Preliminary survey of possible cooling methods for hypersonic aircraft

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Title:
Preliminary survey of possible cooling methods for hypersonic aircraft
Series Title:
NACA RM
Physical Description:
38 p. : ill. ; 28 cm.
Language:
English
Creator:
Esgar, Jack B
Hickel, Robert O
Stepka, Francis S
Lewis Research Center
United States -- National Advisory Committee for Aeronautics
Publisher:
NACA
Place of Publication:
Washington, D.C
Publication Date:

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Subjects / Keywords:
Hypersonic planes   ( lcsh )
Airplanes -- Motors -- Cooling   ( lcsh )
Aerodynamics -- Research   ( lcsh )
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federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Abstract:
Abstract: Many methods of cooling the structure of an aircraft capable of flight speeds up to 18,000 feet per second were studied. Water and hydrogen stored in the liquid state appear very promising as both coolants and heat sinks. The storage and circulation of hydrogen throughout the aircraft need not be a hazard. Cooling the outer skin of the aircraft in high-equilibrium-temperature regions could probably be avoided by using a material such as silicon carbide. The internal structure could be cooled by use of a thin layer of balsa wood saturated with water. In this way tanks for storage of coolant would be avoided.
Bibliography:
Includes bibliographic references (p. 22-24).
Additional Physical Form:
Also available in electronic format.
Statement of Responsibility:
by Jack B. Esgar, Robert O. Hickel, and Francis S. Stepka.
General Note:
"Report date January 7, 1958."
General Note:
"Declassified February 8, 1960."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003810781
oclc - 135378011
System ID:
AA00009182:00001


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ILI








NACA RM E57L19


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


RESEARCH MEMORANDUM


PRELIMINARY SURVEY OF POSSIBLE COOLING METHODS

FOR HYPERSONIC AIRCRAFT

By Jack B. Esgar, Robert 0. Hickel,
and Francis S. Stepka


SUMMARY

An investigation was conducted to determine the relative advantages
and limitations of a number of fluids for use as either heat sinks or
coolants for hypersonic aircraft. At flight speeds on the order of
18,000 feet per second it may be necessary to provide cooling for almost
the entire aircraft structure. The cooling problem can be divided into
two main categories: (1) high-level cooling in which the outside surface
of the aircraft must be cooled to keep the surface from melting or oxi-
dizing, and (2) low-level cooling in which the outside surface is allowed
to reach equilibrium temperature but the internal support structure must
be cooled to give it adequate strength. In general, cooling is much more
difficult in the high-level regions.

Several heat sinks are probably feasible for hypersonic aircraft.
Water appears best on a storage volume basis, and liquid hydrogen is best
on a weight basis. Although lithium and lithium hydride have better heat-
absorption capacities than either water or hydrogen, they do not appear
to be suitable as heat sinks at the present time because of difficulties
in handling molten lithium.

In the high-level regions suitable coolants would be hydrogen, water,
sodium, or a sodiLum-potassium mixture. Surface cooling could possibly be
eliminated in these regions by using uncooled surfaces made of silicon
carbide, impregnated graphite, or tungsten with a silicide coating. In
the low-level regions practically any coolant could be used for convec-
tion cooling, but water and hydrogen are particularly promising because
they are also good heat sinks.

It may be possible to eliminate the necessity of taking useful vol-
ume from the aircraft for heat-sink storage by imbedding balsa saturated
with water within the aircraft structure for low-level cooling and by
using uncooled surfaces of high-temperature materials in the high-level
regions.








NACA RM E57L19


INTRODUCTION

At hypersonic flight speeds aerodynamic heating may require that
structural members be cooled in some way. This paper presents results
of a preliminary study to determine feasible methods of cooling an air-
craft structure at sustained flight speeds up to at least 18,000 feet per
second at altitudes on the order of 180,000 feet. The aim of this study
is to find reliable methods of cooling that are low in weight and volume,
flexible enough to provide for variations in cooling requirements, not
unduly complicated, and not hazardous to personnel.

A number of studies, such as references 1 to 6, have considered air-
craft or aircraft-equipment cooling systems. Probably the most feasible
method of cooling at hypersonic speeds within the Earth's atmosphere is
to carry a heat sink or an expendable coolant within the aircraft. Air
cannot be taken aboard for direct use as a coolant because of the ex-
tremely high temperatures the air attains from ram compression. Refrig-
eration systems do not appear practical, because they also require a heat
sink at a reasonable temperature level, and the heat load to the sink is
increased because of the inherent inefficiencies in the refrigeration
system. It appears logical, therefore, to study the heat-transfer char-
acteristics and heat capacities of a number of fluids for possible use
as coolants and/or heat sinks.

At this stage of the investigation it is not expedient to make a
design study of cooling for a particular aircraft, because configurations
vary substantially. The configuration will affect the heat load, the
volume available for the coolant system, and the problems associated with
local hot spots. Instead, this study considers the problems associated
with a range of heat fluxes that should encompass the range required for
aircraft of the general class that would be capable of flight within the
Earth's atmosphere at speeds up to at least 18,000 feet per second. Al-
though much of the discussion is directed towards the cooling of wings,
the same methods should be equally applicable to the fuselage.


THE HEAT LOAD

A detailed study of aerodynamic heating was not made for this paper,
but some knowledge of the range of heat-transfer rates and equilibrium
surface temperatures to be encountered in hypersonic aircraft is required
in order to make an intelligent investigation of cooling methods. Heat-
transfer rates and equilibrium temperature are influenced by speed, alti-
tude, and aircraft configuration. The effects of speed and altitude on
equilibrium temperature are shown in reference 1. It will generally be
necessary to fly as high as possible at a given flight speed to reduce
aerodynamic heating, but the altitude attainable is also a function of








NACA RM E57L19


speed. As speed is increased, dynamic pressure is increased, and in
addition the lift that must be developed by the wings is decreased owing
to increasing centrifugal force. Both effects permit flight at higher
altitudes. When satellite velocity (about 26,000 ft/sec) is reached, no
vring lift is required. A study of the results in reference 1 reveals
that aerodynamic heating for a glide missile is most severe for flight
speeds between about 18,000 and 22,000 feet per second. This is the type
of vehicle that would require cooling devices such as those studied
herein.

The higher the wing loading, the more severe the aerodynamic heating
will be, because attainable altitudes for a given flight speed will be
lower. To determine an approximate idea of the magnitude of the heat
fluxes (heat-transfer rate per unit area) and equilibrium temperatures
to be encountered, heat-transfer calculations were made for a wing with
a 780 sweep angle and a 0.25-inch leading-edge radius, flying at 18,000
feet per second at 180,000 feet with a wing loading of 20 pounds per
square foot and at an angle of attack of 50. The calculation methods of
references 7 to 9 were used. The results of these calculations, which
are presented in figure 1, are not to be considered authoritative but
rather are estimates for a class of aircraft on which part of the present
analysis is based. High heat-transfer rates are encountered for only the
first inch of the leading edge, and beyond 6 inches on the lower surface
and 1 inch on the upper surface (measured along the surface normal to the
leading edge) the equilibrium temperature is less than 18000 F for a sur-
face emissivity of 0.9.

There are several schools of thought concerning the best type of
structure in the regions where the equilibrium temperature is 18000 F or
less. Some of the possible types of construction are (1) building the
structure of materials such as a molybdenum alloy that can safely with-
stand this temperature if a suitable oxidation-resistant coating can be
found, (2) building the structure of high-temperature alloys such as
those developed for gas-turbine engines and applying enough internal
cooling to reduce the structure to a safe operating temperature (e.g.,
12000 to 16000 F), and (3) utilizing essentially an uncooled outer skin,
a layer of insulation, and an internally cooled support structure as il-
lustrated by the low-level configuration in figure 2(b). In addition to
the aforementioned components in the third type of construction, rein-
forcement channels such as illustrated in figure 2(b) would have to be
placed periodically throughout the structure in order to fasten the outer
skin to the cooled corrugated component and to add rigidity to the outer
skin. The degree to which the internal structure would have to be cooled
would be a function of the material used for the structure. Approximate
temperatures for structures might be 2500 F for aluminum, 5000 F for tita-
nium alloy, and 12000 F for a high-temperature alloy such as Inconel X.








NACA RM E57L19


The study herein will consider primarily the cooled structures. Only
7 inches of the wing (measured normal to the stagnation line) near the
leading edge requires surface cooling. This will be called the "high-
level" cooling region. In the regions where the equilibrium temperature
is 18000 F or less ("low-level" cooling regions), only the interior struc-
tural members of the wing will require cooling. The temperature to which
the structure must be cooled depends on the materials. The heat flux to
these regions is quite low less than 1/2 Btu/(sq ft)(sec) for a 1-inch-
thick layer of low-conductivity insulation such as Thermoflex.


THE HEAT SINK

A variety of elements and compounds shown in reference 1 are capable
of absorbing a large quantity of heat if vaporized, but usually the va-
porization temperature is higher than the 18000 F that is considered to
be the maximum temperature for the aircraft. Figure 3 shows the heat-
absorption capacity of several elements and compounds that may be suit-
able for an aircraft heat sink. The heat-absorption capacity is consid-
ered to be zero at the coolant storage temperature that appears prac-
ticable for the aircraft. For helium or hydrogen this temperature is
the boiling temperature at a pressure of 1 atmosphere. Approximately
1000 F is the storage temperature for the other heat-sink materials. The
gases helium and hydrogen obtain their heat capacity from a high specific
heat and a large possible temperature change. Water and the light metals,
sodium and lithium, obtain most of their heat capacity through vaporiza-
tion, as shown by the vertical lines on the figure. The temperature at
which this vaporization occurs is influenced by the pressure level. Va-
porization is shown for two pressure levels for each of the light metals
and water. The compounds methanol CH30H and lithium hydride LiH ob-
tain most of their heat capacity through dissociation. The dissociation
occurs over a range of temperature and is a function of pressure level
(ref. 10). The rate of dissociation is also influenced by the concentra-
tion of the dissociation products. Complete dissociation would require
separation of the products from the original compound. This might be
quite difficult for methanol, since both the original compound and the
products would be in the gaseous state.

Figure 3(a) shows that on a weight basis the potential heat capacity
of water is lower than that of any of the other heat-sink materials con-
sidered. Since volume is also an important consideration in an aircraft,
the heat capacity on a volume basis is shown in figure 3(b). Water has
the highest density of all of the heat-sink materials considered; conse-
quently, it shows up better on a volume basis than on a weight basis.
At 15000 F the volume of water required to absorb a given quantity of
heat is only about one-sixth of the volume required for helium and about
one-fourth of that required for hydrogen, but the weight of hydrogen re-
quired is less than one-third that of water.








NACA RM E57L19


The liquid metals and lithium hydride show up very favorably in
figure 3 based on heat-absorption capabilities, but there are a number
of difficulties associated with handling liquid metals at high temper-
atures refss. 11 and 12). Lithium hydride dissociates to molten lithium
I and hydrogen; therefore, its handling properties will be similar to those
for lithium. Lithium readily absorbs nitrogen and forms lithium nitride,
which is extremely difficult to remove. In the molten form lithium ni-
tride might almost be called the universal solvent, as it will dissolve
practically all known containers at high temperatures. Further research
is required on suitable methods of handling molten lithium before serious
consideration can be given to its use in hypersonic aircraft. Sodium
does not present such a difficult problem. If a system is kept clean
and, in particular, free of oxygen, sodium can be handled very well up
to temperatures of at least 12000 F. Mass transfer may be a problem at
higher temperatures where sodium vapoiizes. For flights of short dura-
tion this may not be a problem, but oxygen leaking into a system could
cause oxide plugging if the sodium has to flow through any small openings.

Helium and hydrogen (particularly hydrogen) appear quite promising
as heat sinks if the volume required is not excessive for the aircraft
in question. The storage of a liquefied gas in aircraft appears to be
feasible. Figure 4 shows some of the details of a possible method of
tank construction. The tank is formed of concentric shells of stainless
steel with Styrofoam insulation between the shells. The insulation,
which should be about 2 to 3 inches thick, has a density of about 2.5
pounds per cubic foot. The volume where this insulation is located could
be evacuated to reduce heat transfer. The insulation serves as a separa-
tor between the two shells. As a result, the load due to the vacuum on
the outer shell is quite low and the shell can be quite thin. A thick-
ness of from 0.010 to 0.020 inch seems sufficient. The inner shell should
be thick enough to withstand the internal pressure, which should probably
be about 50 pounds per square inch absolute in order to provide sufficient
pressure to circulate the gas if it is also used as the coolant.

With the proper operating procedure, carrying the hydrogen aboard
the aircraft should not be a hazard to personnel. Purging and explosion
problems associated with circulation of hydrogen throughout the coolant
passages in the aircraft are believed to be easily controllable and will
be discussed later.

Methanol shows a very good heat capacity on either a volume or weight
basis (fig. 3). A nickel catalyst is required for the dissociation re-
action, and the dissociation products are carbon monoxide and hydrogen
gas. A further investigation of the dissociation reaction of methanol
concerning rates, vapor pressures, and removal of dissociation products
from the methanol gas would probably be warranted.








NACA RM E57L19


The gas or liquid used as the heat sink may or may not be used as
the circulating coolant. For simplicity it would be desirable for the
heat-sink fluid to serve also as the circulating coolant, because it
would then be unnecessary to carry and store more than one fluid. This
would also eliminate the complexity, weight, and volume occupied by a
heat exchanger.


HEAT-TRANSFER CHARACTERISTICS OF SEVERAL COOLANTS

Configurations Investigated

If the entire structure of an aircraft is to be cooled, combining
the heat-transfer and structural members of the airframe is desirable.
One method of doing this is to use a corrugated structure similar to
that shown in figure 2. This structure can be lightweight and rigid and
at the same time provide cooling passages in the corrugations. For this
study no attempt was made to optimize the configuration on either a heat-
transfer or a structural basis. Probably some compromise will ultimately
be required, because a single configuration is quite unlikely to be best
for both uses. The configuration for this analysis was arbitrarily chosen.
Studies that have been made on similar types of coolant-passage geometries
for turbine blades (ref. 13) have shown that there is considerable free-
dom in choosing configurations; therefore, the configuration chosen is
probably adequate at the present stage of investigation and will be suit-
able for comparing the relative merits of various coolants.

Figure 2 shows configurations for high-level, low-level, and leading-
edge cooling. In the high-level and leading-edge regions (figs. 2(a)
and (c), respectively), cooling is applied directly to the outside skin
in order to reduce its temperature to 18000 F. For simplicity the con-
figuration directly at the leading edge is similar to that for the high-
level region behind the leading edge, and the dimensions of the corruga-
tions (triangular passages) are such that they fit into a leading edge
having a 0.25-inch radius.

For the low-level regions (fig. 2(b)), the equilibrium temperature
of the skin is 18000 F or less; therefore, direct cooling of the skin is
not required. The outer skin probably would not be sheet material but
would be a honeycomb or corrugated structure to provide added rigidity.
The temperature that the interior structure can withstand will be con-
siderably less than 18000 F; therefore, structural cooling is still re-
quired. A 1-inch-thick insulation blanket of Thermoflex (thermal con-
ductivity equal to 1.1 (Btu)(in.)/(sq ft)(hr)(oF)) would considerably re-
duce the quantity 6f heat transferred to the interior structure. The
corrugation configuration was assumed to be the same as used in the high-
level regions. The permissible temperature level at the corrugations
would depend on the materials used in the structure. Three approximate








NACA RM E57L19


average temperature levels were considered: 2500 F for an aluminum
structure, 5000 F for a titanium structure, and 12000 F for a high-
temperature-material structure such as Inconel X.


Assumptions and Conditions for Calculations

As previously pointed out, the fluids used for coolants need not be
the same as those used for heat sinks, although many of the heat-sink
fluids are also excellent coolants, and the over-all system can be sim-
plified if one fluid is used for both purposes. The following coolants
were considered for the high-level and low-level regions: Air (closed
system, not taken aboard from atmosphere), helium, hydrogen, water,
steam, and NaK (a mixture of sodium and potassium having a melting point
as low as 120 F depending upon the relative amounts of sodium and potas-
sium). Although sodium was considered as a heat-sink material, NaK was
considered as the coolant because of its lower melting point. Its heat-
transfer properties do not differ greatly from those of sodium, so that,
as a first approximation, results presented for NaK are also applicable
to sodium.

High-level and leading-edge cooling. The calculations for high-
level and leading-edge cooling were made over a range of heat fluxes from
1 to 150 Btu/(sq ft)(sec), which more than covers the range shown in fig-
ure 1. The general calculation procedures were as follows: With gaseous
coolants and with a given heat flux and coolant inlet temperature (cool-
ant temperatures from 00 to 10000 F were assumed), the coolant flow was
set to give a wall temperature of 15000 F at the coolant inlet. Varia-
tions in wall temperature, coolant temperature, coolant pressure loss,
coolant Mach number, and pumping power were calculated as a function of
distance from the inlet using methods similar to those in reference 13
and using average heat-transfer coefficients for triangular passages from
reference 14. The inlet pressure was assumed to be 50 pounds per square
inch absolute.

When NaK was used as a coolant, the wall temperature was assumed to
be the same as the coolant temperature because of the extremely high
heat-transfer coefficients that are obtained with liquid-metal cooling.
As mentioned previously, sodium (or NaK) can be handled easier if its
temperature is kept at less than about 12000 F. In addition, NaK boils
at about 14400 F at 1 atmosphere of pressure, so that the surface tem-
peratures will have to be lower when NaK is used as a coolant than when
the coolant is a gas. For these calculations the flow rate of NaK was
set by allowing its temperature to rise from 12000 to 13500 F in 5 feet
of coolant-passage length. The pressure loss and pumping power were cal-
culated for these conditions.

With water as the coolant, the wall temperature was assumed to ap-
proach the water temperature of 2810 F (saturation temperature at 50








NACA RM E57L19


lb/sq ft abs). The coolant-flow rate was set to allow 10 percent of the
water to vaporize in 5 feet of coolant-passage length. The pressure loss
and pumping power were calculated for these conditions.

Low-level cooling. The heat fluxes used for the low-level-cooling
calculations were determined by the amount of heat that would be conducted
through the 1 inch of Thermoflex insulation for a surface temperature of
18000 F and the assumed internal structure temperature (a function of
material used). For gases the flow rate was that which resulted in an
inlet Mach number of 0.05 and the coolant temperature for an inlet wall
temperature of the internal structure of 2000, 4000, or 11000 F for alum-
inum, titanium, or Inconel X structures, respectively. These assumed
inlet temperatures are lower than the average temperatures previously
mentioned, in order to allow for heating of both the coolant and the pas-
sage wall along the passage length. Variations in wall temperature, cool-
ant temperature, coolant pressure loss, coolant Mach number, and pumping
power were then calculated as a function of distance from the inlet. The
inlet pressure was assumed to be 50 pounds per square inch absolute.

Calculations using NaK as a coolant were similar to the high-level
cooling calculations, except that wall temperatures and heat fluxes were
lower. The NaK temperature was allowed to increase 1000 F in 5 feet of
passage length.

Calculations using water as the coolant were also similar to those
for high-level cooling except that the heat flux was lower.


Comparison of Coolants

The question arises whether it is more advantageous to pass the
coolant through all the passages in the corrugated structure or only
through those passages adjacent to the outside surface. The cooling may
be adequate with either procedure, but the pressure losses are less if
the coolant passes through all passages because of the larger flow area.
As a result, the comparison of coolants shown in figure 5 is for the cool-
ant passing through all passages. Heat-transfer and pressure-loss results
are shown for two heat fluxes. For NaK and water, however, heat fluxes
will be higher than for the gaseous coolants for a given flight condition
because of the necessity of reducing the skin to a lower temperature, as
previously discussed. This lower temperature results in higher aerody-
namic heat transfer to the skin and less radiation from the skin. The
heat fluxes in figure 5 were therefore adjusted (increased) for NaK and
water to make the comparison fair for all coolants.

Figure 5(a) compares the coolants for a heat flux of 150 Btu/(sq ft)
(sec), which would be about the maximum that could be expected at the
stagnation line of the wing. At a pressure level of 50 pounds per square








NACA RM E57L19


inch absolute it was not possible to pass enough air or steam through
the cooling passages to obtain adequate cooling, and consequently air
and steam are not shown in figure 5(a). Calculations showed that it was
not possible to cool with a heat flux of 100 Btu/(sq ft)(sec) with these
two coolants either. With hydrogen and NaK the maximum length of passage
that could be cooled was a little over 3 feet because of wall temperature
rise with hydrogen and pressure-drop limitations with NaK. Helium could
cool a little less than 1- feet of passage before a pressure-drop limit
was reached.

If higher inlet pressures had been used, helium and NaK would have
been better coolants, because they were limited by pressure loss and/or
Mach number before they were limited by excessive temperature rise of
the wall along the passage length. It was also interesting to find that,
when helium and hydrogen were used as coolants, the wall temperature 1
foot from the inlet was actually less than the inlet wall temperature
because of a transition from laminar to turbulent flow, which improved
the heat transfer. This wall temperature reduction, which was small, is
not shown by the figure. The inlet coolant temperature for hydrogen and
helium was 00 F for the results shown. Water, which was allowed to
vaporize partially, was the only coolant with enough heat capacity per
unit volume to permit cooling a passage 5 feet long.

The power requirements shown are those required to restore the
coolant to 50 pounds per square inch absolute at the end of the passage
length considered and for the weight flow required for cooling a surface
1 foot wide. By restoring the pressure the coolant could be circulated
again. This would result in better utilization of the heat capacity if
the coolant were also the heat sink. If the coolant were not the heat
sink, the coolant system would be a closed cycle with heat rejected to
a heat sink by means of a heat exchanger. In this case the coolant would
be recirculated after rejecting heat to the heat sink, and the pressure
would definitely have to be restored. Because of the high pumping power
requirements and the complication of installing a gas pumping system, the
coolants hydrogen and helium, which are also heat sinks, would probably
be exhausted at a low pressure and would not be pumped back to a high
pressure for further circulation to better utilize the heat capacity.
For the liquid coolants such as NaK and water, which would require re-
circulation, the power requirements are quite low.

From figure 1 it can be seen that within an inch of the leading edge
the heat flux has dropped to a value of less than 10 Btu/(sq ft)(sec),
so that only a very small portion of the aircraft will have the difficult
cooling problem considered in-figure 5(a). In figure 5(b) the heat-
transfer and pressure characteristics of the coolants are shown for a
heat flux of 10 Btu/(sq ft)(sec). The temperature rises, pressure drops,
and coolant Mach numbers required for adequate cooling are probably








NACA RM E57L19


satisfactory for passage lengths up to at least 5 feet for all coolants
except air. With air as the coolant, wall temperatures may become ex-
cessive because of rapid coolant temperature rise. At this heat flux
the flow in the coolant passages was found to be laminar for all coolants.
For laminar flow the heat-transfer coefficient is essentially constant
regardless of flow rate within the passages. As a result the wall tem-
perature is controlled by controlling the coolant temperature. The re-
sults shown in figure 5(b) are for an inlet Mach number of 0.1 and an
inlet coolant temperature on the order of 2000 F less than the wall tem-
perature. If very cold coolant were used, the structure would be over-
cooled, regardless of the flow rate, near the entrance of the passages.
This overcooling can result in a waste of heat-sink capacity and in pos-
sible thermal stress problems. A possible method of at least partially
overcoming this problem will be discussed later.


LEADING-EDGE AND HIGH-LEVEL COOLING

Suitable High-Level Coolants

From the study of heat-transfer characteristics of the coolants con-
sidered in this investigation, it was found that air or steam would be
almost completely unsatisfactory for heat fluxes of 100 Btu/(sq ft)(sec)
or higher. Water is the best coolant if 10 percent of the water is
allowed to vaporize along the length of the coolant passage. The pres-
sure drop and pumping power for water are almost negligible, and quite
long passages could be cooled. The next best coolant is NaK, followed
by hydrogen and helium.

Freezing could be a serious problem with water or NaK as the coolant.
Freezing could be encountered on the ground or at high altitudes before
reaching speeds where aerodynamic heating becomes a problem. Possible
solutions to the freezing problem are circulating the coolant in a heated
state during freezing conditions, or adding propylene glycol to water to
provide protection to approximately -650 F. Absolute protection against
freezing would be required, because local blockage by frozen coolant
could quickly result in overheating or burnout in some other portion of
the system during hypersonic flight.

The freezing problem could be overcome by using hydrogen as the
coolant in the high-level regions. With hydrogen, the cooling passage
lengths would have to be chosen so that there were no limitations due to
pressure loss or excessive wall temperature rise. For a heat flux of
150 Btu/(sq ft)(sec), a length of about 3 feet appears to be practical
for a supply pressure of 50 pounds per square inch absolute. Higher
supply pressures, which would probably require use of a liquid-hydrogen
pump, would help a hydrogen-cooled system, both by permitting higher
pressure drops and by causing a higher heat capacity per unit volume of
hydrogen.








NACA RM E57L19


Another difficulty encountered with hydrogen as a coolant and heat
sink is the utilization of the full heat capacity that is, heating the
hydrogen to a temperature near the wall temperature. Using the full
heat capacity would require pumping so that the hydrogen could be recircu-
lated, and the power requirements for this might be exorbitant. It
would probably be more economical to discharge the hydrogen without com-
pletely using its heat capacity rather than to pump it. For the major
portion of the high-level region where the heat flux is less than 10
Btu/(sq ft)(sec), quite long coolant passages could be utilized, but
here again difficulty would be encountered in utilizing the full heat
capacity of the coolant without pumping.


Operating Procedure with Hydrogen Coolant

One of the principal factors affecting the use of hydrogen as a
coolant is the safety problem. Aside from the fact that hydrogen has
very wide flammability limits, it is probably no more dangerous than the
natural gas used for heating houses. By taking a few precautions it
should be possible to circulate hydrogen through the aircraft structure
without creating a hazard to the pilot.

One problem that must be overcome is the possible occurrence of a
damaging explosion when hydrogen and air become mixed within the coolant
passages of the aircraft structure. For most applications with hydrogen,
the passages are purged with an inert gas such as helium before the hy-
drogen is introduced and after it is shut off. In the cooled aircraft
structure, purging with helium may not be feasible because of complica-
tion, weight, and volume of the purging equipment. If the pressure with-
in the structure were sufficiently low, however, detonation pressures
would be tolerable, if a detonation did occur. In addition, the prob-
ability of a detonation is quite low for structure temperatures less
than 7500 F at the time the hydrogen and air are mixed refss. 15 to 18).
Detonation pressure ratios are a maximum of about 60 (ref. 19). At the
altitudes that the aircraft will fly, the base pressures can be low
enough that if detonation does occur the detonation pressure within the
coolant passages will be 2 atmospheres or less. Since these passages
should probably be designed to stand over 3 atmospheres, the occurrence
of detonation would not damage the structure.

Figure 6 shows minimum altitudes as a function of flight Mach number
where it would be safe to charge or bleed the hydrogen cooling sys-
tem without danger to the structure if a detonation occurred. The figure
is based on bleeding the system to the base pressure that would occur
at the rear of the fuselage or at the trailing edge of the wing. The
data for the figure are from unpublished experimental results of Feshotko
and Cortwright at the HIACA Lewis laboratory. As long as the hydrogen
coolant is turned on or off and purged with air at altitudes above the







NACA RM E57L19


minimum shown on the curves, there should be no hazard from a possible
detonation. Flight above the minimum altitudes shown in figure 6 for
corresponding Mach numbers seems feasible, and structural cooling would
not be required at these conditions. These conditions should, therefore,
be safe for turning the hydrogen flow on or off.


Utilization of Hydrogen Heat Capacity

Figure 3(a) showed that on a weight basis hydrogen is a superior
heat sink at almost any temperature level, but it would be desirable to
raise the temperature of hydrogen as high as possible in order to utilize
its full heat capacity. This requires the use of hydrogen over a tem-
perature range from its boiling point of about 400 R (-4200 F) up to as
high a temperature as possible. Two problems are immediately obvious:
First, unless the proper design is used, some areas of the structure are
liable to be much overcooled because of the very low hydrogen temperature
as it comes from the storage tank. This overcooling wastes heat-sink
capacity and may cause structural problems resulting from thermal expan-
sions and contractions. Second, full utilization of the hydrogen heat
capacity probably is not possible unless pumping is used. In the low-
level region the low structure temperatures limit the temperature to
which hydrogen can be heated; and in the high-level region pressure-drop
limitations are a problem, as previously discussed.

There are at least two possible solutions to the problem of over-
cooling. One solution is the use of counterflow cooling of the corru-
gated structure as illustrated in figure 7. The metal used for the cor-
rugations forms a natural divider, so that counterflow can be easily ob-
tained if coolant is introduced to the passages above the corrugated
divider on one end and to the passages below the corrugated divider on
the other end. This method of cooling should reduce metal temperature
variations along the length of the cooling passages and eliminate drastic
overcooling. The highest metal temperature will be near the midpoint of
the passage. Overheating could occur at this location if the passages
were too long.

A second solution to the overcooling problem is the use of a closed-
circuit cooling system and a heat exchanger to reject heat to the heat
sink. The coolant in the closed circuit would have to be either hydrogen
or helium to preclude freezing in the heat exchanger. The temperature
of the coolant in the closed circuit could be high enough, however, that
overcooling of the structure would not. occur. A serious disadvantage of
this type of system is the weight and complication of the heat exchanger
and pump required for circulating the secondary coolant. Therefore,
counterflow cooling is probably a better solution.








NACA RM E57L19


This discussion of heat-capacity utilization is applicable to all
materials that must be heated to high temperatures to exploit their full
heat capacity. Further complications result if the material must be
vaporized in order to obtain the major portion of the heat capacity. In
these cases mixtures of both liquid and vapor must be handled, and under
some conditions careful design is required to avoid "slugging" or burnout
from vapor pockets.


Upper Wing Surface as a Heat Sink

Figure 1 shows that the equilibrium temperature of the upper wing
surface is less than 10000 F at distances from the leading edge greater
than about 30 inches. This area could possibly be used as a heat sink
for part of the high-level cooling region. Assuming that the coolant is
a gas, such as helium or hydrogen, which would have a temperature inter-
mediate between the high-level region and the heat-sink surface temper-
ature, it might be possible to operate with a heat-sink surface temper-
ature of about 15000 F and radiate heat picked up in the high-level
region to the atmosphere. Calculations indicate that, if the area shown
in figure 8(a) were used as a heat sink, between 30 and 50 percent (depend-
ing on surface emissivity) of the heat picked up from the wing high-level
regions could be dissipated in the heat sink for an aircraft about 75
feet long.

Circulation of the coolant from the high-level regions to the heat
sink requires additional ducting, a pump capable of operation at about
16500 F, a power source for the pump, and a modification to the wing
structure to provide an additional corrugated structure on the outside
surface for the radiating heat exchanger, as shown in figure 8(b).
Whether this added complication and weight increase would offset the
savings in weight in the fluid heat sink is not known. It appears
doubtful, however, that use of the upper wing surface as a heat sink
for high-level cooling would be warranted.


Effect of Coatings on Heat Flux

In areas where there is a high heat flux, large temperature reduc-
tions across relatively thin coatings are possible if the coating has a
sufficiently low thermal conductivity. Stabilized zirconia coatings
(Rokide "Z") are capable of withstanding very high temperatures. Calcu-
lations were made to determine the effectiveness of such coatings in the
high-level regions. The results are shown in figure 9 for coating thick-
nesses of 0.025 and 0.050 inch from a metal temperature under the coating
of 18000 F. The figure shows that temperature drops up to 16000 F are
possible through the coating. The resulting higher surface temperature
on the outside of the coating would reduce the convection heat transfer








NACA RM E57L19


and increase the heat loss by radiation for a given surface emissivity.
The emissivity of zirconia coatings is only about 0.2 at high tempera-
tures; therefore the coatings are not as effective in radiating heat as
other common materials. The high-level heat fluxes that must be removed
by cooling, shown in figure 9, are actually greater with the coating
than without the coating for either thickness considered except for about
the first 0.4 inch of the leading edge. It therefore appears that coat-
ings will not effectively reduce high-level heat flux unless coatings
can be found that have both low thermal conductivity and high emissivity.
Such coatings capable of withstanding the very high temperatures required
are unknown to the authors.


Transpiration Cooling

It is well known that transpiration cooling is theoretically the
most effective method of cooling known at the present time. The analyt-
ical study of reference 20 shows how much coolant could be saved if tran-
spiration cooling were used rather than convection cooling. The savings
were appreciable with air as a coolant, but considerably smaller with
helium as coolant. Experience and calculations refss. 21 and 22) have
shown that attaining the ideal coolant flows for transpiration cooling
is extremely difficult, particularly if there are pressure gradients on
the surface. A relatively complex metering device for the coolant will
be required in order to obtain the proper flows at various locations for
the range of altitudes that must be encountered, and the permeability of
the surface must be controlled very carefully more carefully than has
been found practicable up to the present time. If these factors are not
all carefully controlled, coolant flows can become greater than required
for convection cooling, and local areas can overheat. Past experience
in attempting to transpiration-cool turbine blades, which have many of
the same cooling problems as wing leading edges, indicates that it would
not be wise to attempt to use transpiration cooling in the high-level
region at the present state of development.


Uncooled Materials

Figure 1 shows that, for a surface emissivity of 0.9, which is pos-
sible for several materials, the equilibrium temperature is 34000 F at
the stagnation line and drops rapidly farther back along the leading
edge. Conduction would probably lower the equilibrium temperature at
the stagnation line. Several materials could probably be used without
cooling at these temperature levels. Possible materials are tungsten,
molybdenum, graphite, and silicon carbide. Reference 23 discusses re-
search on coatings for molybdenum and tungsten. Information on graphite








NACA RM E57L19


is given in reference 24. The properties of these materials are summa-
rized in the following table:


Material Coating Maximum Specific Tensile Sur-
temperature, gravity strength face
oF at 35000 F, emis-
psi sivity

Molybdenum MoSi2 Coating, 3200 10.2 2,800
Mo melts, 4750

Tungsten Silicide Coating, 3800 10 to 19.3 10,000 ?
W melts, 6150

Graphite SiC Coating, 2900 1.6 to 1.7 4,000 0.9
to 3300
Impregnated, 4000
C sublimes, 6600

Silicon None Oxidizes, 2900 3.2 12,000 at 0.9
carbide to 3360 27500 F
Decomposes, 4500


The materials that appear most promising in this table are tungsten,
graphite, and silicon carbide. The surface emissivity of coated tung-
sten is not known by the authors. If it is low, the equilibrium temper-
ature at the leading edge can be considerably higher than 35400 F. The
specific gravity of tungsten is a function of sintering and pressing
techniques. Graphite has excellent thermal shock properties; its
strength increases with temperature up to about 50000 F; and its thermal
conductivity is high up to 5 times that of steel at room temperature,
but decreasing with increasing temperature. Conduction would undoubtedly
lower the stagnation-line equilibrium temperature for a graphite leading
edge. The silicon carbide coating for graphite starts oxidizing at about
29000 F. The rate of decomposition is relatively slow at higher temper-
atures, but if the coating is thin the protection may soon be lost. Ref-
erence 25 states that graphite impregnated with refractory metal resists
oxidizing atmospheres at temperatures well over 40000 F without dimen-
sional change.

Recent developments with silicon carbide have made it a very prom-
ising material for use on uncooled leading edges. When about 7 percent
molybdenum disilicide is added to almost pure silicon carbide, very good
oxidation resistance has been reported at 32700 F and fair resistance at
33600 F. Pure silicon carbide starts oxidizing at about 29000 F, and
oxidation is quite severe at 33000 F. The material has good thermal shock









NACA RM E57L19


properties and a thermal conductivity about half that of graphite, and
its strength at elevated temperature should be adequate.

Coatings for molybdenum have not been very satisfactory since, if
a pin hole develops, the oxide of molybdenum, which is a vapor, will
escape and the structure will disintegrate. In addition, the material
has less high-temperature strength than tungsten, graphite, or silicon
carbide, and the coating is good only to 32000 F; therefore, molybdenum
is not considered a suitable uncooled material for the aircraft.

The use of an uncooledd" leading edge still does not entirely elim-
inate the cooling problem in this region of the aircraft. This uncooled
member will have to be attached to a lower-temperature cooled structural
member and will transfer heat by conduction at the points of attachment
and by radiation in other areas. This radiant heat transfer from the
uncooled member to the cooled members can be much less than 10 percent
of the heat that would be transferred by convection to a cooled leading
edge. In order to obtain low heat-transfer rates, low-emissivity coat-
ings such as stabilized zirconia (Rokide "Z") should be used on the
cooled portion of the structure behind the uncooled skin.


LOW-LEVEL COOLING

Several cooling methods are probably feasible in the low-level
region of the aircraft. For a double-wall insulated structure the cool-
ing could be by convection or by a built-in heat sink as shown in figure
10. A convection-cooled structure would have the advantages that the
coolant could be turned on only when needed, thereby saving in the quan-
tity of coolant required, and that the structure could be cooled to al-
most any desired temperature. It has the disadvantages that additional
ducting would probably be required and a pumping device may be necessary
for circulating the coolant if the tank pressure is not sufficient. In
addition, lightweight structures that may be adequate for carrying the
structural loads may not be suitable for very high internal coolant pres-
sures. Coolant sealing could also be a problem.

The built-in heat sink is advantageous in that it is simple, that
no tanks are required for coolant storage, and that more uniform cooling
probably results. Its disadvantages are that renewing or replacing the
heat sink between flights may be a problem, that there is less control
over the temperature to which the structure can be cooled, and that the
cooling cannot be turned off or on at will. A further possible difficulty
with water as the.heat sink is that the internal structure will be cooled
more than it has to be because the low pressures encountered at altitude
will result in boiling temperatures less than 1000 F unless some pressur-
ization is used within the wing. Quite possibly the increase in structural








NACA RM E57L19


weight to permit pressurization would much more than offset the small
weight increase in water required for overcooling.

Reducing the quantity of beat-sink fluid required with the built-
in heat sink might be possible if the structure is capable of with-
standing a higher temperature than the vaporizing temperature of the
heat sink. In this case the quantity of fluid in the heat sink could
be controlled so that it would provide adequate cooling in the early,
high-speed part of the flight. In the later portion of the flight, when
the flight speed and equilibrium temperatures have decreased, cooling
may no longer be needed. By proper planning, the heat-sink fluid can be
expended by that time with a saving in the quantity needed.

Other cooling methods could be used with single-wall uninsulated
structures. For these structures, the material used in the structure
should be capable of withstanding as high temperature as possible to re-
duce the heat load. Some convection cooling, probably with a gas because
of the high temperature levels, could be used for controlling structure
temperature. This would be similar to the method used for high-level
cooling. Another possibility would be the use of full-depth honeycombs
extending from lower to upper wing surfaces. Radiation from the lower
surface to the much cooler upper surface would permit the use of un-
cooled structures over much of the wing surface if the stricture could
withstand temperatures up to about 16000 F. This temperature appears
feasible for some applications where the structure loads are small if
the structure is made of a nickel- or cobalt-base high-temperature alloy.


Convection Cooling

With the type of structure shown in figure 10 there are no real
problems in heat removal by convection cooling with any of the coolants
considered in the low-level regions. In general, calculations showed
pressure losses and coolant velocities can be quite low. The permissible
length of coolant passages and the coolant temperature rises can be con-
trolled to a considerable extent by the mass-flow rates. Generally, the
structure temperature will be very close to the coolant temperature. As
mentioned previously, however, the utilization of heat-sink capacity is
a problem with low-level cooling, because the heat sink cannot be heated
to a higher temperature than the permissible structure temperature in
the low-level region. In addition, the fact that the wall temperature
will be very close to the coolant temperature can result in drastic over-
cooling in some areas, particularly if hydrogen or helium is used as the
coolant. This overcooling can be avoided by using counterflow-. cooling,
as discussed previously and illustrated in figure 7.

Because of the difficulty in fully utilizing heat sinks such as
hydrogen, helium, or the liquid metals, water is a promising heat sink








NACA RM E57L19


for the low-level regions. Water could be used for convection cooling
by circulating the water through the structure at approximately boiling
temperature at a rate that causes about 10 percent of the water to vapor-
ize through the circuit. The vapor would be bled off and the remaining
water would be recirculated.

A coolant system that circulates a liquid coolant suffers the dis-
advantage of considerable dead weight from the residual coolant that is
necessary to fill the system. Since the system must remain full to the
end of the time that cooling is required in order to ensure circulation,
this residual coolant cannot be used for a heat sink. Since a large
area must be cooled, this can amount to almost as much weight in some
cases as the weight of water that would be required for a heat sink for
all the low-level regions on the airplane. If water is used as the heat
sink, the complication of tanks and pumps for recirculating the water
can possibly be eliminated by storing the water at the location where
the heat is to be absorbed by the built-in heat sink, as will be dis-
cussed later.

The discussion up to this point on convection cooling in the low-
level region has dealt with a double-wall insulated structure as shown
in figure 10. For an uninsulated structure the cooling problem would be
no worse than that illustrated in figure 5(b). The structure could be
cooled with any of the coolants considered. Water or NaK would probably
cool the structure more than necessary, and the total heat loads would
be high. A gaseous coolant is therefore indicated. The best coolant
would be hydrogen, and next best would be helium. Although the heat
loads would be higher for the single-wall than for the double-wall con-
struction because of higher permissible surface temperatures in the
double-wall construction, the heat capacity of a heat sink like hydrogen
could be more fully utilized because it could be heated to a higher tem-
perature. As a result, the weight of heat-sink fluid may not be any
higher for the single-wall construction. To determine which type of con-
figuration is the best for cooling would require a heat-transfer analysis
based on the structural configurations considered for a particular appli-
cation and for the desired coolant and heat-sink fluid.


Cooling with Built-In Heat Sink

A heat-sink cooling system (fig. 10(b)) should be able to maintain a
reasonably constant temperature by utilizing the heat of vaporization of
a liquid. The temperature level can be controlled by choice of the heat-
sink fluid and by the pressure level maintained in the system. For the
fluids that w.jere considered in this analysis, water would be suitable
for temperatures up to about 2120 F, sodium for temperatures from about
10000 to 16000 F, and lithium at temperatures in excess of 16000 F. A
further requirement for heat-sink cooling is that the heat-sink material








NACA RM E57L19


be capable of proper distribution over the areas that must be cooled,
and that its distribution not be adversely affected by acceleration
forces that might be experienced in the airplane. With all factors con-
sidered, water appears the most practical fluid for heat-sink cooling at
this time.

'The next question is how to hold the water in the proper locations
under all flight conditions. Probably the most practical way of storing
the water is to let it be absorbed in some porous material that can be
distributed throughout the airframe. Most materials that are capable of
absorbing large quantities of water will lose most of the water under
acceleration forces of a few g's. An exception, however, is saturated
balsa. Experiments conducted at the NACA Lewis laboratory have shown
that balsa wood can be saturated to a specific gravity greater than
unity if submerged in boiling water under 15 pounds gage pressure. The
weight of the saturated balsa is up to 10 times its dry weight. Centri-
fuge tests have shown a water loss of less than 2 percent after 5 min-
utes at 5 g's. Most of this loss was probably due to evaporation rather
than centrifugal-force effects. Reat-transfer tests in an apparatus that
simulated the structure shown in figure 10 revealed that temperatures
could be adequately controlled with saturated balsa as the heat sink.
For a flight about 1 hour long with about 40 minutes of aerodynamic heat-
ing at the conditions considered in determining the heat load for figure
1 (18,000 ft/sec at an altitude of about 180,000 ft), the required balsa
thickness would vary from about 0.1 to 0.3 inch, depending on the sur-
face equilibrium temperature, which is a function of distance from the
leading edge.

Certain difficulties are involved in the use of saturated balsa as
a heat sink. It tends to distort upon drying; if exposed to air under
the heating conditions encountered in this application it will char after
drying; and resaturation may be difficult within the aircraft because it
would require circulating pressurized hot water throughout the entire
heat-sink area. Because of these difficulties balsa is probably not the
best carrier for water in a built-in heat sink, and some other material
may be better. This preliminary study has shown, however, that cooling
with a built-in heat sink is feasible and that materials can be found
that can absorb adequate amounts of xater and not lose this water under
acceleration forces up to at least 5 g's.


Weight Effects

In order to determine the optimum material to use in the main struc-
ture of the aircraft, it is necessary to know the effect of structural
temperature on the weight of the cooling system. In general, structures
designed for higher temperatures will be heavier, and the cooling systems
will be lighter; consequently, the weight of structure has to be balanced








NACA RM E57L19


against the weight of the cooling system. Several weight comparisons
are made for a double-wall insulated cooling system in figure 11 for
hydrogen and water as heat sinks. In this figure the weight of the cool-
ing system is given per square foot of aircraft surface, and it is taken
as the weight of 3-pound-per-cubic-foot Thermoflex insulation plus the
weight of heat-sink material required to absorb the heat that would have
to be removed during 40 minutes of aerodynamic heating. The cooling-
system weights obtained by this analysis are probably optimistic. No
tank or ducting weight was included for hydrogen, and it was assumed
that the hydrogen could be heated to the same temperature as the struc-
ture. This assumption gives the maximum heat capacity possible. For
water a 10-percent dead weight was included in the calculations to ac-
count for balsa wood used to hold the water. Experience may show, how-
ever, that another 5- or 10-percent increase in weight may be necessary
because of difficulty in obtaining complete drying of the balsa during
flight.

Figure 11(a) shows the effect of insulation thickness for an outer-
skin temperature of 18000 F. As would be expected, the weight is less
with hydrogen than with water as the heat sink because of the greater
heat capacity of hydrogen. In addition, water could not be used for
structure temperatures higher than about 2120 F for reasonable internal
wing and fuselage pressures because water absorbs its heat by evapora-
tion. With hydrogen, the higher the structure temperature, the more
heat can be absorbed because of higher temperature rises to the heat
sink. This is doubly effective because the heat flux is decreased owing
to smaller temperature gradients through the insulation at higher struc-
ture temperatures. Figure 11(a) shows that it would be detrimental to
increase the insulation thickness from 1 inch to 2 inches for hydrogen,
because the increased weight of insulation more than outweighs the re-
duction in hydrogen required for cooling. If water is used as the heat
sink, however, it would be advantageous to increase the insulation thick-
ness, if possible, to reduce weight.

Figure 11(b) shows how insulation thickness could be decreased in
the aircraft as structural temperature is increased and a constant heat
load is maintained. The weight saving shown as structure temperature is
increased and insulating thickness is decreased is due primarily to added
heat capacity in the hydrogen at high structure temperatures. The effect
of insulation weight reduction is very small. Comparison of figures
11(a) and (b) shows that 1 inch of insulation thickness approaches an
optimum for hydrogen as the heat sink. For a given structure temperature
the coolant-system weights for 1 inch of insulation are slightly less
than those for thinner insulation, but the effect of thickness is quite
small. This indicates that, if hydrogen is used as the heat sink, there
can be considerable freedom in the choice of insulation thickness.








NACA RM E57L19


Figure 11(c) shows the effect of skin temperature on coolant-system
weight. The equilibrium temperatures of the skin will vary with distance
from the leading edge, flight speed, and altitude. The skin temperature
has a significant effect on the weight of water required for a heat sink
but only a small effect if hydrogen is the heat sink.

Increasing the structure temperature from 2120 up to 12000 F could
result in a maximum reduction in hydrogen cooling-system weight of about
1/2 pound per square foot of surface area (fig. 11). An analysis of the
structure is needed to determine whether this weight saving would warrant
building the structure of high-temperature materials. In addition, from
0.3 to 1.3 pounds per square foot of aircraft surface can be saved by
using a hydrogen cooling system rather than a water cooling system, de-
pending on structure temperature and insulation thickness. This weight
saving has to be balanced against the volume that would be required for
storing hydrogen. Essentially no useful volume is taken by the water
if it is in balsa imbedded within the aircraft structure.


CONCLUSIONS

From this preliminary investigation, the following conclusions ca-i
be drawn:

1. Several heat sinks are probably feasible for hypersonic air-
craft. Water and liquid hydrogen appear to be most practical, but
sodium also looks promising. At the present time lithium and lithiuiur
hydride do not appear to be suitable as heat sinks because of difficulties
in handling molten lithium.

2. It appears that storage and circulation of hydrogen in an air-
craft need not be a hazard; therefore, hydrogen would be better as a
heat sink than helium, which would have similar handling and storage;
problems but less heat capacity than hydrogen.

3. Cooling or temperature-reduction devices that were studied and
that do not appear to be practical at the present time in high-level
regions are: (1) use of the upper wing surface as a heat sinR, because
of weight and complication of ducting and pumps required, K.2) lo--therrral-
conductivity coatings, because known coatings that may be applicable have
such low emissivities that they do not reduce the heat flu-x to the cool-
ant, and (5) transpiration cooling, because coolant-flou. distributions
that are presently obtainable in regions .-where there are large pressure
gradients are so far from ideal that coolant requirements may be no
smaller than for convection cooling.

4. Uncooled surfaces nade of silicon carbide, inpr"gnated ,raphir.
or tungsten with a silicide coating appear to be feasible in the hi _h-
level regions of the airplane, so that high-level cooling could possiblyc
be eliminated.








NACA RM E57L19


5. If it is not desirable to use uncooled surfaces in the high-level
regions, cooling could be accomplished with hydrogen, water, sodium, or
a sodium-potassium mixture. A counterflow cooling system should be used,
particularly with hydrogen, to reduce temperature variations and prevent
overcooling in some areas. Hydrogen has the advantage that pumps and
heat exchangers would not be required.

6. Practically any coolant could be used for convection cooling in
low-level regions, but water and hydrogen are particularly promising be-
cause they are also good heat sinks. A cooling-system weight analysis
showed that increasing the structure temperature from 2120 to 12000 F
could result in a maximum reduction in hydrogen cooling-system weight of
about 1/2 pound per square foot of surface area. In addition, the hydro-
gen system can be from 0.3 to 1.3 pounds per square foot lighter than
the water system, depending on structure temperature and insulation
thickness.

7. The reduced weight of a hydrogen cooling system has to be bal-
anced against the volume required for hydrogen storage. It appears pos-
sible to eliminate the necessity of taking useful volume from the air-
craft for heat-sink storage by imbedding balsa saturated with water with-
in the aircraft structure for low-level cooling and by using uncooled
surfaces of high-temperature materials in the high-level regions.


Lewis Flight Propulsion Laboratory
National Advisory Committee for Aeronautics
Cleveland, Ohio, January 7, 1958


REFERENCES

1. Masson, D. J., and Gazley, Carl, Jr.: Surface-Protection and Cooling
Systems for High-Speed Flight. Preprint No. 638, Inst. Aero. Sci.,
June 1956.

2. Dukes, W. H., and Schnitt, A.: Structural Design for Aerodynamic
Heating. Pt. I Design Information. TR 55-305, Wright Air Dev.
Center, Wright-Patterson Air Force Base, Oct. 1955. (Contract
AF-33(616)-2581.)

3. Dukes, W. H., and Schnitt, A.: Structural Design for Aerodynamic
Heating. Pt. II Analytical Studies. TR 55-305, Wright Air Dev.
Center, Wright-Patterson Air Force Base, Oct. 1955. (Contract
AF-33(616)-2581.)








NACA RM E57L19


4. Geankoplis, C. J., Kay, W. B., Lemmon, A. W., and Robinson, W.: Heat-
Transfer Fluids for Aircraft-Equipment Cooling Systems. TR 54-66,
Wright Air Dev. Center, Wright-Patterson Air Force Base, Feb. 1954.
(Contract AF-33(616)-147.)

5. Zimmerman, R. H., and Robinson, W.: Equipment Cooling Systems for
Aircraft. Pt. 1 Summary of Cooling System Study. TR 54-359,
Wright Air Dev. Center, Wright-Patterson Air Force Base, Sept. 1954.
(Contract AF-33(616)-147.)

6. Zimmerman, R. H., Robinson, W., Hornung, K. G., and Krauss, W. E.:
Equipment Cooling Systems for Aircraft. Pt. 3 Cooling Systems
Evaluation. TR 54-359, Wright Air Dev. Center, Wright-Patterson
Air Force Base, Sept. 1954. (Contract AF-33(616)-147.)

7. Reshotko, Eli, and Beck-with, Ivan E.: Compressible Laminar Boundary
Layer over a Yawed Infinite Cylinder with Heat Transfer and Arbi-
trary Prandtl Number. NACA TN 3986, 1957.

8. Eckert, E. R. G.: Engineering Relations for Friction and Heat Trans-
fer to Surfaces in High Velocity Flow. Jour. Aero. Sci., vol. 22,
no. 8, Aug. 1955, pp. 585-587.

9. Goodwin, Glen, Creager, Marcus 0., and Winkler, Ernest L.: Investi-
gation of Local Heat-Transfer and Pressure Drag Characteristics of
a Yawed Circular Cylinder at Supersonic Speeds. NACA RM A55H51,
1956.

10. Gibb, Thomas R. P., Jr., and Messer, C. E.: A Survey Report on
Lithium Hydride. NYO-3957, United States Atomic Energy Comm., Oak
Ridge Nat. Lab., May 1954. (Contract AT(30-1(-1410).)

11. Anon.: Liquid-Metals Handbook. Second ed., Atomic Energy Comm. and
Bur. Ships, Dept. Navy, 1952. (NAVEXOS, P-733.)

12. Hoffman, E. E., and Manly, W. D.: Comparison of Sodium, Lithium, and
Lead as Heat Transfer Media from a Corrosion Standpoint. Preprint
No. 74, Am. Inst. Chem. Eng., 1955.

13. Esgar, Jack B., Schum, Eugene F., and Curren, Arthur N.: Effect of
Chord Size on Weight and Cooling Characteristics of Air-Cooled
Turbine Blades. NACA TN 3923, 1957.

14. Lowdermilk, Warren H., Weiland, Walter F., Jr., and Livingood, John
N. B.: Measurement of Heat-Transfer and Friction Coefficients for
Flow of Air in Noncircular Ducts at High Surface Temperatures.
NACA RM E53J07, 1954.








NACA RM E57L19


15. Stephan, Elmer F., and Peoples, R. S.: A Literature Survey on the
Explosions of Mixtures of Gases Containing Hydrogen. Battelle
Memorial'Inst., Columbus (Ohio), July 1954.

16. Lewis, Bernard, and Von Elbe, Guenther: Combustion, Flames and Ex-
plosions of Gases. Academic Press, Inc., 1951.

17. Dixon, Harold Baily, and Croft, James Murray: The Firing of Gases
by Adiabatic Compression. Pt. II The Ignition-Points of Mixtures
Containing Electrolytic Gas. Jour. Chem. Soc. Trans. (London),
vol. 107, 1914, pp. 2036-2053.

18. Drell, Isadore L., and Belles, Frank E.: Survey of Hydrogen Combus-
tion Properties. NACA RM E57D24, 1957.

19. Ordin, Paul M.: Hydrogen-Oxygen Explosions in Exhaust Ducting.
NACA TN 3935, 1957.

20. Rubesin, Morris W.: The Influence of Surface Injection on Heat
Transfer and Skin Friction Associated with High-Speed Turbulent
Boundary Layer. NACA RM A55L13, 1956.

21. Esgar, Jack B.: An Analytical Method for Evaluating Factors Affect-
ing Application of Transpiration Cooling to Gas Turbine Blades.
NACA RM E52G01, 1952.

22. Esgar, Jack B., and Richards, Hadley T.: Evaluation of Effects of
Random Permeability Variations on Transpiration-Cooled Surfaces.
NACA RM E53G16, 1953.

23. Fitzer, E.: Materials of Maximum Creep and Oxidation Resistance Pro-
duced by Coating Tungsten and Molybdenum with Vapor-Deposited
Silicon. Trans. 2956, Henry Brutcher Tech. Trans., Altadena
(Calif.). (Trans. from Berg- und Huttenmannische Monatshefte, vol.
97, no. 5, 1952, pp. 81-92.)

24. Hogerton, John F., and Grass, Robert C., eds.: The Reactor Handbook.
Vol. 3 Materials. Sec. I General Properties. AECD-3647,
United States Atomic Energy Comm., Oak Ridge Nat. Lab., Mar. 1955.

25. Promisel, N. E.: Recent Development in High-Temperature Alloys.
Machine Design, vol. 29, no. 15, July 25, 1957.











NACA RM E57L19


_____ ^-ity


6
0 d
.r?


UI



I a
I
~141




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d a-








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-










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Temperature, F

I s) RcBt content ojn ra *esrilty tami.

Figure 3. Total great content of reverail Lterials Ior a rarge of temperatur.m. BSct content
assumed zero at storage temperature; pressure, i.0 atmciphcr' unliei norctd CiLterLusL.


NACA RM E57L19


- LCI0









28 NACA RM E57U9






o




a a





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4j0 0I






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NACA RM E57L19


Evacuated Styrofoam
insulation;
density = 2.5 lb/cu ft


2 to 3 in.


Stainless-steel
outer shell
0.010 to 0.020
in. thick -


Figure 4. Possible configuration for an insulated,
vacuum-jacketed storage tank for liquid helium or
hydrogen.









NACA RM E57L19


Example of structural
element being considered


x, ft

O s5
CZ3
E3 1

*Heat flux
adjusted
for lower-
surface temp.


f 1600
-Inner
wall


1200 -

00
0
S 00c
o
O (U
S0 -




400 -
-4
cx


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,4


600
-c
00


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0-400




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-


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-9-


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16C


o
P.
42


CM
4-4

0 40
o
0


7


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Li


(a) Heat flux, 150 Btu/(sq ft)(sec).

Figure 5. Comparison of several coolants.


40-


n -


.J'JI


PrH
0

@2 ,0
Z1 E

00
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Cd
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'l-lvl
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0
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43


1-










NACA RM E57L19


Example of structural
element being considered


x, ft


3



Heat flux
adjusted for
lower-surface
temp.


Inner /
Sall- -


U



u C
cc



uci

a -0
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600

400


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ar








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u -


(b) Beat flux, 10 Btu/(sq ft)(sec).

Figure 5. Concluded. Comparison of several coolants.








NACA RM E57L19


80X103


Safe


Fuselage rear-end bleed


Wing trailing-
edge bleed





Unsafe








1 2 3 4 5 (
Flight Mach number

Figure 6. Minimum altitude for purging cooling system for
hydrogen detonation pressure less than 2 atmospheres.








NACA RM E57L19


Figure 7. Example of counterflow cooling of a corrugated
structure.







NACA RM E57L19


Area for radiant
heat sink


(a) Top view of aircraft.


High-level
coolant -


+ + t 4


-Heat rejected
to atmosphere
+ 4 4 by radiation


Heat rejected ..*:*.:
to low-level .....
coolant


Reinforcement ..... :
channel


a:


. ... .. ....................... .. .. .

Thermoflex
insulation
Low-level
.....:.:: .::: ......... .................... ... Low -lev e l
.-:-:-:-:::.: .. :::: : ::.:..:::. .:::::::--:- -:-: coolant


(b) Cross section of structure in radiant heat-sink area.

Figure 8. Example of using the upper surface of hypersonic
aircraft as a radiant heat sink for high-level coolant.


_______


JyyyyyyyV







NACA RM E57L19


120




*.S 80



a,
^^ 40





0


Distance along wing surface from stagnation line, in.
(distance measured normal to leading edge)

Figure 9. Variation of temperature drop and heat flux through
zirconia coatings applied to leading-edge region of wing of a
hypersonic aircraft. Flight speed, 18,000 feet per second;
altitude, 180,000 feet; wing loading, 20 pounds per square
foot; angle of attack, 50; sweep angle, 780; leading-edge
radius, 0.25 inch.


Coating, in. Emissivity

None 0.9
------0.025 .2
------ .050 .2 ---





-_J= ---*- -- -- -- -- -- -- --







NACA RM E57L19


N\ A\7AKAk7\&'\tHeat sink 0.5 In. max.


neycombs or corrugations

(a) Convection-cooled (b) Heat-sink-cooled
structure. structure.

Figure 10. Possible low-level cooling methods.











NACA RM E57L19


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