| NextTechs present research has led to the development of a novel thermal protection system (TPS) concept for spacecraft. Recent work has shown that lightweight carbon-carbon (C/C) composites produced by pressure-assisted co-polymerization of carbon fibers, hydrocarbons and fullerenes have excellent mechanical properties and high thermal stability that makes them ideal for thermal protection applications. Epoxy Matrices Currently used thermal protection systems utilize epoxy for the composite matrix. Epoxy matrices are known to be unstable and consequently can decompose during reentry. Another problem with currently used TP systems is the adhesives that are used to bond the composite to the metal panels of spacecraft. A case in point is the Space Shuttle where adhesives are used to bond refractory tiles over large areas of the exterior surface. These adhesives can degrade and lead to devastating consequences. Carbon-Carbon Composite 
Figure 1. Thermal protection system (TPS) concept for spacecraft Problems Solved Problems solved by the innovative design: · The carbon-carbon composite structure consists of a three dimensional (3-D) carbon fiber grid and a carbon matrix that is capable of protecting the structural parts of spacecraft from erosion by plasmas in the range of 2000-7000°C. · The C/C structure is anchored to the perforated metallic substrate (skin of the craft) by Z-directional carbon fibers (or strips). The metallic substrate can be cooled by a heat exchanging fluid. · For atmospheric application oxidation resistant layers can be applied to the outer surface of the carbon-carbon composite. These optional coatings offer thermal protection for long periods and have application for aircraft and re-entry vehicles. The protective layers prevent carbon from oxidizing when subjected to high temperature in the red-white incandescent heat range at atmospheric air pressure. Innovative Features The innovative features include: · The elimination of porosity is achieved by high pressure carbonization. Porosity causes mechanical weakness and high ablation rates by plasmas. · The “Z” oriented carbon fibers in the composite exposes the “hard” direction of the carbon structure that is more resistant to plasmas. · The C/C composite is mechanically and chemically bonded to the spacecraft skin. The bonding results by anchoring the ends of the carbon fibers into the perforations of the metal panel. · Additional protection is afforded by drawing heat away from the underside of the metal panel by circulating cooling fluids that keeps the metal layer below 200°C. The organic fiber/epoxy coating is bonded to the underside of the metal panel. This increases the mechanical strength of the TPS and diminishes the thermal flux to the interior of the spacecraft. · The optional oxidation resistant coating for aircraft and re-entry vehicles widens the area of application into the red-white incandescent heat range. |