The present invention relates to a single-board processor card configured for use in a 1U CubeSat payload form-factor multi-purpose architecture, including: a field-programmable-gate-array (FPGA) which is reconfigurable in flight; wherein a configuration memory of the FPGA can be scrubbed in flight to correct errors or upsets; and a radiation-hardened monitor (RHM) which provides radiation mitigation and system monitoring of the single-board processor card, and which reconfigures said FPGA during flight, scrubs the configuration memory, and monitors a health of the FPGA. The 1U CubeSat payload form-factor multi-purpose architecture includes a backplane having a plurality of slots, one of the plurality of slots which accommodates the single-board processor card, wherein the backplane routes signals to a plurality of standard-sized processor cards, interchangeably disposed in any of the plurality of slots.

The invention relates to a field coil (18, 20), in particular for a satellite hall-effect plasma thruster, said field coil (18, 20) comprising a core (22) on which a conductor (24) is wound, characterized in that the conductor comprises an inorganic insulation cable (26) impregnated with a high-temperature-resistant silicone coating (32).

A spacecraft sunshade is provided. The sunshade includes a surface that is maintained in a sun facing orientation. Adjustments to a position of the sunshade are made in a plane that is transverse to a line of sight to the sun, in order to block sunlight from being directly incident on an instrument associated with the spacecraft. The sunshade can include photovoltaic elements on the sun-facing surface of the sunshade. In addition, the sunshade can be formed from an opaque material, and further from a material that absorbs heat from the sun and reradiate that heat to the instrument. The sunshade can perform stray light blocking, electrical power generation, and radiational heating functions.

A system and method for protecting an electronics module on a spacecraft in space are described. The system includes a non-radiation hardened electronics module electrically connected to a power supply, with a switch connected between the power supply and the electronics module. The switch can disconnect the electronics module from the power supply in response to an event signal. A sensor which is capable of detecting a solar proton event is connected to the switch. The sensor emits the event signal upon detection of the solar proton event.

Disclosed is a circulation type space-based solar power system, the system including: one or more solar modules; a conveyor belt on which the solar modules are attached, whereby the solar modules move between a solar power generating position and a recovery position, the solar modules receiving sunlight to generate solar power in the solar power generating position, and not receiving sunlight in the recovery position; a driver moving the conveyor belt; and a protective plate blocking cosmic rays incident to the solar modules located in the recovery position. The system can generate solar power for a long time by moving the solar modules between the solar power generating position and the recovery position. While some of the solar modules generate solar power, the remaining solar modules having damage are recovered.

Software-based solutions may mitigate physical damage to multi-layer networks, such as neural networks having shortcut (residual) connections. An example includes: providing a multi-layer network comprising a plurality of nodes; for each of a plurality of training cases: determining a set of dropout nodes, based at least on a damage model having a probability of a node being selected for dropout that is based at least on a target operating environment of the multi-layer network, wherein the probability of a node being selected is spatially correlated; and training the multi-layer network with the determined set of dropout nodes disabled (with a different set of dropout nodes for different training cases). In some examples the damage model involves expected physical radiation damage to a computing device hosting the multi-layer network, such as on board an aircraft or an earth-orbiting satellite. Thus, multiple degrees of expected damage may be addressed.

Apparatus, systems, and methods for protecting a vehicle from a radiation source (e.g., the sun) are provided. One apparatus includes a set of first wires and a set of second wires located proximate to the set of first wires. The set of first wires maintains a positive voltage and the set of second wires maintains a negative voltage. The set of first wire and the set of second wires are arranged to generate an electrostatic field (ESF) between the vehicle and the radiation source. A system includes a spacecraft and a field generator that generates an ESF between the spacecraft and a radiation source. A method includes tracking a location of a spacecraft relative to a radiation source and generating an ESF between the spacecraft and the radiation 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 spacecraft radiation shield system (2) is disclosed for improving the protection from ionising radiation from the external environment and providing an improved freedom of orientation to the spacecraft. The spacecraft radiation shield system includes: at least two magnets arranged in a magnetic multipole (6), a magnetometer (14), and an adjustable magnet (10). The magnetometer (14) is configured to measure the magnetic field experienced at the spacecraft, and the magnetic field orientation of the adjustable magnet (10) is controlled in response to the measured magnetic field thereby controlling the direction and magnitude of the overall dipole moment of the system

A spacecraft structure for transporting propellant to be consumed by a thruster includes storing the propellant in the spacecraft in a solid state during at least a portion of a take-off procedure and supplying the propellant to the thruster in a liquid or vaporous state when the spacecraft is in space.