Some effects of heat transfer at Mach number 2.0 at stagnation temperatures between 2,310° and 3,500° R on a magnesium f...

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Title:
Some effects of heat transfer at Mach number 2.0 at stagnation temperatures between 2,310° and 3,500° R on a magnesium fin with several leading-edge modifications
Series Title:
NACA RM
Physical Description:
29 p. : ill. ; 28 cm.
Language:
English
Creator:
Bland, W. M ( William M )
Bressette, Walter E
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
NACA
Place of Publication:
Washington, D.C
Publication Date:

Subjects

Subjects / Keywords:
Heat -- Transmission   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Magnesium -- Electrometallurgy   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Abstract:
Abstract: Four models of a thin magnesium fin, with the leading edge swept back 35°, have been tested in the preflight high-temperature jet of the Langley Pilotless Aircraft Research Station at Wallops Island, Va., at a Mach number of 2.0 for various stagnation temperatures between 2,310° and 3,500° R. This exploratory investigation was made to determine some effects of aerodynamic heating at high stagnation temperatures on the leading edges of fins and to determine the relative effectiveness of several leading-edge protective schemes.
Bibliography:
Includes bibliographic references (p. 11).
Additional Physical Form:
Also available in electronic format.
Statement of Responsibility:
by William M. Bland, Jr., and Walter E. Bressette.
General Note:
"Report date April 18, 1957."
General Note:
"Declassified December 13, 1957."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003853254
oclc - 154234111
System ID:
AA00009185:00001


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ARCH MEMORANDU




XFEC.TS OF 'HEAT TRANSFER AT MACH NUMBER 2.0
Tr AOw TEMPERATURES BETWEEN 2,3100 AND
NMlAGNESIUM FIN WITH SEVERAL
I $AtYING-EDGE MODIFICATIONS

p t X1a Tr., and Walter 'E. Bressette

P e I3y 3Aezvnautical Laboratory
Lrzgley Field, Va.

UNIVABY OF FLORIDA
|DO|i NS. DEPARW
I2QMAST SCIENCE UBRARY
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FL 32611-7011 USA


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IIACA EM L57C14

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

RESEARCH MEMORANDUM


SOME EFFECTS OF HEAT TRANSFER AT MACH NUMBER 2.0

AT STAGNATION TEMPERATURES BETWEEN 2,3100 AND

3,5000 R ON A lAGrrESILJR1 FIN WITH SEVERAL

LEADING-EDGE MODIFICATIONS

By William M. Bland, Jr., and Walter E. Bressette


SUMMARY


Four models of a thin magnesium fin, with the leading edge swept
back 350, of a type used to stabilize the first stages of rocket-
propelled multistage hypersonic models have been tested in the pre-
flight high-temperature jet of the Langley Pilotless Aircraft Research
Station at Wallops Island, Va. This exploratory investigation was made
to determine some effects of aerodynamic heating at high stagnation
temperatures on the leading edges of fins and to determine the relative
effectiveness of several leading-edge protective schemes.

Results of these tests, which were conducted at Mach number 2.0 for
various stagnation temperatures between 2,3100 and 3,5000 R, indicated
that under similar test conditions a magnesium fin with a blunt leading
edge suffered much less damage than one with a very sharp leading edge
even when only the mass remaining after blunting is considered. Also,
wrapping sheet Inconel around the leading edge proved to be a very
effective scheme for protecting the leading-edge region. Elementary
calculations appeared reasonably capable, though conservative, of pre-
dicting the time for melting to occur on the Inconel leading edge.


INTRODUCTION


Problems associated with flight at supersonic speeds have been
investigated in free flight by the Langley Pilotless Aircraft Research
Division with multistage rocket-propelled models. Conventional fins
have been used to stabilize the model-booster combinations at relatively
low supersonic speeds. These, fins had sharp leading edges to decrease
drag and were made of magnesium to decrease the weight. With the advent






NACA RM L57C14


of research at hypersonic speeds with rocket-propelled models, the
velocities attained by the fin-stabilized model-booster combinations
have been increased at low altitudes until the aerodynamic heating has
become severe enough to be damaging. Recently a large two-stage fin-
stabilized model, launched from the ground at about 550 above the
horizontal, was accelerated to a Mach number of 2.2 in 4.8 seconds by
the first-stage rocket motor. After a short interval of decelerating
flight, the model was accelerated by its rocket motor from a Mach num-
ber of 1.1 for 2.75 seconds until it unexpectedly underwent an abrupt
change in flight path that resulted in model destruction at a Mach
number of 4.7 and an altitude of approximately 15,500 feet. Fin fail-
ure could have caused the abrupt change in flight path. Subsequent
heating calculations were made using the actual flight-path conditions.
These calculations indicated that the fin leading edges could have
reached the melting temperature of magnesium about 1.5 seconds after
the beginning of the second period of acceleration and that the temper-
ature of the magnesium, 5 inches behind the leading edge, could have
risen about 5000 F by the time of model failure. The results of these
calculations thus indicated that the model was probably lost by failure
of the fins.

In order to determine some effects of aerodynamic heating at high
stagnation temperatures on the leading edge of the fins and to determine
the relative effectiveness of several leading-edge protective schemes,
an investigation has been initiated by the Langley Pilotless Aircraft
Research Division. The first phase of the investigation, as reported
herein, was conducted by testing a series of four uninstrumented models
of a booster fin in a jet at Mach number 2.0 in which the stagnation
temperature could be varied from 1,2000 to 4,0000 R. In some cases the
jet conditions were adjusted so that heating conditions at the fin
leading edge were similar to those encountered in free flight by the
model previously discussed. These tests were conducted at high stagna-
tion temperatures in the preflight jet of the Langley Pilotless Aircraft
Research Station at Wallops Island, Va.


MODELS


The plan form chosen for this exploratory investigation simulated
the outboard leading-edge portion of a thin booster fin with the rela-
tively small leading-edge half-wedge angle of 30 and the leading edge
swept back 550. Four models (see fig. 1) of this plan form were fabri-
cated from magnesium plate. One, the basic fin, represented the
leading-edge region of the full-scale booster fin. The models were
modified in the leading-edge region as follows:






NACA RM L57C4lk 5


Model 1 . . . Basic fin

Model 2 . . Blunt leading edge

Model 3 . Leading edge wrapped with -inch-thick
32
Inconel and blunted as model 2

Model 4 Magnesium leading edge replaced by one machined
from stainless steel and blunted as model 2


The models were not instrumented.


TEST PROCEDURE


The investigation was conducted by exposing the models at a Mach
number of 2.0 in the 12-inch-diameter preflight high-temperature jet.
Each model was mounted on a stand that would insert and withdraw it from
the jet once desired flow conditions had been established. The motion of
the stand was such that a model traversed about one-half the jet stream
while being rotated to the test position and while being withdrawn.
Approximately 0.4 second was spent traversing the jet stream in either
direction. Model 2 is shown erected to the testing position, in the
center of the jet, in figure 2. The black chordwise line indicates the
center of the jet in the vertical plane. A more detailed description
of the operation and characteristics of the high-temperature jet is pre-
sented in the appendix.

Motion pictures of the model and of an electric clock were taken
from one side and from overhead during each test at approximately
128 frames per second. These films provided the only source of data
from these tests other than jet operating conditions. From these films
were obtained the elapsed time each model was in the testing position
and, where applicable, the time at which leading-edge damage was first
observed.


TESTS AMD CALCULATIONS


Tests

General.- Calculated stream conditions along the center line of
the jet ahead of the model position are presented in figure 5 for dif-
ferent center-line stagnation temperatures. The stagnation temperatures
referred to are average values along the center line of the jet for the





NACA RM L57TC4


time of a test. Variances from the average are quoted for each test in
table I.

The tunnel was operated so that the stream static pressure along
the center line at the jet exit was 0.78 times the ambient pressure.
This resulted in a total pressure of 11,500 pounds per square foot behind
a detached shock which is ahead of the 550 sweptback leading edge. An
equivalent pressure would be obtained in free flight at Mach numbers 2.6
and 4.0 at altitudes of 20,000 and 40,000 feet, respectively.

Since the jet static pressure was less than ambient, shock diamonds
were formed near the exit and extended downstream to intersect several
inches behind the leading edges of the models. Information concerning
the shock cone is included in the appendix.

Test times given are for the interval of time a model was exposed
to the jet in the testing position. Other pertinent information con-
cerning the tests is included in table I.

Model 1, basic fin.- The basic magnesium fin with the very sharp
leading edge (1/64-inch radius) was inserted in the jet with the stag-
nation temperature at 2,5900 R. Melting of the wing leading edge was
observed to start near the jet center line at 0.6 second. Damage after
exposure for 2.5 seconds was extensive as shown in figure 4. As a mat-
ter of interest, the calculated heat input to the fin leading edge during
the 2.5-second test was of the same order as that calculated for the
flight condition discussed in the "Introduction."

Model 2, blunt leading edge.- The magnesium fin with the leading
edge blunted to a 1/16-inch radius was inserted in the jet with the
stagnation temperature at 2,5100 R. Melting of the wing leading edge was
observed to start slightly above the jet center line at 1.9 seconds which
is considerably later than the time melting was observed to start on the
basic fin. Damage after exposure for 2.5 seconds was relatively small as
shown in figure 5. The benefit derived by blunting the leading edge is
shown quite clearly by a comparison between figures 4 and 5. The black
line extending from root to tip just behind the leading edge of model 1
(basic fin) in figure 4, which shows the relative position of the
leading edge of model 2 (fig. 5), is indicative of the amount of material
removed from the basic fin to arrive at model 2. Thus for approximately
similar test conditions the fin with the sharp leading edge sustained
more damage than the fin with the blunt leading edge even when only the
reduced mass of model 2 is considered. That is, damage to the fin with
the sharp leading edge extended farther behind a line representing the
leading edge of model 2 than the damage to model 2. The stagnation tem-
peratures of these tests, 2,5900 and 2,5100 R, compare approximately with
the stagnation temperatures that would be obtained in flight at Mach num-
bers 5.0 and 4.9, respectively, at an altitude of 40,000 feet. (See
table I.)






NACA RM L57C14 5


Model 3, blunt leading edge wrapped with Inconel.- The fin with
the 1/52-inch-thick Inconel wrapped around the leading edge was exposed
in the jet at conditions slightly more severe than those of the previous
tests (2.5 seconds at a stagnation temperature of 2,5400 R) without dam-
age. In subsequent tests the fin was exposed for 2.4 seconds at 2,9100 R
and for 2.5 seconds at 5,2200 R. The Inconel was held in place by rivets
as indicated in figure 1. Prior to the tests, considerable concern was
expressed on the possible effectiveness of this type of attachment because
of the different coefficients of thermal expansion of Inconel and magne-
sium. Examinations after each of these tests disclosed only minor effects
such as Inconel discoloration, deformation of the rivet heads, and some
Inconel buckling between the rivets. The extent of each of these effects
of heating increased as the stagnation temperature increased. No evidence
of damage to the exposed magnesium surfaces was observed during or after
these three tests.

Model 5 was finally tested in the jet with the stagnation temperature
at 5,5000 R, the maximum stagnation temperature available at the time of
the investigation. After exposure for about 2.5 seconds the Inconel
melted at the leading edge near the jet center line and the magnesium
appeared to ignite under the Inconel and immediately behind the Inconel on
the side of the fin. Total time in the testing position was 53.2 seconds.
Damage to the leading edge and the rest of the fin is shown in figure 6.

Model 4, blunt leading edge made of stainless steel.- This model,
which had the magnesium replaced by stainless steel at the leading edge
and for a considerable distance behind the leading edge (see fig. 1) was
tested at a stagnation temperature of 5,5000 R. It was tested at only
this stagnation temperature because experience with model 5 indicated
that model 4 would survive exposure at the lower stagnation temperatures.
At about 2.5 seconds, as in the test of model 5 at the same stagnation
temperature, the magnesium was observed to ignite near the jet center
line immediately to the rear of the stainless steel, which was red hot
where it joined the magnesium. Total time in the testing position was
5.7 seconds. The stainless-steel leading edge was undamaged during the
test; however, considerable damage was sustained by the magnesium behind
the stainless-steel section as can be seen in figure 7. The stagnation
temperature of 5,5000 B of the last test of model 5 and of the test of
model 4 is comparable to the stagnation temperature that would be
obtained at a Mach number of 6.5 at an altitude of 40,000 feet.

Some idea of the relative effectiveness of the three leading-edge
protective schemes can be obtained from figure 8 and from table I.

The first damage observed during some of the tests was melting of
the magnesium; in other tests ignition of the magnesium was the first
damage abser-ied. According to reference 1, magnesium could be expected
to ignite near the melting temperature; therefore, the appropriate test
times for which ignition was first observed are taken as times for
melting to begin.





NACA RM L57C14


Calculations

Heating calculations at the leading edge, on the 1/52-inch-thick
Inconel, and at a station 2.5 inches behind the leading edge, on
0.17-inch-thick magnesium, have been made for model 5 at a stagnation
temperature of 3,5000 R by using simple heat balance relations. Gen-
eral assumptions made in performing these calculations are as follows:

1. No temperature gradients along the surface or through the material

2. No radiation

Assumptions made in performing the calculations at the leading edge are as
follows:

1. Flow is laminar

2. Adiabatic wall temperature equal to the stagnation temperature

5. Effective thickness of Inconel was taken as 74.8 percent of sheet
thickness. This resulted from dividing the volume of the material by the
surface area to obtain an average thickness.

4. No conduction to the magnesium enclosed by the Inconel

Assumptions made in performing the calculations behind the leading edge
are as follows:

1. Turbulent flow existed from the leading edge

2. Van Driest's values of the turbulent flat-plate skin friction
were applicable

5. Reynold's analogy constant was 0.6.

4. Recovery factor was equal to the cube root of the Prandtl number
based on wall temperature.

Average aerodynamic heat-transfer coefficients for the leading-edge calcu-
lations were calculated by the method of reference 2 for a two-dimensional
body.

Calculated wall temperatures at the leading edge and at a station
2.5 inches behind the leading edge are shown in figure 9. The temperature
calculations at the leading edge on the Inconel appear to be conservative,
that is, melting was calculated to occur at 2.0 seconds while actual fin
failure was observed to occur at the later time of 2.5 seconds. This con-
servatism in the calculation of the temperature of Inconel is influenced
by the assumption of no conduction and by the assumption of no radiation.






NACA RM L57C14


It should also be noted that, although this conservative calculation
failed to predict the time of model failure from Inconel melting by a
considerable fraction of the total test time, the actual error in heat
input was only about 10 percent of the total heat required to raise
Inconel to its melting temperature.

The calculated melting time for the magnesium (2.5 in. behind the
leading edge) was nearly 4.4 seconds. This, when compared with the
observed fin failure time of about 2.5 seconds supports the observation
made in the previous section that first failure on the fin occurred on
the Inconel.

Calculated heating rates are presented in figure 10 to give some
idea of the severity of these tests. These heating rates are somewhat
disproportionate; that is, under the artificial test conditions of the
high-temperature jet, the heating rate at the leading edge may not be
related to the heating rate behind the leading edge on the wing surface
in the same manner as it would be in free flight. For instance, at a
stagnation temperature of 5,5000 R, the calculated heating rate at the
leading edge is less than that calculated for altitudes of 50,000 feet
and under. On the other hand, the calculated heating rate on the sur-
face of the fin 2.5 inches behind the leading edge is less than that
calculated for altitude of 50,000 feet and under. Thus, for many tests
in the jet it is possible that the heating behind the leading edge may
become too severe when some leading-edge conditions are reproduced.
This discrepancy exists because the Mach number in the jet cannot be
varied to complete the simulation of the leading-edge conditions.


CONCLUSIONS


Results of seven tests of four models of a magnesium fin, three
with modifications designed to alleviate heating effects in the leading-
edge region, in a high-temperature jet at Mach number 2.0 indicate the
following conclusions:

1. Under similar test conditions, a magnesium fin with a blunt
leading edge suffered much less damage than one with a very sharp leading
edge even when only the mass remaining -after blunting was considered.

2. Wrapping Inconel around the leading edge, while a very simple
modification, proved to be a very effective scheme for protecting the
leading-edge region.






8 NACA EM L57C1k4


5. Elementary calculations appeared reasonably capable, though con-
servative, of predicting the time for melting of an Inconel leading edge.


Langley Aeronautical Laboratory,
National Advisory Committee for Aeronautics,
Langley Field, Va., March 4, 1957.






NACA RM L57C4lk


APPENDIX A


HIGH-TEMPERATURE JET


In order to study high-temperature effects on components of missiles
expected to obtain hypersonic speeds, it became necessary to develop a
ground-test jet capable of producing high-temperature flow associated
with hypersonic speeds. This high-temperature jet was obtained by ducting
air from the storage spheres of the preflight test facilities of the
Langley Pilotless Aircraft Research Division (ref. 5) through a fuel spray
and flame-holder donut-type burner where ethylene vapor fuel (C2H. is
injected into the airstream. The resulting combustible mixture is then
ignited and burned in a combustion chamber with the products of combustion
exhausted into the atmosphere through a convergent-divergent exit nozzle
at Mach number 2. A schematic drawing showing the internal character-
istics of the jet ducting and the Mach cone is presented in figure 11.
Also shown in figure 11 is the operation of the swing mechanism used for
inserting the test models into the hot jet after the jet has reached
steady-state conditions. The stagnation temperature of the jet exhaust
can be varied from the preheat values of the air supply by regulating
the fuel supply, whereas the static pressure at the nozzle exit is con-
trolled by regulation of the total pressure upstream of the burner. The
calculated variation in temperature with fuel supplied is presented in
figure 12. The adiabatic equilibrium flame temperature after burning
ethylene (C2H) fuel with fuel-air ratio was computed from data presented
in reference 4. Calibration of the exhaust-temperature profiles across
the nozzle exit at various injection values of fuel and air supply was
obtained up to a value of 2,7000 R by a temperature survey rake stationed
at 450 across the exit. Typical stagnation-temperature profiles obtained
with the temperature rake are presented in figure 15. Figure 15 shows
that the stagnation temperature is not constant across the nozzle exit
of the jet and that the maximum for any test is near the center line of
the jet. Also as the center-line stagnation temperature increased, the
temperature gradient from the center to the nozzle wall increased. For
center-line stagnation temperature near 2,500 R, the temperature gra-
dient was about 1000 per inch near the center line. Because center-line
stagnation temperature above 2,7000 R could not be measured in the cali-
bration of the jet, any center-line value above 2,7000 R must be esti-
mated by extrapolation of the values of the stagnation temperature and
fuel-air ratio.

Presented in figure 14 is the percent by weight of the ethylene
exhaust-gas products for fuel-air ratios less than the stoichiometric
value. The values of the exhaust-gas products were computed by the
method presented in appendix B of reference 5 with the assumption that





NACA RM L57C14


the hydrogen-carbon fuel (C2H4) is completely converted to carbon dioxide
and water vapor. It can be seen in figure 14 that the percent by weight
of Nitrogen in the gas is reduced by only 6 percent from a value of
77 percent for air to a value of 71 percent for the exhaust at the stoi-
chiometric value. The fact that the nitrogen content of the exhaust is
large and fairly consistent with the nitrogen content of air indicates
that the heat transfer from the high-temperature exhaust should be closely
similar to the heat transfer at high temperatures from air. However,
the reduction in oxygen with the resultant increase in carbon dioxide and
water may alter substantially the surface chemistry phenomena at high
temperature.

It is also necessary to have some idea of the thermodynamic prop-
erties of the exhaust mixture. Although these thermodynamic properties,
such as gas constant R, specific heat at constant pressure cp, and
ratio of specific heats 7, might vary slightly for different runs, a
reasonable engineering estimate can be computed by a weighted averaging
process for the values of temperature with fuel-air ratio as presented
in figure 15 by using data presented in references 4 and 5 for the ther-
modynamic properties of each of the exhaust products with temperature.
The computed values for cP and 7 are presented in figure 15 whereas
figure 5 of reference 5 shows, for the combustion of ethylene fuel which
has a hydrogen-carbon ratio of 0.168, that R remains approximately
ft-lb
55.5 --- over a range of fuel-air ratio from 0 to the stoichiometric
lb OF
value. Also presented in figure 15 are the cp and 7 variations with
temperature for air as obtained from references 6 and 7. It can be seen
in figure 15 that the thermodynamic properties of the exhaust gas are
very similar to those of air.

Thus, from the results presented in this appendix, it can be expected
that tests of aerodynamic shapes in the high-temperature jet simulate
tests in the atmosphere under similar temperature conditions.






NACA EM L57C14


REFERENCES


1. Hill, Paul R., Adamson, David, Foland, Douglas H., and Bressette,
Walter E.: High-Temperature Oxidation and Ignition of Metals.
NACA RM L55L25b, 1956.

2. Goodwin, Glen: Heat-Transfer Characteristics of Blunt Two- and Three-
Dimensional Bodies at Supersonic Speeds. NACA RM A55L15a, 1956.

35. Faget, Maxime A., Watson, Raymond S., and Bartlett, Walter A., Jr.:
Free-Jet Tests of a 6.5-Inch-Diameter Ram-Jet Engine at Mach Numbers
of 1.81 and 2.00. NACA EM L50L06, 1951.

4. Fricke, Edwin F.: Statistical Thermodynamics Applied to Chemical
Kinetics of Combustion. Rept. No. EDR-22-407, Republic Aviation
Corp., Oct. 1, 1947.

5. Pinkel, Benjamin, and Turner, L. Richard: Thermodynamic Data for the
Computation of the Performance of Exhaust-Gas Turbines. NACA
WR E-25, 1944. (Formerly NACA ARR 4B25.)

6. Hilsenrath, Joseph, Beckett, Charles W., et al.: Tables of Thermal
Properties of Gases. NBS Cir. 564, U. S. Dept. Commerce, 1955.

7. Ames Research Staff: Equations, Tables, and Charts for Compressible
Flow. NACA Rep. 1155, 1955. (Supersedes NACA TN 1428.)
























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