A power supply device (104-1) includes a substrate (20) on which an electric component (25) is mounted, a chassis (10) having a chassis surface (11) and a threaded part (10a), a chassis-side resin part (91) connected to a back surface (20a) and the chassis surface (11), a fixation screw (29), and an insulating member (60). The fixation screw (29) fixes both the electric component (25) and the substrate (20) to the chassis (10) by screw-coupling an end part (29c) of the fixation screw (29), exposed in a direction toward the chassis (10) from an open hole formed through the insulating member (60), to the threaded part (10a) of the chassis (10). In addition, the fixation screw (29) brings the electric component (25) and the chassis (10) into electrical noncontact with each other by being placed in an open hole of the insulating member (60).

Systems and methods for describing, simulating and/or optimizing spaceborne systems and missions. Configurations for spaceborne systems are generated and validated based on simulation output.

A spacecraft is disclosed, comprising a deployable spacecraft body (110) comprising a plurality of sub-systems (321-324) for controlling operations of the spacecraft, and a plurality of panels (101, 102) and a plurality of hinges (112-115) each connecting adjacent ones of the plurality of panels, the hinges being arranged to permit the plurality of panels to be folded into a stowed configuration and unfolded into a deployed configuration, wherein the plurality of sub-systems are fixed to and supported by one or more of the plurality of panels. By forming the body of the spacecraft from a deployable structure, the overall size of the spacecraft can be significantly reduced in the stowed configuration. In some embodiments, a plurality of the spacecraft in the stowed configuration can be combined into a modular spacecraft assembly prior to launch, with data and power connections between the plurality of stowed spacecraft being used to transfer power from, and data to, a payload monitoring unit on the launch vehicle.

A satellite system operates at altitudes between 180 km and 350 km relying on vehicles including an engine to counteract atmospheric drag to maintain near-constant orbit dynamics. The system operates at altitudes that are substantially lower than traditional satellites, reducing size, weight and cost of the vehicles and their constituent subsystems such as optical imagers, radars, and radio links. The system can include a large number of lower cost, mass, and altitude vehicles, enabling revisit times substantially shorter than previous satellite systems. The vehicles spend their orbit at low altitude, high atmospheric density conditions that have heretofore been virtually impossible to consider for stable orbits. Short revisit times at low altitudes enable near-real time imaging at high resolution and low cost. At such altitudes, the system has no impact on space junk issues of traditional LEO orbits, and is self-cleaning in that space junk or disabled craft will de-orbit.

The present invention provides 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 of the optical system, thereby defining an interior cavity of the optical system; at least one control system, the control system comprising at least one physical element configured for performing one or more functions for enabling operation of the vehicle; and at least one holding assembly for holding the at least one control system, the holding assembly comprising a folding mechanism configured and operable 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 of the folding mechanism, the control system that is held by the holding assembly is at least partially located in the interior cavity of the optical system for stowage, and in the deployed position of the folding mechanism, the control system that is held by the holding assembly is located outside the interior cavity and outside the field of view of the optical system, thereby allowing operation of the optical system.

A method for coupling an energy supplying device to an energy consuming device in space and to a system including an energy supplying device, an energy consuming device and a mobile assembly unit residing in space. The energy supplying device is operable in space and includes solar panel units to convert received light into electrical energy, first coupling members to couple the solar panel units with each other, and a second coupling member to electrically couple the energy supplying device with the energy consuming device. The energy supplying device is configured to supply electrical energy to the energy consuming device in space. A configuration of the solar panel units is changeable in space according to an operational requirement of the energy consuming device.

A method for providing thermal balance of spacecraft electronics is provided. The spacecraft includes two or more electronic units wherein each electronic unit is capable of performing the same spacecraft operational task. The method for balancing the temperature of spacecraft electronics further includes providing each of the two or more electronic units with a temperature sensor for determining the temperature of that electronics unit. The electronic units and their respective temperature sensors are connected to a controller. In the event that the controller determines that the temperature of an activated first electronics unit has reached or exceeded a predetermined threshold, and the controller has determined that the temperature of a second deactivated electronics unit is below a predetermined threshold, the controller automatically deactivates the first electronics unit and activates the second electronics unit to perform the task previously being performed by the first electronics unit. This process continues automatically.

A method for providing thermal balance of spacecraft electronics is provided. The spacecraft includes two or more electronic units wherein each electronic unit is capable of performing the same spacecraft operational task. The method for balancing the temperature of spacecraft electronics further includes providing each of the two or more electronic units with a temperature sensor for determining the temperature of that electronics unit. The electronic units and their respective temperature sensors are connected to a controller. In the event that the controller determines that the temperature of an activated first electronics unit has reached or exceeded a predetermined threshold, and the controller has determined that the temperature of a second deactivated electronics unit is below a predetermined threshold, the controller automatically deactivates the first electronics unit and activates the second electronics unit to perform the task previously being performed by the first electronics unit. This process continues automatically.

Establishing and growth of a lunar or planetary surface base involves continuing to use landing spacecraft as docked modules of the base for habitation and work. A first spacecraft is landed at a specified surface site then doubles as first module of the base. A second (and later third and subsequent) spacecraft is landed at the site a safe distance from the existing base modules then moved over the surface into a side-by-side position to dock with selected base modules. At least some of the landing, surface transport, and operational electric power is supplied by micro-fusion using ambient cosmic rays and muons interacting with deuterium-containing particle fuel material to generate energetic reaction products.

A flying capacitor multilevel (FCML) converter including a gate driver circuit comprising a DC-DC flyback converter having a plurality of isolated outputs. In various examples, the FCML circuit further includes a first control circuit connected to the FCML circuit determining the load current associated with a desired power output from the load; and determining a desired output voltage associated with the load current; a second control circuit that drives an inductor current (I.sub.L) through the inductor so that the output applies an output voltage comprising the desired output voltage; and a third control circuit obtaining a comparison of an average of the inductor current (I.sub.L) through the inductor with a predetermined reference current (I.sub.LREF) and setting the duty cycle so that the average does not exceed the predetermined reference current. Also described is the converter driving a load comprising a plasma and a propulsion system comprising the converter.