Ward, D. H.

English

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wm-TUNNEL TESTS 
COORD. MO. ~ ~ - 2 4 7  
By Donald E. Ward 
Langley Aeronautical. Laborat 
Langley Field, Va. 
Restriction/Classification 
Cancelled
https://ntrs.nasa.gov/search.jsp?R=19630043871 2020-03-11T17:06:12+00:00Z
NACA RM SL57J15a 
NATIONAL ADVISORY COMMITTEX FOR AERONAUTICS 
RESEARCH MEMORANDUM 
for the 
U. S. Air Force 
WIND-TUNNEL TESTS OFA SERIES OF PRACTICE BOMBS 
COORi NO. AF-247 
By Donald H. Ward 
Zero-lift drag data were obtained for several practice bomb con- 
figurations at Mach numbers from 0.60 to 1.10. The seven configura- 
tions tested were different combinations of interchangeable noses and 
tail cones with fins, and some of these configurations had different model 
surface conditions. The tests were made in the Langley 8-foot transonic 
tunnel with the Reynolds number varying from 5.190 x 106 to 6.452 x 106 
during the investigation. 
The results of this investigation indicate that the surface condi- 
tions of the model had no significant effect on the model drag. The 
thick tail fins which were tested contributed a large portion of the 
drag throughout the Mach number range; whereas, a set of thin tail fins 
contributed greatly to the drag only at Mach numbers below 0.95. Two 
of tne nose shapes tested contributed essentially the same amount to 
the drag of the model; however, the third nose, which was hemispherical 
in shape, greatly increased the model drag throughout the Mach number 
range and changed the shape of the drag curve at Mach numbers from 
0.80 to 0.95. 
INTRODUCTION 
An investigation was conducted in the Langley 8-foot transonic 
tunnel to determine the zero-lift drag characteristics of a series of 
Republic Aviation Corporation practice bombs. Three interchangeable 
nose shapes and a tail cone with two sets of fins were tested to deter- 
mine the effect of nose shape and of tail fins-on the drag of the model. 
NACA RM SL57J15a ahi-=--- 2 
a 
'0 
Tests were also made to determine the effect of body surface and tail- 
cone roughness on the model drag. These tests were made at Mach num- 
bers rsnging from 0.60 to 1.10 during which the Reynolds number varied 
from 5.190 x 10~ to 6.452 x 10~~ The tunnel stagnation pressure was 
1 atmosphere. 
SYMBOLS 
M Mach number 
CD total drag coefficient, y 
A 
9 
P 
V 
R 
'b 
maximum frontal area of model 
dynamic pressure, $ pv2 
density 
velocity 
Reynolds number based on body length 
static pressure at base of model 
PO3 free-stream static pressure 
%,b base pressure coefficient, 
'b - 'co 
9 
APPARATUS ANDMETRODS 
Tunnel 
The Langley 8-foot transonic tunnel has a dodecagonal slotted test 
' section which permits continuous testing through the Mach number range 
from 0.60 to 1.10. Details of the tunnel design are presented in 
reference 1. Maximum deviation of the free-stream Mach number is indi- 
cated to be no more than kO.003 from the average in the model test region. 
The models were sting supported as show-r in figure 1 and were attached 
to the model support system sholm in f'rgure 1 of reference 1. 
. 
1 . 1 
b. s 
. 
,b 
I ’ 
. 
\ 
+ 
F , 
,l 9 
I 
I 
I 
* , 
*I 
. 
. 
NACA RM SL57Jl5a 
Models 
The models tested consisted of one basic body with three interchange- 
able tail cones (-WY-with fins) and three interchangeable noses. Fig- 
ure 1 shows the model assembled with nose A and tail C!, and figure 2 is 
a photograph of the three noses snd three tail cones. The test configura- 
tions are given in table I. Coordinates for all of the noses and the 
basic body are given in figure 3(a) and the tail-fin dimensions are given 
in figure 3(b). The nose and basic-body sections were made of aluminum 
and were finished 11as cast" with a maximum surface roughness of 250 root- 
mean-square microi&hes. The tail cones snd fins were made of aluminum 
alloy with a smooth machined finish. All fins tested had a sweepback 
angle of 3500 The thickness ratio for the thin fins was 2;percent and 
for the thick fins was 4 percent. 
The models were mounted on a three-component strain-gage balance 
from which the eag data were recorded, and a static orifice was located 
at the base of the model for measuring base pressures. 
Tests 
In this investigation, each configuration given in table I was 
testkd at Mach numbers of approximately 0.60, 0070, 0.80, Oag09 0.925> 
0.95, 0.975,, 1.00, and T.10 with the model at 0' angle of attack. 
Drag data and base-pressure data w$re recorded at each Mach number. 
The tunnel stagnation 
B 
ressure was 1 atmosphere and the stagnation 
temperature varied $rom 114 F to 155' F. The Reynolds number, based 
on body length, varied from 50190 x 10 6 to.6.452 x 10 6 during the investi- 
gation. Figure 4 shows the variation of Reynolds number with Mach num- 
ber for the models tested. 
RESULTS AND DISCUSSION 
-. 
The drag data are presented in coefficient form in figures 5 to 7- 
The drag coefficient CD is based on the area of a 3-inch-diameter 
circle which is the maximum cross-sectional frontal area of the models 
tested. ThFse coefficients have been corrected to give free-stream 
'conditions at the base of the model, The base pressures used for this 
correction are presented in figure 8. Based on the balance accuracy, 
the drag coefficient is esttiated to be accurate within +_0.005 at a 
Mach number of 0.80. ' 
NACA RI4 SL57J15a 
Effect of Nose Shape on Drag 
0 
0 
0 
10 e 0 
,eo 
Figure 5 presents the drag curves which show the effect of nose IL 
shape on the drag of the model. A comparison of the curves in fig- r: 
ure 5(a) indicates that there is no significant change in the drag 4 
characteristics of the practice bomb if nose B is used to replace the 13 
conventional practice bomb nose (nose A). 
Figure 5(b) presents the drag curves for configurations 5 and 6. 
A comparison of these curves shows the effect of replacing the conven- 
tional practice bomb nose (nose A) with a blunt nose (nose C). The 
blunt nose increases the drag throughout the Mach number range; however, 
this increase is not significant at Mach numbers 0.70 and below. At 
Mach numbers from 0.80 to 0.95 the drag of the model with the blunt 
nose is considerably higher and with a different slope to the drag curve. 
From Mach numbers 0.975 to 1.10 the slopes of the two curves appear to 
be about the same although the blunt nose causes an increase in the drag 
coefficient of approximately 0.07 in this speed range. 
Effect of Model Roughness on Drag 
The effects of body surface roughness on the model drag may be 
seen by comparing the curves in figure 6(a). Configuration 6 was tested 
in the "as castW condition and configuration 7 was the same model which 
had been sprayed with glazing putty, fine sanded, waxed, and polished. 
Since the difference in the drag of these two models is less than the 
accuracy of the balance, the difference is considered to be insignificant. 
The drag curves for configurations 2 and 3 may be seen in fig- 
ure 6(b). The effect on the drag of configuration 2 due to adding No, 60 
Carborundum grains to the tail cone may be seen by comparing these curves. 
Configuration 3 is similar to configuration 2 except for the addition of 
the Carborundum. Considering the accuracy of the results, there is no 
significant change in the model drag due to the Carborundum on the tail 
cone. 
Effect of Tail Fins on Drag 
The effect of tail fins on the model drag is shown by comparing 
the curves in figure 7a The increase in drag due to the thin fins of 
tail B can be seen by comparing the curves in figure 7(a). Configura- 
tion 1 is the basi,c body with nose A and tail B (thin fins) and con- 
figuration 6 is a similar model with no tail fins (tail A). The thin 
fins caused an increase in drag coefficient in the subsonic range up to 
Mach number 0.95 of approximately 0.04. From a Mach number range of 
0.95 to 1.10 the increment of drag increase due to the thin fins 
decreases from 0.04 to approximately 0.01. 
NACA RM SL37J13a 5 
Drag data for configurations 4 and 5 are presented in figure 7(b). 
These configurations were essentially the same except for the thick tail 
fins which were used on configuration 4. A comparison of the curves 
in figure 7(b) shows the increase in drag of the model with the thick 
tail fins (tail C). This comparison shows that the thick fins increase 
the drag coefficient of the mo&el by approximately 0.03 throughout the 
subsonic range and by 0.06 at Mach number 1.10. 
CONCLUDING REMARKS 
An investigation was conducted in the Langley 8-foot transonic 
tunnel to determine the zero-lift drag characteristics of a series of 
Republic Aviation Corporation practice bombs. The results of -I&Is 
investigation are as follows: 
1. The conventional practice bomb nose and the conical nose con- 
tributed essentially the same amount to the drag of the practice bomb, 
whereas a blunt nose greatly increased the drag at Mach numbers above 
0.70 with a particularly large increase at a Mach number of 0.90. 
2. The surface roughnesses tested had no significant effect on the 
drag of the model. 
3. The thin tail fins contributed a large part of the drag at Mach 
numbers below 0.93; however, the additional drag due to the thin fins 
' at supersonic speeds was small. 
4. The thick tail fins contributed a large part of the drag through- 
out the Mach number range. 
Langley Aeronautical Laboratory, 
National Advisory Committee for Aeronautics, 
Langley Field, Va,, 
Donald H. 
dwq 
Ward 
Approved: Aeronautical Research Engineer 
rh Chief of 
1. Ritchie, Virgil S., and Pearson, Albin 0.: Calibration of the Slotted 
Test Section of the Langley 8-Foot Transonic Tunnel and Preliminary 
Experimental Investigation of Doundary-Reflected Disturbances. NACA 
RM L5lJa4, 1932. 
.......,.I, . . .A . ,.. .A ,...r.,._ :r. a..- o;;, .; , .,. “-r.-- :, , 
NACA FM SL57J15a 
TABLE I 
!I!!IST CONFIGURATIONS 
6 
Configuration Nose Tail Remarks 
1 A B Tested as manufacture6 
; 
B B Tested as manufactured 
B B Tail cone heavily coated 
with No. 60 carborun- 
dum grains 
4 C C Tested as manufactured 
2 
C A Tested as manufactured 
A A Tested as manufactured 
7 A A Model surface smooth* 
*Model surface was sprayed with glazing putty, fine sanded, 
waxed, and polished. 


V Basic body 
Basic-body CoordltI~tes 
shti00 x riaaius, r stdi0n x Radius, r station x maim, 
4.665 1.500 14.015 1.319 10 7 2 .829 .220 ;j.g; 
g.;g 1.488 1.0% 3.&895 
"2;: 
1.454 lg.160 
13.201' 1.397 :E 
Nose 000raiw38 
Nooe A Nose B Nose C 
itdon x rinaiut3, I stbion x maiuo, F stdion x hahe, I 
0 o:jE$ 0 
.lO .12 
:2 22 .24 
L.10 ;.07; :t," 
oaz3a O;E o;$g 
:5g 
&a 1.032 
:% 
$5; ::;; 
$3 :% 
:; 
1.00 ?2i 
1:432 1.40 9 1.0-p .208 t:,"o 1:311 . 44
Z:Z 1.470 99 2.34 82 1.3oo 2.00 
4.365 1.500 3: 
50 1.386 419
3.00 1.444 
1.470 3.50 
1.499 
(a) Basic body and noses. 
Figure 3.- Details of models. 
Notor All dimmelons in 
im..hen except c.8 noted. 
NACA RM SL'j7J15a 
0 -7~78r--- -- 
qO5 -7.02 
Details of tail fins for tail 6 
JO 
J- 1 y* 
3-r=? 
Details of tail fins for tail C . 
(b) Tail fins. 
Figure 3.- Concluded. 
_. _ - - _~ _._ - 
0 NACA R&I SL57J15a 11 
, 
5,o 0 .8 9 
Mach m.mber,M 
. 
Figure 4.- Variation of Reynolds number based on body length with Mach 
number. 
NACA RM SL57J15a 12 
.40 
35 
30 
-35 v-v 
n I 
z I . I IA 
I 
I 
$ I 
.I0 
.05 
.8 .9 
Mach number,M 
(a) Configurations 1 and 2. 
Figure 5.- Variation of zero-lift drag coefficient with Mach number for 
test configurations with different nose shapes. 
-. - _.._ . . _~ z-.. 
\ 
NACA FM SL57J15a 13 
-45 
-40 
,35 
“-5 .6 7 .8 .9 
Mach number ,M 
I.0 1.1 I.2 
(b) Configurations 5 and 6. 
Figure 5.- Concluded. 
.  -~ 
.~ _~._. __ . I  
NACA FW SL57J15a 14 aa 1 e ISe 
-40 
0” 
.25 
. 
-E a, .- 
$j 20 
s 
F 
n .I5 
.I0 
.05 
n 
.6 .7 .8 .9 1.0 I.1 1.2 
Mach number,M 
(a) Configurations 6 and'7. 
Figure 6.- Variation of zero-lift drag coefficient with Mach number for 
test configurations with different surface roughnesses. 
.~ ~. -..--.- -- --- 
0 NACA RI4 SL57J15a 15 0 0 
c  
Configuration Nose Tail 
02 I3 
03 B E(tail cone coated with 
. No. 60 carborundum grains) 
3 
2 
0” 
.I 
.O 
0 
:5 
I- I I I I I I I B I I I 0 
5 
‘5 
nl I I I I I I I I I I .I I 
v.5 .6 7 xl .9 
M&h number,M 
I.0 1.1 I.2 
(b) Confi&rations 2 and 3. 
Figure 6. - Concluded. 
I NACA RM SL57J15a 16 
-40 
. 7 
. 
Conf iquiation Nose Tai I 
35 
r-t-i 
0 -1 
•I 6 AA !i 
.3 
.2 
,2 
.I 
15 
10 
5 
Mach number$ “.- _ 
(a) Configurations 1 and 6. 
Figure 7.- Vsxiation of zero-lift drag coefficient with Mach number for 
test configurations with different tail fins. 
r 
NACA RM SL57J15a 
1 F’ 
Configuration Nose, Tail 
0” h 
-$ a25 
;E 
% 
s 
H -20 
n 
.6 .7 9 
Ii% num~er,M 
I.0 I.1 I.2 
(b) Configurations 4 and 5. 
Figure 7.- Concluded. 
i 
ee 
D a 
Bb 
NACA RM SL57Jl'ja 18 
lee 
* 
1.e 
00 
I e 
e 
I 
eea 
..8 -9 I.0 
Mach number, M  
Figure 8.- Variation of base pressure coefficient with Mach number. 
NACA R&I SL57J15a 
lClT.ND--TESTS OFASERIES OFPRACTICEBOMBS 
COORD. NO. AF-247 
By Donald H. Ward 
ABSTRACT 
Zero-lift drag data were obtained in the Langley 8-foot transonic 
wind tunnel for several practice bomb confimations at Mach nu&ers 
from 0.60 to 1.10. Seven configurations were tested with different 
nose shapes and with different tail cones with fins. Body and tail- 
cone surface conditions were varied fos some tests. The Reynolds num- 
ber varied from 5.190 x 106 to 6.453 x 106 during the investigation, 
INDEX HEADINGS 
Bodies - Surface Conditions 1.3.2.4 
Tail-Body Combinations - Missiles 1.7.2.1.2 
Missiles, S$ecific Types 1.7.2.2 
t--. ._ . _ - - .~ ). .>.a 
~~~~~!~~~.~~ ,- .‘1 
~. _r 



Wind tunnel tests of a series of practice bombs

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