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 lightning protection device for placing on a structure for protecting, and including a first coating including at least one layer of conductive paint; and a second coating deposited on the surface coating and including a material that is thermally insulating and electrically conductive.

The present invention is a method for seaming a polymeric material. The method involves forming an interface between a plurality of separate polymeric sheets of materials. Next, heat and pressure are applied to facilitate the diffusion of the polymer molecules at the interface. A diffusion weld is made when the polymer molecules diffuse across the interface.

A self-supporting space environment protection device and method of manufacturing the device are provided. The device includes a self-supporting polymer-based body manufactured for installation on a component of a space-based system and composed of a space radiation resistant polymer-based material. The device further includes a low-absorptivity finish applied to an external surface of the self-supporting polymer-based body, the low-absorptivity finish having a solar absorptivity () value lower than the space radiation resistant polymer-based material, and a low-emissivity finish applied to the internal surface of the self-supporting polymer-based body, the low-emissivity finish having an infrared emissivity () value lower than the space radiation resistant polymer-based material.

Disclosed is a micrometeoroid and debris protection system (MDPS) for a thermal control system on a spacecraft. The MDPS comprises a radiator face-sheet, a truss attached to the radiator face-sheet, and a thermally transparent bumper disposed on the truss. The thermally transparent bumper shields the radiator face-sheet from micrometeoroids and debris and enables thermal transfer from the radiator face-sheet through the thermally transparent bumper.

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 spacecraft thermal control system is disclosed that includes a software reconfigurable radiator assembly capable of dynamically switching between high and low absorptance states to regulate thermal load. The system comprises a reflective display layer, such as an electrophoretic material, integrated into a multilayer radiator structure. A controller modulates the thermos-optical properties of the radiator based on pre-programmed or real-time inputs. The system allows for pre-integration into a satellite platform prior to receipt of mission-specific parameters and supports in-orbit thermal reconfiguration. The system can adapt to match changing orbits and mission environments, enabling flexible and responsive thermal management across different phases of operation. Additional features may include heat storage elements and thermal switches to manage payload-specific thermal requirements.

The invention relates to a satellite comprising: a satellite body including at least one radiative surface configured to emit satellite heat to space; and an attitude determination and control ADCS system for controlling an orientation of the satellite travelling in orbit around Earth, wherein the ADCS system is configured to change the orientation of the satellite with respect to the sun between a first position in which the radiative surface points away from the sun and a second position in which the radiative surface is exposed to the sun. The invention further relates to a method of controlling thermal radiation heat transfer of a satellite.

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.