Multifunctional Ablative Thermal Protection System

The invention includes an exposed surface cap with a specially formulated coating, an insulator base adjacent to the cap with another specially formulated coating, and one or more pins that extend from the cap through the insulator base to tie the cap and base together through ceramic bonding and mechanical attachment. The cap and insulator base have corresponding depressions and projections that mate and allow for differences in thermal expansion of the cap and base. The cap includes a high-temperature, low density, carbonaceous, fibrous material whose surface is optionally treated with a High Efficiency Tantalum-based Ceramic Composite (HETC) formulation, the fibrous material being drawn from the group consisting of silicon carbide foam and similar porous, high temperature materials. The insulator base and pin(s) contain similar material. The mechanical design is arranged so that thermal expansion differences in the component materials (e.g., cap and insulator base) are easily tolerated. It is applicable to both sharp and blunt leading edge vehicles. This extends the possible application of fibrous insulation to the wing leading edge and/or nose cap on a hypersonic vehicle. The lightweight system comprises a treated carbonaceous cap composed of Refractory Oxidation-resistant Ceramic Carbon Insulation (ROCCI), which provides dimensional stability to the outer mold line, while the fibrous base material provides maximum thermal insulation for the vehicle structure. The composite has graded surface treatments applied by impregnation to both the cap and base. These treatments enable it to survive in an aero-convectively heated environment of high-speed planetary entry. The exact cap and base materials are chosen in combination with modified surface treatments and a specially formulated surface coating, taking into account the duration of exposure and expected surface temperatures for the particular application.

Going farther, faster and hotter in space means innovating how NASA constructs the materials used for heat shields. For HEEET, this results in the use of dual-layer, three-dimensional, woven materials capable of reducing entry loads and lowering the mass of heat shields by up to 40%. The outer layer, exposed to a harsh environment during atmospheric entry, consists of a fine, dense weave using carbon yarns. The inner layer is a low-density, thermally insulating weave consisting of a special yarn that blends together carbon and flame-resistant phenolic materials. Heat shield designers can adjust the thickness of the inner layer to keep temperatures low enough to protect against the extreme heat of entering an atmosphere, allowing the heat shield to be bonded onto the structure of the spacecraft itself. The outer and inner layers are woven together in three dimensions, mechanically interlocking them so they cannot come apart. To create this material, manufacturers employ a 3-D weaving process that is similar to that used to weave a 2-D cloth or a rug. For HEEET, computer-controlled looms precisely place the yarns to make this kind of complex three-dimensional weave possible. The materials are woven into flat panels that are formed to fit the shape of the capsule forebody. Then the panels are infused with a low-density version of phenolic material that holds the yarns together and fills the space between them in the weave, resulting in a sturdy final structure. As the size of each finished piece of HEEET material is limited by the size of the loom used to weave the material, the HEEET heat shield is made out of a series of tiles. At the points where each tile connects, the gaps are filled through inventive designs to bond the tiles together.

The Multilayered Fire Protection system uses technology from the space craft flexible heat shield for future planetary missions. By optimizing this material for the fire environment, utilizing heat shield test methods, and experimenting with different materials, the NASA team developed a multilayered fire protection system. This system includes an outer textile layer which reflects over 90 percent of the radiant heat, an insulated layer which protects against convective heat and hot gases, and a non-porous film layer which is a gas barrier layer.

The invention includes an exposed surface cap having a specially formulated coating, an insulator base adjacent to the cap with another specially formulated coating, and one or more pins that extend from the cap through the insulator base to tie the cap and base together through ceramic bonding and mechanical attachment. The cap and insulator base have corresponding depressions and projections that mate and allow for differences in thermal expansion of the cap and base. The cap includes a high temperature, low density, carbonaceous, fibrous material whose surface is optionally treated with a HETC formulation, the fibrous material being drawn from the group consisting of silicon carbide foam and similar porous, high temperature materials. The insulator base and pin(s) contain similar material. The mechanical design is arranged so that thermal expansion differences in the component materials (e.g., cap and insulator base) are easily tolerated, and is applicable to both sharp and blunt leading edge vehicles. This extends the possible application of fibrous insulation to the wing leading edge and/or nose cap on a hypersonic vehicle. The lightweight system comprises a treated carbonaceous cap composed of Refractory Oxidation-resistant Ceramic Carbon Insulation (ROCCI), which provides dimensional stability to the outer mold line, while the fibrous base material provides maximum thermal insulation for the vehicle structure. The composite has graded surface treatments applied by impregnation to both the cap and base. These treatments enable it to survive in an aero-convectively heated environment of high-speed planetary entry. The exact cap and base materials are chosen in combination with the surface treatments, taking into account the duration of exposure and expected surface temperatures for the particular application.

There has been much interest in developing polymeric nanocomposites with ultrahigh thermal conductivities, such as with exfoliated graphite or with carbon nanotubes. These materials exhibit thermal conductivity of 3,000 W/mK measured experimentally and up to 6,600 W/mK predicted from theoretical calculations. However, when added to polymers, the expected thermal conductivity enhancement is not realized due to poor interfacial thermal transfer. This technology is a method of forming carbon-based fillers to be incorporated into highly thermal conductive nanocomposite materials. Formation methods include treatment of an expanded graphite with an alcohol/water mixture followed by further exfoliation of the graphite to form extremely thin carbon nanosheets that are on the order of between about 2 and about 10 nanometers in thickness. The carbon nanosheets can be functionalized and incorporated as fillers in polymer nanocomposites with extremely high thermal conductivities.