Sandahl, Carl A

Strass, H Kurt

English

NASA Technical Reports Server

j 
I 
,I 
I 
RM L50D17 
NACA 
RESEARCH MEMORANDUM 
ROLLING EFFECTIVENESS OF A THIN TAPERED WING HAVING 
PARTIAL-SPAN AILERONS AS DETERMINED 
BY ROCKET- POWERED TEST VEHICLES 
By Carl A. Sandahl and H. Kurt Strass 
Langley Aeronautical Laboratory 
Langley Air Force Base, Va. 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
May 23, 1950 
https://ntrs.nasa.gov/search.jsp?R=19930086271 2020-06-17T13:14:11+00:00Z
.. 
• 
NACA RM L50D17 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
ROLLING EFFECTIVENESS OF A THIN TAPERED WING HAVING 
P ARTIAIr SPAN AILERONS AS DETERMINED 
BY ROCKET-POWERED TEST VEHICLES 
By Carl A. Sandahl and H. Kurt Strass 
SUMMARY 
The rolling effectiveness of a thin, tapered, and essentially 
unswept wing having outboard partial-span plain ailerons has been 
determined by means of rocket-powered test vehicles. The rolling 
power decreased abruptly in the Mach number range from 0.85 to 0.95 
and more gradually at higher speeds. At the maximum Mach number 
of the tests (1.95), the rolling effectiveness was only 20 percent 
of that at a Mach number of 0.85. Good agreement was obtained with 
the theory of NACA TN 1890 which takes into account wing elasticity. 
INTRODUCTION 
The rolling effectiveness of a thin, tapered, and essentially 
unswept wing having outboard partial-span plain ailerons has been 
determined over the Mach number range from 0.85 to 1.9, by means 
of rocket-propelled test vehicles using the technique described in 
reference 1. The variation of total drag coefficient with Mach 
number was also obtained. In order to obtain an indication of 
aeroelastic effects, wings of two different stiffnesses were used: 
one was made of solid steel and the other of solid duralumin. The 
results of these tests and a comparison of the measured variation of 
rolling effectiveness with Mach number with that calculated by the 
theoretical method of reference 2, . which takes into account wing 
elastiCity, are presented herein. 
2 
EE. 
2V 
m 
e 
p 
5 
R 
M 
c 
NACA RM L50D17 
SYMBOLS 
wing-tip helix angle, radians 
total drag coefficient based on exposed wing area (1.563 sq ft) 
wing-torsional-stiffness parameter, inch-pounds per radian 
couple applied near wing tip in plane parallel to model center 
line and normal to chord plane, inch-pounds 
local wing-twist angle produced by m measured in plane 
parallel to that of m, radians 
ambient static pressure, pounds per square foot 
standard sea-level static pressure (2116 Ib per sq ft) 
deflection of each aileron, degrees 
Reynolds number based on mean exposed wing chord of 0.642 foot 
Mach number 
wing chord parallel to model center line 
sweepback angle of 50-percent chord line 
TEST VEHICLES AND TESTS 
The general arrangement of the test vehicles is shown in figures 1 
and 2. The details of the wing-aileron arrangement are shown in 
figure 3. Two models were flown: in one, the wing panels were made of 
duralumin and in the other, of steel. The measured spanwise variation 
of wing torsional stiffness obtained with the two materials is shown in 
figure 4. The wings had an aspect ratio of 2.91 obtained by extending 
the leading and trailing edges to the model center l i ne, a taper ratio 
of 0.4, 80 of sweepback of the 50-percent chord line , and modified 
hexagonal airfoil sections of 0.045 thickness ratio. The ailerons had 
unswept hinge lines and chords equal to 25 percent of the local wing 
chord. The aileron deflection was set at the desired value of 50 during 
construction and was nonadjustable. The actual measured aileron deflec-
tions are listed in figure 5. 
NACA RM L50D17 3 
The test vehicles were accelerated by a two-stage rocket-propulsion 
system to a Mach number of about 1.9. During coasting flight, following 
burnout of the rocket motor, time histories of the rolling velocity pro-
duced by the ailerons (obtained with spinsonde ,equipment, reference 3) 
and the flight-path velocity (obtained with Doppler radar) were recorded. 
These data, in conjunction with atmospheric measurements obtained with 
radiosondes, permitted the evaluation of the rolling effectiveness 
pb/2V parameter 5 as a function of Mach number. The variation of the 
drag coefficient of the test vehicles with Mach number was obtained by 
differentiation of the flight-path-velocity time history. The scale of 
the tests is indicated by the curve of Reynolds number against Mach 
number shown in figure 5. A more complete description of the method is 
given in reference 1. 
ACCURACY 
The error in the determination of the quantity pb/2V for anyone 
5 
model is estimated to be within ±o.0004 and ±0.0002 at the lowest and 
highest test Mach numbers, respectively. However, experience has shown 
that the Pbf2V obtained from nominally identical models may vary, 
because of small physical differences in the models, by as much as 
±0.OOO5 and ±0.0003 at the lowest and highest test Mach numbers, 
respectively. 
The errors in the drag coefficient and the Mach number are esti-
mated to be within ±0.OO5 and ±0.002 at the lowest and highest test Mach 
numbers, respectively. 
The values of pb/2V 5 obtained during flight deviated slightly from 
steady-state values because the models experienced a continuous rolling 
acceleration or deceleration. The deviation reached a maximum of about 
10 percent in the Mach number range from 0.93 to 0.97. The deviation 
was negligible over the remainder of the Mach number range investigated. 
The results have not been corrected for this effect. 
RESULTS AND DISCUSSION 
The results of the investigation are given in figure 5 as curves of 
the rolling effectiveness parameter Pb(2V and total drag coefficient ~ 
as functions of Mach number. It should be noted that the quantity pb/2 5 
4 NACA RM L50D17 
is simply the ratio of pb/2v to 5 for the particular aileron deflec-
tions tested. The variation of the pressure ratio plpo with Mach 
number obtained during the test flights has also been included to permit 
the correction of the measured values of rolling effectiveness for the 
effects of wing twisting. 
As shown in figure 5, the rolling effectiveness was reduced abruptly 
in the Mach number range from 0 . 85 to about 0.95. The rolling effective-
ness at the highest supersonic Mach numbers was only about 20 percent of 
that at M = 0.85. The rolling effectiveness of the steel-wing config-
uration was only slightly higher than that of the duralumin configura-
tion indicating that the results for both configurations apuroach those 
which would be obtained with infinitely rigid wings. 
In figure 6, the measured values of rolling effectiveness are com-
pared with those obtained from reference 2 . In making the calculations, 
it was necessary to make the reasonable assumption that the wing was 
unswept at the 50-percent chord line. The theory satisfactorily predicts 
the rolling effectiveness of the configurations tested. 
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-Span 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. Tucker, Warren A., and Nelson, Robert L.: The Effect of Torsional 
Flexibility on the Rolling Characteristics at Supersonic Speeds 
of Tapered Unswept Wings. NACA TN 1890, 1949. 
3. Harris, Orville R.: Determination of the Rate of Roll of Pilotless 
Aircraft Research Models by Means of Polarized Radio Waves. 
NACA TN 2023, 1950. 
~-.-----
• 
• 
NACA RM L50D17 
.3.2.S rocket motor 
s-inch diame fer 
~'---------- 43. 19 -----------.1 
I 
12.24 
~--------------~600------------~~ 
Fuse/age 
Sfation 
0 
2.5"0 
500 
7..50 
10.00 
12.S0 
1.500 
17..5"0 
Or-dinan>s 
Diameter 
0 
1.22 
2 .30 
3.16 
..3.92 
4.5'2 
4-.88 
500 Model vlPwed 
from rear 
~ 
Figure 1.- General arrangement of test vehicles. All dimensions are 
in inches . 
5 
J 
NACA RM L50D17 7 
. 
Q) 
rl 
U 
or-! 
,.q 
Q) 
> 
• 
.p 
Ul 
Q) 
.p 
rl 
cd 
U 
-H $ 
. 
(\J 
Q) 
8 
bO 
-H 
Ji.< 
• 
. I 
_J 
• 
• 
.. 
NACA RM L50D17 
---S'3£HljIH~fm· ~''H.::;lE?~'7'?==:;e;,;;;;;;;::::::=::::=T.L* <=5 
0ectioo A-A 
~4.92----i 
Tip of revolu fiOn"~/.46~ J-I.461 
Jl 
3 .90 
~J 
9. 74 
2 .S 
t ~------ I0.475----------i-l 
\ f471 18.495 R. 
,,777 TIll 7r/lr~11 T2722 
I I II u fin z z 1/ OJJIliuJu I Z I I 
Section B-B 
Figure 3.- Wing plan-form details. All dimensions ar e i n inches . 
9 
10 NACA RM L50D17 
I 
.20 xlO -4-
Wing 
50Jid dura/uml n 
./6 
- - Solid steel 
/ 
'/ 
/ 
/ V 
/ V 
V 
~ 
I 
~ 
" ~
.12 ~ Q. 
§ 
'6 ~ 
.08 
'" / 
/ 'i~~ 
.04 
o 
o 
-c 
". 
/' ~ 
/ 
", 
","""-
~ ~ ~---
~ 
---
1-- ~ l---
--.--- I I 
I 2 3 4 5 6 
Dis tonce from side of ruse/age, tn. 
Figure 4.- Measured spanwise variation of reciprocal of wing torsional 
stiffness parameter. 
~----. 
.. 
• 
I 
~ 
". 
I ., 
- ~- ~-- --- ---.~ 
NACA RM L50D17 11 
/.0 
.6 
x 
6 
7? 
4 
2. 
.04 
o 
.0/2 
.008 
p-D/zV 
c5 
.004 
1-- - I-- ~ :.--
---
f-- - I----I----
-
~ ~ ~ - -::::: 
~ f-~ ~ 
-
~ ~ 
-
~ R is based upon a mean 
,-----~ exposf'd chord of .6 4 2 ft 
:r---
/ ~ 1== -
./ 
~ ~. 
\L 
~ 1 
v ~ 
1. 0 
~ r'::.:, 
~ 
12 
Model Wing <5 
I 
---- 2-
~ 
-.-
14 
M 
Solid dtlra/tlmm 4. 4 ° 
Solid .s feel 4.6 0 
-
- -
~ 
I I 
16 1.8 2.0 
Figure 5.- Test results. 
12 
.008 
pb/zv r-.... 
s 
.004 
o 
/.0 
,008 
pb/2V 
r---S 
o 
1.0 
~ ~ 
..." 
1.2 
~ r-.... 
-
12 
NACA RM L50D17 
Experln1enfol 
--- Theory ( rI9/d) 
----- Theory (/lexlbleJ 
~-
~ 
-= 
18 2.0 
/VI 
(a) Steel wings. 
.. 
Expenmenlal 
--- Theory ( rigid) 
- - ---- Theory (flex/ble) 
~--
-; --I-
-I- '1--
-I-
-
~ 
I I 
1.4 1.6 18 2.0 
M 
(b) Duralumin wings. 
Figure 6.- Comparison of experimental and theoretical results from 
reference 2. 
L I 
NACA - Langley Fi e ld , Va. I 
_J 


Rolling Effectiveness of a Thin Tapered Wing Having Partial-span Ailerons as Determined by Rocket-powered Test Vehicles

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server