Heat transfer and pressure measurement on a 5-inch hemispherical concave nose at a Mach number of 2.0

MISSING IMAGE

Material Information

Title:
Heat transfer and pressure measurement on a 5-inch hemispherical concave nose at a Mach number of 2.0
Series Title:
NACA RM
Physical Description:
20 p. : ill. ; 28 cm.
Language:
English
Creator:
Markley, J. Thomas
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
NACA
Place of Publication:
Washington, D.C
Publication Date:

Subjects

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

Notes

Abstract:
Abstract: The concave-nose stagnation-point heating is 40 percent of that of a hemisphere nose shape having the same diameter. At angles of attack of ±5° and ± 10° there is no increase in heating of the nose. Total pressures behind the shock were experienced up to 60° on the concave part for all angles of attack. The tests were made under sea-level conditions for a Reynolds number per foot of about 14 x 10⁶.
Bibliography:
Includes bibliographic references (p. 9).
Additional Physical Form:
Also available in electronic format.
Statement of Responsibility:
by J. Thomas Markley.
General Note:
"Report date July 17, 1958."
General Note:
"Copy 478."
General Note:
"Classification changed to unclassified authority: NASA Technical Publications Announcements #14 Effective Date: February 8, 1960 WHL."--stamped on front cover

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003834297
oclc - 150453812
System ID:
AA00009184:00001


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NACA RM L58Cl4a CONFIDENTIAL

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

RESEARCH MEMORANDUM


HEAT TRANSFER AND PRESSURE MEASUREMENT ON A

5-INCH HEMISPHERICAL CONCAVE NOSE

AT A MACH NUMBER OF 2.0*

By J. Thomas Markley


SUI I4ARY


A 5-inch-diameter hemispherical concave nose was tested at a Mach
number of 2.0 in a free jet to determine heat transfer and pressure dis-
tribution. The tests were made under sea-level conditions for a Reynolds
number per foot of about 14 x 10 .

The concave-nose stagnation-point heating is 40 percent of that of
a hemisphere nose shape having the same diameter. At angles of attack
of 50 and 100 there is no increase in the heat-transfer coefficient
of the nose. However, some increase in heat-transfer coefficient is
shown for the afterbody section of the model for windward angles of
attack. Pressures measured up to 600 on the concave part of the model
were equal to total pressure behind the shock at all angles of attack.


INTRODUCTION


The Langley Pilotless Aircraft Research Division is currently
investigating blunt nose shapes for application to the design of super-
sonic missiles. Blunt nose shapes have considerably less heat transfer
than a pointed nose tip. The hemisphere and flat-face nose shapes have
been tested extensively. It has been shown that the stagnation-point
heat transfer to a flat face is one-half, or less, that to the stagna-
tion point of a hemisphere. Several investigators have suggested that
a concave nose shape would probably have even less stagnation-point
heating than the flat face. Reference 1 presented heat-transfer coeffi-
cients for several blunt shapes with modest depressions at the nose; how-
ever, no beneficial effects of these depressions were noted. Other tests
of concave nose shapes like those reported in reference 2, which includes
tests directed toward the study of heating in concave hemispherical
depressions, have indicated stagnation heating rates considerably less


*Title, Unclassified.


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NACA RM L58Cl4a


than for the hemisphere. In order to evaluate the heat-transfer coeffi-
cient of a concave nose shape in more detail, at the lip and at angle of
attack as well as at the stagnation point, a concave nose has been tested
and the results are reported herein.


SY1TBOLS


a angle of attack, deg

ey specific heat of skin, Btu/lb-0F

Ow mass density of skin, lb/cu ft

h local aerodynamic heat-transfer coefficient, Btu/(sec)(sq ft)(CF)


hhh
stag of hemisphere

hstag stagnation point heat-transfer coefficient, Btu/(sec)(sq ft)(0F)

M Mach number

Npr Prandtl number


pt total pressure ahcad of shock, lb/sq ft

p local static pressure, lb,sq ft

p free-strean static pressure, lb/sq ft


pt,2 total pressure behind normal shock, lb/sq ft

S distance along surface from center line, in.


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S1


tw

T
Taw

Tt


angle between the model surface and the
deg


free-stream direction,


time, sec

local temperature behind normal shock

angle from vertical reference plane, deg


APPARATUS AIID TESTING


All tests were conducted at the preflight
Pilotless Aircraft Research Station at Wallops
down jet has true sea-level conditions.


jet of the Langley
Island, Va. This blow-


The model was located with its center line on the center line of
the jet with its face 2 inches downstream of the nozzle. Figure 1 shows
the model before it was swung into the jet stream. The model was moved
downstream so that the picture could be taken. The model was within the
Mach diamond. Shadowgraph pictures were made during all tests and are
shown in figures 2 to 4 for angles of attack of 00, 50, and 100. From
these figures no interaction with the Mach diamond of the jet can be
seen.

The pressures were measured by using Statham gages which are accu-
rate to within 1 percent. The temperatures are measured by thermo-
couples which have the junction box at the base of the stand. The ref-
erence junction temperature is measured by reading a thermometer located
in a box which is free of any wind currents.


CONFIDENTIAL


maximum distance along surface from center line to lip, in.

skin thickness, ft


adiabatic wall temperature, CR

free-stream stagnation temperature, OR

wall temperature, OR

static temperature ahead of shock, OR

recovery factor


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NACA RM L58C14a


MODEL


The dimensions of the model are shown in figure 5. The model was
made of Inconel, of a nominal 0.050-inch thickness, but because of the
spinning process used in the construction, the thickness of the skin
varied. These variations are tabulated in figure 5. No support was
given the skin other than the model shape itself.

Instrumentation consisted of chromel-alumel thermocouples
(No. 50 gage) welded to the interior of the skin. A ray of thermo-
couples and pressure orifices were located 1800 from each other so that
when the thermocouples were windward for a test at angle of attack, the
pressure orifices were leeward by the same angle. There were 15 thermo-
couples and 11 pressure orifices located at positions shown in figure 5.
The inside diameter of the pressure orifices was 0.050 and the tubing
length to the Statham gages was 5 feet.

The surface roughness of the model was about 15 microinches as meas-
ured by a Physicists Research Co. Profilometer, Model No. 11, Type 9, for
the initial tests; however, during the ensuing runs the roughness increased
from 15 to an estimated value of 50 microinches.


TEST CONDITIONS


The model was tested at a Mach number of 2.0 at angles of attack of
00, t50, and 100. All tests were made in a free jet with a 27- by
27-inch nozzle which allowed testing at constant sea-level pressure and
temperature for 8 seconds. Reynolds number of the test based on body
diameter was 6.4 x 100. The model was injected into the airstream and
was on center line approximately 0.1 second after steady-flow conditions
of the jet had been reached.


DATA REDUCTION


The aerodynamic heat-transfer coefficients were calculated from
data measured during the transient heating of the model at the earliest
possible time, which was 0.1 second after the model was on center line.
At this time the estimated radiative and conductive heat loss to the


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rACA RM L58Cl4a CONFIDENTIAL 5


air behind the model was negligible. The heat-transfer coefficients
were calculated by using the equation


h pwcwtw dT,
Taw Tw dT


The time rate of change of wall temperature was obtained from plots
of the wall temperature as a function of time. The heat-transfer coeffi-
cient was then evaluated by using a mass density for Inconel of
518 lb/cu ft and a specific heat of 0.11 as given in reference 5. Meas-
ured values of skin thickness were used in all calculations.

In the case where the thermocouple was located internally on the
lip, the following equation was used to take in the variations in surface
area:


R2 dT
Qwc t ( 2 dT
h =
Taw Tw


where R1 is the inside radius of the lip and R2 is the outside radius.
This expression gives an effective thickness of the Inconel; as the result
of the equation, the average thickness is obtained by dividing the volume
of the material by the surface area. The value for the effective thickness
was 75 percent of the actual measured thickness.

The adiabatic wall temperature was obtained by using the equation



Taw Tt[r(i T! +)L
Tt) Tz
aw t 1r Tt t



where ir = (tNpr)l The fact that (Npr) 12 varies over the tempera-
T
ture range makes little difference since the minimum ratio of -1 on the
Tt
body was 0.9. The equation used to calculate the heating rates assumes
constant temperature through the wall.


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NACA RM L58C14a


RESULTS AND DISCUSSION


Pressure Distribution

Figures 6 and 7 show the pressure distribution for the model at
angles of attack of 00, f50, and 100. The local measured pressure was
divided by the maximum measured pressure to obtain the ratio of local
pressure to total pressure behind the shock as presented in figures 6
and 7. Figure 6 shows the windward pressure distribution plotted against
the distance along the surface from the stagnation point. For all angles
of attack, the three pressure stations up to and including the 600 sta-
tion, which was the last station at which measurements were taken on the
concave part, measured total pressure behind the shock. The dotted line
represents a fairing of the data for the 00 angle-of-attack test. The
five pressure-measuring orifices on the lip recorded free-stream static
pressures for the 00 angle-of-attack test. All pressure gages on the
lip were of low range 15 lb/sq in.

Figure 7 represents leeward pressure distribution plotted against
surface distance from the stagnation point. As in figure 5, the dotted
line represents a fairing of the data for the 00 angle-of-attack test.
The leeward tests also show total pressure behind the shock for the
three measuring stations up to 600 on the concave part. Figures 6 and 7
show that for these test conditions the model experienced total pressure
behind the shock up to and including the 600 station which was the last
station at which measurements were taken between 600 and the lip.


Heat Transfer

Figure 8 shows the wall-temperature distribution for the model at
angles of attack of 00, 50, and 100. The faired Line indicated by
0 seconds represents the wall temperature at which time the heat-transfer
coefficients were obtained as represented in figures 9 and 10.

Figures 9 and 10 show the heat-transfer distribution for the model
at angles of attack of 00, 50, and 100. Figure 9 presents the heat
transfer for the windward test at 50 and 100. The dotted Line is the
fairing for the data at an angle of attack of 00. This figure shows that
out to the station where S = 5.25 inches (750) the heat transfer remains
fairly constant. The station immediately inside the lip iT7) experienced
the highest heat transfer, but it is important to note that this heat
transfer remains constant with windward angle of attack. The lip at a
windward angle of 100 experienced heat-transfer coefficients only 22 per-
cent higher than those at an angle of attack of 00. The lowest heat-
transfer coefficients are on the cylinder which is in a region of low


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NACA RM L58Cl4a


pressure. The heat-transfer coefficients on the cylinder are the highest
when the model is at windward angles of attack of 50 and 100. The maxi-
mum heat-transfer coefficient for the test at an angle of attack of 5o
is 21 percent higher than that for the test at 00, and the maximum for
the test at 100 is 51 percent higher.

Figure 10 presents the heat-transfer coefficients for the Leeward
test at 50 and 100. The dotted line connects the points from the test
at an angle of attack of 00. This figure shows that the heat-transfer
coefficients for the leeward test, up to station S = 5.25 inches (75o)
are the same as for the windward test. The highest heat-transfer coeffi-
cient experienced in the leeward tests is on the inside of the lip (T ,
the same as in the windward tests. On the cylinder of the model the
heat transfer is lower than that of the 00 angle-of-attack test. The
lip, at a leeward angle of attack of 100, experienced 50 percent lower
heat-transfer coefficients than those of the 00 angle-of-attack test.

In figure 11, heating rates obtained for the model at an angle of
attack of 00 are compared with those obtained experimentally and calcu-
lated by laminar theory for both a hemisphere nose and a flat nose. This
comparison has been made by presenting the ratio of the local heat-
transfer coefficient to the stagnation-point heat-transfer coefficient
for a hemisphere of the same diameter. The calculations were made by
the laminar theory presented in reference 4 for both the hemisphere and
flat noses. The experimental data for the hemisphere (ref. 5) and the
flat noses were obtained in the preflight jet under sea-level conditions
at M1 = 2.0. .The test results indicated transition at about S/S1 = 0.35.
Comparison of these values with those for the concave nose shape at 00
angle of attack indicates that the Local heat-transfer coefficients on
the concave nose are lower than both theoretical and experimental values
for either the flat or the hemisphere nose, until S/SL = 0.825 or
S = 5.25 inches. The heat-transfer coefficient at the stagnation point
is approximately 40 percent of that on the hemisphere.

Another comparison may be made on the basis of total heat input.
In this case it is important to remember that the hemisphere and the con-
cave nose have twice the surface area as that of the flat-face model. The
following chart shows the comparison in total heat input for a Mach number
of 2.0; however, reference 2 shows a different relationship in total
heating for higher Mach numbers:


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8 CONFIDENTIAL


Nose shape Data obtained by Total heat input,
Nose shape Data obtained by Btu/sec

Hemisphere Laminar theory 574
Hemisphere Experiment 957
Flat face Experiment and 255
laminar theory
Concave Experiment 556

Comparison of the experimental heat-transfer distribution of a hemi-
sphere where transition occurs at S/Sl = 0.355 shows that the total
heating is 1.8 times higher than experimental values for the concave nose.
Since the freestream Reynolds numbers for the tests of the concave and the
hemisphere noses were about the same, this comparison indicates that the
concave nose might be extremely worthwhile in conditions where transition
would be expected on a hemispherical nose.

The nose shape reported herein was investigated by the Langley
Pilotless Aircraft Research Division on a two-stage rocket-propelled
model at Mach numbers between 5 and 7. These unpublished data indicated
stagnation-point values of only one-tenth to one-twentieth of those of
the hemisphere. Other tests as discussed in reference 6 indicate two
types of flow about the nose, steady and unsteady flow. Both types of
flow were observed under the same flow conditions and no reason could be
given to explain the two types of flow. The heat-transfer coefficients
measured in reference 6 for the unsteady flow were approximately 6 to
7 times the coefficients for the steady flow. The steady-flow coeffi-
cients in the tests of reference 6 varied from 20 percent to 50 percent
of the values at the stagnation point of a hemisphere. The differences
between the apparently steady flow results of the present test, the tests
of reference 6, and the flight tests have not been explained as yet.

CONCLUSIONS

From tests made in the preflight jet of the Langley Pilotless Air-
craft Research Division (at its testing station at Wallops Island, Va.)
at a Mach number of 2 and sea-level Reynolds numbers on a concave nose,
the following results are evident:

1. Pressures measured up to 600 on the concave part of the model
were equal to the total pressure behind the shock at angles of attack of
00, 50, and 100. The lip of the model experienced free-stream static
pressures at 00 angle of attack.

2. The heat-transfer coefficient at the stagnation point at 00 angle
of attack is approximately 40 percent of that on the same size hemisphere.


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NACA RM L58C14a






MJACA RM L58Cl4a


3. The highest local heat-transfer coefficient was measured immedi-
ately inside the lip; this heating was of the same magnitude for all angles
of attack. The heat-transfer coefficient on the lip at an angle of attack
windward of 100 was lower than that experienced immediately inside the lip.
4. Comparing the concave nose and the hemisphere, which were tested
at the same free-stream Reynolds number, on the basis of total heat input,
the hemisphere was heating 1.8 times higher. This comparison indicates
that the concave nose might be extremely worthwhile in conditions where
transition would be expected on a hemispherical nose. However, other
tests on concave noses indicate the possibility of unsteady flow condi-
tions in the cup which give a large increase to the total heat input. At
present, the conditions under which these unsteady flows are obtained are
not understood.


Langley Aeronautical Laboratory,
National Advisory Committee for Aeronautics,
Langley Field, Va., February 27, 1956.



REFEREIJCES

1. Carter, Howard S., and Bressette, Walter E.: Heat Transfer and Pres-
sure Distribution on Six Blunt Hoses at a Mach [lumber of 2. 11ACA
RM L57C18, 1957.
2. Hopko, Russell UI., and Strass, H. Kurt: Some Experimental Heating
Data on Convex and Concave Hemispherical Hose Shapes and Hemispheri-
cal Depressions on a 500 Blunted Hose Cone. 11ACA RM L5SA17a, 1953.
5. Lucks, C. F., Bing, J. F., Matolich, J., Deem, H. W., and Thompson,
H. B.: The Experimental Measurement of Thermal Conductivities,
Specific Heats, and Densities of Metallic, Transparent, and Pro-
tective Materials Part II. AF Tech. Pep. Io. 6145 (Contract
No. AF 55(058)-20558), Battelle Memorial Inst., July 1952.
4. Lees, Lester: Laminar Heat Transfer Over Blunt-Ilosed Bodies at
Hypersonic Flight Speeds. Jet Propulsion, vol. 26, no. 4,
Apr. 1956, pp. 259-269.
5. Chauvin, Leo T., and Maloney, Joseph P.: Experimental Convective
Heat Transfer to a 4-Inch and 6-Inch Hemisphere at Mach IIumbers
From 1.62 to 5.04. UIACA RM L55LOSa, 195 4.
6. Cooper, Morton, Beckwith, Ivan E., Jones, Jim J., and Gallagher,
James J.: Heat-Transfer Measurements on a Concave Hemispherical
[lose Shape With Unsteady-Flow Effects at Mach Hlumbers of 1.98 and
4.95. IIACA RM1 L58D25a, 1958.


COIFIDEIJTIAL


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Figure 2.- Shadowgraph of model at angle of attack of 0 .


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Figure 5.- Shadowgraph of model at angle of attack of 50.


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Figure 4.- Shadowgraph of model at angle of attack of 100. L-58-1612


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NACA RM L58C14a


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T 1T|tO T| T T12DIT1


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Inside radius of lip.O050


All measurements in inches

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ttilcon. le slong surface fr'm thickness. 0, deg station along surface from 0, deg
k center line, c, in. .center line,
S, in. tw, in.
1 0 .0c8 ,a 1 1 0 00
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14 5.. .061 90
15 5.49 .0'1 o70

inriermE'-.uple and pressure pi~ckup3 are numtered in order from center line.


Figure 5.- Sketch of model and station locations.


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NACA RM L58C14a


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O All angles

og 50 Windward pt =83.5 lb/sq in.

l 0a Windward pt =83.7 lb/sq in

& 00 Pt84.0 Ib/sq in.


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3.U 4.U

Distance along surface, S


Figure 6.- Pressure distribution at windward angle of attack.


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Figure 7.- Pressure distribution at leeward angle of attack.


CONFI DENTAL


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