NATIONAL ADVISOht,·· ·C

""

FOR AERONAUTICS

T ECHNICAL NOT E 3300

INVE STIGATION OF LIFT , DRAG, AND PITCHING MOMENT

OF A 60 0 DELTA-WllJG - BODY COMBINATION

(AGARD CALIBRATION MODEL B) rn THE

LANGLEY 9 -INCH SUPERSONIC TUNNEL

By August F. Bro mm, Jr.

Langley Ae r onautic al Laboratory
Langley Field, Va.

Washington
September 1954

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NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

TECHNICAL NOTE 3300

INVESTIGATION OF LIFT, DRAG, AND PITCHING MOMENT

OF A 60 0 DELTA-WING--BODY COMBINATION

(AGARD CALJBRATION MODEL B) IN THE

LANGLEY 9-INCH SUPERSONIC TUNNEL

By August F. Bromm, Jr.

SUMMARY

The lift, drag, and pitching-moment characteristics of the AGARD

Calibration Model B as determined in the Langley 9-inch supersonic tunnel
are presented at Mach numbers of 1.62, 1.94, and 2.41 and at a Reynolds
number, based on body length, of approximately 3.0 x 106 • The zero-lift
drag data compared favorably with available data and were in the proper
sequence for the effects of Reynolds number.

INTRODUCTION

During the early period of development of subsonic wind tunnels,

important discrepancies in data from different testing facilities were
found. Many of these difficulties were resolved by improved techniques,
equipment, and data corrections. In order to reduce further the uncer-
tainty of comparison of data from different sources, a program of testing
the same model in the primary test facilities of the world was instituted
(ref. 1). As a result of these tests, the subsonic wind tunnel has become
a reliable source of information; any discrepancies which remain are
fairly well understood. Now, the same problem has arisen with the super-
sonic wind tunnels which have been built in recent years, and interest
has been expressed in a test program for supersonic facilities similar
to that for the subsonic facilities.

It was decided at the Rome meeting of the Advisory Group for

Aeronautical Research and Development (AGARD) of the North Atlantic
Treaty Organization in December 1952 to encourage such a program of
tests in supersonic wind tunnels.

The first configuration selected for this purpose (AGARD Calibration

Model A) was a slender body of revolution. This configuration was

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- - - - - -- -

2 NACA TN 3300

designed by the National Advisory Committee for Aeronautics and tested

in earlier corr elation tests of its own facilities. It is probably better
known as the NACA RM- 10 research missile. Reference 2 is a presentation
of the zero- lift drag data for this configuration measured in several
NACA wind tunnels and in flight. A second AGARD configuration was also
selected at the Rome meeting and was designated AGARD Calibration Model B.
It is a new configuration consisting of a wing-body combination. The
specifications for both AGARD models may be found in reference 3.

The purpose of the present paper is to present the results of tests

of this new configuration (AGARD Calibration Model B) in the Langley
9-inch supersonic tunnel. The measurements included lift, drag, and
pitching moment over an angle-of-attack range of ±6°. The zero-lift base
drag of this model was also measured. Tests were conducted at a Reynolds
number, based on body length, of approximately 3.0 x 106 at Mach numbers
of 1.62, 1 . 94, and 2.41.

SYMBOLS

drag coefficient, D
qS

Cn . minimum drag coefficient (zero lift and zero base drag)

-mln

base-drag coefficient, Sb
Pb -
S

rise in drag coefficient above minimum, CD - Cn

""1Il.in

lift coefficient, ..L..

qS

Pitching moment
pitching-moment coefficient,
qSc

c wing root chord, measured along body center line

c mean aerodynamic chord, two-thirds root chord

D drag
 NACA TN 3300 3
•
d diameter of body

L lift

(L/D)max maximum lift-drag ratio

M Mach number

base-pressure coefficient

dynamic pressure

R Reynolds number, based on body length

r radius of body at any station x

s total wing area

total base area

maximum wing thickness

x distance from nose along body axis

distance from nose to center of pressure, body diameters

angle of attack, deg

APPARATUS

Wind Tunnel

The Langley 9-inch supersonic tunnel is a cont~uous-operation

closed-circuit tunnel in which the pressure, temperature, and humidity
of the enclosed air can be regulated. 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 relatively large area settling chamber ahead
of the supersonic nozzle. The turbulence level of the tunnel is con-
sidered lOW, based on past turbulence-level measurements.

Models

A drawing illustrating the construction details of the AGARD

Calibration Model B and giving the pertinent dimensions is shown in

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4 NACA TN 3300
•
figure 1. A photograph of the unassembled model is shown in figure 2(a).
The model is a wing-body combination with a fineness ratio of 8.5. The
lifting surface is a 60 0 delta wing with a span four times the body diam-
eter and has a symmetrical circular-arc section with a thickness ratio
of 0.04 based on the streamwise chord. The body is a body of revolution
having a cylindrical afterbody and a nose profile determined by the
following equation which is obtained from the more general equation in
reference 4:

(1)

The model was sting supported and had a sting-windshield arrangement

as shown in figures 1 and 2(b). The straight portion of the sting wind-
shield is about 2 body diameters in length; this is one-half a body diam-
eter longer than is specified in reference 3. This modification seems
justified because previous shroud-interference tests (turbulent boundary
layer) have indicated that the critical length is about 2 body diameters
in the Mach number range of these tests. The ratio of sting-windshield
diameter to base diameter for the force tests conforms to the AGARD
specifications, although this ratio is known to border on the critical.
For the base-pressure tests, the ratio of sting diameter to model diam-
eter was 0.375; and the ratio of the length of sting of constant diameter
behind the model base to model diameter was 3.6. Accordingly, the sting
effects of these tests are considered to be negligible. (See, for example,
refs. 5 and 6.)

Four probes mounted as shown in figures 2(b) and 2(c) were used in
the force tests to sense the pressure acting on the annulus of the model.
This pressure and the pressure within the balance-enclosing box were
employed to reduce the drag and the pitching moment to the condition of
base pressure equal to stream pressure. A hollow, cylindrical sting
vented to the area just inside the base of the model vas employed in the
base-pressure tests to sense tbe base pressures.

TESTS

All tests were conducted at a Reynolds number of approximately

3.0 x 10 6 , based on body length, or 0.60 x 10 6 , based on the mean aero-
dynamic chord, and at Mach numbers of 1.62, 1.94, and 2.41. The force
tests were made over an angle-of-attack range of ±6°, and the base-
pressure tests were conducted at an angle of attack of 0 0 • Fixed-
transition tests were made with strips 3/16 inch wide by approximately
2/100 inch thick affixed as shown in figure 1. Measurements of lift,
 NACA TN 3300 5
I •
drag, and pitching moment were made by means of an external six-component
self-balancing mechanical balance. An optical system employing a small
I mirror mounted in the rear of the model was used to measure the angles
of attack.
I
PRECISION OF DATA

The precision of the results has been evaluated by estimating the

uncertainties in the balance measurements involved in a given quantity
and combining these errors by a method based on the theory of least
squares. A summary of these estimates follows:

Lift coefficient, CL" ±0.0004

Drag coefficient, CD' ±0.001
Base drag coefficient, CDb' ±0.002
Pitching-moment coeffiCient, ±0.002
Angle of attack, a ±0.01
Mach number, M • • • • • • • ±0.01

RESULTS AND DISCUSSION

The basic data are presented in the form of lift, drag, and pitching-
moment coefficients, and the coefficients are based on the total wing
area. Pitching-moment coefficients are based on the mean aerodynamic
chord of the total wing and are referred to a point on the body axis two-
thirds of the root chord from the apex of the wing. These data are pre-
sented for the condition of zero base drag in figure 3.

The parameters Cr.,

""'1:l,
Cm,
CD. ,and Cn. are presented in fig-
mln
a -0
ure 4 as a function of Mach number. The slope of the lift curve decreases
with Mach number as would be expected, and the slope of the pitching-
moment curve increases with Mach number. The minimum-drag (zero-base-
drag condition) values and the base-drag values for the model decrease
slightly with Mach number. Figure 5 shows the movement of the center of
pressure to be forward with Mach number.

The application of transition strips to the model has very little

effect on C~ but has the effect of slightly increasing C at
lla
M = 1.62 and slightly decreasing ~ at M = 2.41. The largest effect
of the transition strips is seen to be on the minimum drag. Figure 4
.. shows that the fixed-transition drag values are 75 to 90 percent greater
6 NAeA TN 3300

than the clean-model values. The increase in minimum drag may, for the
most part, be attributed to an increase in skin-friction drag; a small
portion may be due to an increase of pressure drag caused by the transi-
tion strips.

Figure 6 presents the variation of the lift-drag ratios with angle

of attack. The curves of figure 6 were extrapolated to maximum values,
and the extrapolated curves are shown in figure 7; this extrapolation
appears justified because the experimental data seem to be very near a
maximum. The maximum lift-drag ratio decreases slightly as the Mach
number increases . A decrease of about 20 percent in the lift-drag ratios
is experienced when transition strips are applied. This decrease is due
primarily to the increase in drag which accompanies the use of fixed
transition.

Figure 8 presents the variation of drag rise due to lift with Mach
number. The values of drag rise for both the clean-model condition and
the fixed- transition case at each Mach number were obtained by plotting
LCD
against angle of attack. These curves had an approximately zero
2
CL
slope except at very low angles of attack; therefore a single value for
ea.ch curve is presented. The only data available for comparison with the
present results are the zero- lift drag data of reference 7 which include
base drag. (In ref. 7 AGARD Calibration Models A and B are referred to
as AGARD Models 1 and 2, respectively.) For comparison purposes, the
data of reference 7 have been extrapolated and are presented in figure 9.
The drag coefficients for both the fixed-transition and the clean-model
conditions are presented with and without base drag (fig. 9). The base-
drag values are obtained from the separate base-drag tests mentioned
previously. The present results and the extrapolation of the results of
reference 7 compare favorably and are in the proper sequence for the
effects of Reynolds number upon the skin-friction drag.

CONCLUDING REMARKS

The lift, drag, and pitching- moment characteristics of the AGARD

Calibration Model B are presented at Mach numbers of 1.62, 1.94, and 2.41
and at a Reynolds number, based on body length, of approximately 3.0 x 10 6 •
The zero - lift drag data of tne present tests compared favorably with
available data and were in the proper sequence for the effects of
Reynolds number .

Langley Aeronautical Laboratory,

National Advisory Committee for Aeronautics,
Langley Field, Va . , July 27, 1954.
 NACA TN 3300 7
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REFERENCES

1. Anon.: Report on Aerofoil Tests at National Physical Laboratory and

Royal Aircraft Establishment. R. & M. No. 954, British A.R.C.,
May 1925.

2. Evans, Albert J.: The Zero-Lift Drag of a Slender Body of Revolution

(NACA RM-10 Research Model) As Determined From Tests in Several
Wind Tunnels and in Flight at Supersonic Speeds. NACA TN 2944, 1953.

3. Anon.: Specifications for AGARD Wind Tunnel Calibration Models.

AGARD Memo. AG4/M3 (Paris).

4. Roy, Maurice: Tuyeres, Trompes, Fusees et Projectiles - Problemes

Divers de Dynamique des Fluides aux Grandes Vitesses. Pub. No. 203,
Pub. Sci. et Tech. du Ministere de l'Air (Paris), 1947.

5. Perkins, Edward W.: Experimental Investigation of the Effects of

Support Interference on the Drag of Bodies of Revolution at a Mach
Number of 1.5. NACA TN 2292) 1951.

6. Chapman, Dean R.: An Analysis of Base Pressure at Supersonic

Velocities and Comparison With Experiment. NACA Rep. 1051, 1951.
(Supersedes NACA TN 2137.)

7. Piland) Robert 0.: The Zero-Lift Drag of a 60 0 Delta-Wing--Body

Combination (AGARD Model 2) Obtained From Free-Flight Tests Between
Mach Numbers of 0.8 and 1.7. NACA TN 3081, 1954 .

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1- - -- - - - - - - - - . -- - - - - --

Cb

A Location of transition
strips when installed

3.2
-
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---,
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···:·!:t;:.::... II I

2.40 1.5011 2.771 ·1 1.122 .. , 1.6

6.8

Figure 1.- Drawing of AGARD Calibration Model B. Equation of nose contour: ~

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r " ~~ - ~(~)2 + 5\(~)~. All dimensions are in inches unless otherwise

1-3
~
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noted. \.>I
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r NACA TN 3300 9
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rl
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rl
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rl
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(b) Model and sting windshield assembled on sting.

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( c ) End view of sting windshield and pressure t ubes . L-85660 ~
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Figure 2.- Concluded. o

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 NACA TN. 3300 11
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.2 4
/
/
/ ~/
.20
/
/
.16 f--- f---
o CL ;,f/
o Co //
f--- f--- OC /
m /
. 12 logg ed symb ols
I;/' "
r--- I-- denote fix ed
tran siti on ~.
.04 .08
II'" ,....-:l ~ v-
.0 2 .04
Ii/ ~ 6::;::: V cY

;(. ~ ~ ~
'> HV
,/~
;:0-
f.,.-- ~

- .02 - .0 4 ,/ 08
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- .04 -.08 07
if- fy

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lr~ ..-
lit-
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- .,..-' ~ 05
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./

-... --;~
V 03
-' p.--

02

01

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-6 -5 -4 -3 -2 -I o 2 3 4 5 6
Angle of attock, a , deg

(a) M::: 1. 6 2.

Figure 3.- Variation of lift, drag, and pitching moment wit h a ngle of
attack. C~::: O.

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 12 NACA TN 3300
•

24
/j~
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-jjC

. 16
)
o CL jf'
- -
o Co j
. 12 - - OCm ,/D'
- - Flagged symb ols
.0 4 .08 - - denote fixed ..... / /~
~
transition / ./

.04 V ~~ ~~
.02
V y
- ~
f""" r<-
Cm 0 CL 0 ~~
~ y-
J:~ /
- .02 -.04 >=
pl r -/
V V
~~
-.04 - .08
)y j~
V-
-.12
i 07
-//
1 1{/ ,'<
- .16 06
~ /-/ lc/
- .20 ~ t:.. . tv-' 05
(/ ~,
"x --..,. -< ~// /
V ~ --- 1-, [t- - ~- ---{ IJ-- - - I tl---
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r-..... ~ k"
~ '-.... ./
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........... / 03
t---ttJ. .d:y
02

01

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-6 -5 -4 -3 -2 -I o 2 3 4 5 6
Angle of attack, a, deg

(b ) M = 1.94 .

Figur e 3 .- Continued .

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 NACA TN 3300 . 13

.20
V
.16 IV
V
.06 .121--- I---
o CL /1<-
o Co ./ ....
I--- I---
<> em ,/ t<- ~ V V
.04 .oS t"logged . symbols
I-- I-- denote fix ed !(' ....... ~ ~
transition #~ ~ ~
.02 .04
( Ld V
"-
~~
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- .02 - .04 /J Q/
~/

- .04 - .OS
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07

\ . - .12 06
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~ tr/
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[~- --I
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04
,7
V
03
-V-
V

02

0 1

o
-6 ~5 -4 -3 -2 - I o 2 3 4 5 6
Angle of attock, a, deg

(c) M = 2 .41.

Figure 3.- Concluded.

_J
14 NACA TN 3300

o Clean model
o Fixed transition

.08 .00 8

.06 .00 ....

C b ~
- ~ b:1~
C La ':}.. _v-<l
C ma p
.04 t-' ~ r--=.. .00 4
""\ D
.02 .002

0
1. 6 2.0 2.4
o 1.6 2.0 2.4

M M

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CO min
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~- - n '-'
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.02 .02 :=::r --...:...r p
o 1.6 2.0 2.4
0
1.6 2.0 2.4

M M

Figure 4.- Variation of the aerodynamic characteristics with Mach number.

• I
 NACA TN 3300 l5
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l.

a I

7 r-
Moment reference
~

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\ \ '" ~
~
\:.
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h.
Ir--::J"- - I- h \
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--1
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5
~
\
\
4
\
\

3
o Clean model
o Fixed transition
2
\ \

\
\,
o 1.6 2.0 2.4 2.8
M

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Figure 5. - Variation of center of pressure with Mach number.

I •

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16 NACA TN 3300

•
o Clean model
o Fixe d transition
- - Extrapolation
5 - r-
r-
,- r---
~V
4
V I-- t--- t---
~ p-
3
V .ill----'
I /
~

2
If'
/ .,/
/ .,., V
(a)M= 1.62
./'

o l!'

5
I - f-- f-
V
4 V
V ,...--1 tJ-=:::: i - r--
I - t-- r--

~fT
3
V ,/

LID V ./
V'

2 / V - I
J. ) IT'
II /
/ /IJ (b) M=1.94

o~

-- f-- I -
4 ~
//

3
/' f---i r- - f-- r-

L "",b----'
2
V ..tV
V .,/
/ V
LL If (c) M= 2.41

rY
o 2 3 4 5 6 7 8 9 10

Angle of attack, a, de g

Figur e 6 .- Variation of LID wit h angl e of at t a ck .

 -~

• •

~
~
f-3
Z
o Clean model \>l
o Fixed transition \>l
o
o
5 1.0
p
- t::::--.
~ P a
4
0-
- - ro- - r--
- -- I:J I
...... :)
3 '--- I--- -
--+--+--- - - -
6

(LID) max .6 CD V V ~ P
C L2
D- t-- v ~
v-
-. I---
2
VV
2

o 1.6 2.0 2.4

0
1.6 2.0 2.4
M M

Figure 7.- Variation of (L/D)max with Figure 8.- Variat ion of the drag rise
Mach number. due to lift with Mach number.

f-'
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L_____________"_____ __--- -- - - - --- -----

r- - ---- - ---- - .- - -- ---- --- - --- ---- - -- - - --

I-'
OJ
t. )
0 Clean model, R~ 3xlO 6
0 Fixed transition
.0 :3 <> CI~an madel, base drag added
t::,. Fixed transition, base drag added
6
3 - - - Extrapolation of ref. 7 , R:::::.12xI0
.0

.0 r
o ~
u
.0 ::>
r---.
-
C
~

----
---- "----
N '----
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-
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(")

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