ba (a)	 _. = O.183. 
Figure 6. - Variation of the lateral control characteristics with angle of 
attack. Inboard ailerons. 
NACA RM L9LO9a	 11 
L.i 
ci 
.4-
cz 
E
- 
60(0 
o-20 
0/0 
20 
-4	 0	 4	 8	 12	 16 
Angle of attac/çcc,deg
q) 
'I-
 q) 
ci 
.4-
q) 
E 
EQ1 
-8 20 24 28 
(b) ba  = 0.242. 
Figure 6.- Concluded.
18
	
NACA RM L9LO9a 
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I ,
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to n 
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NACA RM L9LO9a
	
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NACA PM L9LO9a 
0 
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NACA-Langley - 7-2-54 - 125 


The effect of aileron span and spanwise location on the low-speed lateral control characteristics of an untapered wing of aspect ratio 2.09 and 45 degree sweepback

The effect of aileron span and spanwise location on the low-speed lateral control characteristics of an untapered wing of aspect ratio 2.09 and 45 degree sweepback

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