A structural spacecraft component comprising internal microstructure; wherein said microstructure comprises a plurality of parallel layers and a plurality of spacers that connect adjacent parallel layers; wherein said structural spacecraft component is a product of an additive manufacturing process.

A cooling assembly includes walls extending around and defining an enclosed vapor chamber that holds a working fluid. An interior porous wick structure is disposed inside the chamber and lines interior surfaces of the walls. The wick structure includes pores that hold a liquid phase of the working fluid. The cooling assembly also includes an exterior porous wick structure lining exterior surfaces of the walls outside of the vapor chamber. The exterior wick structure includes pores that hold a liquid phase of a cooling fluid outside the vapor chamber. The interior wick structure holds the liquid working fluid until heat from an external heat source vaporizes the working fluid inside the vapor chamber. The exterior wick structure holds the liquid fluid outside the vapor chamber until heat from inside the vapor chamber vaporizes the liquid cooling fluid in the exterior wick structure for transferring heat away from the heat source.

A solar rejection system includes an enclosure for housing a sensor, and a movable sunshade. The housing has an opening or aperture for admitting light to the sensor, and the sunshade is moved as needed to prevent harmful solar illumination of the sensor. The sunshade may be a flat panel. The sunshade panel is mounted to a hinge that is located on one side of a large diameter bearing that allows the shade to be rotated around the aperture of the sensor to always prevent the sun from illuminating the aperture. The hinge allows the shade to be tilted to either allow the sensor to see further off axis without obscuration or to block the sun when it moves in front of the sensor. Full closure of the sunshade on its hinge allows it to also function as an aperture door, blocking the opening or aperture.

A composite structure comprising an organic resin and carbon fibers comprises planar structure graphene nanosheets embedded in the resin. This structure combining good properties in terms of mechanical resilience, thermal conductivity and electrical conductivity can advantageously be used for thermal dissipation devices, as solar generator substrate or else as housing of electronic components, carried on board satellites.

Space craft comprising a body, at least one pair of supporting arms, a first device mounted on a first supporting arm and a second device mounted on a second supporting arm. The first arm is rotatably mounted on the body of the craft about an axis of rotation. The second arm is fixed to the body, and in which craft of the first device and the second device at least one is offset from the axis of rotation of the first arm. The pair of supporting arms further comprises a hollow module for the rotatable mounting of the first arm on the body. The mounting module comprising an opening through which the axis of rotation and the second supporting arm pass.

A method for attaching a heat-emitting device and a capillary heat pipe to a panel of a spacecraft wall is disclosed including the steps of: a) positioning a capillary heat pipe on a portion of the panel; attaching female attachment bodies to the panel, the female attachment bodies protruding relative to the capillary heat pipe; c) placing a thermally-conductive and self-curing paste over a portion of the capillary heat pipe or over a heat-emitting device; d) placing a heat-emitting device on the thermally-conductive and self-curing paste and on the female attachment bodies, said heat-emitting device bearing against and being in direct contact with the female attachment bodies, and e) attaching the heat-emitting device and said capillary heat pipe to the panel by attaching male attachment members to the female attachment bodies.

A spacecraft is disclosed having at least three flat side walls, at least one main communication antenna, including a radiating element having a central axis of radiation (AC-AC), a movable arm configured to move between a deployed position and a folded position, a reflector suitable for reflecting or receiving radiofrequency waves in a direction of emission (DE). The radiating element is fixed to a side wall so that the central axis of radiation (AC-AC) is arranged perpendicularly to the side wall, and the movable arm is shaped so that an offset angle (β) of between 25° and 65° is formed between the side wall and the direction of emission (DE), when the movable arm is in a deployed position.

An insert for a condensing heat exchanger, having: a body extending aft from a forward end to an aft end, and defining: a body exterior surface; a forward segment that extends aft from the forward end of the insert to a first axial location between the forward and aft ends of the insert, along the forward segment the body exterior surface is without openings; a middle segment that extends aft from the first axial location to a second axial location, along the middle segment the body exterior surface is cylindrical; and an aft segment that extends aft from the second axial location to the aft end of the insert, along the aft segment the body exterior surface of the body is cylindrical and defines axially extending grooves, and the grooves are spaced apart from each other and extend forward from the aft end of the insert to the middle segment.

A heat rejection system that employs temperature sensitive shape memory materials to control the heat rejection capacity of a vehicle to maintain a safe vehicle temperature. The technology provides for a wide range of heat rejection rates by actuation of the orientation or position of a heat rejection panel which impacts effective properties of the heat rejection system in response to temperature. When employed as a radiator for crewed spacecraft thermal control this permits the use of higher freezing point, non-toxic thermal working fluids in single-loop thermal control systems for crewed vehicles in space and other extraterrestrial environments.

An apparatus includes a thermally conductive interface assembly including a first component associated with a first interface surface and a second component associated with a second interface surface. The apparatus also includes a shape memory alloy component coupled to the thermally conductive interface assembly and configured to move one or more components of the thermally conductive interface assembly between a first state and a second state based on a temperature of the shape memory alloy component. In the first state, the first interface surface is in physical contact with the second interface surface, and in the second state, a gap is defined between the first interface surface and the second interface surface.