US20070252045A1 - Materials for self-transpiring hot skins for hypersonic vehicles or reusable space vehicles - Google Patents
Materials for self-transpiring hot skins for hypersonic vehicles or reusable space vehicles Download PDFInfo
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- US20070252045A1 US20070252045A1 US11/380,450 US38045006A US2007252045A1 US 20070252045 A1 US20070252045 A1 US 20070252045A1 US 38045006 A US38045006 A US 38045006A US 2007252045 A1 US2007252045 A1 US 2007252045A1
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- support structure
- front face
- outer layer
- back face
- hot skin
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- 239000000463 material Substances 0.000 title claims abstract description 35
- 239000002131 composite material Substances 0.000 claims abstract description 18
- 239000000919 ceramic Substances 0.000 claims abstract description 11
- 238000001816 cooling Methods 0.000 claims abstract description 7
- 238000006243 chemical reaction Methods 0.000 claims description 16
- 239000000203 mixture Substances 0.000 claims description 14
- AKJVMGQSGCSQBU-UHFFFAOYSA-N zinc azanidylidenezinc Chemical compound [Zn++].[N-]=[Zn].[N-]=[Zn] AKJVMGQSGCSQBU-UHFFFAOYSA-N 0.000 claims description 13
- BIXHRBFZLLFBFL-UHFFFAOYSA-N germanium nitride Chemical compound N#[Ge]N([Ge]#N)[Ge]#N BIXHRBFZLLFBFL-UHFFFAOYSA-N 0.000 claims description 12
- YBMRDBCBODYGJE-UHFFFAOYSA-N germanium oxide Inorganic materials O=[Ge]=O YBMRDBCBODYGJE-UHFFFAOYSA-N 0.000 claims description 12
- PVADDRMAFCOOPC-UHFFFAOYSA-N oxogermanium Chemical compound [Ge]=O PVADDRMAFCOOPC-UHFFFAOYSA-N 0.000 claims description 11
- 229910010271 silicon carbide Inorganic materials 0.000 claims description 9
- 239000011159 matrix material Substances 0.000 claims description 7
- HBMJWWWQQXIZIP-UHFFFAOYSA-N silicon carbide Chemical compound [Si+]#[C-] HBMJWWWQQXIZIP-UHFFFAOYSA-N 0.000 claims description 7
- 239000007787 solid Substances 0.000 claims description 5
- VYPSYNLAJGMNEJ-UHFFFAOYSA-N Silicium dioxide Chemical compound O=[Si]=O VYPSYNLAJGMNEJ-UHFFFAOYSA-N 0.000 claims description 4
- 239000000853 adhesive Substances 0.000 claims description 4
- 230000001070 adhesive effect Effects 0.000 claims description 4
- 229910052799 carbon Inorganic materials 0.000 claims description 4
- 239000011153 ceramic matrix composite Substances 0.000 claims description 4
- 238000010438 heat treatment Methods 0.000 claims description 4
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 claims description 3
- 229910010293 ceramic material Inorganic materials 0.000 claims description 3
- 229920000642 polymer Polymers 0.000 claims description 3
- 229910002790 Si2N2O Inorganic materials 0.000 claims description 2
- CREMABGTGYGIQB-UHFFFAOYSA-N carbon carbon Chemical compound C.C CREMABGTGYGIQB-UHFFFAOYSA-N 0.000 claims description 2
- 239000011203 carbon fibre reinforced carbon Substances 0.000 claims description 2
- 229910052681 coesite Inorganic materials 0.000 claims description 2
- 229910052906 cristobalite Inorganic materials 0.000 claims description 2
- 239000013078 crystal Substances 0.000 claims description 2
- 239000000377 silicon dioxide Substances 0.000 claims description 2
- 229910052682 stishovite Inorganic materials 0.000 claims description 2
- 239000000126 substance Substances 0.000 claims description 2
- 229910052905 tridymite Inorganic materials 0.000 claims description 2
- 239000011810 insulating material Substances 0.000 claims 8
- 238000000034 method Methods 0.000 claims 6
- 230000008878 coupling Effects 0.000 claims 5
- 238000010168 coupling process Methods 0.000 claims 5
- 238000005859 coupling reaction Methods 0.000 claims 5
- 229910052581 Si3N4 Inorganic materials 0.000 claims 2
- 230000004907 flux Effects 0.000 abstract description 16
- 230000003993 interaction Effects 0.000 abstract 1
- 239000007789 gas Substances 0.000 description 16
- 239000000843 powder Substances 0.000 description 7
- 239000011343 solid material Substances 0.000 description 7
- 230000008016 vaporization Effects 0.000 description 5
- 238000009834 vaporization Methods 0.000 description 5
- 238000000354 decomposition reaction Methods 0.000 description 4
- 238000002411 thermogravimetry Methods 0.000 description 4
- IJGRMHOSHXDMSA-UHFFFAOYSA-N Atomic nitrogen Chemical compound N#N IJGRMHOSHXDMSA-UHFFFAOYSA-N 0.000 description 3
- 230000008901 benefit Effects 0.000 description 3
- 238000002844 melting Methods 0.000 description 3
- 230000008018 melting Effects 0.000 description 3
- 230000001419 dependent effect Effects 0.000 description 2
- 238000013021 overheating Methods 0.000 description 2
- 238000000859 sublimation Methods 0.000 description 2
- 230000008022 sublimation Effects 0.000 description 2
- RZVHIXYEVGDQDX-UHFFFAOYSA-N 9,10-anthraquinone Chemical compound C1=CC=C2C(=O)C3=CC=CC=C3C(=O)C2=C1 RZVHIXYEVGDQDX-UHFFFAOYSA-N 0.000 description 1
- 229920000877 Melamine resin Polymers 0.000 description 1
- HCHKCACWOHOZIP-UHFFFAOYSA-N Zinc Chemical compound [Zn] HCHKCACWOHOZIP-UHFFFAOYSA-N 0.000 description 1
- 230000002411 adverse Effects 0.000 description 1
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 description 1
- -1 carbon nitrides Chemical class 0.000 description 1
- 238000010276 construction Methods 0.000 description 1
- 229910001873 dinitrogen Inorganic materials 0.000 description 1
- 230000010006 flight Effects 0.000 description 1
- 229910052732 germanium Inorganic materials 0.000 description 1
- GNPVGFCGXDBREM-UHFFFAOYSA-N germanium atom Chemical compound [Ge] GNPVGFCGXDBREM-UHFFFAOYSA-N 0.000 description 1
- 239000012774 insulation material Substances 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 230000007246 mechanism Effects 0.000 description 1
- 150000007974 melamines Chemical class 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 150000004767 nitrides Chemical class 0.000 description 1
- 229910052757 nitrogen Inorganic materials 0.000 description 1
- 239000001301 oxygen Substances 0.000 description 1
- 229910052760 oxygen Inorganic materials 0.000 description 1
- 238000012856 packing Methods 0.000 description 1
- 239000002245 particle Substances 0.000 description 1
- 239000011148 porous material Substances 0.000 description 1
- 230000004044 response Effects 0.000 description 1
- 239000006104 solid solution Substances 0.000 description 1
- 230000005068 transpiration Effects 0.000 description 1
- 239000013585 weight reducing agent Substances 0.000 description 1
- 229910052725 zinc Inorganic materials 0.000 description 1
- 239000011701 zinc Substances 0.000 description 1
Images
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/52—Protection, safety or emergency devices; Survival aids
- B64G1/58—Thermal protection, e.g. heat shields
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/14—Space shuttles
Definitions
- 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.
- hypersonic or reusable space vehicles capable of reaching speeds as high as Mach 12 .
- examples of such vehicles include, for example, missiles, hypersonic cruise vehicles, and spacecraft such as the space shuttle.
- 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.
- 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.
- 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.
- 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
- 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.
- 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.
- 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;
- 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.
- 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 .
- 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 .
- 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.
- 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.
- CMC ceramic matrix composite
- One such CMC material is a carbon fiber-reinforced silicon carbide matrix composite (or “C—SiC”).
- C—SiC carbon fiber-reinforced silicon carbide matrix composite
- 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 o ) of up to 1600 degrees Celsius.
- 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 .
- 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 .
- 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 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 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.
- 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 .
- Zn 3 N 2 zinc nitride
- reaction (1) in an inert environment: Zn 3 N 2 ⁇ 3Zn (g) +N 2(g) (600-1000 degrees Celsius, ⁇ H: 400 kJ/mole) (1)
- TGA Thermal gravimetric analysis
- 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.
- 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
- the actual response of the zinc nitride material system 34 is also dependent upon the physical characteristics of the zinc nitride material system.
- 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.
- the nature of the powder packing will affect the conduction of heat within the powder and thus the reaction rates.
- Another preferred material system 34 that satisfies these requirements based on thermodynamic calculations consists of a mixture of germanium nitride (Ge 3 N 4 ) and germanium oxide (GeO 2 ), with the following series of reactions (2), (3) and (4) (in an inert environment): Ge 3 N 4 ⁇ 3Ge+2N 2(g) (600-1000 degrees Celsius, ⁇ H: 500 kJ/mole) (2) Ge+GeO 2 ⁇ 2GeO (g) (850-1000 degrees Celsius, ⁇ H: 400 kJ/mole) (3) 2GeO2 ⁇ 2GeO (g) +O 2(g) (1400-1800 degrees Celsius, ⁇ H: 450 kJ/mole) (4)
- TGA Thermal gravimetric analysis
- 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.
- Si 3 N 4 , Si 2 N 2 O, and Si 3 N 4 +SiO 2 behave similarly to the germanium cases, but at significantly higher temperatures.
- mixed crystals of the type ZnGeN 2 and ZnSiN 2 are known and could have some utility in covering large temperature ranges. Mixtures (e.g., Zn 3 N 2 and Si 3 N 4 ) in which one component (Zn 3 N 2 ) decomposes at low temperatures and the other (Si 3 N 4 ) decomposes at higher temperature could also be useful.
- 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.
- 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.
- 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.
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Abstract
Description
- 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.
- 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.
- 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 ˜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.
- 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.
- 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.
- 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
- 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.
- 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.
-
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 -
FIG. 2 is a section view of the portion of the hypersonic or reusable space vehicle ofFIG. 1 during a short period of high heat flux. - Referring now to
FIG. 1 , aregion 18 of a hypersonic orreusable space vehicle 20 is depicted and includes a ceramic hot skinouter layer 22 coupled to anoptional insulating layer 24, both of which are coupled over theouter support structure 38 of thevehicle 20. Theouter layer 22 andoptional insulating layer 24 together function to provide thermal protection for thevehicle support structure 38 and vehicle components during flight, although alternatively theouter layer 22 may provide adequate thermal protection in systems not requiring aninsulating layer 24. - The hot skin
outer layer 22 includes aback face 26 and afront face 28 coupled together using a series of connectingportions 30. Theback face 26,front face 28, and connectingportions 30 together define one ormore cavity structures 32. Thus, the hot skin outer layer itself qualifies as an insulative protection layer. Theback face 26,front face 28, and connectingportions 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. - 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. Thefront face 28 has a controlled porosity and has an upper temperature capability (To) of up to 1600 degrees Celsius. - 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. Theinsulating layer 24 has lower temperature resistant capabilities than the overlying hotouter skin layer 22 and so is an optional layer that is utilized to optimize the thermal protection aspect of the entire thermal protection system. Theback face 26 of thehot skin 22 is preferably coupled to the insulatinglayer 24 using ahigh temperature adhesive 36 such as a preceramic polymer that forms a composite with heat treatment. In alternative preferred arrangements, theback face 26 could simply be coupled directly to theunderlying support structure 38 of thevehicle 20 by mechanical means and theinsulating layer 24 simply inserted between theunderlying structure 38 andback face 26. - Coupled within each of the one or
more cavity structures 32 is asolid material system 34 that provides ablative (i.e. transpiration cooling) thermal protection to theouter layer 22 during a short period of high heat flux within theregion 18. - As best shown in
FIG. 2 , thesolid material system 34 undergoes a reaction that generates gas (shown as arrows 40) when the temperature of thevehicle 20 nears the upper temperature capability To of the ceramic hot skinouter layer 22 in theregion 18. The generation ofgas 40 occurs as thesolid 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 generatedgas 40 flows through the porous structure of thefront face 28 of theceramic skin 22 and cools thefront face 28 below the upper temperature capability To during these short periods of abnormally high heat flux. This protects the integrity of the hot skinouter layer 22 and thevehicle support structure 38. The range of useful vaporization temperatures for systems utilizing a C—SiC ceramic hot skinouter layer 22 is expected to be between about 1000 and 1500 degrees Celsius. - The generation of
gas 40 that occurs during this high heat flux event is the result of a chemical reaction of thesolid material system 34. This reaction generates thegas 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 thesolid material system 34. - One
preferred material system 34 that satisfies these requirements based on thermodynamic calculations is zinc nitride (Zn3N2), with the following reaction (1) (in an inert environment):
Zn3N2→3Zn(g)+N2(g) (600-1000 degrees Celsius, ΔH: 400 kJ/mole) (1) - Thermal gravimetric analysis (TGA) has confirmed that the decomposition of zinc nitride into nitrogen gas (N2(g)) and zinc vapor (3Zn(g)) begins at around 600 degree Celsius leading to complete mass loss at around 1350 degrees Celsius.
- 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 thickerfront 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 thefront face 28 will affect the flow rate of thegas 40 through thefront face 28, with a more porous material allowing a larger flow ofgas 40 within thefront face 28 - 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 zincnitride material system 34 within thecavity 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. - Another
preferred material system 34 that satisfies these requirements based on thermodynamic calculations consists of a mixture of germanium nitride (Ge3N4) and germanium oxide (GeO2), with the following series of reactions (2), (3) and (4) (in an inert environment):
Ge3N4→3Ge+2N2(g) (600-1000 degrees Celsius, ΔH: 500 kJ/mole) (2)
Ge+GeO2→2GeO(g) (850-1000 degrees Celsius, ΔH: 400 kJ/mole) (3)
2GeO2→2GeO(g)+O2(g) (1400-1800 degrees Celsius, ΔH: 450 kJ/mole) (4) - 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.
- 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 ofgas 40 according to reaction mechanisms (1)-(4) described above, over a customized temperature range from about 600 to 1600 degrees Celsius. - The similar systems Si3N4, Si2N2O, and Si3N4+SiO2 behave similarly to the germanium cases, but at significantly higher temperatures. In addition, mixed crystals of the type ZnGeN2 and ZnSiN2 are known and could have some utility in covering large temperature ranges. Mixtures (e.g., Zn3N2 and Si3N4) in which one component (Zn3N2) decomposes at low temperatures and the other (Si3N4) decomposes at higher temperature could also be useful.
- 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, thesolid material system 34 may be replaced (possibly via introduction through a portal in thehot skin 22 or porous facesheet) for subsequent flights. - 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.
- 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.
Claims (26)
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US11/380,450 US7281688B1 (en) | 2006-04-27 | 2006-04-27 | Materials for self-transpiring hot skins for hypersonic vehicles or reusable space vehicles |
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US11/380,450 US7281688B1 (en) | 2006-04-27 | 2006-04-27 | Materials for self-transpiring hot skins for hypersonic vehicles or reusable space vehicles |
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US7281688B1 US7281688B1 (en) | 2007-10-16 |
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