To manage propellant in a spacecraft, the method of this disclosure includes storing propellant in a tank as a mixture of liquid and gas; transferring the propellant out of the tank; converting the mixture of liquid and gas propellant into a single phase, where the single phase is either liquid or gaseous; and supplying the single phase of the propellant to a thruster.

Retrofittable satellite systems for an in-orbit host satellite comprising an enhancement module for adding a capability to the in-orbit host satellite, modifying the function of the in-orbit host satellite, and/or extending the function of the in-orbit host satellite. The in-orbit, retrofittable satellite system comprises a transfer vehicle for transferring the enhancement module from a first to a second location and a service vehicle for receiving the enhancement module from the transfer vehicle and installing the enhancement module on the in-orbit host satellite. In-orbit space situational awareness systems, comprising one or more in-orbit host satellites having one or more enhancement modules attached thereto, the enhancement modules comprising sensors such as satellite spatial location/position sensors, range sensors, navigation sensors, and/or proximity sensors for detecting other objects in-orbit, their location, speed, acceleration, orbital trajectory or the like, wherein the enhancement modules communicate to create a mesh network between the satellites.

An electrical power system has a dual battery configuration that enables sufficient power supply for a spacecraft bus and a payload module being carried by the spacecraft. During a sunlight power mode, power is drawn from a solar array of the bus to power a low-discharge payload of the spacecraft and a high-discharge payload of a payload module. During the sunlight power mode, a low rate discharge battery and a high rate discharge battery are charged by a battery charge management unit of the spacecraft bus. During an eclipse power mode, the low rate discharge battery powers the low-discharge payload of the spacecraft and the high rate discharge battery powers the high-discharge payload of the payload module. The high-rate discharge battery may also be used to power the high-rate discharge payload in the sunlight power mode to meet its high current demands to meet a flexible mission operations.

A radiator system uses an innovative passive control scheme in combination with dependable mechanical design features to meet or exceed the requirements for orbital applications. The disclosed radiator system is unique because we target an extremely high turndown ratio of 200:1 with an entirely passive two-phase pumped loop using ammonia as the working fluid. Sections of the radiator will selectively freeze to assist the turndown, and the mechanical design of the radiator can handle the high pressures experienced during such freezing and thawing events.

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.

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.

A satellite is disclosed, including a body and a communication device attached to the body. The body has an additively manufactured external wall structure at least partially forming an enclosed compartment, and the communication device is configured to receive and transmit data while in space.

A self-contained payload module and method of deployment of a payload includes a housing that is configured to engage a propulsive payload adapter hub, at least one deployable payload that is arranged in an interior cavity of the housing and deployable using the propulsive payload adaptor hub, a payload electronics system arranged in the interior cavity, a thermal control system in communication with the at least one payload and arranged in the housing, and at least one antenna that is arranged on the housing and configured for wideband communication.

An apparatus comprising a satellite includes a first radiator panel on a first side of a central body of the satellite and a second radiator panel on a second side of the central body is provided. The first radiator panel is thermally attached to the second radiator panel. A third side of the central body is positioned between the first side and the second side. A fourth side of the central body is positioned between the first side and the second side, opposing the third side. The first radiator panel is thermally attached to the second radiator panel by one or more coupling heat pipes, the coupling heat pipes exposed to an exterior environment of the satellite.

A spacecraft includes a semi-rigid solar array and a main body structure, the main body structure configured as a convex polyhedron and including an aft and a forward face disposed opposite to the aft face and at least four side faces disposed between and approximately orthogonal to the aft face and the forward face. The solar array includes a number of panels linked together with flexible couplings. In an undeployed configuration, panels of the solar array cover at least two adjacent side faces, the flexible couplings providing an articulable joint approximately aligned with a line along which the two adjacent side faces are joined and connecting a first panel of the solar array and a second panel of the solar array, the first panel being proximal to a first side face and the second panel being proximal to a second side face.