Fields, Edison M

Strass, H Kurt

English

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RM L9L19 
RESEARCH MEMORANDUM 
FLIGHT ThlVESTIGATION OF THE EFFECT OF THICKENING THE 
AILERON TRAILING EDGE ON CONTROL EFFECTIVENESS FOR 
SWEPTBACK TAPERED WINGS HAVING SHARP-
AND ROUND-NOSE SECTIONS 
By H. Kurt Strass and Edison M. Fields 
Langley Aeronautical Laboratory 
Langley Air Force Base, Va. 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
May 2, 1950 
https://ntrs.nasa.gov/search.jsp?R=19930086249 2020-06-17T13:14:48+00:00Z
• 
NACA RM L9L19 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
FLIGHT INVESTIGATION OF THE EFFECT OF THICKENING THE 
AILERON TRAILING EDGE ON CONTROL EFFECTIVENESS FOR 
SWEPTBACK TAPERED WINGS RAVING SRARP-
AND ROUND -NOSE SECTIONS 
By H. Kurt Strass and Edison M. Fields 
SUMMARY 
An investigation has been made, by means of rocket-propelled test 
vehicles, of the rolling effectiveness at transonic and supersonic 
speeds on a 38.10 sweptback tapered wing having NACA 0010-64 airfoil 
sections and equipped with partial - span ailerons of true contour and 
modified to a flat-side blunt trailing -edge shape . The results are 
compared with previous results for a similar modification to a wing 
having circular-arc sections. The results indicated that, for both 
airfoil sections with a control deflection of 50, the original true-
contour ailerons, which had trailing- edge angles of 210 for the 
circular-arc section and 170 for the round -nose section, showed a sharp 
loss in rolling effectiveness between Mach numbers of 0 . 9 and 1.0 with 
a reversal of control at a Mach number of approximately 0.95. Modifying 
the aileron by th~ckening the trailing edge to one-half the aileron 
thickness at the hinge line, which reduced the trailing-edge angle to 
approximately 80 , eliminated the control reversal and in general 
improved the effectiveness in the transonic range for both airfoil 
sections, although at some sacrifice in drag . 
INTRODUCTION 
The Langley Pilotless Aircraft Research Division is conducting an 
investigation of wing-aileron control a t subsonic, transonic, and 
supersonic speeds by means of rocket -propelled test vehicles in free 
flight. Investigation of an aileron on a tapered sweptback wing having 
a circular-arc section with a trailing- edge angle of 21 0 indicated a 
2 NACA RM L9L19 
control reversal near Mach number 1.0 which was eliminated by modifying 
the aileron to a flat-side blunt trailing -edge shape which effectively 
reduced the trailing- edge angle. This phenomenon has been investigated 
further with a round-nose section and the results are presented herein 
and compared with previous results. 
The tests were made with 0 .171 - chord outboard half-span ailerons 
with true contour and also with the trailing edge increased to one-half 
the thickness at the hinge axis. The wing had an aspect ratio of 3 . 61, 
a taper ratio of 0 . 455, a sweepback angle of 38 .10 measured at the 
quarter - chord line, and an NACA 0010 - 64 airfoil section normal to the 
44.8-percent - chord line. Drag as well as control effectiveness was 
measured . 
A 
b 
M 
m 
p 
pb 
2V 
SYMBOLS 
aspect ratio (~:) 
diameter of circle swept by wing tips (2 . 23 ft) 
mean aerodynamic chord parallel to model center line (0.625 ft) 
total drag coefficient based on total exposed wing area, 8 
wing - chord -plane incidence angle, degrees 
free - stream Mach number 
couple, applied at wing tip in a plane parallel to model center 
line and normal to wing-chord plane, inch-pounds 
model rate of roll, positive if clockwise when viewed from rear, 
radians per second 
rolling effectiveness parameter 
R Reynolds number based on c 
8 total exposed wing area (1. 563 sq ft) 
81 area of two panels taken to center line (1. 378 sq ft) 
V velocity of model along flight path, feet per second 
0a deflection of each aileron from Wing -chord plane, measured in a 
plane parallel to model center line, degrees 
e angle of wing twist (produced by m) at any section along wing 
span in a plane parallel to free stream and normal to wing-
chord plane, radians 
• 
NACA RM L9L19 
e 
m 
wing-torsional-stiffness parameter in a plane parallel to 
model center line, radians per inch-pound 
taper ratio 
TEST VEHICLE AND TESTS 
The geometrical arrangement of the test vehicles used in this 
investigation is given in table I and figures 1 and 2. The model 
finish consisted of several coats of clear lacquer rubbed almost glass-
smooth. The torsional-rigidity characteristics for the wing plan form 
tested are presented in figure 3. The parameter e was obtained by 
m 
applying a known couple at the tip of the wing and then measuring the 
variation of twist along the span. The torsional rigidity as given in 
figure 3 has been shown by unpublished data to be suff iciently rigid to 
limit t he loss of control effectiveness to a maximum of 20 percent at a 
Mach number of 1.8. 
3 
The test vehicles were of the same simple and relatively inexpen-
sive construction as has been used previously and are described iL 
reference 1. The measured values of aileron deflection and wing inci-
dence are estimated to be within ±O.lo and ±0.05°, respectively, of the 
actual values. The test vehicles were propelled to approximately M = 1.7 
by a two-stage rocket-propulsion system. All the data were obtained 
during the coasting period following burnout of the sustainer motor. 
The flight-path velocity was measured by the use of the CW Doppler radar 
technique and the rolling velocity by the use of spinsonde radio equip-
ment. Atmospheric data were obtained from radiosonde records. Drag 
information was obtained by the differentiation of the flight-path 
velocity curve against time. A more complete description of the tech-
nique is given in references 2 and 3. 
ACCURACY 
The estimated accuracies for the parameters measured in this 
invest i gation are as follows: 
pb/2V 
---5-- (due to limitations on model construction and 
measurement accuracy) 
pb/2V ( 5 due to limitation in instrumentation) 
~T 
M 
±0.00025 
±0.0001 
±0.0020 
±0.010 
4 NACA RM L9L19 
CORRECTIONS 
Because of the limitations in constructional accuracy, the 
resulting small deviations from the design values were measured and 
the data were corrected to unit aileron setting and zero wing incidence 
to allow direct comparison between models. The r olling effectiveness 
t pb parame er 2V was assumed to vary linearly with aileron deflection. 
Errors in incidence were corrected by assuming that a steady-state 
rolling condition existed when the rolling moment caused by incidence 
was equal to the damping moment. In addition, the lift-curve slope at 
any position along the span due to incidence was assumed to be constant 
and equal to the lift-curve slope due to damping. The resulting general 
e quat i on is as follows: 
where 
b. pb 
2V 
2 iw (1 + 2"- ) 
57. 3 1 + 3"- 0. 0281iw 
iw measured wing incidence in degrees 
"- taper ratio, tip chord to extended chord at model center 
line (0.455 ) 
The correction for incidence was assumed to be constant throughout t he 
entire Mach number range. This correction has been verified exper i -
mentally for an unswept rectangular plan form of aspect ratio 3.7 except 
for the region around M ~ 0.95 where a small abrupt di scontinuity 
occurs. (See fig. 4.) 
The effect of the moment of inertia about the roll axis on the 
measured value s of ~ was assumed to be negligible. (See reference l~ 
The pb/2V values from reference 1 when plotted in thi s paper were 
corrected to 0a = 50 streamwise to facilitate dire ct comparison. 
RESULTS AND DISCUSSION 
The variation of Reynolds number with Mach number i s shown in 
figure 5. A comparison of the control effectiveness of the two types of 
ailerons is presented in figure 6 as curves of the control effectiveness 
parameter pb/2V against Mach number. The original true-contour aileron 
exhi bited a very abrupt loss of control accompanied by a slight reversal 
in the range between M ~ 0 .9 and 1. 0 . Thi s condition wa s previously 
encountered on a wing of similar geometric characteristics but with a 
NACA RM L9Ll9 5 
different airfoil section (lO -percent circular arc). By increasing the 
thickness at the trailing edge to one-half the thickness at the hinge 
axis (reference 1), the aileron control was improved. The same 
technique was used in this investigation and yielded approximately the 
same results. The improvement in control is confined largely to the 
speed region below M ~ 1.2. In figure 7, a direct comparison is made 
of the results from the two series of tests. For the true-contour 
aileron, the curves of pb/2V from the present tests and from refer-
5 
ence 1 agreed well except for a limited region between M ~ 0.9 and 1.2 
where the aileron from the present tests was more effective. The blunt 
trailing-edge control used in the present tests was slightly more 
effective throughout most of the speed range than was the same type of 
control from reference 1 . In general the data seem to indicate that 
the effectiveness of a given type of control is only slightly affected 
by the thickness distribution of the forward portion of the airfoil. 
The use of the blunt trailing -edge aileron caused a large increase in 
total drag coefficient below the speed of sound; however, the drag-
coefficient increase above the speed of sound was less (see fig. 8). 
CONCLUDING REMARKS 
A free-flight investigation at transonic and supersonic speeds of 
the rolling effectiveness of ailerons on a sweptback tapered wing 
having a round NACA 0010-64 section with a trailing-edge angle of 170 
indicated that a large abrupt decrease in aileron effectiveness 
accompanied by reversal occurred in the transonic region. The loss of 
effectiveness was similar to that reported previously with a circular-
arc section having a trailing-edge angle of 210. 
An effective means of reducing this loss of control for either 
airfoil section was the use of a flat-side aileron, the contour of which 
was modified by thickening the trailing edge to a value of one-half the 
thickness at the hinge line . 
The drag of the wing was appreciably increased by the modification 
to the aileron at subsonic speeds, although at speeds greater than a 
Mach number of 1.0 the drag increase was small . 
Langley Aeronautical Laboratory 
National Advisory Committee for Aeronautics 
Langley Air Force Base, Va. 
6 NACA RM L9L19 
REFERENCES 
1. Sandahl, Carl A.: Free -Flight Investigation at Transonic and Super -
sonic Speeds of the Rolling Effectivene s s of Several Aileron 
Configurations on a Tapered Wing Having 42.70 Sweepback. NACA 
RM L8K23 , 1949 . 
2 . Sandahl, Carl A.: Free-Flight Investigation of Control Effectiveness 
of Full-Span, O. 2-Chord Plai n Ailer ons at High Sub s oni ~, Trans onic, 
and Supersonic Speeds to Determine Some Effects of Wing Sweepback, 
Taper, Aspect Ratio, and Section Thickness Ratio. NACA RM L7F 30, 
1947. 
3 . Sandahl, Carl A., and Marino, Alfred A.: Free-Flight Investigation 
of Control Effectiveness of Full -Span O.2- Chord Plain Ailerons 
at High Subsonic, Transonic, and Supersonic Speeds to Determi ne 
Some Effects of Section Thickness and Wing Sweep back. NACA 
RM L7D02, 1947. 
NACA RM L9L19 
TABLE I 
GEOMETRIC CHARACTERISTICS 
Aspect ratio . . . . • . . . . 
Total exposed wing area, square feet 
Taper ratio . . . . 
Mean aerodynamic chord, c, feet 
Sweepback of wing L.E., degrees 
Sweepback of wi ng c/4, degrees 
Sweepback of wi ng T.E., degrees 
Ratio of ail eron span t o wing span 
Average moment of i nert i a about roll axis, 
slug-feet2 . . . . . . . . . . . . . . . 
Location of hinge axi s , percent free-stream 
chord . . • . . . . . . . . . . • . 
Airfoil section normal t o 44 .8-percent free-stream 
chord ................... . 
True-contour aileron t r a iling-edge angle, normal 
to 44 .8 -per cent fre e - stream chord, degrees 
Modified ail er on trail i ng-edge angle, normal to 
44 . 8 -percent f ree- stream chord, degrees 
aObtained by extending wing to center line of fuselage. 
7 
3 . 61 
1. 563 
aO . 455 
0 . 625 
41. 6 
38 .1 
25 .2 
aO . 5 
0 . 0768 
82. 9 
NACA 0010 - 64 
17 
7 .8 

NACA RM L9L19 9 
Figure 1.- PhJtograph of typical model. 
l 
- - - -- - -- - - . - ~------... ---~--
NACA RM L9L19 11 
5'10Iame~r 
~-5plnsonde 
r-----------------
1-----L ________ _ 
Rocket motor ---' I 
56.0" ----------~---i-_1 
• 
,/ Rear-end vi(2W 
(a) Location of wings on bod'!. 
Flqure 2.-Genero/ arrongem@nt of models. 
12 NACA RM L9L19 
1-----9.634 ---~~~~~~2=-----T 
L ~od/nq edqp of 
Inlay ol.l5 chord 
line ~ 
foo.-- --9163 
10.865 
~ TrOI!/I7Q 
/0....---1---.-
edge of /17lof:! 
ot hinqe line, 
.829 chord 
4.1(J3 
Fuselage 
Model C"{Jl7ter line 
All d/mensiol75 il7 inche5 
Chord meosurements ,Dorallel to model t:8nter 1117t? 
" 
c l/3211Veneer tiTrue-
/~ contour 
q020 steel ~? ____ :;..-..~ Q;/erof7 
/n/ot) ~~ ~ 
Zvtodified Section A-A 
(6) Winq p/onform and profile. oilpron 
Figure 2. - Concluded. 
NACA RM L9L19 
~ 
I 
.8 x 10-4 
,6 
f-
o 
o 
/ V 
./ V 
~ / 
-
---
-
-- ~ I I 
2 4 8 
Distance from fuse/age.; In. 
Fiqure 3. - Variation of tors/onal-5tiffneS5 
parameter 8/177 with spon. 
13 
/0 
14 
.04 
p6/2V 
. 02 
o 
A: 3.7) A =1 
AIr -
flow 
Experimental 
curves 
NACA RM L9L19 
(]JIm: .970 
® 1m =.49° 
.6 .8 10 /.2 /,4 /.6 /.8 
M ~ 
F/qure 4. - LOl77porlsol7 (for unswept p/Cll7torm) of 
theorC'z'kol al7d CJxper;;nel7z'ol pb/eV 
volues due to incidence of wil7g. 
.. 
• 
.' 
NACA RM L9L19 
6 
5 
4 
Q: 
\...."1 
Qj 
~ 3 § 
~ 
'ti 2 ~ 
\:: ~ C( 
/ 
o 
v 
V 
V 
/ 
V 
V 
V 
/V 
V 
/' 
/ 
/ 
'/ 
~ 
I 
.6 .8 10 12 14 1.6 
M 
F/qure 5. - Variation of Re(jno/d.s numbe>r with 
Mach number. 
I 
15 
1.8 
16 
.O(J 
.06 
.04 
pb/2V 
.02 
o 
- .02 
- r--
-.... 
.......... 
............. ~ ......... 
.......... 
"" ~ ~
~ -1 \ \ 
\ '\ / 
1\ 1\ I 
\V 
I 
\ 
I 
\/ 
\f 
NACA RM L9L19 
c- ~ 
/' oa=5.o/~ iw = 0.07 0 
/' da=4.73~ iw;: 0.04° 
'\ 
c- ~ 
~y--oa=4.e3~ (~= -0.05° 
"'" 0 i? '-- r- --
/ = 
.6 ,8 10 1.2 
M 
14 I. 6 /(3 
(0) Before correcting to iw;: 0 0 Qnd 
c!Q = unit deflection. 
Figure 6. - Variation of rol/inq effBc tiv8n t'f/55 
with Mach number. 
~ 
,~ 
NACA RM L9L19 
.0/6 
.0/2 
,008 
pfy?V 
0;; 
.004 
o 
- .004 
.6 
~ 
~ 
~ \ 
~ , I r c-\ 
\1\/ \< V 
\ ~ /jC 
I , , t---
\ ~ --::c - t- -r-- _ I ~ 1/ 
VJ 
.8 10 12 14 
M 
( b) After correcting to flU = 0° and 
do == unit derleclion. 
F/qure 6. - C onc/uded, 
17 
~ 
~ 
I.::::- ::::::---:::: 
-
/6 /(J 
18 NACA RM L9L19 
.012 
.008 
pb/2v 
a 
.004 
o 
t-. ~ 
~ 
'\ 
II 
\ 
I 
I 
\/ 
IVI 
1/ 
" 
C 17° ~ 
From reference / 
~<:::: 21
0 
~ 
---
- -
V [... ./ - r---. ~ 
./ 
", 
/ 
~ 
-.004 
.6 .8 10 12 14 16 
M 
(0) Comparison of true-contour aileron 
ellectiv~l7~ss Of) circulor-orc ond 
NACA 0010-64 oirfoils . 
Figure 7. - Comparison of oileron effect/vc?nc?ss witlt 
tl;c;t of reference I 
I 
18 
NACA RM L9Ll9 
0/6 
.012 
.008 
pb/2V 
0' 
.004 
o 
.6 
19 
I--~ 
r- ___ 1\ ~/o I- _ rC '\. 
\ rU 
1\/ ~\ From rererence / ~G r--c:= \ I " ~ \ I '-
r-- r--
1"--_ 
--
f- __ 
.8 10 /,2 14 
(b) Comparison of wedge-t!Jpe aileron 
effectiveness 017 circular-orc and 
IVACA 0010-64 oirrdls. 
16 /.8 
Flqure 7. - Concluded. 
20 NACA RM L9L19 
,08 
.06 
CDT 
.04 
.02 
f-- -
o 
.6 
-~ 
-
-/ 
.8 
--1/-r~ b= - -1---_ - -I v....--- - -
II / ~f-
II ( 
I 
I 
I 
1;1\ 
v; "" ~c ~ ~ 
"" ""'- c ~ 
10 12. /4 16 
M 
Figure 8. - VOrlation of total drag coC'fficient with 
Mach number. 
NACA - Lan gley Field , Yd. 
/.8 
" I 
r 
NACA RM L9L19 
016 
. Ole 
.008 
pb/2V 
6' 
.004 
o 
.6 
19 
----~ 
- ~ ~ ~/o -- rc-
'\. 
\ 
\\/ '\ From reference 1 $0 r<:::: \ I ~~ \ I 
" 
---
r-
......... -r-
---
.8 10 I.e 1.4 
(0) COII/por/son OT wedge-type olleron 
eff6>ctivt?ness On circular-arc and 
IVACA 0010-64 olrfdl5. 
16 1.8 
F?qure 7. - Concluded. 
20 NACA RM L9L19 
,08 
.06 
CDT 
.02 
--
o 
.6 
-
-~ 
= 
;::::/ 
.8 
~ 
1/ _ ~~ t - -~ --~-IV _-
-
1/11 ~r-
( 
I 
1/ / 
;1\ 
1/"" ~c ~ ~ 
"" ~c ~ 
/.0 12 14 16 
M 
Figure 8. - Variation of totol drag coefficient LLlith 
Mach number. 
NACA - La ngley Field, Yd . 
/.8 
1 
• 
-. 
, I 


Flight Investigation of the Effect of Thickening the Aileron Trailing Edge on Control Effectiveness for Sweptback Tapered Wings Having Sharp-and Round-nose Sections

Flight Investigation of the Effect of Thickening the Aileron Trailing Edge on Control Effectiveness for Sweptback Tapered Wings Having Sharp-and Round-nose Sections

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