Investigation of Reynolds number effects for a series of cone-cylinder bodies at Mach number of 1.62, 1.93, and 2.41

MISSING IMAGE

Material Information

Title:
Investigation of Reynolds number effects for a series of cone-cylinder bodies at Mach number of 1.62, 1.93, and 2.41
Series Title:
NACA RM
Physical Description:
20 p. : ill. ; 28 cm.
Language:
English
Creator:
Grigsby, Carl E
Ogburn, Edmund L
United States -- National Advisory Committee for Aeronautics
Publisher:
NACA
Place of Publication:
Washington, D.C
Publication Date:

Subjects

Subjects / Keywords:
Skin friction (Aerodynamics)   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Abstract:
Abstract: An investigation of the Reynolds number for transition and the skin-friction drag at zero lift of eight cone-cylinder bodies having varying fineness ratios has been made at Mach numbers of 1.62, 1.93, and 2.41 over a Reynolds number range from 0.3 x 10⁶ to 10 x 10⁶. The accuracy of the skin-friction data was not sufficient to permit any general conclusions to be drawn. The Reynolds number for transition was found to be dependent upon both the tunnel stagnation pressure and Mach number.
Bibliography:
Includes bibliographic references (p. 10-11).
Statement of Responsibility:
by Carl E. Grigsby and Edmund L. Ogburn.
General Note:
"Report date August 18, 1953."
General Note:
"Classification changed to Unclassified Authority J.W. Crowley Change No. 3071 August 17, 1955."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003808078
oclc - 128322298
sobekcm - AA00006165_00001
System ID:
AA00006165:00001

Full Text
i.p "&r ...SE9UR.T Y IN FO.RIAAT^ION





















SCOE-CYLINDER BODIES AT MACH NUMBERS
SF 1.6, 1.9, AD 2.Copy
l Aeoa:' tc. RM L53H21


SLangley Field, V a.
p.,
















UNIVERSITY OF FLORIDARANDUM
...;'DOCUMENTS DEPAR''" '
,' .,,,o -.i ,,, '3 ..... .,













..'....., & CONE-CYLINDER BODIES AT MACH NUMBERS

i O1 MF 16ARON SENCE BD RA

RO. BOX 11







AINESVLLE, FL 32611-7011 USA
ii 1E.CH AM ERONAEUM T DC US
4...;. t ..,u nd .n ..









'j: *' t naaa toa n iiIubct* a g tH u t a l thm Uflttd BaieN within the menag
a038: ::. : :6 u S Aa m I M M Ul lm an or MlWOC of which In W
Ai I "






T "LONao RYNOADVISORY COMMITTEE AEE







i2 WASHtN GTON
October 7, 1953
,',,:,,,,.,~ ; ,'i ge A rn uialL brtr
d ..:'
,,( ~ ~ ~ ~ ~~~' i:" "",',: ': ', .,i ,:;::: ", ",,, age.Fed a


: ,: :'; ,; i :, ;. : .


I










NACA RM L55H21


!IATI,,IIAL ADVISORY COTIMIMTE FOR .A'E: [i...TICS


FESEAF C'H MEMORANDUM


IIJVESTIGATI,2If OF -EYTI':LDS NUMBER EFFECTS. FOR A SEF E;

OF CONE-CYLINDER PBODLES AT MACH [E.rlEFP

OF 1.62, 1.95, AND 2.41

By Carl E. Grigsby and Edmund L. Ogburn


SUMMARY


An investigation of the Reynolds number for transition and the skin-
friction drag at zero lift of eight cone-cylinder bodies having various
fineness ratios has been made at Mach numbers of 1.62, 1.95, and 2.41
over a Reynolds number rla'e from 0.5 x 10 to 10 x 10 The accuracy
of the skin-friction data was not sufficient to permit any general con-
clusions to be drawn. The Reynolds number for transition was found to
be dependent upon both the tunnel stagnation pressure and Mach number.


INTRCDUCTIOj


Considerable interest is shown currently in the aerodynamic charac-
teristics of bodies of revolution at supersonic speeds and special atten-
tion is shown to the Reynolds number for transition and to the effects
of Reynolds number upon the skin-friction dra:-. In references 1 and 2
results are presented of an investigation of the friction drat and
boundary-layer transition on cone-cylinder bodies over a range of Mach
number. In this investigation, variations in Reynolds number were made
by lengthening the cylindrical portion of the bodies. References 5 to 6
have also presented a considerable amount of aerodynamic data on a series
of bodies having near-parabolic and conical noses and cylindrical after-
bodies. In these tests, variations in Reynolds number were accomplished
by changes in tunnel stagnation pressure. These investigations illus-
trate two techniques for obtaining the effect of Reynolds number upon
skin friction.

In reference 3 data are presented which indicate a dependence of the
Reynolds number for transition upon stagnation pressure. It was suggested
that changes in tunnel turbulence level were responsible for this effect.


CONFIDENTIAL


CONFIDENTIAL







NACA RM L55H21


Additional data for bodies of revolution indicating this same phenomenon
have been published in references 2 and 6 with results for several hollow
cylinders presented in reference 7. Although the temperature was held
nearly constant for the wind-tunnel tests, it was not clear that similar
results could not have been obtained by variations in temperature, in
which case the results would be more properly expressed as a function of
Reynolds number per unit length. The need for further research on this
phenomenon is apparent.

The purpose of the present investigation was to determine the effects
of Mach number and sti attion pressure upon the Reynolds number for tran-
sition, and to obtain the zero-lift skin-friction drag from measurements
of total drag, base drag, and forebody pressure drag for a series of
eight cone-cylinder models of varying fineness ratio. Some objections
have been raised about the use of cone-cylinder bodies for skin-friction
investigations because of the severe adverse pressure gradient and the
possibility of local separation at the juncture of the cone and the cyl-
inder. These objections are based on the belief that this local separa-
tion or the adverse pressure gradient or both would make the results of
questionable value in assessing theoretical predictions. Although there
is some justification for objections on this basis, there is also suffi-
cient reason to investigate these bodies in that they are employed in
several current and proposed missiles and have the advantage of simplified
construction.

The tests were conducted at Mach numbers of 1.62, 1.95, and 2.41 over
a Reynolds number range from about 0.3 x 106 to 10 x 10 for the condition
of zero heat transfer.


SYMBOLS


Amax maximum cross-sectional area of body (equal to AB)

Aw wetted area of body (surface area forward of base)


AB base area

Total drag
CDT total-drag coefficient,
qAmax

AB
CDB base-drag coefficient, PB PB
CmTlax


CONFIDENTIAL


Ci FID DiTIAL ,L






NACA RM L55H21


pL \
I d /r \
CDp foreb:L,- pressure-drag coefficient, j P d dx
O dx\rmax


Cf skin-friction coefficient. AB CDT Dp + CDB


L body length

r local body radius

rmax maximum body radius

P2 Ps
P pressure coefficient, P
q


PB base pressure coefficient

PO stagnation pressure

Ps free-stream static pressure

PZ local static pressure


q free-stream dynamic pressure

M free-stream Mach number

R Reynolds number

RL free-stream Reynolds number based on model length

RT transition Reynolds number based on axial length to transition
point

Tn stagnation temperature


CONFIDENTIAL


CONFIDENTIAL







NACA RM L55H21


APPARATUS AND TE'S'T

Wind Tunnel


The Langley 9-inch supersonic tunnel is a continuous-operation,
closed-circuit type of wind tunnel in which the pressure, temperature,
and humidity of the enclosed air can be re _-lated. 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 tun-
nel is considered low, based on the turbulence-level measurements pre-
sented in reference 8. A schlieren optical system is provided for
qualitative-flow observations.


Models

A sketch illustration, the models and sting support and giving the
pertinent dimensions is shown in figure 1, and a photograph of the models
is shown in figure 2. The eight models varied in fineness ratio in incre-
ments of 1.0 from 2.0 to 9.0. All models for the force tests were made
of magnesium and were available from the investi.-tion of reference 9.
The surface roughness of these models was about 14 rms microinches. At
the beginning of each run the model was polished with a metal polish and
carefully wiped with chamois to preserve a uniformity of surface condi-
tions during the tests. The hollow sting which served as a conduit for
the strain-gage wires was sealed at the support end and vented to the
chamber within the model. The pressure in the hollow sting was measured
and was assumed to be the average pressure in the chamber within the
model.

A special model constructed of steel having a surface roughness of
8 rms microinches, and otherwise identical with model 8, was employed
for the detailed schlieren observations of transition and for the pressure-
distribution tests. Pressure orifices were located in the conical nose
on the 00, 900, 1800, and 2700 meridian planes. As in the other models,
the hollow sting served as a conduit for the pressure tubes and was sealed
at the suppJrt end.

The tests were conducted at Mach numbers of 1.62, 1.93, and 2.41 and

over a Reynolds number rarLce from about 0.5 x 10 to 10 x 106. The starna-
tion temperature was 1000 l 50F and data were obtained only for equilib-
rium temperature conditions. Throughout the tests the dewpoint was kept
sufficiently low to insure negligible effects of condensation. A condi-
tion of zero pitch and yaw was maintained as closely as possible.


C'OIFIDETITIAL


(CF I DEPiTIAL







NACA RM L55H21


The first phase of the investigation consisted of detailed schlieren
observations of the boundary layer for the visual determination of transi-
tion Reynolds numbers of model 8 (steel). This model was later used to
measure the pressure distribution over the conical portion of the body.
The effect of the tunnel static-pressure distribution upon the forebody
pressure drs, was found to be negligible.

The second phase of the investigation comprised the measurements of
total drse and base drag over the Reynolds number ra .-e at each test Mach
number. The magTLeslium models were used for these tests. It will be
noted from figure 1 that the strain-rave balance protruded from the rear
of models 1 and 2 and caused an interference in the base-pressure measure-
ments. Additional base-pressure measurements were made without the bal-
ance, and the total dra,- measurements were corrected by the difference in
the two base-pressure measurements. Additional unknown tare forces may
still exist on models 1 and 2; however, these forces are believed to be
small, especially for mrdel 2.


Precision of Data

All models were maintained within +0.150 of zero pitch and yaw with
respect to the tunnel sidewalls and center line, respectively. Previous
measurements of the flow an :ularity in the tunnel test section have shown
negligible deviations. The estimated accuracies of the test variables
and measured coefficients are given in the subsequent table. Values are
given for a tunnel stagnation pressure of 30 in. Hi-. The accuracies of
the coefficients are functions of the stav-nati-n pressure and increase
with decreasing sta,-nati n pressure.

Mach number, M . . ... ..... + 0.01
Reynolds number, R, per in. . ... .0.004 x 106
Total-drag coefficient, CDT . . .. 0.005
Forebody pressure-drag coefficient, CDp . .. .0.002
Base-drag coefficient, CDB . ... 0.0[



RESUzLTS AND DISCUSSI:IJ

Total and Base Dra-;


The total-drag coefficients for all models are shown for varying
Reynolds number at M = 2.41 in figure 5. These data are typical of
the results obtained at the other test Mach numbers. The cori-resp-.nding
base-drag coefficients are shown in figure 4. For model 8, the reflected


CONFIDENTIAL


CONFIDENTIAL







NACA RM L53H21


nose shock entered the wake at a position such that the base drag was
affected (see ref. 10). r:'e variation of both the base- and total-dra,,
coefficients with Reynolds number is typical of the variation shown in
previous results for this type of confi -irati:n. The effects of both
model fineness ratio and Mach number upon base pressure for these con-
figurations have been discussed in reference 9.


Forebody Pressure Drag

Typical pressure distributions over the conical portion of model 8
at M = 2.41 are shown in figure 5. These distributions at each Reynolds
number were integrated to obtain the forebody pressure-drag coefficients
shown in figure 6. It can be seen that the forebody pressure dra., is rela-
tively independent of Reynolds number at the Mach numbers tested. The
experimental results are also compared with the values from the tables of
solutions to the theory of Taylor and Maccoll given in reference 11. The
experimental results are about 6 percent higher than theoretical results
at M = 1.62, in good agreement at M = 1.;3 and about 4 percent lower
at M = 2.41.


Skin-Friction C efficient

The skin-friction data results left much to be desired with regard
to accuracy and scatter of the results; consequently, only typical results
at M = 2.41 will be presented. The skin-friction coefficients were
obtained in the following manner:



Cf = CD CD + CDB (1)



The results at M = 2.41 are shown in figure 7. Also shown in figure 7
are the following theoretical results for laminar flow: the flat-plate
incompressible result of Blasius, the compressible result of Chapman and
Rubesin, and the flat-plate values corrected to the cone-cylinder by the
formula given in reference 2.



Cf = Cf f2 ( + a)(s + a (2)
flat plate s + 2a


CONFIDENTIAL


CONFIDENTIAL







NACA RM L53H21


where s is the slant height of the cone and a is the length of the
cylindrical afterbody. This formula follows from the transformation by
Mangler and does not consider changes in pressure along the body. The
incompressible, turbulent, skin-friction coefficient is also presented
together with the extended Frankl and Voishel theory. These theoretical
predictions for turbulent flow are presented only as a matter of refer-
ence since there are no comparable experimental results.

The experimental results can be seen to exhibit considerable scatter,
particularly in the transition range where the values of skin-friction
coefficient are smallest. No general conclusion can be drawn from the
results about the effects of varying model fineness ratio upon skin
friction.


Reynolds Number for Transition

From theoretical considerations, it is well known that, for airfoils
at subsonic speeds, the Reynolds number for transition is a function of
wing Reynolds number. This dependency upon wing Reynolds number is a
consequence of the favorable pressure gradient existing over the forward
position of the airfoil. Configurations having zero pressure gradient,
such as flat plates, have transition Reynolds numbers which are invariant
with wing Reynolds numbers. Thus, it is surprising when the results in
reference 7 for hollow cylinders at supersonic speeds show transition
Reynolds numbers which increase with increasing stagnation pressure
(increasing stream Reynolds number). Since, for a given Mach number,
Reynolds number is a function of temperature and pressure, it was not
clear that similar results could not have been obtained by variations
in stagnation temperature in which case the transition Reynolds numbers
would have been shown as a function of Reynolds number per unit length.
However, unpublished data of the transition Reynolds number on a 100 cone
from the Langley 9-inch supersonic tunnel have indicated that decreasing
the stagnation temperature (increasing stream Reynolds number) gave
slightly lower transition Reynolds numbers, whereas increasing the stag-
nation pressure (increasing stream Reynolds number) gave higher transi-
tion Reynolds numbers. Thus, it appears that the effect cannot be iso-
lated as a function of Reynolds nurLber per unit length, but is a function
of some parameter which is influenced by changes in stagnation pressure.

Schlieren photographs of model 8 (steel) were obtained for several
stagnation pressures at each Mach number; typical results at M = 1.62
are presented in figure 8. Points of transition were measured at each
stagnation pressure from the photographs and the ccrresp-ndinr transition
Reynolds numbers were determined. These transition Reynolds numbers are
shown in figure 9 t:--ether with a compilation of data from other sources
which include results for several bodies of revolution, a cone and two
hollow cylinders refss. 2, 3, 6, 7, and unpublished results). The


COrJFIDENrTIAL


C-UITFIDEiTLAL







8 C':i~I IEEIITIAL NACA RM L35H21


ballistic-rsar,- results of reference 2 are plotted with ambient pressure
as the abscissa. The wind-tunnel results shown in figure 9 represent
equilibrium temperature conditions. The relative turbulence levels of
the various tunnels are not known, and the possible effects of stagnation
pressure upon these turbulence levels and upon other tunnel conditions
such as Mach number and stream angularities and disturbances, cannot be
determined. In the pressurized ballistic-range tests (ref. 2), any effects
of tunnel turbulence are presumably excluded, although it is possible that
heat-transfer effects and effects of slight oscillations in angle of attack
are present. The results for the bodies of revolution contain effects of
varying pressure gradient over the cylindrical afterbody, and it also
appears that consideration must be given to the length of the adverse pres-
sure gradient as well as to the value of the pressure gradient.

However, in spite of the variety of the test conditions and tech-
niques represented in the summary of data, a definite increase in Reynolds
number for transition with increasing sta.nati.rn pressure is evident. The
present results showed an increase with increasing stagnation pressure
ranging from about 3 x 106 at 30 in. Hg to about 5 x 106 at 120 in. Hg.
It is also interesting to note that the results shown for the cone and for
the hollow cylinders which have essentially zero pressure gradient are in
substantial r '-er--:nt with the results for the bodies of revolution. Up
to the present time, no satisfactory explanati n has been found for this
phenomenon, but it is evident that comparisons of wind-tunnel-transition
results or attempts to apply these results to free flight must take into
consideration this phenomenon.

The variation in Reynolds number for transition at the base with Mach
number as determined from schlieren photographs is presented in figure 10
together with a summary of results for cone-cylinder bodies of revolution
refss. 1, 2, 5, 6, 10, and 12 to 15). The average surface roughness for
these configurations ranges from about 8 to 20 rms microinches. Each
point represents a single value of stagnation pressure; some effect of
stagnation pressure as discussed previously may be seen in the present
results where the low-fineness-ratio bodies have the largest values of
transition Reynolds number. In view of the number of factors which may
influence transition and which may occur as variables in the present com-
pilation, it is not surprising that the results show considerable scatter.
However, it may be seen that, in general, the variation of Reynolds number
for transition with Mach number is to increase with increasing Mach number
and, then, reach a peak in a range of Mach number from about 2.0 to 2.5
and, thereafter, decrease with further increases in Mach number. This
decrease in transition Reynolds number with Mach number is consistent with
theoretical results for the stability of the laminar boundary layer in com-
pressible flow (see, for example, ref. 16). It might be noted that higher
Reynolds numbers for transition have been obtained at the higher Mach num-
bers where boundary-layer c.- lln;, was present. For example, a transition


CONFIDENTIAL






NACA RM L53H21


Reynolds number of about 8.5 x 106 has been obtained on a hollow cylinder
in the Langley 11-inch hypersonic tunnel at a Mach number of 6.9.


CONCLUDING REMARKS


An investigation of the Reynolds number for transition and the skin-
friction drag at zero lift of eight cone-cylinder bodies having varying
fineness ratios has been made at Mach numbers of 1.62, 1.95, and 2.41
over a Reynolds number range from 0.5 x 10 to 10 x 10 The accuracy of
the skin-friction data was not sufficient to permit any general conclu-
sions to be drawn.

The Reynolds number for transition was indicated to be a function
of some parameter which is influenced by changes in stagnation pressure.
For the present results, the transition Reynolds number increased with
increasing stagnation pressure r-,ing from about 3 x 106 at 0 in. Hg
to about 5 x 10 at 120 in. Hg. Up to the present time, no satisfactory
explanation has been found for this phenomenon. Analysis of the present
tests results together with the results of other cone-cylinder bodies of
revolution having zero-heat transfer showed that the Reynolds number for
transition at the base increased with increasing Mach number and reached
a peak at a Mach number from about 2.0 to 2.5, and, thereafter, decreased
with further increases in Mach number.


Langley Aeronautical Laboratory,
National Advisory Committee for Aeronautics,
Langley Field, Va., August 18, 1955.


CONFIDENTIAL


COirF IDElITIAL







NACA RM L55H21


REFERENCES


1. Potter, J. L.: Friction Drag and Transition Reynolds Number on Bodies
of Revolution at Supersonic Speeds. NAVORD Rep. 2150, U. S. Naval
Ord. Lab., White Oak, Md., Aug. 20, 1951.

2. Potter, J. L.: New Experimental Investigations of Friction Drag and
Boundary Layer Transition on Bodies of Revolution at Supersonic
Speeds. NAVORD Rep. 2571, U. S. Naval Ord. Lab. (White Oak, Md.),
Apr. 24, 1952.

5. Jack, John R., and Burgess, Warren C.: Aerodynamics of Slender Bodies
at Mach Number of 5.12 and Reynolds Numbers From 2 x 106 to 15 x 106.
I Body of Revolution With Near-Parabolic Forebody and Cylindrical
Afterbody. NACA FM E51H15, 1951.

4. Jack, John R., and Gould, Lawrence I.: Aerodynamics of Slender Bodies
at Mach Number of 5.12 and Reynolds Numbers From 2 x 10 to 15 x 106.
II. Aerodynamic Load Distributions of Series of Five Bodies Having
Conical Nose and Cylindrical Afterbodies. NACA RM E52C10, 1952.

5. Jack, John R.: Aerodynamic Characteristics of a Slender Cone-Cylinder
Body of Revolution at a Mach Number of 3.85. NACA RM E51H17, 1951.

0. Jack, John R.: Aerodynamics of Slender Bodies at Mach Number of 5.12
and Reynolds Numbers From 2 x 106 to 15 x 106. III Boundary Layer
and Force Measurements on a Slender Cone-Cylinder Body of Revolution.
NACA RM E53B05, 1953.

7. Brinich, Paul F., and Diaconis, Nick S.: Boundary-Layer Development
and Skin Friction at Mach Number 3.05. NACA TN 2742, 1952.

8. Love, Euierne S., Coletti, Donald E., and Bromm, August F., Jr.: Inves-
tigation of the Variation With Reynolds Number of the Base, Wave,
and Skin-Friction Drs- of a Parabolic Body of Revolution (NACA RM-10)
at Mach Numbers of 1.62, 1.93, and 2.41 in the Langley 9-Inch Super-
sonic Tunnel. NACA RM L52H21, 1952.

9. Love, E-i-ene S.: The Base Pressure at Supersonic Speeds on Two-
Dimensional Airfoils and Bodies of Revolution (With and Without Fins)
Having Turbulent Boundary Layers. NACA RM L55C02, 1955.

10. 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.


CONFIDENTIAL


CONFIDENTIAL






NACA RM L553H21


11. Staff of the Computing Section, Center of Analysis (Under Direction
of Zdenek Kopal): Tables of Supersonic Flow Around Cones. Tech.
Rep. No. 1, M.I.T., 1947.

12. Chapman, Dean R.: An Analysis of Base Pressure at Supersonic Veloc-
ities and Comparison With Experiment. NACA Rep. 1051, 1951.
(Supersedes NACA TN 2157.)

13. Bogdonoff, Seymour M.: A Preliminary Study of Reynolds Number Effects
on Base Pressure at M = 2.95. Jour. Aero. Sci., vol. 19, no. 5,
Mar. 1952, pp. 201-206.

14. Kurzweg, H. H.: Interrelationship Between Boundary Layer and Base
Pressure. Jour. Aero. Sci., vol. 18, no. 11, Nov. 1951,
PP. 743-748.

15. Eber, G. R.: Recent Investigation of Temperature Recovery and Heat
Transmission on Cones and Cylinders in Axial Flow in the N.0.L.
Aeroballistics Wind Tunnel. Jour. Aero. Sci., vol. 19, no. 1,
Jan. 1952, pp. 1-6 and 14.

16. Anon.: Bi-Monthly Survey of the Project Hermes. No. 47, Gen. Elec.
Co., Nov.-Dec. 1949.


C01iIFIDEIDTIAL


CONFIDENTIAL







NACA RM L55H21


(n
m

Li)


cr


-tc

'00



I 1
ci) f






z


- (1J


, co--)

o a


CONFIDENTIAL


C(.l!iFIDEnlT. IAL







NACA RM L53H21


N-


CONFIDENTIAL


CONFIDENTIAL







14 COTFIDEiTIL NACA RM L55H21


36




i, | .. .. i ""
0 2 3 0 I 2 3 0 2 3 4
Model 1 Model 2 Model 3
36


32 --
32' .'- _.




S241 i- -
2 2 3 4 5 0 I 2 3 4 5 6
Mo 3el 4 Model


,r ^ jrI -, ..
S236---- -- -- ---- -- -- --- t .


-4.

^24--------------------------------
CDT 28 -


.24

Model 6









24 -- -- -- .-.




16 ---- --

Molel 7
248 -. --


2 ---- --. -.
NACA

200 2 3 4 5 6 7 8 9 10 I 12
RL x 0


Figure 5.- Variation of total-drag coefficient with Reynolds number
at M = 2.41.


C 2'IFIT IE TTI AL








NACA RM L55H21


l2


Molel 1


!'obel 2


M, el 4
.16





08

.04



0 I 2 3 4 5
)Moel 6


30 I 2 3 4














I : 3 4 5
eolel 3







'odel 5



-- -. -_


^^ J^^


0


2 3
"o lel 7


5 6


RL x 0-6
"odel 8


Figure 4.- Variation of base-drag coefficient with Reynolds number
at M = 2.41.


CONFIDENTIAL


' '


I


CONFIDENTIAL







NACA RM L55H21


R= I 66x 10"


3 4 5

( rx 2
rmax


II ACA

I I I


7 8 9


Viewed from
rear of model





o 0
o 90
O 180
A 270
- Theory,
0 ref II
0


Figure 5.- Typical pressure distributions over conical portion of cone-
cylinder body. M = 2.41.


0 O _


95-


IJ-i--


TIL


-i


0 1.2 14 16 18


Exp.
SM= 1.62
- D M=1.93
SM = 241
ACA, Theory,
0-. ref. II
20 2.2 24x 106


Reynolds number based on length of conical portion of body


Figure 6.- Variation of forebody pressure-drag coefficient with
Reynolds number.


CONFIDENTIAL


o.

--
,

8


Q,

.I-


=E I I


0 I


CDp


Az


I I I i i


CONFIDENTIAL


24-
r


_ l __l"


'1~ [P~ P ~q~


i I I








NACA RM L53H21


C'?I FFIDEI TiAL ,


io




0)
a
r*
















Cd
tJ
o
r-1
c)






I
'-




to




*H
t-4


0-I
T-'-

1-H


CONFIDENTIAL







NACA RM L55H21


It






4
*i


I
'L. #


CONFIDENTIAL


-MR






t



C



I '


A7
4cC


I-I
ri













r-l
0
O

II


43

4








cO
CO
ri
di
S-


I I
..
cr




A


CONFIDENTIAL







NACA RM L55H21


'uo!HsuDi jol jaqwunu spiouAea


CONFIDENTIAL


9-01 x -l


:0I[FIDEIITIAL








CONFIDENTIAL


NACA RM L55H21


I -----'------,e- --- '---o- --- --I--I--I--1
(n c ^ 0 C C ,
-(n 0)0 0 0
aa oa) 0
") o3 C'J- 0 ->V- (> A D Iz
iC __ __ --


-c OJ 0 t 4' -
0 o 1










SC A

c>














_ ()c I___


m



H 0

ca














00)
S(U



.0
c ad
*I-l

a *i


*0 0)





, *O
C (H
--4





0 P

0 )
CH"
0 -


9_01 x -ly 'uoi!suDJl jo jaqwunu splou/ae


CONFIDENTIAL


NACA-Langley 10-7-53 325


O








---O


8.=C S <8

0 0

0 00 0 ":O

S C. | so -





I jo P cd0 .

w,1n L- o C a.

0 0 0 0
~ 1 C,
.l a)sg o







U -b ., 11 S
a, o
0q 0 Z, r r
W00Z 0 Ed : r





0 C L ) 0 w 0 c.1
00 C ,-'0 0 0)

g.- m w4 0 Z


> E Z: V o Z
00 ,g i, ,= ,E =
.5










1 ) o -I 1 l0 0
= c'
r.K w 0 O-O


z C C. 4. b. Ct .

= !0 2 E o zC
lriw Cd H '.2 W









U O- a fccx~bflgg
H- .co o 2
CdJIs' a













00 mr 00


z CM CO m 'd Ca CM

C')OM -t C'13- -
2w
4[ o ,4 -.,


-2 s "

5 i -i H ta ,3 C

Uuo 2 2 oo.<
w m -ooz
-' CM O 1 -


CT Cd zo
2^ ^' u ; v c.^ lI3'
-w

S~ g ^ I9 -g t.| <
:3 i 11 I
g .a h j 03 CQUUOZ
zto z~e 1
0 < 60 : 05
PL : 0 o t- a 0


m m mu ':



I^S S- g^ 3^ ^
oi' Ea a a H -s '.- i c^
.2 'S _2 o



^e^~~ a gxsda5's^
^a^^.l Sb,-m g~
00o m w -O S E" o



|;. :> Z^ g Lo u5 19



i N 0 ll wj
U3 .i -;; 5






^ig~t ^IJ~gs w
H U bp 11 C
0.0 0




u P, u z
,~~*2~o POoE



0w 0P 0 0 6
>,6 0 0
.0 o0b >, c a 0 w0



W 'a w a pd r w '
a) cd




cP:O >1 cd W zE
r- rn -0;9 rz (D >
w Ua m md ri E
U bb r :
1z0 CCR1C 0


m.r r~ Q V- 9 36
u- IUb. 3CCUI.0
~o~g ~ C a ~
X.~ 3 ~.~ ~~0 *U0


\ -
z


0
u









z~ 0 ;

ra wT u^
f z4 .- -

1' F.( z g >4y

0 \2
4 P4

U~2 .2 .2 Z: ',c .
Q m U 0 C OZ




^g= MC -i w^m ^

i us | .2s m


2 2 -
s a .0 w '.. 0


04t -04n~
05rtbL 'aIs l^ w


9) 0

W' c C .4. 0
zg g 0 Q) Pm :03,




2d o
, 4. 0 o, 0 ^ -s -gg _C s..



o 0 0 = 1-0 ; c- Q0 a
z> ') r C', En a
wA Z


c a 4- M o *' 0C'
>,, 0 'O .'j'D
2 0


'0 0 C4-. .

,~2 I~ 5 0 n <
z cdC4 4Ci

.,.' 10. rdu0


0 00 UOOZ
U.2 .2 0 00 ro- U O <
C') c- 04C~ -


c.~ C')O~ ')


S o o c^o ; oz




0 .. 'oo
a ^ s~~~i-g.





uf 0o 0e -0 = lOOO
0 0C ;. W<
4 ~ m ooz


r)
W


L

0 0






n ~
,0oz
% .)
s'oa i=






0m 0
o 0 .|
> E_ ^- S
< ws3
<^ s'
ZI E
M x >1



>j|< 0 S

z z 5&


0 p o
4 2 *-.4.C


4ca
no. .0)-


0 C
o-ii -



ml 0. C",i-o
00
r o a iu



44 W WO W 2w.. ; h
'0. 4;-)-
r-n->
r w 2 3 3 0 a)
0 r F.0. ;

4 c.) A a- C4




So- c, 4"S ">c ';
a .o'? g 2 C; '




C)< cXN.=< C') 3ao 0.


CO c- C 0 '
z C" C COCO C

R r~

ooo oJ g 1oC- ....

u.2; .2 2 4 or
CM)

*- CM c) ;m ,-;


ui~U co


(I' z


=mm"Mo5
0m
4C 0 St2


iiin




0 0
*0 Oc rI 0
E gzo


s


2 = >f 0 Ot


2 M
)CR


0- r
a I=



O Cl nii
>,0 11 Cj3 >,g 0


04
a. C 0 a ~
"- C-o @



oT'0 CI a.S "o a-
a It 04 4
= ,o 36


>~ L..) ~


ac 0 E c 0 1.
< = 0 u c a' 0... a.


z
z Pk
z
0
u


izi
<2

z
0
u











































































































I'




SECURITY INFORMATION

CONFIDENTIAL

UNIVERSITY OF FLORIDA

3 1262 08106 531 9




UNIVERSITY OF FLORIDA
DOCUMENTS DEPARTMENT
120 MARSTON SCIENCE LIBRARY
P.O. BOX 117011
GAINESVILLE, FL 32611-7011 USA





































CONFIDENTIAL


T'"i







































jl
S.ii






i
I


,ii

r



























.,ig