Hopko, R. N.

Kinard, W. H.

English

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https://ntrs.nasa.gov/search.jsp?R=19690087813 2020-03-16T23:56:41+00:00Z
NATIONAL ADVISORY COMMImE FOR AERONAUTICS 
.RESEARCH MENORANDUM 
f o r  the 
U. S. A i r  Force 
DRAG AT MODEL TRIM LIFT OF A l/l>-SCAI;E 
CONVAIR B-58 SUPERSONIC BOMBER 
\ 
By Russell N. Hopko and W i l l i a m  H. Kinard 
SUMMARY 
An investigation has been made by the Langley P i lo t less  Aircraft  
Research Division u t i l i z ing  a l / l>-scale rocket -propelled model of the 
Convair B-58 supersonic bomber. The drag at  model t r i m  l i f t  was obtained 
a t  Mach numbers between 0.85 and 2.0 a t  corresponding Reynolds number per 
6 6 foot  of 3.5 x 10 and 13.7 x 10 , respectively. The resu l t s  of the pres- 
ent investigation are  compared with unpublished data obtained from several 
f a c i l i t i e s ,  WADC 10-foot tunnel, Ames 6- by 6-foot supersonic tunnel and 
the Iangley 16-foot transonic tunnel. A comparison of the drag at  tran- 
sonic speeds and a t  approximately the same Reynolds numbers showed excel- 
l en t  agreement, A drag coefficient of 0.028 a t  a Mach number of 2.0 was 
obtained a t  zero- l i f t  conditions. 
INTRODUCTION 
A t  the request of the U, S. A i r  Force, the  Langley P i lo t less  Aircraft  
Research Division has undertaken a f l i gh t  test  program t o  determine the 
drag near zero 1if.t of the Convair B-58 composite airplane. 
portion of the B-58 supersonic bomber consists of two parts;  the basic 
inhabited airframe which is designated the  return camponent, and an 
expendable pod which is an air-to-surface m i s s i l e .  The complete vehicle 
with pod attached on a short pylon i s  designated the  composite airplane. 
The vehicle 
The return component consists of a 60° modified de l ta  wing a t  3O of 
area rule" fuselage, and a swept tapered ve r t i ca l  
incidence incorporating 0.19-local-semispan leading-edge camber, a Mach 
number 2.0 supersonic 
t a i l  mounted atop the fuselage aft of the wing t r a i l i n g  edge. 
nacelles are pylon mounted underneath the Wing. 
( 1  
Four 
Restriction/Classification Cancelled
investigation was conducted at the Langley P i lo t less  Aircraft  Research 
Station a t  Wallops Island, Va. 
SYMBOLS 
A cross-sectional area 
A 1  longitudinal acceleration, f t /sec2 
63 
CC chord-Force coefficient , Chord Orce 
acceleration due t o  gravity, f t /sec2 
ss 
Drag drag coefficient, -
qs CD 
C D I  in te rna l  drag coefficient, 
external drag coefficient, C% - C D ~  - C% 
PO s t a t i c  pressure, lb/sq f t  
'b base pressure, lb/sq f t  
9 dynamic pressure, lb/sq f t  
R Reynolds number 
3 
S 
sb 
t o t a l  wing area including body in-ercept, 6.86 sq f t  
area of nacelle base (four nacelles) 
t time 
V velocity, f t / sec  
W model weight, l b  
7 angle between instantaneous f l i gh t  path and the horizontal, 
deg 
MODEL 
The general arrangement of the model i s  shown i n  figure 1 and a 
photograph of the model is shown i n  figure 2. Other pertinent physical 
characterist ics are presented i n  tables  I, 11, and 111. The wing, con- 
structed mainly of steel, has a diamond plan form with 60° sweep of the  
leading edge and a -IOo sweep of the t r a i l i n g  edge. 
t ion  3.33 the wing has an NACA ooO4.08-63 a i r f o i l  section; a t  the root 
it has an NACA 0003.46-64,069 section. 
dihedral of 2°1394fs'f outboard of s ta t ion 3.767. 
designed fo r  a l i f t  coefficient of 0.22 a t  a Mach number of 1.414, 
elevon deflection was Oo. 
Outboard of sta- 
The wing has 3 O  of incidence and 
The camber has been 
The 
The pod wing and canard are of similar plan form t o  the  wing of the 
return component and have NAGA 0004.5-64 a i r f o i l  sections. 
faces were at Oo deflection f o r  the present investigation. 
These sur- 
The ve r t i ca l  t a i l  is  a swept tapered surface with an NACA 0005-64 air- 
f o i l  section. The leading-edge sweep i s  52' and the taper r a t i o  is 0.324. 
The pod t a i l  is a swept tapered surface with an NACA 0005-64 section. 
The leading edge is  swept 600 and the taper r a t i o  is 0.35. 
The model was constructed by 
Et. Worth, Texas. 
4 
The models were internal ly  instrumented by the Iangley Aeronautical 
Laboratory of the National Advisory Committee f o r  Aeronautics with an 
eight-channel telemeter which transmitted the following information: 
longitudinal acceleration ( two instruments), normal acceleration, trans- 
verse acceleration, t o t a l  pressure (two instruments) , s t a t i c  pressure, 
and base pressure. 
inboard nacelle by using four pressure or i f ices  manifolded together and 
connected t o  a pressure pickup instrument. 
uni t  was used t o  determine the space posit ion of the model i n  f l i g h t .  
The velocity was obtained w i t h  a CW Doppler velocimeter and a rawinsonde 
provided atmospheric conditions and winds a l o f t  veloci t ies  throughout 
the a l t i t ude  range transversed by the model i n  flight. 
The base pressure measurements were made on one 
A modified SCR-384 radar 
Propulsion 
The model a t ta ined a maxirmun Mach number of approximately 2.0 with 
an M-5 Ja to  (Nike booster), 
the model and data were  obtained during coasting f l i gh t .  
the model i n  launching posit ion i s  shown i n  figure 3 .  
After burnout the booster drag separated from 
A photograph of 
DATA FBDUCTION 
Ground Radar 
Drag coefficients were obtained during model flight by evaluating 
the following expression 
V is the v e l  
a t o  t 
d f o r  
T e l  
erometer data w e r e  u i n  the following 
equation 
cc = A1 -w/s 
g q  
a Y b  " a  
* a  * 
? NACA RM SL56G23 P 5 
A similar expression was used t o  evaluate the normal and side-force 
coefficients using the normal and transverse accelerations, respectively. 
"he base drag coefficients were determined from 
ACCURACY 
The accuracies i n  coefficient form for  the Mach number, drag, and 
normal-force data are estimated t o  be 
RESUITS AND DISCUSSION 
The Reynolds numbers per foot are  given i n  figure 4. 
drag coefficients for the  configuration are shown i n  figure 5. These 
drag coefficients were determined by both CW Doppler radar and telem- 
etered accelerations. 
M = 2 t o  M = L 5  f o r  t h i s  investigation; the data range of the telem- 
eter was from M = 2 t o  M = 0.85. Excellent agreement was obtained 
between the  Doppler and telemeter data. Base pressures were measured on 
one inboard nacelle and these data reduced t o  drag coefficient are shown 
i n  figure 6. 
both the  inboard and outboard nacelles, obtained i n  the Langley 16-foot 
transonic t m e l  and the Ames 6- by 6-foot supersonic tunnel. Inasmuch 
as the outboard nacelle base pressures were not measured i n  fl ight*and 
because the  comparison of the f l i g h t  and tunnel base pressure measure- 
ments f o r  the inboard nacelle shows excellent agreement, the  outboard 
nacelle base pressure measurements obtained i n  the Langley 16-foot tran- 
sonic tunnel were employed i n  evaluating the external drag data f o r  t he  
present rocket model. 
The t o t a l  
The data range Of the  Doppler radar wits from 
A l s o  shown i n  figure 6 are base pressure measurements on 
N o  measurements of the internal  drag were made with the fl ight model. 
The in te rna l  drag measurements obtained i n  the Langley 16-foot transonic 
tunnel on both the inboard and outboard nacelles a re  shown i n  figure 7. 
Also shown are in te rna l  drag measurements obtained i n  the  Langley 4- by 
&-foot supersonic pressure tunnel on one outboard nacelle of a similar 
6 
nacelle configuration. 
investigation was set a t  the M = 0.9 
flow r a t i o  of approximately 0.9. 
spikes were  set for  approximately the f l i gh t  condition at M = 2 w i t h  a 
respective mass-flow r a t i o  of 1.0 and the mass-flow ra t io s  of subsonic 
speeds were approximately 0.7. Therefore, calculated values of the  inter-  
na l  drag using one-dimensional-flow theory and a l so  reference 4 w e r e  used 
t o  determine the external drag i n  t h i s  investigation. 
The inlet spike f o r  the  16-foot-transonic-tunnel 
cruise position giving a mass- 
For the present investigation the i n l e t  
- s k n  A comparison of available external drag data is made i n  figure 8,& 
rag coefficient a t  model t r i m  l i f t  (shown i n  f ig .  9) and essent ia l ly  
zero sideslip,  as determined from the  transverse accelerations,,is pre- 
sented and compared with data obtained a t  WADC 10-foot tunnel, Langley 
l6-foot transonic tunnel, and Ames 6- by 6-foot supersonic tunnel; the 
data of the  latter configuration were obtained with the inboard nacelles 
para l le l  t o  the wing chord and the outboard nacelles a t  -5O t o  the  wing 
chord. The configuration tes ted i n  the Langley 16-foot transonic tunnel 
was similar t o  the configuration of t h i s  investigation. A comparison of 
the external drag obtained i n  t h i s  investigation with the resu l t s  obtained 
i n  the 16-foot transonic tunnel a t  approximately the  same Reynolds numbers 
showed excellent agreement. 
t rans i t ion  and a AC, 
t r i p  was estimated. T h i s  increment i n  drag coefficient has not been 
subtracted from the data obtained i n  the Ames 6- by 6-foot supersonic 
tunnel presented herein. 
The 6- by 6-foot tests were made with fixed 
of approximately 0.0014 due t o  the boundary-layer 
A nondimensional cross-sectional area diagram of the present con- 
figuration is  shown i n  figure 10. 
CONCLUDING RESIARKS 
The drag at  model trim l i f t  of the Convair B-58 supersonic bomber 
was obtained at  Mach numbers between 0.85 and 2.0 at  corresponding 
Reynolds number per foot of 3.5 x 106 and 13.7 X 106. 
of the model at trim l i f t  has been compared with data obtained i n  the  
Langley l6-foot transonic tunnel. 
speeds and a t  approximatelythe same Reynolds number showed excellent 
agreement. 
The external drag 
A comparison of the drag at  transonic 
A drag coefficient of 0.028 at  a Mach number of 2.0 was 
Aeronautical Research Scientist 
a1 
esearch Division 
REFERENCES 
1. Hall, James R., and Hopko, Russell N. : 
Low Lift of Rocket-Powered Models of the Convair MX-1626 Airplane 
at Mach Numbers From 0.7 to 1.5. 
Drag and Static Stability at 
NACA RM SL53FOga, U. S. Air Force, 
1953 
2. Swihart, JohnM,, and Foss, Wi E., Jr.: Transonic Aerodynamic 
and Trim Characteristics of 
NACA RM L5512D, 1956. 
i-Engine Delta-Wing Airplane Model. 
8 
T 
P 
rl 
v 
d B 
% 
QJ 
W 
% a 
d 9 
i 
E 
& 
0 
NACA RM SL56G23 
10 
PHYSICAL CHARACTERISTICS OF THE MODEL 
Wing: 
Area (included). sq i n  . . . . . . . . . . . . . . . . . .  
Aspect r a t i o  . . . . . . . . . . . . . . . . . . . . . . .  
Mean aerodynamic chord. i n  . . . . . . . . . . . . . . . .  
Sweepback of leading edge. deg 
Trail ing edge sweep. deg . . . . . . . . . . . . . . . . .  
A i r f o i l  section . . . . . . . . . . . . . . . . . . .  NACA 
Span.in . . . . . . . . . . . . . . . . . . . . . . . . .  
. . . . . . . . . . . . .  
Incidence. deg . . . . . . . . . . . . . . . . . . . . .  
. . 987.26 . . 45.49 
e . 2.096 . . 28.94 . .  60 
. e  -10 
+3 
io& . 08-63 
Vertical ta i l :  
A r e a . s q i n  . . . . . . . . . . . . . . . . . . . . . . . . . .  102.4 
Aspec t r a t io  . . . . . . . . . . . . . . . . . . . . . . . . .  2.64 
Sweepback of leading edge, deg . . . . . . . . . . . . . . . .  52 
A i r f o i l  section . . . . . . . . . . . . . . . . . . . .  NACA 0005-64 
Tape r ra t io  . . . . . . . . . . . . . . . . . . . . . . . . .  0.324 
Pod wing: 
Area (included). sq i n  . . . . . . . . . . . . . . . . . . . .  89.60 
Span. i n  . . . . . . . . . . . . . . . . . . . . . . . . . . .  13.70 
Aspect r a t i o  . . . . . . . . . . . . . . . . . . . . . . . . .  2.10 
A i r f o i l  section . . . . . . . . . . . . . . . . . . .  NACA 0004.5-64 
Pod canard: 
Area (included). sq i n  . . . . . . . . . . . . . . . . . . . .  29.44 
~ p a n . i n  . . . . . . . . . . . . . . . . . . . . . . . . . . .  7.86 
Aspectrat io  . . . . . . . . . . . . . . . . . . . . . . . . .  2.10 
A i r f o i l  section . . . . . . . . . . . . . . . . . . .  NACA 0004.5-64 
) . s q i n  . . . . . . . . . . . . . . . . . . . .  19 . . . . . . . . . . . . . . . . . . . . . . . . . .  4.10 . . . . . . . . . . . . . . . . . . . . . . . . .  1.73 . . . . . . . . . . . . . . . . . . . . . . . . .  0.35 
ion . . . . . . . . . . . . . . . . . . .  NACA 0004.5-64 
sweep . . . . . . . . . . . . . . . . . . . . . .  60 
NACA RM SL56G23 
\ 
rl 
a, a 
0 
Ei 
a, 
-P 
4 
@ 
0 
0 
% 
rn 
rl 
2 
rl 
a, 
a0 
% 
Ei 
I 
r j  
CHORD PLANE 
E AIRPLANE AND SYM. 
WING STA. 0 
UPPER ORD. 
WING SPAN 
- -  I 
0, see table I.) 
inued 
NACA RM SL56G23 
OUTBID 
LOOKING Fw' 
G CONTOUR f! 
HORD PLANE S P A N  S T A .  5.333 
W I N G  CONTOUR 
2 
OUTBID CHORD P L A N E  
S E C T I O N  G G 
E 
S P A N  STA. 5.333 
LWING CON' "/ , \ I 
O U T 3  ' 3  
LOOKING FW'D 
S E C T I O N  H H / 
/:-FACE 
MAIM LANDING 
I GEAR F A I R I N G  
RAE, "R" 
T Y P I C A L  ' S E C T I O N  
I O F  DRAG S T R U T  F A I R I N G  
\ 
I 
3 
I 
H 
-! 
I 
fairing and main landing-gear fairing. 
Figure 1.- Continued. 
NACA RM SL56G23 
cn 
N 
cn 
% 
2 
I 
I rn 
NACA RM SL56G23 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
I 
co m 
6-4 cu 
e 
'u. 
r-- 
m 
d 
5 m * 
cn 
ul 
e 
u 
2 
n 
a, 
W 
0 
e 
NACA RM SL56G23 
1 NAC, INTERNAL I 
(f) Nacelle geometry. 
Figure 1.- Continued, 
NAG. STA, I RAD. 
NACA RM SL36G23 
WL- Y I \- I 
STA. 
17.826 
STA. 
23 e333 
STA. 
76.467 
a STRAIGHT LINE BETWEGN OMINATES FEARTNG 
THIS SYMBOL 
is distance above WL 6.667 I 
(h) Fuselage geometry. 
Figure 1.- Continued. 
FUSELAGE 
I I I I I 
l a  is distance above V?L 6,667 
(i) Fuselage-wing f i l l e t  geometry. 
Figure 1.- Continued. 
NACA RM SL56G23 
I 
d 
k 
rl 
cd 
I d -9 k 8 
n 
'rJ 
W 
FA IR I N G  
VERTICAL F 
Figure 1.- Continued. 
RADOME GEOMETRY 
VERT. TAIL 7 
RADOME STRUT 
( 2 )  Radome geometry. 
Figure 1.- Continued. 
0 
0 
0 cu 
NACA RM SL56G23 
4 
0 
\ 
\ 
POD 
I CHORE 
c l  C 
SECTION C-C SF,CTION D-D 
DIST. I ORD. DIST.  I ORD. 
ORD 
- 
T Y P I C A L  SECTION NACA OOO4.5-64 
CANARD AT ZERO INCIDENCE AND DIHEDRAL 
ROOT CHORD 2.200 INCHES BELOW PARTING 
PLANE NACA OOOh.5-6k AIRFOIL SECTION 
c I I I I I 
J I I I I I 
(0) Pod canard geometry. 
Figure 1.- Continued. 
NACA RM ~ ~ 5 6 ~ 2 3  
a w 0 
F4 
0 
0 
m 
m 
* 
NACA RM SL56G23 
POD S B  
A 
A 16. ooo 
\ 
w 60 
\ 
- 
\ 
- 
\ 
- 
\ 
\ 
\ 
\ 
_ -  
1,868 I 
- 
TYPICAL SECTION NACA 0004.5-64 
(9) Pod wing geometry. 
Figure 1,- Concluded. 
NACA RM SL56G23 
k 
0 
NACA RM SL’j6G23 
I; 
Figure 3.- Model and booster i n  launch position. 
NACA RM SL56G23 
NACA RM SL56G23 
ru 
0 
2 
9 9 
co 
0 0 
$ 1  I 3 
NACA RM SL56G23 
a3 
0 
9 
5 
9 
0 
NACA RM SL56G23 
NACA RM SL56G.23 
NACA RM SL'j6G23 
NACA RM SL56G23 
INDEX 
Subject 
Airplanes - Specific Ty-pes 
Airplanes - Performance 
Number 
1.7.1.2 
1.7.1.3 
ABSTRACT 
An investigation has been made by the Langley P i lo t less  Aircraft 
Research Division u t i l i z ing  a 1/15-scale rocket -propelled model of the 
Convair B-58 supersonic bomber. The drag at  model t r i m  l i f t  was obtained 
a t  Mach numbers between 0.85 and 2.0 a t  corresponding Reynolds number per 
foot of 3.5 x 10 6 and 13.7 x lo6, respectively. The r e su l t s  of the  pres- 
ent investigation are  compared with unpublished data obtained from several 
f a c i l i t i e s ,  WADC 10-foot tunnel, Ames 6- by &foot supersonic tunnel and 
the Langley 16-foot transonic tunnel, A comparison of the drag a t  tran- 
sonic speeds and a t  approximately the same Reynolds numbers showed excel- 
lent  agreement. A drag coefficient of 0.028 a t  a Mach number of 2.0 was 
obtained a t  zero- l i f t  conditions e 
Restriction/Classification Cancelled


Drag at model trim lift of a 1/15-scale Convair B-58 supersonic bomber

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