Clark, Frank L

Mcdearmon, Russell W

English

NASA Technical Reports Server

II 
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t 
I 
l 
RM L56B15 
RESEARCH MEMORANDUM 
WIND-TUNNEL INVESTIGATION OF THE DAMPING IN ROLL 
OF THE BELL X-IE RESEARCH AIRPLANE AND ITS 
COMPONENTS AT SUPERSONIC SPEEDS 
By Russell W. McDearmon and Frank L. Clark 
Langley Aeronautical Laboratory 
Langley Field, Va. 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
April 26, 1956 
Declassified October 28, 1960 
https://ntrs.nasa.gov/search.jsp?R=19930089155 2020-06-17T08:37:36+00:00Z
NACA RM L56B15 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH M:EMORANDUM 
WllID-'lUNNEL INVESTIGATION OF THE DAMPING IN ROLL 
OF THE BELL X-1E RESEARCH AIRPLANE AND ITS 
COMPONENTS AT SUPERSONIC SPEEDS 
By Russell W. McDearmon and Frank L. Clark 
Experimental values of the damping in roll at zero angle of attack 
of the Bell X-lE research airplane and various combinations of its com-
ponents have been obtained at Mach numbers of 1.62, 1.94, 2.22, 2.41, 
and 2.62. 
The damping in roll of the complete model was of the order predicted 
by theory. Very slight and gradual decreases in the damping in roll were 
obtained as the Mach number was increased from 1.62 to 2.41, followed by 
a somewhat more abrupt decrease as the Mach number was increased from 
2.41 to 2 .62 . The wing was the predominant contributor to the damping 
in roll throughout the Mach number range of the tests. The dorsal and 
ventral fins had little effect on the damping in roll of the complete 
model. 
INTRODUCTION 
The only wind- tunn71 data available on the aerodynamic character-
istics of the Bell X- LE research airplane have been in the subsonic and 
transonic speed ranges. In order to supply information at supersonic 
speeds, a general program of investigations is being undertaken in the 
Langley 9 - inch supersonic tunnel to determine same of the dynamic and 
static stability characteristics of the X- lEo 
In the present investigation, the damping in roll at zero angle of 
attack of the complete airplane apd various combinations of its components 
was obtained at Mach numbers of 1 . 62, 1 .94, 2 . 22, 2 .41, and 2 .62 . Included 
were determinations of the effects of the dorsal and ventral fins on the 
damping in roll. In this report, the term "dorsal fin" includes the 
~--- - -
" 
2 NACA RM L56B15 
canopy and the conduit which extends rearward from the canopy and is 
faired smoothq into the vertical tail. The term "ventral fin" refers 
to the smaller conduit which extends along a major portion of the under-
side of the body . Comparisons were made with some theoretical predictions . 
b 
Cz 
Cz P 
M 
p 
pb/2V 
q 
S 
v 
SYMBOLS 
wing span, ft 
rolling-:moment coefficient, 
.. 
damping- in- roll derivative, 
rolling moment, ft - lb 
free - stream Mach number 
Ocz 
d pb 
2V 
rolling angular velocity, radians/sec 
wing- tip helix angle, radians 
free - stream dynamic pressure, lb/sq ft 
total wing area, including portion submerged in body, sq ft 
free - stream velocity, ft/sec 
Configurati on identification : 
BW body and wing 
BV bod.0r and v~rtical~tail 
BVE body, vertical tail, and 'horizontal tail 
BWV body, wing, and vertical tail 
BWVH body , wing, vertical tail, and horizontal tail 
NACA RM L56B15 3 
APPARA'IUS 
Wind Tunnel 
All investigations were conducted in the Langley 9-inch supersonic 
tunnel which is a closed-circuit) continuous-operation type in which 
stream pressure) temperature) and humidity can be controlled at all 
times during tunnel operation . Different test Mach numbers are provided 
by interchangeable nozzle blocks which form test sections approximately 
9 inches square. Eleven fine-mesh turbulence-damping screens are 
installed in the settling chamber ahead of the supersonic nozzle. The 
. . l 
turbulence level of the tunnel is considered low) based on past turbulence-
level measurements. 
Modela) Support} and Rolling-Moment Balance 
A drawing of the complete 1/62-scale model of the X-lE is presented 
in figure 1. The sting was an integral part of the model body. In order 
to use a sting of sufficient strength to withstand the forces which would 
be encountered in testing} it was necessary to alter the shape of the 
rear portion of the body} as shown in figure 1. The effect of this alter-
ation on C1p was believed to be negligible. 
In order to attain high rotational speeds with a minimum of model 
vibration in testing} it was necessary that the models be carefully mass-
balanced and be lightweight} c~ensurate with strength requirements. 
Three identical bodies were constructed} one for the BWVH and BWV config-
urations) one for the BW configuration) and one for the BY and BVH con-
figurations. Three bodies were required so that BWVH) BW} and BVH could 
be mass -balanced separately by inserting lead weights at various positions 
along each body. Configurations BWV and BY were not mass-balanced sepa-
rately) since the effect of the horizontal tail on the mass-balancing 
was negligible. One wing) one vertical tail) and one vertical-tail--
horizontal- tail unit were made . Each of these was removable) so that 
it co~d be installed 'on the desired body . The nose portions of the 
bodies were made of aluminum. The remaining portions of the bodies} the 
integral stings) and the wings were made of steel. The tail panels and 
the dorsal and ventral fins were molded from plastic materials. When 
the dorsal ~d ventral fins were removed) the body became a body of revo-
lution) and the vertical ta~l was faired smoothly into the body) as shown 
in figure 1. 
Transition strips of aluminum oxide particles were placed on th~ 
components of all models. These strips were approximately 0.006 inch 
thick and were located as shown in figure 1 . 
4 NACA RM L56B15 
Since the tail panels were molded from plastic materials, they m~ 
have experienced slight bending or twisting when tested. However, the 
resulting aeroelastic effect on the contributions of the tail panels 
to C2 is believed to have been small. p 
Photographs of the damping-in-roll test apparatus are presented in 
figure 2 . (The model shown in fig. 2(a) is the Bell X-lA model used in 
the investigation of ref . 1. In all other respects the tunnel setup 
used in r ef. 1 and that used in the present investigation are identical.) 
The model sting was inser ted into the spindle of the rolling-moment 
balance and s ecured by a Woodruff key and setscrews. The spindle was 
rotated by means of gears and an electric motor outside the tunnel. The 
rolling ve l ocity was measured with a Stroboconn frequency indicator which 
was modi fied to indicat e -revolutions per minute by means of a generator 
attached to the r ear of the spindle. The rolling moments were measured 
by s t r ain gages on the spindl e and were transmitted through slip rings 
and brushes t o a Brown s elf-balanc ing potentiometer outside the tunnel. 
TE3TS 
The damping in roll at zero angle of attack was obtained at Mach 
numbers of 1 . 62, 1. 94, 2.22 , 2.41, and 2 .62 for the confi gurat ions list ed 
in the fo llowing tab le: 
Configur ation Dorsal fin Ventral fin 
BWE On On 
BWVR Off Off 
BWVR On Off 
BWV On On 
BW On On 
BW Off Off 
BV On On 
BVH On On 
The test Reynol ds number range for the BWVR confi gur at ions was 
from 0. 33 X 106 to 0 .62 X 106 , based on the mean aerodynamic chor d of 
the wing . However, all t est s were conducted with t ransition s t rips on 
the components to create a turbulent boundary layer over most of t he 
model and thereby more closely simulate full-scale conditions. The 
effectiveness of similar transition s t rips i n creat i ng a turbul ent bound-
ary l ayer m~ be seen in reference 2. 
NACA RM L56B15 5 
PREDISION 
The precision of the data has been determined by estimating the 
accuracies of the measured quantities and evaluating their effects on 
the coefficient Cz and the parameter pb/2V. The probable error in 
the strain-gage indication produced the following errors in CZ: 
Er-ror i n C z for configuration -
M 
BWVR, BWV, BW BVH, BV 
1.62 ±C. 00030 t o . 00015 
1.94 ±.ooo34 1;.00018 
2.22 t . 00041 ±.OO020 
2 .41 t.00026 t .00022 
2. 62 ±.00028 ±.00025 
Error in t he measurement of the rolling velocity caused a maximum 
error in pb/2V of to.00009. The surveyed variation of each of the 
f ree-stream Mach numbers was approximately to.Ol, which produced a maxi-
mum error in pb/2V of ±()'00010. Thus t he maximum t otal error in pb/2V 
was 1;0.00019 . 
Model alinement was maintained to within ±C.lo of zero pi t ch and 
yaw with respect to the tunnel center line. 
The r olling-moment balance was s t atically c.;tlibrated, before and 
at intervals during the testing , to ascert ai n that there were no changes 
in the strain-gage constant. Throughout t he tests, the moisture content 
in the tunnel was kept sufficiently low to insure t hat t he effects of 
condensation were negligible. 
IDSULTS AND DISCUSSION 
The variations of rolling-moment coef ficient wit h wing- tip helix 
angle for the various configurations are presented in f igure 3. In gen-
eral, the variations were linear. 
The Contributions of the Airplane Components to CLp 
The variations with Mach number of CZp for the complet e model 
and its components are pres ent ed i n f i gure 4. The values of CZp were 
6 NACA RM L56Bl5 
obtained by taking the slopes of the variations of C1 with pb/2V 
presented in figure 3 . 
For the complete model} a very slight and gradual decrease in the 
damping in roll was obtained as the Mach number was increased from 1.62 
to 2 .41} followed by a somewhat more abrupt decrease as the Mach number 
was increased from 2.41 to 2.62. The dorsal and ventral fins had little 
effect on the damping of the complete model. 
The wing was the predominant contributor to C rp throughout the 
Mach number range of the tests} although the contributions of the tail 
panels to C1p and interference effects of the dorsal and ventral fins 
were in some instances significant} especial~ near M ~ 1.94. It is 
interesting to note that at M = 1.62 and 1.94 the contribution of the 
horizontal tail was marked~ affected by the wing; the addition of the 
horizontal tail to BWV increased the damping} but the addition of the 
horizontal tail to BV decreased the damping. At Mach numbers greater 
than 1.94} the addition of the horizontal tail to BV had very little 
effect on C1p . 
Comparisons of the Experimental Values of With 
Some Theoretical Predictions 
The experimental variations of C lp wi th M are compared with 
some theoretical variations for the complete model and its components 
in figure 5. The theoretical predictions were obtained by the method 
employed in reference 1. This method consisted of predictions by linear 
theory of C lp for the wing and tail p8.Fl.els (shown individual~ in 
figs . 5 (b) and 5(c)) plus approximations of the effects of the wing flow 
f ield on the tail panels. The effect of the interference field from the 
b ody on Clp was neglected. 
The general levels of Crp obtained experimental~ for all the 
configurations were of the order predicted by theory . However} in the 
Mach number region from 2.22 to 2.41 the experimental damping in roll 
was greater than that predicted theoretical~ for all configurations 
containing the wing. This contrasts with the results obtained in refer-
ence 1 for the Bell X-lA researc~ airplane. In reference lJ the experi-
mental damping of the BWVH and BWV confi gurations was considerab~ less 
than that predicted by theory in the Mach number range from 2.22 to 2.41. 
NACA RM L56B15 7 
Although theory underestimated the damping at Mach numbers of 2.22 
and 2 .41 for all configurations containing the wing, at a Mach number 
of 1 . 62, theory overestimated the dampi ng . Also, from the trend estab-
lished by the test results, it appears that at Mach numbers greater than 
2 . 62, theory will again overestimate the damping. 
CONcnmnm REMARKS 
Wind- tunnel investigations of the damping in roll at zero angle of 
attack of the Bell X-lE research airplane and various combinations of 
its components were made at Mach numbers of 1.62, 1 . 94, 2.22, 2.41, 
and 2 . 62 . 
The damping in roll of the complete model was of the order predicted 
by theory. Very slight and gradual decreases in the damping in roll were 
obtained as the Mach number was increased from 1.62 to 2 .41, followed by 
a somewhat more abrupt decrease as the Mach number was increased from 
2.41 to 2 .62. The wing was the predominant contributor to the damping 
in roll throughout the Mach number range of the tests. The dorsal and 
ventral fins had little effect on the damping in roll of the complete 
model. 
Langley Aeronautical Laboratory, 
National Advisory Committee for Aeronautics, 
Langley Field, Va . , February 2, 1956. 
:i 
8 NACA RM L56B15 
REFERENCES 
1. McDearmon, Russell W., and Clark, Frank L.: Wind-Tunnel Investigation 
of the Damping in Roll of the Bell X-lA Research Airplane and Its 
Components at Supersonic Speeds. NACA RM L55I19, 1956. 
2. Fallis, William B.: On Distributed Roughness as a Means of Fixing 
Transition at High Supersonic Speeds. Jour. Aero. Sci. (Readers' 
Forum), vol. 22, no. 5, M~ 1955, p. 339. 
..:. 
NACA RM L56B15 
co 
CO 
CO 
I 
-r-
f 
-t» C\J 0 ~ ;:: N ..t 
! 
Wing : 
Area 
Span 
Aspect ratio 
Section 
Root incidence 
Tip incidence 
Horizontal toil : 
Area 
Section 
Vertical toil : 
Area 
Section 
4 .864 sq. in. 
4 .410 in. 
4 
modified" 64A004 
2° 
2° 
. 250~ 
0 .974 sq. in. 
65-008 
0 .958 sq. in. 
65-008 
~----5. 568 - - -----'>,---+l 
Fairing of vertical toil 
after dorsal fin removed 
Sting 
Q 
<t ~Represents model 
~epresents ful l- scale airplane 
I Dor~al~fin :=~~, '::3-,-
A 
Section A-A Ventral fin 
1<-- ----- 6.0001-------..,., 
Note : all dimensions are in inches. 
Figure 1.- Drawing of the complete model. 
9 
1--- -
(a) Setup in tunnel (top nozzle block removed). 
Figure 2.- Photographs of the damping-in-roll test apparatus and a 
typical model. 
L- 89408 
5 
~ 
:x> 
~ 
(;; 
&i 
t; 
NACA RM L56B15 11 
• 
• 
• 
• 
• • • • 
o , 2 a 
INCH[S 
• • 
• • 
• • 
(b) Interior of balance . L- 89409 
Figure 2 .- Concluded . 
- -----~~- ------~---
I 
! 
L 
12 
.018 
.016 
.014 
.012 
.010 
.008 
.006 
.004 
.00 2 
j 
V 
o 
Confi gurati on 
0 BWVH 
0 BWVH 
<> BWVH 
Bt 
A ~ 
Ie 
I rv 
) r 
V 
.008 .016 
Dorsal 
On 
Off 
On 
A 
kW A/ 
bY 
.024 
pb 
2V 
Fin Ventral Fin 
On 
Off 
Off 
c: ~ Ifj V 
g W 
~ 
.032 .040 
(a) M = 1. 62. 
NACA RM L56B15 
-
.048 
Figure 3. - Variations of rolling-moment coefficient with wing- tip helix 
angle of the complete model and its components at zero angle of attack . 
Flagged symbols indicate check points. 
- I 
NACA RM L56B15 13 
.0 18 
1 
Configuration Dorsal Fin Ventral Fin 
6. BWV On On 
I---
V' BW On On .016 
I> BW Off Off 
<J BV On -On 
[7 BVH On On 
.014 
~ 
v: 
I~ V .012 
j ;y 
h ~ .010 
kr~ 
A / .008 
~ f 
i* ~ 
.006 
~ 
I .004 
I f' 
V 
.<}- H J ...c. I--
.002 
L ..<J--r--<r ,l. F. ~ ~-y ~ -{7-7 -<J [17 ":l 
-
~ VV ..". V 
o .008 .016 .024 .032 .040 .048 
pb 
2V 
(a) Concluded. 
Figure 3.- Continued. 
- . 
14 NACA RM L56B15 
.018 
Configuration Dorsa l Fin Ventral Fin 
.016 o BWVH On On I--
o BWVH Off Off 
<> BWVH On Off 
.014 
.012 
(] 
)~ 
~ f .010 
~ V 
g ~ .008 
h ~ 
~ .006 
F~ 
) f;I" .004 
~ ~ 
/ .002 
/ ---
0 .008 .016 .024 .032 .040 .048 
pb 
2V 
(b) M = 1. 94. 
Figure 3.- Continued. 
NACA RM L56B15 15 
.0 18 
Configuration Dorsa l Fin Ventral Fin 
.016 6 BWV On On '----
'V BW On On 
t> BW Off Off r----
<l BV On On 
C7 BVH On On 
.0 14 
.0 12 
.0 10 ~ 
I~ I' 
16 Y" .008 
~ ~ 
£ ~ 
.006 
~ r 
j ~ .004 
frY 
V ~ 31-~ .002 
~ ~ ~ ~~ 1<:' IV J f2D 
-----= -<r::: 
12 
-v - v 
0 .008 .016 .024 .032 .04 0 .0 4 8 
pb 
2V 
(b) Concl uded. 
Figure 3.- Continued . 
- -------- - --_. 
16 NACA RM L56B15 
.018 
Configuration Dorsa l Fin Ventra l Fin 
0 BWVH On On i-------
0 BWVH Off Off .01 6 
0 BWVH On Off 
.0 14 
.01 2 
<.; ~~ 
.0 10 
~ ~ 
~ ~ .008 
~. / 
~ V .006 
J ~ 
JJ ~ 
.004 
~ ~ 
V .002 
V 
0 .008 .016 .024 .032 .040 .048 
pb 
2V 
(c) M = 2. 22. 
Figure 3.- Continued. 
I 
~v 
I ~ NACA RM L56B15 
.0 18 
.016 
.014 
.0 12 
.010 
.008 
.006 
.004 
.002 
I~ 
o 
Configuration Dorsal Fin 
6,. BWV On 
\7 BW On 
[> BW Off 
<J BV On 
[7 BVH On 
~ v 
A ~) 
~p ~ ~ 
~ V 
~ '/ 
I~ ? 
~ ~ 
/. ~ 
V 
.008 
..... In. 
- ,:., ~ 
.016 
.a. 
..,)<1 
.024 
pb 
2V 
~ 
(c) Concluded. 
'J 
.032 
Figure 3.- Continued. 
17 
Ventral Fin 
On f--
On 
Off f---On 
On 
.040 .048 
----- ----- --~--' 
18 NACA RM L56B15 
.0 18 
Configuration Dorsal Fin Ventral Fin 
.016 0 BWVH On On -
D BWVH Off Off 
0 BWVH On Off 
.014 
.012 
.010 
~ 
~ ¢ .008 
It ~ 
~ ~ 
.006 
~ fP 
~ ~ .004 
~ 
V .002 
1/ 
0 .008 .016 .024 .032 .040 .048 
pb 
2V 
(d) M= 2.4l. 
Figure 3.- Continued. 
• 
NACA RM L56B15 
.018 
I 
Configuration Dorsal Fin 
.016 I:::, BWV On 
V' BW On 
I> BW Off 
<l BV On 
[7 BVH On 
.014 
.012 
.010 
d 
~ Y 
.008 
;/ ~ 
;} ~ .006 
.004 qj ~ 
~ ~ 
~l 
V .002 
lL ,,,..... ;1c. . ,... ""'- ~ JV ' <:.J <J".I' 
o .008 .016 .024 
pb 
2V 
.032 
(d) Concluded. 
Figure 3.- Continued. 
19 
Ventral Fin 
On r--
On 
Off 
-On 
On 
.040 .048 
20 
.018 
.01 6 
.0 14 
.012 
.010 
.008 
.006 
.004 
.002 
V 
o 
Confi gurati on Dorsal Fin 
0 BWVH On 
0 BWVH Off 
<> BWVH On 
A ~ 
~ ~ 
~ ~ 
,~ W 
~ ~ 
/ ~ 
.0 0 8 .016 .024 
pb 
2 V 
(e ) M = 2.62. 
.032 
Figure 3.- Continued. 
NACA RM L56B15 
Vent ra l Fin 
On 
-
Off 
Off 
.040 .048 
'._--' 
---- -~ 
.. 
NACA RM L56B15 
.018 
.016 
.014 
.012 
.010 
.008 
.006 
.004 
.002 
L 
o 
-~ -- - -
Configuration Dorsal Fin 
6. 
\l 
I> 
<l 
[7 
~ 
/ V 
---= :M'::iJ 
.008 
BWV 
-BW 
BW 
BV 
BVH 
A 
~ ~ 
~ 
fZ. -D ~ 
<J<l -q-'-£ 
.016 
On 
On 
Off 
On 
On 
J} l? 
A, ~ ~ 
~v 
~ ~ v ..... 
.024 
pb 
2V 
'-J ~ 
.032 
(e) Concluded. 
Figure 3.- Concluded . 
Ventral Fin 
On 
On 
Off 
On 
On 
.040 
-~-- - - - -
-
t---
.048 
21 
\ 
I 
\ 
22 
.9 
Configuration 
0 BWVH 
0 BWVH .8 o BWVH 
t:, BWV 
V BW 
.7 [> BW 
<J BV 
[7 BVH 
.6 
.5 
.4 
la 
~ ~ ~ ~ 
-.:::::",. 
----
I.rI I-
.......-
V 
r----:::::: ..,:... 
-v- ~ 
.3 
.2 
.1 f 
I<} ~ 
1[7- V 
1.8 2.0 
Dorsal Fin Ventral Fin 
On On 
Off Off 
On Off 
On On 
On On 
Off Off 
On On 
On On 
/.\. 
= ~ til.. 
~ .>~ ~ ,;. 
2.2 
M 
L., 
1Y 
2.4 
~ ~ ~ 
-w 
--
Kz' 
~ 
2.6 
NACA RM L56B15 
I--
I--
I--
2.8 
Figure 4.- Variation with~ch number of the damping in roll of the com-
plet~ model and its components at zero angle of attack. 
• I 
-, 
NACA RM L56B15 
.9 
.8 
.7 
.6 
.5 
.4 
" " 
....... 
l8-. 
~ t---
.3 
.2 
.1 
Configuration 
0 BWVH 
0 BWVH 
<> BWVH 
---Theory 
r--_, 
C\. 
--1-::= ~ 
LJ 
-- --....... 
1.8 2.0 
Dorsal Fin Ventral Fin 
On 0" 
Off Off 
On Off 
b 
1<=::1 ~ ~ -- 1--__ 
'" 
2.2 
M 
1-_ 
~ 
2.4 
~ ~ 
2.6 
(a) BWVH. 
2.8 
Figure 5.- Variations with Mach number of the experimental and theoreti-
cal damping in roll at zero angle of attack. 
23 
24 
.6 
.5 
.4 
~ 
~ ' '-....... ,-A ~ 
:s::--.3 
----.-::. 
.2 
.1 
1.8 
.2 
.1 
- --
:; 
1.8 
NACA RM L56B15 
Experiment: 
Confi guroti on Dorsal Fin Ventral Fin 
D. BWV On On 
\} BW On On 
[> BW Off Off 
Theory: 
- - - BWV 
---- BW 
~ 
"- v ~ V- ~ ~ ~ " v 
«1 
V 
;----
2.0 
-= !...<.. 
-
--
2.2 
M 
-~ 
r--
--
2.4 
(b) BWand BWV. 
Experiment: 
~ 
""-~ 
:::::::", ~ 
v 
2.6 
Configurat ion Dorsal Fin Ventral Fin 
<J BV On On 
[7 BVH On On 
Theory : 
- ----BV 
--- BVH 
~- t;~~t-4 I IP" I Y I \" . ! 
2.0 2.2 2.4 2.6 2.8 
M 
(c ) BV and BVIl. 
Figure 5· - Concluded. 
NA A - La ngley Field , Va. 
- + - --
I 
I 
I 
~ 
---.J 


Wind-tunnel Investigation of the Damping in Roll of the Bell X-1E Research Airplane and Its Components at Supersonic Speeds

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