초록 ▼

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 c

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.

대표청구항 ▼

What is claimed is: 1. 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 coupl

What is claimed is: 1. 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. 2. The thermal protection system of claim 1, wherein said hot skin outer layer is mechanically attached to said outer support structure. 3. The thermal protection system of claim 1, further comprising an insulating material layer coupled between said hot skin outer layer and the outer support structure. 4. The thermal protection system of claim 3, wherein said insulating material is coupled to said back face of said hot skin outer layer with a high temperature ceramic adhesive. 5. The thermal protection system of claim 3, wherein said insulating material is selected from the group consisting of a thermal blanket and a plurality of ceramic tiles. 6. The thermal protection system of claim 1, wherein said hot skin outer layer comprises a ceramic matrix composite material. 7. The thermal protection system of claim 6, 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. 8. 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. 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 is selected from the group consisting of: Si3N4, Si2N2O, Si3N4 and SiO2, mixed crystals of the type ZnGeN2 and ZnSiN2, 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. 10. The thermal protection system of claim 9, wherein said hot skin outer layer comprises a ceramic matrix composite material. 11. The thermal protection system of claim 10, wherein said ceramic composite material comprises a carbon fiber-reinforced silicon carbide matrix composite. 12. 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. 13. 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. 14. The method of claim 13, 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. 15. The method of claim 13 further comprising coupling an insulating material between said outer skin layer and the support structure. 16. The method of claim 15, 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. 17. The method of claim 13, 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.