Exploratory investigation of transpiration cooling of a 40° double wedge using nitrogen and helium as coolants at stagna...

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Material Information

Title:
Exploratory investigation of transpiration cooling of a 40° double wedge using nitrogen and helium as coolants at stagnation temperatures of 1,295° to 2,910° F
Series Title:
NACA RM
Physical Description:
18 p. : ill. ; 28 cm.
Language:
English
Creator:
Rashis, Bernard
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
NACA
Place of Publication:
Washington, D.C
Publication Date:

Subjects

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

Notes

Abstract:
Abstract: The results presented herein were obtained in a transpiration cooling investigation of a double wedge of 40° included angle having a porous stainless-steel specimen inserted flush with the top surface of the wedge. Nitrogen and helium were used as coolants, and tests were conducted for stagnation temperatures ranging from 1,295° to 2,910° F. The data for both the nitrogen and helium coolants indicated greater cooling effectiveness than predicted by theory and were in good agreement with the results for an 8° cone tested at a stagnation temperature of 600° F.
Bibliography:
Includes bibliographic references (p. 10).
Additional Physical Form:
Also available in electronic format.
Statement of Responsibility:
by Bernard Rashis.
General Note:
"Report date August 1, 1957."
General Note:
"Classification changed to unclassified Authority: NASA Technical Publications Announcements No. 11 Effective date: December 1, 1959 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 - 003853034
oclc - 154195625
System ID:
AA00009191:00001


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CH: MEMORANDUM


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NACA RM2 L57F11 CONIFIDEN~TIAL

NATIONAL ADVISORY COM~MITEE FOR AERONUICS

RESEARCH MIEMORI~IilUM


EXPORATORY IN~VESTICATION OF TRANSPIRATION COOLING OF A

400 DOUBLE W~EDG3E UBICNG NITROGEN ANSD HELIUM AS

COOLANTS AT STAGNATION TEMPRATRE

OF 1,2950 TO 2,9100 F

By Bernard Rashis


SUMMARY


An investigation of transpiration cooling has been conducted in the
preflight jet of the Langley Pilotless Aircraft Research Station at
Wallops Island, Va. Th model consisted of a double wedge of 400 included
angle having a porous stainless-steel specimen inserted flush with the
top surface of the wedge. T~he tests were conducted at a free-strean Mach
numer of 2.0 for stagnation temperatures ranging fran 1,2950 to 2,9100 F.

H~itrogen and helium were used as coolants and tests were conducted
for flow rates ranging from approximately 0.03 to 0.30 percent of the
local weight flow. The data for both the nitrogen and helium coolants
indicated greater cooking effectiveness than predicted by theory and we~re
in good agreement with the results for an 8o cone tested at a stagnation
temperature of 6000 F.

The results indicate that t~he helium coolant, for the same amount
of heat-transfer reduction, requires anly about one-fourh to one~-fifth
the coolant flow weight as the nitrogen coolant.


INTRODUCTION


The development of long-range ballistic missiles and the steadily
increasing flight speeds of other missiles and fighter and bomber air-
planes require a thorough investigation of all the high-temperature
problems associated with their trajectories and speeds. Th destrue-
tive effects of aerodynamic heating for stagnation temperatures comn-
parable to a flight Mach number of 7.0 have already been reported in
reference 1. In hot-air-3et tests, copper and stainless-steel models


CO~F IDjENT IAL






2 C~OlI FIDEllTIAL IIACA RMI LS7FL1


were melted in, slightly less than 12 secondsc. It is apparent that con-
siderable effort must be expe~nded towards improving materials and devel-
oping various methods of cooling.

Transpiration cooling systems have already been investigated refss. 2
to 5) and do indicate considerable merit; however, the results were
obtained for a relatively low stagnation temperature. As an initial
effort in the study of transpiration cooling at high stagnation temper-
atures, a series of tests were conducted in the preflight jet of the
Langley Pilotless Aircraft Research Station at Wallops Island, Va.

The tests were conducted for stagnation temperatures ranging from
1,2950 to 2,9100 F. The coolants used were nitrogen and helium, and the
flow rates ranged fran approximately 0.05 to 0.50 percent of the local
weight flow. The model tested consisted of a double wedge of 400
included angle, in which a porous stainless-steel specimen was inserted
flush with the top surface. The free-stream Mach number of the tests
was 2.0, and the Reynolds numbers, based on the surface conditions and
the distance from the leading edge to the thermocoup~le locations, varied
from 0.578 x 106 to 8.2 x 10 The purposes of this paper are to present
these high-stagnation temperature results and to compare them with the
data of reference 2 and the theories of references j and 4, which apply
to nitrogen and helium coolants, respectively.


SYMIBOLS


b thickness of porous specimen, ft

cf local skin-friction coefficient

ep specific heat, Btu/!lb- F

G weight flow rate, l :bj(sq ft)(sec)

h heat-transfer coefficient, btu;~(sq ft)(sec)(oF)

k therml conductivity, Btu/(ft)(see)(OF)

p pressure, Ib/sq in, abs

t time~, see

T temperature, OF

Q heating rate, Btuf(sq f$)(sec)


COITFIDENTIAL







NACA RM~ L57F11


CONFIDNTAL


distance from leading edge of specimen, in.

ratio of coolant weight flow rate to local weight flow rate

Mach number


NPy


P~randt1 number

Reynol~ds number

Stanton number


Subscripts:


adiabatie wall for no coolant flow


coolant values


bounda~ry-latyer recovery value with coolant flow


local values


stagnlation values


local values outside~ porous surface

free-stream values


theory for GCe = 0


APPARATUS AND PROCEDURE


The present tests were conducted in the preflight jet which is ca~pa-
ble of producing a hot jet having a free-stream Mach number of 2.0 and
a maximum stagnationa temperature of 3,5000 F. A detailed description of
this facility is given in appendix A of reference 6.


figure 1, and a sectional
The test model consisted of
leading edge of the wedge
constructed of 1/2-inch-thick


The test model and stand are shown in
drawing of the model is given in figure 2.
a double edgee of 400 included angle. The
h~ad a 0.38-inch radius, and the walls were
cold-rolled steel.


CONFIDENPTIA






CONFIDENITING


NACA RM L57Fll


The test speecimen was of 1/8-inch-thich porous stainless steel
having approximately 25 percent porosity. It was inserted into the top
surface of the wedge flush with the surface. All joints and connections
were sealed with a high-temperature paste which hardens after curing and
acts as an insulation. The inside surfaces of the wedge were also covered
with this paste. Surface static-pressure orifices were Located as shown
in figure 2. The inside coolant pressure was measured by means of an
orifice located directly under the center of the porous specimen. The
specimen temperatures were obtained fran thermocouples which were spot-
welded to the inside surface of the specimen. The incoming coolant tem-
perature was measured by means of a thiermocouple mounted approximately
one-eighth of an inch below the specimnen, a slight distance from the
pressure tube. The thnnnerocuples were No. 3O gage chromel-alumel wires.
The pressure tubes were constructed of stainless-steel tubing having an
inside diameter of 0.060 inch.

The temperature instrumentation was calibrated before and after
each run. The maima error for the system was t1 percent of the maxi-
mnum range.

For each, test, the model was inserted into the jet stream only
after steady conditions had been achieved for both the jet and the
coolant flow.

The coolant: weight flow rates through the porous material were
obtained from the manfacturer's specified calibration curve, which is
given in figure 3. The manufacturer's curve gives the values of the
volume flow rate of air a~t a temperature of 700 F and an ambient pres-
sure of 1 atmosphere against the pressure drop across the porous moater-
ial. Since the pressure drop was measured for all the tests, the volume
flow rate was read from the curve and converted to weight flow rate by
multiplying the volume flow rate by the corresponding coolant density.

Since manufacturer's specifications were available only for pres-
sure drop values up to 10 pounds per square inch, several measurements
were made at higher pressure-drop values by using flowmeters. These
measurements were made with nitrogen and helium, the volume flow rates
through the flowmters being read off cuves computed from previous
calibration tests of the flowmeters where water was used. Although the
flowme~ter measurements shown in figure 3 are not extremely precise,
the extrapolation of the manufacturer's specification is reasonably
accurate, particularly since only extrapolation of the data to pressure-
drop values of 20.0 pounds per square inch were required.


CONFIDENT IAL






NACA RM L57F11 l


CONFIDENTIAL


ANALYST IS


The general form of the heat-balance equation is



saerodynamic absorbedd by coolant = (Itransie~nt f radiation f conduction
(1)

Only that portion of the data obtained after equilibrium conditions
had been achieved b~y the porous specimen was used in the reduction of the
inta. Calculations of the radiation and conduction terms, assumng a
value of emissivity of 0.8 and values of therml. conductivity that were
60 percent (ref. 7) of the nonporous stainl-ess steel, indicated that these
term were at most of the order of 1.0 to 2.0 Btu/(sq ft)(see). The values
of q absorbed by the coolants ranged from 40 to 225 Btu/(sq ft)(sec);
thus, the radiation and conduction terms -were assumed to be negligible
and were not used in the~ reduction of the data.

Thus the generally heat-balance equation becomes


Rarodync i = Qeabsored by coolant
or

h(T, T = Geop~,/ s T (2)

or, in nondimensional form,

N1St ? E c, Ts To
N~t,0 NSt,0 CaP2 f T a,


where NSt0 ws ""computed fron the local skin-friction values of the
zero-pressure-gradient theory of re~ferene 5, by using the local values
of Mach and Reynolds numbers, the x-distance being taken from the edge
leading edge to the thermocouple location. The! values of NiSt,0 are
obtained from the relattion

Nt0 = 1.24~f (4)


CONFIDENTIAL






ZOHIFjIDENTLAL


MACA RM L57Fl1


The values of T, were comrputed by using values of recovery factors
as computed front the theory of reference 8, which gives the variation of
recovery factor with Mach numer, coolant flow rate, and Reynolds number.
Although the recovery factors varied fran approximately 0.88 for no cool-
ant flow to 0.70 for a coolant flow rate of 0.5 percent of the local
weight flow for the nitrogen coolant, the percentage change in the ratio
NJStlNSt,O was less than J.0 percent. Since the variation in recovery
factors as given. in reference 8 was virtually the same as the experi-
mental. values for air reported in reference 5, the effect of cooling on
the recovery factor for the nitrogen coolant for the present tests can
be considered very minor. The variations in recovery factor for the
helium coolant as measured in reference 4 indicates that the present
data for helium would be even. less affected than the nitrogen data, and
for the low flow rates of the present tests, the theoretical values of
reference 7 which apply to air could be used for the present helium data.

Since the incoming coolant temperature is m~uch lower than the porous-
material temperature, the passage of the coolant through the porous speci-
Iren causes a temperature gradient across the thickness. The temperature
difference between the inside and outside surfaces may be calculated from

NSt,0cp~
rrT Fep,c (1 ekb
av a, 1 St 0ep,l



as given in reference 9. The values of T, used in equation (5) were
obtained by adding the calculated values of M: to the measured local
inside surface temperatures.

The porous specimen used for these tests was specified by the manu-
facturer as having approximately 25 percent porosity.

Since the composition of the combustion products of the jet exhaust
is essentially the same as air, the properties for air were used in the
evaluation of the jet exhaust flow. Tfhe specific heat values for air,
nitrogen and helium, and the viscosity values for air were obtained from
reference 10.

?The `valus of local weight flow, G1, were calculated fron



P, M, T1


CONFIDENTIAL







NACA RM L57TF11 CONFIDENTIAL r7


Figure 4 shows the measured-temperature distribution along the
porous specimen and the measured-~pressure distribution taken on the
solid surface 2.0 inches from the thermocouples. The values shown are
for the nitrogen coolant for a stagnation temperature of 2,9100 F. The
coolant temperature was 49.5o F. The reason for the pressure dropoff is
not definitely known. A check run was made with the wedge shown in fig-
ure 2 modified by the addition of side pieces. This modification made
the top surface of the wedge a rectangular section 12 inches wide by
approxtimatelyr 13 inches long; however, the pressure distribution was
unchanged. It thus appears that the dropoff is not due to the configu-
ration but is due to either a dropoff in the~ static pressure normal to
the jet axis or perhaps to an effect of the jet boundaries. Since the
pressure at the downstream end of the specimen is roughly one-half of
t~he theoretical sharp-wedge pressure, the low measured values cannot be
considered as due to the blunted leading edge. The temperature varia-
tion results from the difference in flow rates caused by the variation
of the pressure drop along the specimen.

Since the quantity

p MZ



was virtually constant for the different values of p the value of
G1 was considered to be constant along the specimen.


RESULS AND DISCUSSION


Figure 5 shows time histories of the measured inside surface tem-
perature for several nitrogen coolant flow rates ~tL a, stagnation tem-
perature of 2,9100 F. The coolant temperature was 49.50 F. The curve
for F = 0.~55 percent is for x = 0.5 inch; the curve for F = 0.1r75
percent is for x = 4.5 inches; the curve for F = 0.255 percent is
for x = 6.5 inches, the x-distance being measured from the leading edge
of the porous specimen. Also shown is the time history of calculated
inside surface temperature for no coolant fl-ow or F = 0 percent. The
x-distance for this curve is 4.5 inches. The curve was calculated by
using a value of NSt,O from reference 5, assuming a turbulent bound-
ary layer, and using the local values of Mach and Reynolds number.

Figure 6 shows the vari~iation of the cooling '"efficiency parameter"

Ts Te
Taw Ilc


CONF IMM~TIAcL






CONFIDENTIAL


NACA RM L57F11


with the nondimensional flow-rate parameter FINSt,0. This parameter
is usefu in, that it correlates the results obtained for different local
weight flow rates. Also show are thne results of reference 2 for the 80
cone for both nitrogen and helium coolants. The good agreement between
them and the present results is clearly indicated.

The flagged symbol in figure 6 represents an average of the results
obtained fran the test made with the helium coolant at a stagnation tem-
perature of 1,2950 F. TIhis procedure was used because the incoming cool-
ant flow rate was in the~ high range of the calibration curve and excessive
cooling of the porous specimen resulted. The surface temperatures aver-
aged 1200 F, and sensitivity of the instrumentation was too low to give
a proper evaluation of the effect of the slight differences in the cool-
ant flow rates.

Figure rl shows the variation of the ratio of the Stanton number to
theoretical Stanton number for no coolant flow with the flow parameter
BINSt,0. Included for comparison are the data of reference 2 and the
theoretical curves computed from, the theories of references 3 and 4.

The present results for the nitrogen coolant indicate slightly
greater reduction of th aerodynamic heat input than values computed from
thne theory of reference 3, which. assumes that the coolant and local prop-
erties ar~e identical.

The present data for the he~lium coolant also indicate slightly
greater cooling than the values computed from the theory of reference 4,
which is a modification of fits theory. The value of surface coolant
concentration was assumed to be 1.

Al-so shown in figure 7 is th curve computed from the theory of
reference 4 for the nitrogen coolant. Comparison of the two theoretical
curves shows that modified film theory does not indicate the same degree
of cooling effectiveness as indicated either by reference 3 or by the
present data.

The present data, and those of reference 2 agree for the nitrogen
coolant. There is some discrepancy between the two sets of data for the
he~lium coolant. It should be noted, however, that both sets of data for
the helium coolant are for small flow rates and any small errors in the
measurement of the flow rates would shift the data either downwards and
to the left or upwards and to the right for the parameters of figure 7.

The present results indicate that the helium coolant requires
approximately one-fourth to one-fifth the coolant flow rate of the nitro-
gen to achieve the same amount of heat-transfer reduction. It should
be noted that this ratio for required coolant flow rates is approximately
the same as the ratios of the heat'icapacities of helium to nitrogen.


CONTFIDENT IAL






NAC RM L[57F11


COurlDjEIITIAL


In figure 8, there is shown th~e variation of the heat capacity with
the final temperature of the coolant for. both nitrogen and ~hel~ium. For
the sam final coolant temperatures, 5000 to 1,0000 F, the helium has
approximately four times the heat capacity of the nitrogen.


COUICLUD~IN REMARKS


An exploratory invstigation has been made of thne transpiration
cooling on the surface of 400 wedge using nitrogen and helium as coolants
at stagnation temperatures of 1,2950 to 2,9100 F for a free-stream Mach
number of 2.0.

Substantial reduction of the aerodynamlic heat input was obtained at
high stagnation temperatures by~ means of transpiration cooling. For a
nitrogen coolant which has essentially the same~ properties as air,
slightly greater reduction of the aerodynamic heat input was obtained
than was predicted by a theory which assumes that the coolan and local
properties are identical. For a helium coolant, slightly greater cooling
was obtained than wazs predicted by a modified film theory.

The present results indicate that the helium coolant is from four
to five times more effective than the nitrogen coolant. This ratio is
approximately the sam as the ratio of the heat capacities of helium
and nitrogen.


Langleyr Aeronautical Laboratory,
National Advisory Comrmittee for Aeronautics,
Langley Field, Va. Mlay 21, 1957.


CONIFIDENJTIAL






CON~FIDENITIAL


HJACA RM L57Fll


REFEREN~CE;S


1. Purser, Paul E., and E~opko, Russell N.: Exploratory Materials and
Missile-Nose-Shape Tests in a 4,0000 3F Supersonic Air Jet. NACA
04 L56JO9, 1956.

2. Chauvin, Leo T., and Carter, Howar S.: Ekloratoiry Tests of Tran-
spiration Cooling on a Porous 80 Cone at Mi = 2.05 Using Nitrogen
Gas, H~eliun Gas, and Water as the Coolants. NACA RM L55C29, 1955.

3. Dorrance, Willian H., and Dore, Frank J.: The Effect of Mass Trans-
fer on the Compressible Turbulent Boundary Layer Skin Friction and
Heat Transfer. Rep. ZA-7-01j, Convair, Aug. 5, 1954.

4. Leadon, B. M., and Scott, C. J.: Measurement of Recovery Factors
and HEeat Tlransfer Coefficients With Transpiration Cooling in a Tur-
bulent Boundary Lay-er at MN = 3 Using Air and Relium as Coolants.
Res. Rep. No. 126, Univ. of Minnesota, Inst. Tech., Dept. Aero. Eng.
(Contract AF 18 (600)-1226), Feb., 19g56.

5. Resin, Morris W., Papp~as, Constantine C., and Okuno, Arthur F.:
The Effect of Fluid Injection on the Compressible Turbulent Bound-
ary Layer Preliminary Tests on Transpiration Cooling of a Flat
Plate at M = 2.7 With Air as the Injected G~as. N~ACA RMJ ASSIl9,
1-955.

6. Bland, Willian M., Jr., and Bressette, Walter E.: Some Effects of
Heat Transfer at Mach Numaber 2.0 at Stagnation Temperatures Between
2,3100 and 5,5000 R: on a Magnmesian Fin With Several Leading-Edge
MVod ifications. NACA FAIL57Cllk, 1957.

7. Evans, Jerry E., Jr.: Thermal Conductivity of 14 Metals and Alloys
up to 11000 F. NACA RM E50L07, 1951.

8. Resin, Morris W.: An Analytical Estiimation of the Effect of
Transpiration Cooling on the Heat-Transfer and Skin-Friction Char-
acteristics of a Com~pressible, Turbulent Boundary Layer. NACA
TTN 5541, 1954.

9. Hyma, Seymour C.: A Note on Transpira;tion Cooling. Jet Propulsion,
vol. 26, no. 9, Sept. 1956, p. 180.

10. Perry, John H., ed.: Chemical Engineers' Handbook. Third ed.,
Mc~raw-Hill Book Co., Inc., 1950.


CONFIDENTIAL







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