A triple-point cathode coating and method wherein electrically conductive NEA diamond particles cast or mixed with the adhesive medium and electrically insulative NEA diamond particles are cast or mixed with the adhesive medium to form a plurality of exposed junctions between electrically conductive diamond particles and electrically insulative diamond particles to reduce any electrical charges on a structure coated with the coating.

Systems for insulating a space vehicle or space borne container from an external environment as well as cryogens from heat sources. Such systems also protect the vehicle from the high dynamic pressures, the high heat loads encountered in atmospheric flight, and provide storage capability that strongly limits, or effectively eliminates, cryogenic boil-off losses once in space. Such systems include an expandable structure having a plurality of contiguously adjacent expandable layers. The layers are connected by a plurality of tension connectors between successive layers. For launch and flight the layers can be restrained in a collapsed position. Whereupon exiting a free stream environment, the layers are expanded where they can lock into place or otherwise remain in an expanded state. The expansion creates separation between the layers with minimal conduction paths providing near theoretically perfect multi-layer insulation and extremely effective debris protection.

A honeycomb thermal insulation structure may comprise a first facesheet, a second facesheet, and a honeycomb core between the first facesheet and the second facesheet. The honeycomb core may include a plurality of honeycomb unit cells each composed of walls having a height and spaced by a distance. The walls of the honeycomb cells may have perforations. The honeycomb thermal insulation structure may further comprise a non-convective gas loaded in the honeycomb unit cells between the walls. A flow of the gas through the perforations may be substantially absent.

A method for computing self-contamination processes of a spacecraft involves receiving a first set of input parameters comprising general definitions of the spacecraft, and a second set of input parameters comprising control parameters for the spacecraft orbital data, physics, numeric. A self-contamination process of the spacecraft is calculated based on the received first and second sets of input data by calculating an analytical solution of a basic equation for calculating a deposit of molecules outgassed from surfaces of the spacecraft in a single computation step with the data processing device. The calculated deposit of molecules is then output.

A method of releasing spacecraft from a rocket includes arranging spacecraft in a stack on a rocket, launching the rocket until it attains orbital velocity, imposing a flat spin on the rocket, and releasing the entire stack from the rocket during the flat spin. In one aspect, the method includes applying a compressive load along a length of the stack, launching the rocket until it attains orbital velocity, imposing a flat spin on the rocket, and releasing the compressive load to allow the entire stack to separate from the rocket during the flat spin.

Coatings and materials that are atomic oxygen resistant and have an atomically smooth surface that can reduce drag are disclosed. The coatings and materials can be used on at least a portion of a spacecraft intended to operate in harsh environments, such as stable Earth orbits at about 100 km to about 350 km.

A method providing conductive coatings is provided. A dopant layer with a plasma deposited conjugated polymer is provided. Conductive, conjugated polymer coatings are also provided.

The present invention is directed to a phosphated epoxy resin comprising at least one terminal group comprising a phosphorous atom covalently bonded to the resin by a carbon-phosphorous bond or by a phosphoester linkage; and at least one carbamate functional group. The present invention is also directed towards aqueous resinous dispersions comprising the phosphated epoxy resin, methods of coating a substrate, coated substrates, and methods of making a phosphated epoxy resin.

A high temperature thermal protection systems for rockets, and associated methods, is disclosed. A representative system includes a launch vehicle having a first end and a second end generally opposite the first end. The launch vehicle is elongated along a vehicle axis extending between the first and second ends and carries a propulsion system having at least one nozzle positioned at the second end of the launch vehicle. A thermal protection apparatus positioned around the nozzle is used to provide cooling and/or insulation to the nozzle during the flight of the launch vehicle. The thermal protection apparatus can include multiple fabric layers and an insulation layer stacked and stitched together. The fabric layers can include metal alloy fibers. In representative systems, the thermal protection apparatus can further include provisions for water that saturates the insulation layer to provide further insulating and/or cooling effects.

Systems and methods are provided for protecting a temperature sensitive object. A system includes a temperature sensitive object and a thermal control material in thermal communication with the temperature sensitive object. The thermal control material has an emissivity that varies as a function of temperature, and includes a substrate comprising a first surface comprising one of a photonic crystal, a metamaterial, a metasurface, and a multilayer film, a solid state phase change material in contact with the surface, and a reflective thin film material at one of a second surface of the substrate, at a surface of the solid state phase change material, and on an opposite side of an optical cavity from the substrate.