U.S. patent application number 11/380450 was filed with the patent office on 2007-11-01 for materials for self-transpiring hot skins for hypersonic vehicles or reusable space vehicles.
Invention is credited to Brian Nelson Cox, Janet B. Davis, Julia Mack, David Bruce Marshall, Peter E. Morgan, Olivier H. Sudre.
Application Number | 20070252045 11/380450 |
Document ID | / |
Family ID | 38577708 |
Filed Date | 2007-11-01 |
United States Patent
Application |
20070252045 |
Kind Code |
A1 |
Cox; Brian Nelson ; et
al. |
November 1, 2007 |
MATERIALS FOR SELF-TRANSPIRING HOT SKINS FOR HYPERSONIC VEHICLES OR
REUSABLE SPACE VEHICLES
Abstract
A self-transpiring hot skin for a hypersonic or reusable space
vehicle that can provide protection to the vehicle during short
periods of abnormally high heat flux (either planned in the flight
profile or an off-nominal event). The hot skin includes a ceramic
composite structure having an internal cavity that is coupled
either to the insulating layer or directly to the support structure
of the hypersonic vehicle. The internal cavity includes a material
system that vaporizes, sublimes or decomposes into a gas when the
temperature exceeds the upper temperature capability of the
composite material. The gas transpires through the outer layer of
the composite material to provide cooling to the outer layer below
the upper temperature capability. Cooling may occur both by
conduction of heat from the composite material to the transpiring
gas and by the interaction of the transpiring gas with the boundary
layer of hypersonic flow over the outer surface, leading to a
reduction of the heat flux entering the surface.
Inventors: |
Cox; Brian Nelson; (Thousand
Oaks, CA) ; Davis; Janet B.; (Thousand Oaks, CA)
; Mack; Julia; (Encino, CA) ; Marshall; David
Bruce; (Thousand Oaks, CA) ; Morgan; Peter E.;
(Thousand Oaks, CA) ; Sudre; Olivier H.; (Thousand
Oaks, CA) |
Correspondence
Address: |
OSTRAGER CHONG FLAHERTY & BROITMAN, P.C.
570 LEXINGTON AVENUE
FLOOR 17
NEW YORK
NY
10022-6894
US
|
Family ID: |
38577708 |
Appl. No.: |
11/380450 |
Filed: |
April 27, 2006 |
Current U.S.
Class: |
244/171.7 |
Current CPC
Class: |
B64G 1/14 20130101; B64G
1/58 20130101 |
Class at
Publication: |
244/171.7 |
International
Class: |
B64G 1/52 20060101
B64G001/52 |
Claims
1. (canceled)
2. The thermal protection system of claim 3, wherein said hot skin
outer layer is mechanically attached to said outer support
structure.
3. A thermal protection system for a hypersonic or reusable space
vehicle having an outer support structure, the thermal protection
system comprising: a hot skin outer layer coupled to the outer
support structure, said hot skin layer comprising a front face and
a back face coupled together by at least one connecting portion,
said back face closely coupled to said outer support structure and
located between said outer support structure and said front face;
an inner cavity defined by said front face, said back face and said
at least one connecting portion; and an ablative material system
contained within said inner cavity, wherein the hot skin outer
layer is a continuous porous structure wherein said front face,
said back face and said at least one connecting portion are
indistinguishable.
4. The thermal protection system of claim 3, further comprising an
insulating material layer coupled between said hot skin outer layer
and the outer support structure.
5. The thermal protection system of claim 4, wherein said
insulating material is coupled to said back face of said hot skin
outer layer with a high temperature ceramic adhesive.
6. The thermal protection system of claim 4, wherein said
insulating material is selected from the group consisting of a
thermal blanket and a plurality of ceramic tiles.
7. The thermal protection system of claim 3, wherein said hot skin
outer layer comprises a ceramic matrix composite material.
8. The thermal protection system of claim 7, wherein said ceramic
matrix composite material is selected from the group consisting of:
a carbon fiber-reinforced silicon carbide matrix composite, a
carbon-carbon matrix composite, a silicon carbide reinforced
silicon carbide matrix composite, and oxide-oxide composites.
9. A thermal protection system for a hypersonic or reusable space
vehicle having an outer support structure, the thermal protection
system comprising: a hot skin outer layer coupled to the outer
support structure, said hot skin layer comprising a front face and
a back face coupled together by at least one connecting portion,
said back face closely coupled to said outer support structure and
located between said outer support structure and said front face;
an inner cavity defined by said front face, said back face and said
at least one connecting portion; and an ablative material system
contained within said inner cavity, wherein said ablative material
system comprises a member selected from the group consisting of:
solid zinc nitride, a mixture of germanium nitride and germanium
oxide, and a mixture of germanium nitride, germanium oxide and zinc
nitride.
10. (canceled)
11. (canceled)
12. A thermal protection system for a hypersonic or reusable space
vehicle having an outer support structure, the thermal protection
system comprising: a hot skin outer layer coupled to the outer
support structure, said hot skin layer comprising a front face and
a back face coupled together by at least one connecting portion,
said back face closely coupled to said outer support structure and
located between said outer support structure and said front face;
an inner cavity defined by said front face, said back face and said
at least one connecting portion; and an ablative material system
contained within said inner cavity, wherein said ablative material
system is selected from the group consisting of: Si.sub.3N.sub.4,
Si.sub.2N.sub.2O, Si.sub.3N.sub.4 and SiO.sub.2, mixed crystals of
the type ZnGeN.sub.2 and ZnSiN.sub.2, and mixtures of two
components, wherein a first of the two components decomposes at a
first temperature and the second of the two components decomposes
at a second temperature higher than the first temperature.
13. (canceled)
14. The thermal protection system of claim 12, wherein said hot
skin outer layer comprises a ceramic matrix composite material.
15. The thermal protection system of claim 14, wherein said ceramic
composite material comprises a carbon fiber-reinforced silicon
carbide matrix composite.
16. An ablative thermal protection system for a hypersonic or
reusable space vehicle, the ablative thermal protection system
comprising: a hot skin outer layer comprising a front face and a
back face coupled together by at least one connecting portion; an
inner cavity defined by said front face, said back face and said at
least one connecting portion; and an ablative material system
contained within said inner cavity, wherein said ablative material
system comprises a member selected from the group consisting of
solid zinc nitride, a mixture of germanium nitride and germanium
oxide and a mixture of germanium nitride, germanium oxide and zinc
nitride.
17. (canceled)
18. (canceled)
19. (canceled)
20. A method for forming an integrated insulating and ablative
thermal protection system for a hypersonic and space reusable
vehicle having a support structure, the method comprising: forming
a hot skin outer layer comprising a front face and a back face
coupled together by at least one connecting portion; coupling said
hot skin outer layer to the support structure such that said back
face is located between the support structure and said front face;
and introducing an ablative material system within an inner cavity
of said hot skin outer layer, said inner cavity defined by said
front face, said back face and said at least one connecting
portion, wherein said ablative material system ablates to generate
a gas that transpires through said front face to cool said front
face when a temperature of said front face exceeds an upper
temperature capability of said front face, wherein introducing an
ablative material system within an inner cavity of said hot skin
outer layer comprises introducing a quantity of a substance
selected from the group consisting of solid zinc nitride, a mixture
of germanium nitride and germanium oxide, and a mixture of
germanium nitride, germanium oxide and zinc nitride within an inner
cavity of said hot skin outer layer, said inner cavity defined by
said front face, said back face and said at least one connecting
portion.
21. (canceled)
22. (canceled)
23. The method of claim 20, wherein introducing an ablative
material system within an inner cavity of said hot skin outer layer
comprises: determining an upper temperature capability of said hot
skin outer layer; selecting a solid ablative material system which
vaporizes, sublimes or decomposes via an endothermic reaction to
form a gas at a temperature less than said upper temperature
capability of said hot skin outer layer, said gas capable of
cooling said hot outer skin to a temperature less than said upper
temperature capability of said hot outer skin layer; and
introducing said ablative material system with an inner cavity of
said hot skin outer layer, said inner cavity defined by said front
face, said back face and said at least one connecting portion.
24. The method of claim 20 further comprising coupling an
insulating material between said outer skin layer and the support
structure.
25. The method of claim 24, wherein coupling said insulating
material comprises: providing an insulating material selected from
the group consisting of an insulating blanket and a plurality of
ceramic tiles; applying a preceramic polymer high temperature
adhesive between said insulating material and said back face; and
heating said preceramic polymer high temperature adhesive to form a
ceramic material, said ceramic material coupling said insulating
material to said back face.
26. The method of claim 20, wherein coupling said hot skin outer
layer to the support structure such that said back face is located
between the support structure and said front face comprises
mechanically fastening a back face of said hot skin outer layer to
the support structure such that said back face is located between
the support structure and said front face.
Description
TECHNICAL FIELD
[0001] The present invention generally relates to thermal
protection systems for hypersonic vehicles or reusable space
vehicles and more specifically to materials for self-transpiring
hot skins for hypersonic vehicles or reusable space vehicles.
BACKGROUND ART
[0002] At present, efforts are being undertaken to develop
hypersonic or reusable space vehicles capable of reaching speeds as
high as Mach12. Examples of such vehicles include, for example,
missiles, hypersonic cruise vehicles, and spacecraft such as the
space shuttle.
[0003] Such hypersonic or reusable space vehicles are, of course,
subject to extreme temperature fluctuations within the vehicle's
envelope of performance. Specifically, the leading edges, flight
control surfaces and a substantial portion of the external surfaces
of such vehicle support structures, or frames, as well as the
internal construction associated with engines necessary to power
the vehicle require that thermal design parameters incorporate
means for ensuring structural survivability during short periods of
high heat flux. Thermal protection systems for hypersonic vehicles
essentially fall into two categories: insulative and ablative.
Insulative systems such as those used on the space shuttle have two
advantages: (i) they are generally lighter in weight than ablative
systems and (ii) they maintain a constant outer vehicle surface,
whereas with ablative systems, recession of the outer surface
occurs thus changing the aerodynamic shape of the vehicle. However,
existing insulative systems are limited in the maximum allowable
temperature (or heat flux) at the outer surface (mostly below
.about.1600 deg. C.), whereas ablative systems can be used to much
higher temperatures (and heat fluxes). There exists a need to
provide adequate thermal protection to hypersonic or reusable space
vehicles in the event of a high heat load event that combines the
most desirable attributes of both the insulative and ablative
thermal protection systems. Such a system ideally also realizes
other positive attributes such as cost and weight reduction.
SUMMARY OF THE INVENTION
[0004] The proposed invention combines the attributes of an
insulative and ablative thermal protection system into a single
integrated thermal protection system for a hypersonic or reusable
space vehicle with the capability of surviving long periods of
moderate heating with short periods of higher heating without
sustaining structural damage due to overheating.
[0005] The present invention provides an integrated
self-transpiring hot skin for a hypersonic or reusable space
vehicle that can provide protection to the vehicle during short
periods of high heat flux.
[0006] The hot skin includes a ceramic composite structure, or hot
skin outer layer, having an internal cavity or cavities that are
coupled to a support structure and coupled to an optional
insulating layer of the hypersonic or space reusable vehicle. The
internal cavities include an ablating material system (i.e. a
system that vaporizes or sublimes or decomposes into a gas) at a
temperature below the upper temperature capability of the composite
material. The gas transpires through the outer layer of the
composite material to provide cooling to the outer layer. Normally
it would be preferred that only direct solid-gas reaction be
allowed, with no melting or reaction melting
[0007] The material system contained within the internal cavity is
an effective solid chemical that undergoes an endothermic reaction
(or possibly even mildly exothermic) in the desired temperature
range to produce gases that can penetrate the porous ceramic
material as it is being generated. One material system that meets
these requirements is zinc nitride. Another material system that
meets these requirements is a mixture of germanium nitride and
germanium oxide. In addition, mixtures of these two systems are
also contemplated and may provide cooling over a customized
temperature range from about 600 to 1600 degrees Celsius. Several
other nitrides or oxynitrides are also contemplated.
[0008] Other features, benefits and advantages of the present
invention will become apparent from the following description of
the invention, when viewed in accordance with the attached drawings
and appended claims.
BRIEF DESCRIPTION OF THE DRAWINGS
[0009] FIG. 1 is section view of a portion of a hypersonic or
reusable space vehicle according to a preferred embodiment of the
present invention operating in normal temperature conditions;
and
[0010] FIG. 2 is a section view of the portion of the hypersonic or
reusable space vehicle of FIG. 1 during a short period of high heat
flux.
BEST MODES FOR CARRYING OUT THE INVENTION
[0011] Referring now to FIG. 1, a region 18 of a hypersonic or
reusable space vehicle 20 is depicted and includes a ceramic hot
skin outer layer 22 coupled to an optional insulating layer 24,
both of which are coupled over the outer support structure 38 of
the vehicle 20. The outer layer 22 and optional insulating layer 24
together function to provide thermal protection for the vehicle
support structure 38 and vehicle components during flight, although
alternatively the outer layer 22 may provide adequate thermal
protection in systems not requiring an insulating layer 24.
[0012] The hot skin outer layer 22 includes a back face 26 and a
front face 28 coupled together using a series of connecting
portions 30. The back face 26, front face 28, and connecting
portions 30 together define one or more cavity structures 32. Thus,
the hot skin outer layer itself qualifies as an insulative
protection layer. The back face 26, front face 28, and connecting
portions 30 of the hot skin may have a variety of geometric
arrangements, including continuous porous structures in which the
front face, back face and connecting portions are not clearly
distinguished.
[0013] The hot skin outer layer 22 is formed of a ceramic matrix
composite ("CMC") material that has high heat resistance and
sufficient durability for use as a thermal protection system in
hypersonic travel. One such CMC material is a carbon
fiber-reinforced silicon carbide matrix composite (or "C--SiC").
Other CMC materials may include a carbon-carbon matrix composite, a
silicon carbide reinforced silicon carbide matrix composite, and
oxide-oxide composites. The front face 28 has a controlled porosity
and has an upper temperature capability (T.sub.o) of up to 1600
degrees Celsius.
[0014] The optional insulating layer 24 is a low thermal
conductivity insulation material such as an insulating blanket or
ceramic tiles that are well known in the art for use to thermally
insulate (protect) reusable space vehicles such as the space
shuttle. The insulating layer 24 has lower temperature resistant
capabilities than the overlying hot outer skin layer 22 and so is
an optional layer that is utilized to optimize the thermal
protection aspect of the entire thermal protection system. The back
face 26 of the hot skin 22 is preferably coupled to the insulating
layer 24 using a high temperature adhesive 36 such as a preceramic
polymer that forms a composite with heat treatment. In alternative
preferred arrangements, the back face 26 could simply be coupled
directly to the underlying support structure 38 of the vehicle 20
by mechanical means and the insulating layer 24 simply inserted
between the underlying structure 38 and back face 26.
[0015] Coupled within each of the one or more cavity structures 32
is a solid material system 34 that provides ablative (i.e.
transpiration cooling) thermal protection to the outer layer 22
during a short period of high heat flux within the region 18.
[0016] As best shown in FIG. 2, the solid material system 34
undergoes a reaction that generates gas (shown as arrows 40) when
the temperature of the vehicle 20 nears the upper temperature
capability T.sub.o of the ceramic hot skin outer layer 22 in the
region 18. The generation of gas 40 occurs as the solid material
system 34 vaporizes, sublimes or decomposes (or, generally
"ablates") in the presence of heat--in this case during a short
period of high heat flux. The generated gas 40 flows through the
porous structure of the front face 28 of the ceramic skin 22 and
cools the front face 28 below the upper temperature capability
T.sub.o during these short periods of abnormally high heat flux.
This protects the integrity of the hot skin outer layer 22 and the
vehicle support structure 38. The range of useful vaporization
temperatures for systems utilizing a C--SiC ceramic hot skin outer
layer 22 is expected to be between about 1000 and 1500 degrees
Celsius.
[0017] The generation of gas 40 that occurs during this high heat
flux event is the result of a chemical reaction of the solid
material system 34. This reaction generates the gas 40 either
through vaporization, sublimation, decomposition (i.e. an ablating
reaction) or reaction with gas from the surrounding atmosphere
without substantial melting depending upon the composition of the
solid material system 34.
[0018] One preferred material system 34 that satisfies these
requirements based on thermodynamic calculations is zinc nitride
(Zn.sub.3N.sub.2), with the following reaction (1) (in an inert
environment): Zn.sub.3N.sub.2.fwdarw.3Zn.sub.(g)+N.sub.2(g)
(600-1000 degrees Celsius, .DELTA.H: 400 kJ/mole) (1)
[0019] Thermal gravimetric analysis (TGA) has confirmed that the
decomposition of zinc nitride into nitrogen gas (N.sub.2(g)) and
zinc vapor (3Zn.sub.(g)) begins at around 600 degree Celsius
leading to complete mass loss at around 1350 degrees Celsius.
[0020] The details of the decomposition, sublimation and
vaporization rates are dependent upon numerous factors, including
the temperature gradients, gas flow restriction within the front
face 28, and ambient environment. A thicker front face 28 likely
will have a larger temperature drop between front and back
surfaces, and hence will require a longer period of high heat flux
to initiate the vaporization reaction. Moreover, the porosity of
the front face 28 will affect the flow rate of the gas 40 through
the front face 28, with a more porous material allowing a larger
flow of gas 40 within the front face 28
[0021] Further, the actual response of the zinc nitride material
system 34 is also dependent upon the physical characteristics of
the zinc nitride material system. For example, the particle size
and powder confinement of the zinc nitride material system 34
within the cavity structure 32 may alter the temperature range of
the vaporization reaction. A more finely ground powder will react
(i.e. generate gas 40) more quickly than a coarser powder.
Similarly, a more confined (i.e. packed) powder will react more
slowly than a less confined powder material. Furthermore, the
nature of the powder packing will affect the conduction of heat
within the powder and thus the reaction rates.
[0022] Another preferred material system 34 that satisfies these
requirements based on thermodynamic calculations consists of a
mixture of germanium nitride (Ge.sub.3N.sub.4) and germanium oxide
(GeO.sub.2), with the following series of reactions (2), (3) and
(4) (in an inert environment):
Ge.sub.3N.sub.4.fwdarw.3Ge+2N.sub.2(g) (600-1000 degrees Celsius,
.DELTA.H: 500 kJ/mole) (2) Ge+GeO.sub.2.fwdarw.2GeO.sub.(g)
(850-1000 degrees Celsius, .DELTA.H: 400 kJ/mole) (3)
2GeO2.fwdarw.2GeO.sub.(g)+O.sub.2(g) (1400-1800 degrees Celsius,
.DELTA.H: 450 kJ/mole) (4)
[0023] Thermal gravimetric analysis (TGA) in an inert atmosphere
has shown that mixtures of germanium nitride and germanium oxide
result in complete decomposition of germanium nitride and reaction
of germanium oxide to yield significant mass loss and the
production of nitrogen, oxygen and GeO gases in an endothermic
event.
[0024] In yet another preferred embodiment of the present
invention, an ablating material system 34 may consist of mixtures
and/or solid solutions of germanium nitride, germanium oxide and
zinc nitride. This embodiment therein provides cooling, via the
generation of gas 40 according to reaction mechanisms (1)-(4)
described above, over a customized temperature range from about 600
to 1600 degrees Celsius.
[0025] The similar systems Si.sub.3N.sub.4, Si.sub.2N.sub.2O, and
Si.sub.3N.sub.4+SiO.sub.2 behave similarly to the germanium cases,
but at significantly higher temperatures. In addition, mixed
crystals of the type ZnGeN.sub.2 and ZnSiN.sub.2 are known and
could have some utility in covering large temperature ranges.
Mixtures (e.g., Zn.sub.3N.sub.2 and Si.sub.3N.sub.4) in which one
component (Zn.sub.3N.sub.2) decomposes at low temperatures and the
other (Si.sub.3N.sub.4) decomposes at higher temperature could also
be useful.
[0026] As the solid material system 34 is a non-regenerable
resource, it is capable of protection for only a limited duration
during a high heat flux event. However, the solid material system
34 may be replaced (possibly via introduction through a portal in
the hot skin 22 or porous facesheet) for subsequent flights.
[0027] The proposed invention combines the attributes of an
insulative and ablative thermal protection system into a single
integrated system for a hypersonic or reusable space vehicle with
the capability of surviving short periods of high heat flux (either
planned in the flight profile or an off-nominal event) without
sustaining structural damage due to overheating. The proposed
invention is expected to be cost effective, and can extend the
range of heat loads for insulative thermal protection systems.
Moreover, by properly selecting the ablative material systems for
the perceived temperature range of a high heat flux event, a
customized thermal protection system can be achieved for a desired
application. While not described in detail, it is specifically
contemplated that other ablative materials, including carbon
nitrides and melamines for example, may be used in conjunction, or
in place of, the preferred embodiments described above in similar
or materially different systems desiring thermal protection from
adverse high heat flux events.
[0028] While the invention has been described in terms of preferred
embodiments, it will be understood, of course, that the invention
is not limited thereto since modifications may be made by those
skilled in the art, particularly in light of the foregoing
teachings.
* * * * *