Fields, E M

Schult, E D

Strass, H Kurt

English

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I 
. j 
RM L50B08 
RESEARCH MEMORANDUM 
FREE-FLIGHT INVESTIGATION AT TRANSONIC AND SUPERSONIC 
SPEEDS OF THE ROLLlliG EFFECTIVENESS OF A PARTIAL-SPAN 
AILERON ON AN INVERSELY TAPERED SWEPT BACK WllJG 
By H. Kurt Strass, E. M. Fields, and E. D. Schult 
Langley Aeronautical Laboratory 
Langley Air Force Base, Va. 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
May 1, 1950 
https://ntrs.nasa.gov/search.jsp?R=19930086267 2020-06-17T13:13:39+00:00Z

J 
NACA EM L50B08 
NATIONAL ADVISORY COMMITTEE FOR AERONAurICS 
RESEARCH MEMORANDUM 
FREE-FLIGHT INVESTIGATION AT TRANSONIC ANIJ SUPERSONIC 
SPEEDS OF THE ROLLING EFFECTIVENESS OF A PARTIAL--SPAN 
AILERON ON AN INVERSELY TAPERED SWEPrBACK WING 
By H. Kurt Strass, E. M. Fields, and E. D. Schult 
SUMMARY 
An investigation of the rolling effectiveness at transonic and 
supersonic speeds of an inversely tapered sweptback wing having a 
10-percent-thick cambered airfoil section and outboard partial-epan 
controls has been made by means of rocket-propelled test vehicles. The 
variation of rolling effectiveness with Mach number was smooth with the 
exception of a small discontinuity which occurred at a Mach number of 
0.95 for the models with the controls deflected 50 or - 50. The control 
effectiveness decreased by approximately 60 percent over the range from 
Mach number 0.8 to 1.6. Positive and negative aileron deflections were 
eQually effective in producing roll throughout the entire speed range. 
The angle of attack for zero lift of the cambered airfoil was indicated 
to decrease smoothly from a negative value of approximately 10 at sub-
sonic speeds to approximately 00 at a Mach number of 1.6. 
INTRODUCTION 
A wing-aileron configuration having inverse taper (A = 1.63), 
aspect ratio of 3.1, 400 sweepback of the 5O-percent free-etream chord 
line, and employing a 10-percent-thick low-drag-type airfoil normal to 
the 50-percent free-etream chord line was tested as a part of the wing-
aileron rolling-effectiveness program now being conducted by the 
Langley Pilotless Aircraft Research Division utilizing rocket-propelled 
test vehicles in free flight. The true-contour flap-type ailerons were 
hinged at the 73.04-percent free-etream chord line and extended over 
38.7 percent of the semispan and simulated sealed controls in t~at there 
was no gap at the hinge line. The investigation was made by means of 
the free-flight techniQue described in references 1 and 2 and data were 
obtained over the Mach number range from approximately 0.7 to 1.70 at 
2 NACA RM L50B08 
Reynolds numbers from 2.5 X 106 to 8.,5 X 106. Six models were tested; 
two with 50 aileron deflection, two with 00 aileron deflection, and two 
with -5° aileron deflection. 
A 
b 
c 
M 
m 
p 
pb/2V 
6~) \2V i 
w 
R 
S 
~o 
e 
SYMBOLS 
diameter of circle swept by wing tips, when body is rotated 
about its longitudinal axis, feet 
chord of wing parallel to free stream, feet 
total-drag coefficient, based on total exposed wing area S 
free-etream Mach number 
concentrated couple, applied near wing tip in plane normal 
to wing-chord plane, inch- pounds 
model rate of roll, radians per second 
wing- tip helix angle, radians 
pb/2V due to wing incidence, radians 
pb/2V due to aileron deflection alone, radians 
Reynolds number 
total exposed wing area (three wing panels), s~uare feet 
area of two wing panels taken to the center line, s~uare 
feet 
wing angle of attack for zero lift, measured from chord 
line, degrees 
angle of wing twist produced by m at any section along 
wing span normal to wing-chord plane, radians 
NACA RM L50B08 
elm wing torsional-etiffness parameter, radians per inch-
pound 
ratio of chord of wing at tip to chord of wing at model 
center line obtained by extending leading and trailing 
edges to model center line 
incidence of wing measured in free-etream direction, 
average for three wings, degrees 
deflection of each aileron measured in free-stream direc-
tion, average for three wings, degrees 
TEST VEHICLES AND TESTS 
The geometrical characteristics of the test vehicles are given in 
table I. The general arrangement of the test vehicles is shown in the 
photograph presented as figure 1 and the sketch in figure 2. The 
section airfoil ordinates of A-A and B-B in figure 2 are given in 
table II. 
3 
The wings of the test models were stiffened by means of 0.020-inch 
steel plates glued into the upper and lower surfaces . In addition, a 
central core of ~-inch aluminum alloy was used to provide additional 
32 
protection to the leading and trailing edges of the wings and to stiffen 
the ailerons to prevent movement under the air loads encountered in 
flight. The torsional characteristics of the test wings using this type 
of construction are presented in figure 3. Based on unpublished data, 
the test wings were sufficiently rigid to restrict the loss in control 
effectiveness to approximately 10 percent of the rigid-wing value at a 
maximum Mach number of 1.70. 
Because of the relatively small moment of inertia about the roll 
axis, the measured values of pb/2V are substantially steady-etate 
values even though the test vehicles experienced an almost continuous 
rolling acceleration or deceleration . The data presented herein have 
not been corrected for inertia effects. Except for abrupt changes 
of pb/2V with Mach number, which usually occur in the region between 
M = 0.90 to 0.97, this correction is estimated to be within 3 percent. 
Due to the fact that the airfoil section used was cambered, it was 
necessary to obtain data for the wing-aileron configuration with the 
control set at 00 in order to separate the effect of camber from the 
effect of the aileron. In addition, data were also obtained at a 
negative aileron setting in order to determine whether the control 
4 NACA RM L50B08 
effectiveness was independent of the direction of aileron setting. The 
wings were set into the body of the. test vehicle in such a manner that 
the lift due to the camber caused the models to rotate in a clockwise 
direction as seen from the r ear. Pos itive aileron deflection was 
considered t o be in a direction that would t end to increase the rotation 
due to camber . 
CORRECTIONS 
Due to limitations in constructional accuracy~ resulting in small 
deviations from design values~ the data were c orrected to a c ommon 
aileron setting and zero incidence to allow direct comparison between 
models. The rolling-effectiveness parameter pb/2V was assumed to 
vary linearly with aileron deflection. Errors in incidence were 
corrected by assuming that a steady-state rolling condition existed 
where the rolling moment caused by incidence was e~ual to the damping 
moment. In addition~ the lift-curve slope at any position along the 
span due to incidence was assumed to be e~ual to the lift-curve slope 
due to damping and the resulting general e~uation is as follows: 
2iw 1 + 2)" 
-- ---- == 0.0253 i w 57.3 1 + 3).. 
The correction for incidence was assumed to be c onstant throughout the 
entire Mach number range and bas been verified experimentally~ except 
for the region M ~ 0.85 to M ~ 1.0. (See reference 3.) Listed in 
table III are the measured values of incidence iw and aileron 
deflection ca . The pb/2V curves presented have been corrected 
to iw == 00 and 0a == 50 by the method just described. 
ACCURACY 
Based upon previous experience the ac curacy of the test results 
is estimated to be within the following limits : 
to limitations on model constructional accuracy) •• 
to limitations on instrumentation). 
pb/2V (due 
pb/2V ( due 
CDr •••• 
M 
±O.0025 
±O.0015 
±O.002 
ill .01 
NACA RM L50BOS 5 
RESULTS 
The Reynolds number range for the tests is shown in figure 4. The 
results of the present investigation are presented in figures 5 and 6. 
The total-drag coefficient CD and the rolling-effectiveness param-
T 
eter pb/2V for aileron deflections of 5°, -50, and 0° are presented 
in figure 5 as a function of Mach number . There was no measurable 
change in total-drag coefficient between the models with 0° and -5° 
aileron deflection although a considerable increase in drag was evident 
for the positive aileron setting in the supersonic range. The varia-
tion of pb/2V with Mach number was smooth with the exception of a 
small discontinuity at M = 0.95 for the models with the controls 
deflected 5° or -5°. 
The effectiveness of the aileron in producing roll is shown in 
figure 6, which gives the incremental change in rolling effectiveness 
for the positive and negative aileron deflection from the neutral 
setting. The data are plotted in the same direction to facilitate 
comparison. It can be seen that the effectiveness was essentially 
independent of direction of aileron deflection over the entire Mach 
number range investigated with a gradual loss in effectiveness of 
60 percent being experienced in going from approximately M = O.S to 
M = 1.6. 
As mentioned previously, the rolling velocity obtained with the 
controls neutral is an indication of the effectiveness of the camber in 
producing lift, or in the present case , rolling moment. From figure 5, 
the variation of pb/2V with Mach number f or models 1 and 2 shows that 
the effectiveness of the camber decreased continuously with increasing 
Mach number until at a Mach number of 1.6, the effectiveness was zero. 
The loss in effectiveness of camber in producing roll can be expressed 
as a shift in the angle of zero lift of the wing by applying the 
incidence correction factor given previously. This application has 
been made and the results are shown in figure 6. The camber is seen to 
be e~uivalent to approximately 1 0 incidence at subsonic speeds decreasing 
to approximately 00 at the highest supersonic speeds. The fact that the 
angle of zero lift changes smoothly throughout the speed range indicates 
that an abrupt change in the angle of trim due to this reason is unlikely 
to occur . 
CONCLUSIONS 
The following conclusions may be drawn from an investigation of 
the rolling effectiveness at transonic and supersonic speeds of an 
6 NACA RM L50B08 
inversely tapered swept back wing having a 10-percent-thick cambered 
air~oil section and outboard partlal-span controls: 
1. The variation of rolling effectiveness with Mach number was 
smooth with the exception of a small discontinuity at a Mach number 
of 0.95 for the test vehicles with positive and negative aileron 
deflections of 50 . There was a 60-percent loss in effectiveness 
between Mach numbers of 0.8 and 1.6. 
2. Positive and negative aileron de~lections of 50 were equally 
effective in producing roll throughout the entire speed range. 
3. The angle of attack for zero lift of the cambered airfoil was 
indicated to decrease smoothly from a negative value of approximately 
10 at subsonic speeds to approximately 00 at a Mach number of 1.6. 
Langley Aeronautical Laboratory 
National Advisory Committee for Aeronautics 
Langley Air Force Base~ Va. 
REFERENCES 
1. Sandahl~ Carl A.~ and Marino~ Alfred A.: Free-Flight Investigation 
of Control Effectiveness of Full-8pan 0.2-Chord Plain Ailerons 
at High Subsonic~ Transonic~ and Supersonic Speeds to Determine 
Some Effects of Section Thickness and Wing Sweepback. NACA 
RM L7D02~ 1947. 
2. Sandahl~ Carl A.: Free-Flight Investigation of the Rolling 
Effectiveness of Several Delta Wing-Aileron Configurations at 
Transonic and Supersonic Speeds . NACA RM L8D16~ 1948. 
3. Strass~ H. Kurt~ and Fields~ Edison M.: Flight Investigation of the 
Effect of Thickening the Aileron Trailing Edge on Control Effec-
tiveness for Sweptback Tapered \{ings Having Sharp- and Round-Nose 
Sections. NACA RM L9L19~ 1950. 
I 
~ 
NACA RM L50B08 
TABLE I 
GEOMETRIC CHARACTERISTICS OF TEST VEHICLES 
Wing semispan, ft . . . 
Total exposed wing area 
Area of two wing panels 
line, 81 , s<l ft 
Aspect ratio, A •••• 
Taper ratiO, ~ ••.• 
(3 wing panels), S, s<l ft •. 
taken to model center 
8weepback of 5O-percent free-etre~ chord, deg 
Position of inboard end of aileron~ percent semispan 
Position of outboard end of aileron, percent semispan • 
Ratio of aileron span to wing span • • • • • • . • • • 
Average moment of inertia about roll axis, slug-ft2 •• 
Airfoil profile normal to 50-percent free-etream chord, 
(Republic Aviation Corp. designation) 
Body length, in. ..... 
Body diameter, in. 
R-4, 
7 
0.99 
1.563 
1.261 
a3.1 
a l .63 
40 
55 
93·7 
0.387 
0.083 
40-1710 x 
56 
5 
Body contour • • • • . . . • • See reference 1 
a Obtained by extending leading and trailing edges to model center line. 
~ 
8 NACA RM L50B08 
TABLE II 
ORDINATES OF AIRFOIL SECTIONS IN FIGURE 2 
[All distances measured in inche~ 
Section A-A Section B-B 
Ordinate Ordinate 
Stationl--------.--------t St ation 
Upper surface Lower surface Upper surface Lower surfac e 
o 
.046 
.069 
.115 
.229 
.458 
.687 
.916 
1.374 
1.832 
2.290 
2 .749 
3.207 
3.665 
4.123 
4.581 
5.039 
5.497 
5·955 
6.413 
6.871 
7·329 
7.788 
8.246 
8.704 
9.162 
o 
.061 
.074 
.096 
.136 
.191 
.232 
.265 
.315 
. 352 
.379 
. 396 
.407 
.411 
.407 
.397 
. 380 
.356 
.327 
.292 
.252 
.208 
.161 
.111 
.058 
.007 
L.E. radius: 0.035 
T.E. radius: 0.007 
o 
.050 
.061 
.079 
.109 
.149 
.178 
.200 
.232 
.254 
.269 
.278 
.282 
.282 
.276 
.'265 
.249 
.230 
.205 
.177 
.146 
.115 
.082 
.050 
.022 
.007 
0 0 
.033 
.050 
.084 
.167 
.335 
.502 
.669 
1.003 
1.338 
1.672 
2.007 
2 .341 
2.676 
3.010 
3.345 
3.679 
4.014 
4.348 
4.682 
5·017 
5.351 
5.686 
6.020 
6.355 
6 .689 
L.E. radius: 
T.E. radius: 
0 
.043 .035 
.052 .043 
.068 .055 
.097 .078 
.138 .108 
.169 .129 
.193 .145 
.230 .169 
.257 .186 
.277 .197 
.289 .203 
.297 .206 
.300 .206 
.298 .202 
.290 .194 
.278 .182 
.260 .168 
.239 .149 
.213 .129 
.184 .107 
.152 .084 
.117 .060 
.081 .036 
.042 .016 
.007 .007 
0.025 
0.007 
NACA RM L50B08 9 
TABLE III 
MEASURED VALUES OF iw ANI) 0a 
MaCiel i W °a 
( Cieg) ( Cieg) 
1 0.05 0 
2 .01 0 
3 -.03 -4.74 
4 -.03 -4.74 
5 .01 5. 04 
6 .02 5.08 
• 
NACA EM L50B08 CONFIDENTIAL 11 
o 2 4 6 8 10 12 
INCHES 
Figure 1.- Photograph of typical test vehicle. 
L-61523 

NACA RM L50B08 13 
.1702 
r----"--1---6.383 ----1--+-0---- 6. 36 7 -----10_+_-
'I 
0.020 steel 
Sti ffener 
.10c 
.70c 
Plan form 
All d/mensions ore Ii? inches 
.5344 
Hinge line 
at .7304c 
2.148 4031 
Fuse/age 
-----+ 
2.50 
t 
.020 .steel st iflener 
covert?d w ith _ ---, 
1/32 vene:: .~ Beech , 3/32 24S-T3 aluminum 1 allot;! 
:~ y:::..; 
I 
.I0c 
Typical section 
Figure 2.- Sketch of test wi ng. Airfoil section ordinates of A-A and B-B 
are given in t abl e I I. 
14 
.4 clJOrd line ---------.. 
---r--
/--r-
o 
o 4 6 
Distol7ce from ruse/age, d) il7. 
NACA RM 150B08 
Plane of tor7qe 
application 
Fuse/aqe 
10 
Figure 3.- Average measured variation of torsional rigi dity with distance 
from fuselage . 
NACA RM L50B08 15 
R 
/0 x/O 6 
8 / 
/' V 
/ V 6 
./" V 
V V 
./ 
V V 
4 
/ V 
2 
~ 
o I I 
.6 .C3 10 12 14 1.6 It} 
M 
Figure 4.- Variation of Reynolds number with Mach number) based on 
an average exposed wing chord of 0.681 foot. 
16 
.08 
o 
.6 
./2 
.08 
p6/2V 
.04 
o 
-.04 
.6 
30nd4-\ L50nd~ ---
...... }. 
jJ ~-
III 
-
.P/ 
-~- ........ 
.8 ;0 1.2 1.4 
M 
Model 
/ 
2 -- - --
3 -----
4 -----
-............. 5 -----
-- ::::"'''''''',::: 6 - - ---~-
--
..... 
5and6J ~ 
" 
"r-. 
.::: 
--
-- -=:-:::: :;-;:::.... ............... 
---~ ~ 
---
---
-..-.. -
-
NACA RM L50B08 
-::.. 
-
/.6 
00 
(deq) 
o 
0 
-5 
-5 
.5 
.5 
-
1.8 
lw (deq) 
o 
0 
0 
0 
0 
0 
-~  r-- '--
-
~ 
~
.8 1.0 1.2 1.4 /.6 I.B 
M 
Figure 5.- Variation of rolling e ffectivene ss par amet er pb/ 2V 
a nd total-drag coeffic ient CDr wi th Mach number . 
, ' 
NACA RM L50B08 
.Od 
o 
.6 
/ 
o 
-2 
.6 
r- ..... 
'~ 
£00: -5"{toword cOl7Jb&~ 
/ ",,,,:: ' ~ r--r--r--
Loa=+5°(O't:LIO'!/ rrom comber) 
I I I J I 
.8 10 12 1.4 
M 
~ V--
.,./ 
l...---~ 
.,./ 
/.0 1.2 1.4 
M 
17 
16 /.8 
-
-
~ 
I I 
/.6 /.8 
Figure 6.- Variation of aileron rolling effectiveness 6(pb/2V)Oa 
and effective angle of attack for zero lift ~o with Mach number. 
NACA - Langley Field, Vd. 



Free-flight Investigation at Transonic and Supersonic Speeds of the Rolling Effectiveness of a Partial-span Aileron on an Inversely Tapered Sweptback Wing

Free-flight Investigation at Transonic and Supersonic Speeds of the Rolling Effectiveness of a Partial-span Aileron on an Inversely Tapered Sweptback Wing

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