Materials for self-transpiring hot skins for hypersonic vehicles or reusable space vehicles

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 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. 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 of FIG. 1 during a short period of high heat flux.

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 .

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.

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 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 .

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 .

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 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.

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 .

One preferred material system 34 that satisfies these requirements based on thermodynamic calculations is zinc nitride (Zn 3 N 2 ), with the following reaction (1) (in an inert environment):

Zn 3 N 2 →3Zn (g) +N 2(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 (N 2(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 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

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.

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, ΔH: 500 k/mole)  (2)

Ge+GeO 2 →2GeO (g) (850-1000 degrees Celsius, ΔH: 400 kJ/mole)  (3)

2GeO 2 →2GeO (g) +O 2(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 of gas 40 according to reaction mechanisms (1)-(4) described above, over a customized temperature range from about 600 to 1600 degrees Celsius.

The similar systems 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. In addition, 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.

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.

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.