A summary of information on support interference at transonic and supersonic speeds

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Title:
A summary of information on support interference at transonic and supersonic speeds
Series Title:
NACA RM
Physical Description:
26 p. : ill. ; 28 cm.
Language:
English
Creator:
Love, Eugene S
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
NACA
Place of Publication:
Washington, D.C
Publication Date:

Subjects

Subjects / Keywords:
Aerodynamics, Transonic   ( lcsh )
Aerodynamics, Supersonic   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Abstract:
Abstract: A compilation of available information on the problem of support interference at transonic and supersonic speeds is presented.
Bibliography:
Includes bibliographic references (p. 11-12).
Statement of Responsibility:
by Eugene S. Love.
General Note:
"Report date October 27, 1953."
General Note:
"Classification changed to unclassified authority Mr. J.W. Crowley Change #3072 Aug. 17, 1956."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003808466
oclc - 129352844
sobekcm - AA00006157_00001
System ID:
AA00006157:00001

Full Text

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Copy 36
RM L53K12


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RESEARCH MEMORANDUM







AA SUMMARY OF INFORMATION ON SUPPORT INTERFERENCE AT

TRANSONIC AND SUPERSONIC SPEEDS
,";H .. ..

















By Eugene S. Love

Langley Aeronautical Laboratory '
.. Langley Field, Va.

NIVERSIY OF FLORIDA cuS AT C
O CUMENTS DEPARTMENT
120 MARSTON SCIENCE UBRARY UK ss&3: M
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N GAINESVILLE. FL 32611-7011 USA AUTHORITY MR. J. w. CRLUY
iCliAlGE "# ?072 -
CLASSFBED DOCUMEWT
Thik mnarlnl contaium inrnation aftie g the Natlianal Defene at te United Staes wilMa the meantag
of th apmage laws, Tit a18, U.S.C., Bca. 79S and 94, the trmanamilan or relation of which I any
tma or m unatharked person Is prbhlited by law.

... NATIONAL ADVISORY COMMITTEE

i TRFOR AERONAUTICS
WASHINGTON
H January 12, 1954


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NACA RM L55K12 CONFIDENTIAL

NATIONAL ADVISORY COMMITTEE FOR AEROAIIUTICS


RESEARCH MEMORANDUM


A SUMMARY OF INFORMATION ON SUPPORT INTERFERENCE AT

TRANSONIC AND SUPERSONIC SPEEDS

By Eugene S. Love


SUMMARY


A compilation has been made of available information on the problem
of support interference at transonic and supersonic speeds. This com-
pilation indicates that at supersonic speeds there are sufficient exper-
imental data to design properly sting supports and shrouds having negli-
gible interference. At transonic speeds the interference problem becomes
most acute, and more experimental information is needed.


INTRODUCTION


As a result of difficulties encountered in wind-tunnel investiga-
tions of particular aircraft configurations at transonic and supersonic
speeds and the ensuing evaluation of these difficulties, the general
availability of existing information on sting support and shroud inter-
ference was found to be lacking. Much of the published information on
the problem of support interference is obscured under report headings
that refer, and properly so, to the primary investigation and is there-
fore difficult to locate. Furthermore, some of the valuable existing
information has, as yet, been unpublished and is at the disposal of
only a few experimenters or test facilities. The purpose of this paper
is to bring together most of the information that could be found in the
belief that such a summary would be of value in the design of supports
having small interference. In addition, this summary might also serve
as a basis toward further study of support interference.


SYMBOLS


M Mach number

D diameter of base of test model


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d sting diameter

1 length of sting having constant diameter (measured
from base)

9 semiapex angle of conical shroud

3 boattail angle at base of test model

PB base pressure ceofficient

R Reynolds number (based on model length)

x moment arm

CD total drag coefficient

C, pitching-moment coefficient

CL lift coefficient


PB base pressure

Ps free-stream static pressure

L body length

a angle of attack

S reference area

q dynamic pressure

'q scale factor

Z section modulus

f bending stress

m bending moment

F force normal to sting axis

CF force coefficient, F/qS


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DISCUSSION


In the discussion to follow, any previously unpublished data will be
presented without reference. The data which have been published will be
presented with a minimum of detail and the reader may consult the asso-
ciated reference if additional information is desired. During the com-
pilation of this summary, numerous discussions were made with personnel
of the three NACA laboratories, in particular of the Langley laboratory,
and with a few representatives of industry. Any general opinions
expressed are a result of these discussions or related correspondence.


Supersonic Speeds

Inteference at zero angle of attack.- Perhaps the best known of the
earlier investigations of support interference at supersonic speeds is
that of Perkins (ref. 1). Tests were made at M = 1.5 of two bodies of
revolution at a = 00, one with a cylindrical afterbody and one having a
boattail base. The investigation covered both laminar and turbulent
boundary layers for variations in R from 0.6 x 106 to 5 x 106. Results
for the model having a cylindrical afterbody are shown in figure 1. (The
curve for R = 0.5 x 106 in the upper left-hand plot has been added from
minor extrapolations to curves given in ref. 1.) The important part that
Reynolds number plays in support interference when the flow ahead of the
base is laminar is well illustrated and points up the necessity for knowl-
edge of the factors affecting wake transition. When the boundary layer
is turbulent ahead of the base, effects of Reynolds number are reduced
noticeably. Results for the model having appreciable boattailing
(p 140) are not included herein, but in general, the effects of sup-
port length and support diameter were negligible for both laminar and
turbulent boundary layers as long as the support length was equal to or
greater than 1.7 body diameters and the support diameter was equal to or
less than 0.4 body diameter.

Chapman, in reference 2, has presented results at M = 1.5, 2.0,
and 2.9 of the effects of sting length and diameter upon PB for sev-
eral configurations for laminar and turbulent boundary layers. These
results are shown in figure 2. From these data, the critical value
of is seen to lie between 2 and 5. Also the desirability of not
D
exceeding about 0.4 in d is evident.
D

An investigation at M = 1.62, 1.93, and 2.41 of the effect of
support diameter for several finned body configurations is reported in
reference 3. For these tests, the fins supported the models and were


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8 percent thick with 450 sweepback; body fineness ratio was 9.17. The
results are shown in figure 3. In all cases the boundary layer was tur-
bulent ahead of the base and was always greater than 6. Other tests
D
reported in reference 5 showed that, provided the boundary layer was
turbulent, the fin effects upon PB were small; therefore the sting
interference for the bodies without fins may be assumed of the same order
as that indicated in figure 3.

Reference 4 presents results for M = 2.75 to 4.98 which show the
d 2
effects upon PB of varying and for both laminar and turbulent
D D
boundary layers. All tests to determine the effects of d were con-
D
ducted with = 6, and tests for effects of were made with d= 0.575.
D D D
These data are shown in figure 4 and indicate no unusual trends or dif-
ference in critical values from those exhibited at the lower supersonic
speeds. There remains some question as to whether the boundary layer
was fully turbulent at M = 4.98 for the turbulent tests.

The results which have been presented thus far deal with the effects
of .-, body shape, and Reynolds number for supersonic speeds. These
D D
results permit the design of a reasonable sting that will have small
interference if shroud effects are negligible. For structural reasons
it is desirable that the value of L be as small as possible, and, if
D
a shroud is employed, such a condition brings the semiapex angle of the
shroud e into consideration. External balance housings and other
devices for positioning the model often require shrouds of appreciable
apex angle. Further, the loads on a model sometimes dictate that a
tapered sting must be used to gain strength. It is important, there-
fore, to know the effect of varying shroud angle and whether or not
stings of small taper may be used without creating interference if d
D
is subcritical. The results of an investigation of this type made by
August F. Bromm in the Langley 9-inch supersonic tunnel are shown in
figure 5 for M = 1.62, 1.95, and 2.41. All results are for D = 0.55
D
and for a turbulent boundary layer with R = 2.5 x 106. (It has been
found that when the boundary layer is turbulent, changes in Reynolds
number have only small effect upon the magnitude of the interference;
see fig. 1, for example.) The data of figure 5 show that in this Mach
number range a tapered sting = 0 must have a taper angle less than 2.50

to eliminate interference, even though d exactly at the base is
D


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subcritical. Another result of these tests is that the critical value
of I for a given Mach number is essentially independent of 8 for
D
values of 8 at least up to 200; this factor aids considerably in the
design of the sting-shroud combination. A comparison of the data for
the three Mach numbers shows that the critical value of I decreases
D
slightly with increasing Mach number: from about 2.25 at M = 1.62 to
about 2 at M = 2.41. From reference 5, the distance from the base of
the body to the base of the trailing shock in terms of is also seen
D
to decrease with increasing M; further the critical values for
D
shroud location are seen to correspond approximately to positions 0.85
base diameters downstream of the base of the trailing shock. The addi-
tion of 0.85 to the curve of figure 36 in reference 5 may, therefore,
serve as a tentative guide in establishing critical values for
D
shrouds having 8 no greater than 200. Such a procedure indicates that
at low supersonic Mach numbers a large increase in 1 critical is to be
D
expected.

Figure 6 presents results obtained in the Langley 4- by 4-foot
supersonic tunnel at M = 1.59 for the NACA RM-10 missile body, which
is a parabolic body of fineness ratio 12.2. Here again the critical
value of is seen to be relatively independent of 8 for values up
D
to 200 in spite of the fact that all the values of d are supercritical
D
(left-hand plot). In the right-hand plot are data showing effects of d
D
for laminar flow. As mentioned previously, an understanding of wake
transition is necessary before proper interpretation can be made for
laminar flow.

Interference at angle of attack.- Reference 6 presents some results
of the variation of I and 8 on the lift, drag, and pitching moment
D
of a finned model of the NACA EM-10 missile at M = 1.62. A sketch of
the model is shown in figure 7(a). The value of was held constant
at 2.72, and the boundary layer was turbulent. The results showed that
increasing d from 0.715 to 0.992 (both supercritical) gave greater non-
linearity in the lift and pitching-moment curves, decreased the lift-
cruve slope through zero lift by about 7 percent, increased the pitching-
moment curve slope through zero lift by about 10 percent (less negative),
and reduced the fore drag at zero lift by approximately 10 percent.
Because of structural limitations, the tests at d = 0.489 were confined
D


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to a = 00. (An external balance was employed in these tests, and d
corresponds to the outside diameter of the cylindrical portion of the
sting shield. The actual diameter of the sting is, of course, much
smaller.)

Figure 7(b) presents a sketch of a model of the X-2 airplane which
was tested in the Langley 4- by 4-foot supersonic tunnel at M = 1.6
for a = 0 to 100. The quarter chord of the horizontal stabilizer of
this model is swept back 410. For these tests the boundary layer ahead
of the base was turbulent. The model was tested with various bent stings
at angles of 00, 30, and 60. (See fig. 7(b.) The results showed that
bending the sting had negligible effect upon the lift, drag, pitching
moment, and stabilizer hinge moment. However, the fact that the bent
stings had no effect does not obviate the condition that both D and I
D D
were supercritical.

Results are available from tests in the Langley 9-inch supersonic
tunnel to determine the interference at large a of a sting designed to
have small interference at a = 00. These tests were made with a body
of fineness ratio 9.3 having a cylindrical afterbody and parabolic nose
and mounted to the tunnel side wall by means of a shielded trunnion.
This installation permitted the model to be rotated through 5600 in the
center of the test section and in a plane parallel to the tunnel side
walls. (Shield was fixed to tunnel wall and did not rotate.) A length
of sting having a value of = 0.557, and sufficiently long to extend
D
beyond the base of the trailing shock, could be inserted in the hollow
base of the body. The results of these tests are shown in figure 8. It
should be emphasized that the absolute magnitude of PB has little sig-
nificance because of the effects the trunnion shield may have upon PB;
however, for the assessment of sting interference, these effects from
the trunnion shield are of no concern, since the pressure field which
the shield creates in the vicinity of the base at any value of a is
obviously the same with and without sting. At M = 1.62 no data are
shown beyond a = 400, since reflected shocks appeared to intersect the
wake close to the base. For the same reason, the values from a = 200
upward should be viewed with some caution. At M = 1.95 and 2.41, all
the data are reliable and free of reflected shocks. It is clear from
these results that sting supports so designed to have small effect
upon PB at a = 00 may be expected to have equally small effects at
values of a up to 600. The same would probably apply at M = 1.62.


Transonic Speeds

Consideration will now be given to information at transonic speeds.
In figure 9 are presented results from free-flight tests reported in


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reference 7. These tests were made with a finless body of fineness
ratio 11. A turbulent boundary layer existed ahead of the base. The
large amount of sting interference that occurs in the transonic speed
range is clearly indicated. Of the available information at transonic
speeds, only these data (no sting) offer a basis for assessing wind-
tunnel results and the magnitude of interference without fin effects.

Figure 10 presents some results obtained in the Langley 8-foot tran-
sonic tunnel for a body nearly parabolic in shape and having a fineness
ratio of about 10. (See ref. 8.) Also shown are results obtained by the
Pilotless Aircraft Research Division in free flight of a similar model,
but having three fins, in an effort to shed some light on the sting-
interference problem at transonic speeds. It will be noted (fig. 10)
that both I and I are supercritical for supersonic speeds and would
D D
be more so in the transonic range. Nevertheless, the free-flight results
do serve to show the large interference from such a sting installation.
The difference between the free-flight and wind-tunnel results with sting
is apparently due to the presence of the fins on the free-flight model.

Recently, the staff of the Ames 2- by 2-foot transonic tunnel has
been conducting a rather extensive program to study the model support
problem. Figure 11 presents results obtained in this facility at
R = 6.2 x 106 (turbulent boundary layer). The configuration consisted
of a body with wing (see sketch, fig. 11). The body had a fineness ratio
of 9.9 and was slightly boattailed. The wing had a 3-percent-thick
biconvex section, an aspect ratio of 5.09, and a taper ratio of 0.59.
It is obvious that the value of d = 0.961 is highly supercritical,
D
but the results give considerable insight into the extreme difficulties
confronting experimenters in the high subsonic and transonic speed range.
The critical value of for this particular value of d does not appear
D D
to be reached except at M 1.1. Application of these results to other
values of should be made with caution, since for subcritical values
D
of the critical value of I indicated for M > 1.1 would be too
Ti D
small.

In reference 9, an investigation has been made at transonic speeds
of the effect of d upon the lift, drag, pitching moment, and base pres-
sure of a model of the D-558-II airplane. These results are shown in
figure 12. All stings utilized in obtaining these results had taper of
the order of 20 to 4 with = 0. Extrapolation of the results to
d
S= 0 would be rather broad in any event, and in view of the results

presented in figures 9 and 10 such an extrapolation would lead to


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considerable error in base drag. Whether or not extrapolative proce-
dures may be used in the transonic range awaits further experimental
work.


Other Information

Experimental results.- In reference 10 some effects of support
interference at high subsonic speeds are reported. In view of what
is now known concerning the relation of model size to slotted test sec-
tion size, use of the data of reference 10 would appear limited. Ref-
erence 11 presents information on support interference at supersonic
speeds with emphasis upon windshield design from the standpoint of best
tunnel design. In reference 12 an experimental investigation was per-
formed to determine the effect on base and forebody pressures of using
a sting modified with varying length splitter plates and fins, instead
of conventional sting, to support a cone-cylinder body of revolution.
The investigation was conducted at M = 5.12 for R ranging from 2 x 106
to 14 x 106 and for values of a from 00 to 90. Results indicated that
for R = 8 x 106 and 14 x 106 there was negligible effect of the splitter
plate modification on base pressure and at R = 2 x 106 there was a
small effect. Positioning the leading edge of the splitter plate at or
ahead of the base made no appreciable change in the influence of the
modifications on base pressure at R = 14 x 106. With the fin-type mod-
ification there was a small increase in base pressure. The same general
configuration was tested at M = 1.91 and reported in reference 15
where the pressure upstream of the base varied in accordance with exact
potential flow theory at zero angle of attack. The pressures were
slightly higher as was expected due to the presence of body boundary
layer. In the investigation of a strut-supported 16-inch ram jet at
M = 1.5 to 2.0 reported in reference 14, a dummy strut (identical to
the original in every way) was attached to the tunnel wall with approxi-
mately 3/16-inch clearance maintained between the strut and the model.
It was found that the interference drag could not be measured by the
tunnel scales and was therefore assumed to be negligible. The dummy
strut was then detached from the tunnel wall and attached to the tunnel
scale so the drag of the model and two struts could be measured. Sub-
tracting the drag value for the model and supporting strut gave the
drag of a single strut and hence model drag could be calculated.

There are scattered bits of information and results of minor inves-
tigations available that have not been mentioned herein. These have
been omitted since they are either covered by the data which are included
or because certain of the variables were so highly supercritical that
the results were of little value.


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General comments.- In the process of gathering material for this
summary, it was observed that there was a very noticeable tendency toward
"overdesign" of sting size as model size is increased which could not be
attributed to dead-weight requirements. If a given small-scale model with
subcritical sting size may be tested without fear of failure, increasing
the sting size out of proportion to the increase in model scale cannot be
justified. This may be shown simply. Assume that a force F normal to
the sting axis is the primary load upon the sting and that it acts through
a moment arm x. The bending moment m produced is


m = Fx = (CFqS)x (1)


At a larger scale factor T,


mT =Fx = 2 = (CFqS)rx = i3m (2)


since the area of the model increases as the square of the scale factor
and the ratio is held constant. If a solid sting of circular cross
D
section is used, the bending stress is


fm = (2m (m)



At a larger scale factor r


f = t2 (4)
(ad) T d d


d
since the ratio is held constant.
D

Thus, the stress is seen to be unaffected by scale factor. Quite
often the larger scale models operate at reduced q as compared with
the smaller scale models; if so, the larger models gain in relative
sting strength.


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For the prevention of sting failure caused by the buffeting that
accompanies starting in supersonic tunnels, several methods are commonly
employed: for example, lowering the density of the air during starting
(closed circuit tunnels) and the use of various temporary struts to brace
the model during the strating cycle after which the struts are withdrawn.
In transonic tunnels, buffeting may remain after the starting cycle.
This feature, coupled with the indication that in the transonic speed
range subcritical values of and may approach absurd magnitudes,
D D
would appear to raise doubt that it will be possible to make interference-
free measurements in this speed range. This would apply particularly to
base drag and factors affecting it, such as fin design and location. At
this time, most transonic experimenters appear to feel that, with judicious
support design, interference effects upon lift and pitching moment may
be reduced to negligible quantities at the sacrifice of measuring real-
istic base drag (base pressure is corrected to some appropriate level
such as stream static or replaced by a reasonable estimate).


CONCLUDING REMARKS


A summary has been prepared of available information on the support-
interference problem at transonic and supersonic speeds. The compilation
of experimental data indicates that at supersonic speeds the design cri-
teria for sting supports and shrouds are fairly well established. At
transonic speeds the problem becomes most acute, and more information is
needed in this speed range.


Langley Aeronautical Laboratory,
National Advisory Committee for Aeronautics,
Langley Field, Va., October 27, 1955.


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REFERENCES


1. 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.

2. Chapman, Dean R.: An Analysis of Base Pressure at Supersonic Veloci-
ties and Comparison with Experiment. NACA TN 2157, 1950.

3. Love, Eugene S., and O'Donnell, Robert M.: Investigations at Super-
sonic Speeds of the Base Pressure on Bodies of Revolution with and
without Sweptback Stabilizing Fins. NACA RM L52J21a, 1952.

4. Reller, John 0., Jr., and Hamaker, Frank M.: An Experimental Inves-
tigation of the Base Pressure Characteristics of Ionor fting Bodies
of Revolution at Mach Numbers from 2.75 to 4.98. N- ;1 RM A52E20,
1952.

5. Love, Eugene S.: The Base Pressure at Supersonic Speeds on Two-
Dimensional Airfoils and Bodies of Revolution (With and Without
Fins) Having Turbulent Boundary Layers. NACA :V L53C02, 1953.

6. Coletti, Donald E.: Investigation of the Aerodynamic Characteristics
of the NACA RM-10 Missile (with Fins) at a Mach Number of 1.62 in
the Langley 9-inch Supersonic Tunnel. NACA RM L52J23a, 1952.

7. Hart, Roger G.: Effects of Stabilizing Fins and a Rear-Support Sting
on the Base Pressure of a Body of Revolution in Free Flight at Mach
Numbers from 0.7 to 1.5. NACA RM L52E06, 1952.

8. Osborne, Robert S., and Mugler, John P., Jr.: Aerodynamic Character-
istics of a 45 Sweptback Wing-Fuselage Combination and the Fuselage
Alone Obtained in the Langley 8-Foot Transonic Tunnel. NACA
RM L52E14, 1952.

9. Osborne, Robert S.: High-Speed Wind-Tunnel Investigation of the
Longitudinal Stability and Control Characteristics of a 1/16-Scale
Model of the D-558-2 Research Airplane at High Subsonic Mach Num-
bers and at a Mach Number of 1.2. NACA RM L9C04, 1949.

10. Aldrich, J. F. L.: Some Sting Support Interference Effects in a
Subsonic Slotted Test Section. USCAL 10-5-1 (Contract Noa(s) 10585-
Item 5), Aeronautical Lab., Univ. of Southern Calif., May 31, 1951.

11. Puckett, Allen E.: Final Report on the Model Supersonic Wind-Tunnel
Project. Armor and Ordnance Rep. No. A-269, OSRD No. 3569, NDRC.
April, 1944.


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NACA EM L55K12


12. Baughman, L. Eugene, and Jack, John R.: Experimental Investigation
of the Effects of Support Interference on the Pressure Distribution
of a Body of Revolution at a Mach Number of 5.12 and Reynolds Num-
bers from 2 x 106 to 14 x 106. NACA RM E55E28, 1953.

13. Cortright, Edgar M., Jr., and Schroeder, Albert H.: Preliminary
Investigation of Effectiveness of Base Bleed in Reducing Drag of
Blunt-Base Bodies in Supersonic Stream. NACA RM E51A26, 1951.

14. Nussdorfer, T., Wilcox, F., and Perchonok, E.: Investigation at
Zero Angle of Attack of a 16-Inch Ram-Jet Engine in 8- by 6-Foot
Supersonic Wind Tunnel. NACA RM E5OL04, 1951.






































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M=2, turbulent __



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Figure 2.- Effects upon the base-pressure coefficient of the ratio of
sting diameter to base diameter and of sting length to base diameter
for laminar and turbulent boundary layers at M = 1.5, 2.0, and 2.9.


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M
D
2(lam.) 7
2(turb.) 5
2.9 3.3


M --
D
1.5 7
2.0 5
2.9 3.3


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NACA RM L55K12 CONFIDENTIAL 15















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d d
D D


(a) Effect of


d I
D; = 6.
D D


8 = 5.25*, 5
M RxlO-6
2.73 3.2
3.49 4.3
4.03 3.6
4.48 2.8
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Turbulent
3.49


4.03
4.48


4.98
0 2 4

D


(b) Effect of -; = 0.375.
DD

Figure 4.- Effects upon the base-pressure coefficient of the ratio of
sting diameter to base diameter and of sting length to base diameter
for laminar and turbulent boundary layers at M = 2.75, 5.49, 4.05,
4.48, and 4.98.


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NACA RM L53K12


28 -

-26 -8





l L =5 d 0.33
200


L550
-24 -------- -- -






V2 --1 5-4----




..





-08 Z
L0
,iNoshroud,-tj- 9





















S02



O tt
-.1---2 4 5
----- ---~-- --25
Ir i -\ \ I1















-.02 -------- 4------- --- I



06 I I

-08k- 2 3 4 5 6


(a) M = 1.62.


Figure 5.- Effects upon the base-pressure coefficient of the ratio of
sting length to base diameter for various semiapex angles of a conical

shroud. Turbulent boundary layer; = 0.35.

CONFI DTAL

CONFIDIETIAL


CONFIDENTIAL






NACA RM L55K12


(b) M = 1.95.

Figure 5.- Continued.
CONFIDENTIAL


CONFIDENTIAL






NACA EM L55K12


2 3 4 5
D


(c) M = 2.41.

Figure 5.- Concluded.
CONFIDENTIAL


CONFIDENTIAL







CONFIDENTIAL


NACA RM L53K12


0
0

\ IIl


.4-.
0
I=
u.
0
.4
c-
9
4-
c
0

*c
.c-

Co
0
z


I


0 CD 8
O O --








0 0




0 0
OO

OO 0
N-~--
ou


CONFIDENTIAL


o wo





0


oa a

*H H
4-4C Cd

0
a)Xa)
i wd



H 0
CH



0 ed
E1 0



(D a)
Scd-P+



CD II
a-)
d 4-> *r d
0 V 0





0 -






F w
r r-A
_> Q tH hO 4-
I" ~a f
P4 -P-r







NACA RM L55K12




































0 0C
t ^'fO <


0 N
I r-
U i
cI.O


CONFIDENTIAL


to.

,vi


9D
a S O
Ic- cIn


CONFIDENTIAL







NACA RM L53K12


With sting







0 No sting
E With sting
0 No sting-transition
strip at base


(a) M= 1.62


PB
-22f


(b) M 1.93


'O 10 20 30 40 50
a, deg
(c) M 241


Figure 8.- Effects of a sting and of a transition strip at the rear of
the body upon the base pressure at angle of attack.


CONFIDENTIAL


4 R .63x 106



-2 PB=-0.01 at a 80 --
.I^l


70 80 90


CONFIDENTIAL








NACA RM L55K12 CONFIDENTIAL

















-cd 0o





















N -
ell


















4-3



0 0



4---)
I I2
____
















a -4

__"__ ____ 0
qla
14-
Sr-l

VO ll '
--- --- -- --- --- --- --- ) i-
S~a -
X 'tk-i f
y M r-1
----- ---- ----- __ ---- ----- ---- ----- ^
I n
0 ^ ^* CTL










C?


CONFIDENTIAL







NACA RM L53K12


8 Fins


Free flight (PARD) 0.755 0387 4.250
Wind tunnel (8'T.T.) 0.755 0387 3.90


I.U I.2


3-450sweep
none


Filure 10.- Effects upon the base-pressure coefficient of a transonic
sting installation on a finned missile in free flight. (Also
comparison with unfinned wind-tunnel results.)







CONFIDENTIAL


$
5.70
57


1 1- 1 -- --


with sting-free flight






/ \with sting-wind tunnel







0-
____ li4 __ __







no sting-free flight





^


CONFIDENTIAL


-g- D


1.8 2.0


A







NACA RM L53K12




a=CP
aoOO
4,


CONFIDENTIAL




a = 8.5*
4I


I 2

,16- -; -

20-




08 ------ -
u04




D 4
D


IL


20

24 ___




C4

08

12-


.16
01- ------ -- -

0
0----1--t------- -- ---

04--------

08-----


0 I 2 3 4 5 6 7 8


Figure 11.- Effects upon the base-pressure coefficient of the ratio of
sting length to base diameter for a = 00 to 15 and for M = 0.60
to 1.50.


CONFIDENTIAL


a= 5








NACA RM L55K12


M

.95
o .6
1.2


-2---



.04


1.2



.95
04 --
.6


08- --------------------
08-



8
1.2


06 --------------------


95
04- --



02-- -
.6 .8


0 2 4 6 8 I.


I V-------- -- ---- -- --- ---- ----



I 08 -//


0---- --- --
I,o _
M


104
.85




100----2-------------2---
.6






.98--



96



094
1.2


92-



90


R8


Figure 12.- Effects upon the force and
the ratio of sting diameter to base
(D-558-II model; a = -20; turbulent


base-pressure measurements of
diameter for M = 0.6 to 1.2.
boundary layer. Tapered sting.)


CONFIDENTIAL


NACA-Langley 1-12-54 325


CONFIDENTIAL








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1)



















































































































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CONFIDENTIAL

UNIVERSITY OF FLORIDA
II II 11111 liii I II
3 1262 08106 533 5


UNIVERSITY OF FLORIDA
LiOCUMNTS DEPARTMENT
' M:) .ARSTON SCIENCE LIBRARY
S0. BOX 117011
GAINESV;LLE, FL 32611-7011 USA






























CONFIDENTIAL


i1
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*:

;



.ii
II
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