A distributed computer system for a spacecraft is disclosed. The system has multiple computer nodes, each controlling a different aspect of a mission of the spacecraft. Each node includes a control circuit(s) that controls a set of components, a router processor, and a programmable processor. The programmable processor of each respective computer node issue commands to the control circuit(s) of the respective computer node to carry out an aspect of the mission associated with the respective computer node. Upon failure of the programmable processor in a particular computer node, a healthy programmable processor send commands to the router processor in the particular computer node The router processor of the particular computer node routes the commands received from the remote programmable processor to the control circuit(s) in the particular computer node to control the set of components to carry out the aspect of the mission associated with particular computer node.

Methods, systems, and devices for satellite operations are described. A system for satellite communications may include a payload, a power system, and a thermal management component. The payload may be configured to provide a service with varying levels of capacity based on a demand profile. The payload may consume electrical energy at a peak rate when a level of demand indicated by the demand profile is above a threshold and at a lower, off-peak rate when a level of demand indicated by the demand profile is below a threshold. The peak rate may exceed a rate at which electrical energy is generated by the power system. The thermal management component may process excess thermal energy generated by the payload when the payload operates at the peak rate. Processing the excess thermal energy may include storing thermal energy while the payload operates at the peak rate.

A universal test port is connected to the different functional sub-systems of a spacecraft, allowing the sub-systems to be tested from a single location of an assembled spacecraft. The universal test port is mounted on an external surface of the spacecraft and configured to connect to the different functional sub-systems (such as power, propulsion, and command and data handling, for example) of the assembled spacecraft, allowing for the streamlining of testing operations by electrical ground system equipment during assembly, integration, and test (AIT) operations and reducing the risk of collateral damage to spacecraft hardware during testing in AIT.

Technology is disclosed herein for a power control and distribution unit (PCDU) of a spacecraft that has a microcontroller to control battery charging from solar arrays. Using a microcontroller within the PCDU reduces the complexity of the PCDU. The microcontroller may be programmable and reprogrammable, which allows the charging of the battery to be adapted to various conditions. For example, the microcontroller can be programmed in accordance with the mission to optimize battery charging for that mission.

A voltage measurement apparatus for a spacecraft thruster includes a thruster interface having a platform and at least one spacer element configured to attach the platform to the thruster. At least one probe is mounted to the thruster interface with the probe configured to engage an anode assembly of the thruster when the thruster interface is attached to the thruster. The apparatus also includes voltage metering circuitry coupled to the at least one probe, the voltage metering circuitry configured to be powered by the spacecraft thruster when the spacecraft thruster is powered on.

Systems, apparatuses, and methods provide for technology that locates one or more masses of a thunderstorm system, as the thunderstorm system starts to organize and before the thunderstorm system spawns a tornado, and controls a transmission of electromagnetic radiation from the space platform to the one or more masses, wherein the transmitted electromagnetic radiation tracks the one or more masses as the thunderstorm system is starting to organize and rotate, and wherein the transmitted electromagnetic radiation prevents the thunderstorm system from rotating and spawning the tornado.

According to certain aspects, an electric-propulsion thruster is used as part of a base or platform which also includes a power converter, having a plurality of inductors and other electrical components, and a printed circuit board (PCB). The PCB includes a layer at which the other electrical components and printed circuit inductor traces, for the plurality of inductors, are secured. The electric-propulsion thruster includes a housing (e.g., as part of the base or platform) providing a cavity and having at least one structurally-rigid side wall along the cavity, where the PCB is integrated with the electric-propulsion thruster for a compact arrangement which can be used to propel the apparatus. Such a compact design might be used as an important part of thruster spacecraft architecture such as micro-satellites (e.g., CubeSats).

A thrust generator is provided for producing thrust to move a spacecraft. The thrust generator includes a housing having a first end and an opposing second end. The first end is associated with a mount for coupling to the spacecraft. The housing further defines a central axis extending through the first end and the second end. The second end defines an annular propulsion outlet. At least one nozzle is positioned proximate the second end. The thrust generator is selectively operable in a first mode in which the thrust generator uses propellant to electrostatically generate thrust via the annular propulsion outlet, and a second mode in which the thrust generator uses propellant to gas-dynamically generate thrust via the at least one nozzle.

A space vehicle comprising an optical system having a field of view, the optical system comprising at least two optical elements spaced from one another along an optical axis, thereby defining an interior cavity; at least one control system comprising at least one physical element configured for performing function(s) for enabling operation of the vehicle; and at least one holding assembly for holding the at least one control system and comprising a folding mechanism configured to move between a folded position corresponding to an inoperative mode of the optical system, and a deployed position corresponding to an operative mode of the optical system, such that in the folded position, the control system is at least partially located in the interior cavity for stowage, and in the deployed position, the control system is located outside the interior cavity and outside the field of view of the optical system, allowing operation of the optical system.

A reconnaissance rover configured for multiple agile and autonomous landings over a small body or moon. The reconnaissance rover comprises a detection unit, a processing unit, a control unit and a drive unit. The detection unit is configured to detect at least an environment in front of the reconnaissance rover, in the direction of a trajectory of the reconnaissance rover over a surface of the small body or moon. The detection unit is further configured to provide environmental data based on the detected environment. The processing unit is configured to update the trajectory based upon the provided environmental data. The control unit interacts with the drive unit to move the reconnaissance rover according to the updated trajectory.