The disclosure relates to a method and apparatus of rotating mass attitude control. The method and apparatus entails rotating a mass to generate thrust. Varying the speed and direction of rotation provides some control of the magnitude and direction of the thrust generated. The method and apparatus of the invention pertinent to an attitude control system for spacecrafts or astromotive vehicles under conditions of zero to low gravity and atmosphere.

A satellite comprises a body and a generally planar structure extending from the body. One or more radio frequency “RF” antennas, an amplification system for RF signals, and a power distribution system for the amplification system are mounted on the generally planar structure. Two or all of the power distribution system, the one or more RF antennas and the amplification system are arranged on respective parallel boards (310, 312, 314) forming part of the generally planar structure (300). One or more of the parallel boards (310, 312, 314) and the components mounted thereon may be connected to another similar board to form, respectively, a larger power distribution system, antenna array or amplification system, for example arranged in a plurality of modules, each comprising at least one antenna, at least one power distribution system and at least one amplifier supported on at least two respective boards. A satellite may be manufactured by assembling the modules to form a generally planar structure, and attaching the planar structure to

A battery arrangement for structurally integrating batteries in a vehicle, in particular an aircraft or spacecraft, includes at least one battery, two supporting cooling plates between which the at least one battery is held on both sides via battery holders, and connecting rods which connect the two cooling plates to one another.

A universal external port is proposed to add new functionality to or replace existing functionality of an already deployed spacecraft (e.g., a satellite in orbit). The universal external port is mounted on an external surface of the spacecraft and configured to connect to different types of external modules that have different functions, without removing components from the spacecraft other than one or more components of the universal external port. A communication interface onboard the spacecraft is configured to wirelessly receive a software patch from an entity remote from the spacecraft (e.g., from a ground terminal or other spacecraft) to program the spacecraft to change operation of the spacecraft to utilize the external module when the external module is connected to the universal external port.

In accordance with various embodiments of the disclosed subject matter, a system and method is configured for scheduling and invoking power sharing among satellites within a constellation of satellites such that energy storage systems at a target satellite may by charged prior to the use of electric propulsion thrust activation or other high electricity demand operations (or such operations contemporaneously augmented) by power beams transmitted from other (source) satellites within the constellation.

A spacecraft propulsion system including a first thruster system including a first power processing unit connected to a first thruster string via a crossover switching unit and connected to a second thruster string via the crossover switching unit. A second thruster system including a second power processing unit connected to the second thruster string via the crossover switching unit and to the first thruster string via the crossover switching unit. A controller is connected to each of the first power processing unit, the second power processing unit, and the crossover switching unit.

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).

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).

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.

Described herein is a power processing unit (PPU) for use with a Hall Effect Thruster (HET) and a Propellant Management Assembly (PMA) of a spacecraft. The PPU comprises an anode and ignitor supply subsystem that provides anode and ignitor signals to an anode and an ignitor circuit of the HET. The PPU also comprises a valve control subsystem that provides valve control signal(s) to valve(s) of the PMA. The anode and ignitor supply subsystem and the valve control subsystem are each coupled to a low voltage (LV) bus of an electrical power subsystem of the spacecraft. The anode and ignitor supply subsystem includes a step-up DC-DC converter having a transformer that steps-up a voltage of the LV bus to a higher voltage used to produce the anode and ignitor signals. The valve control subsystem is devoid of a transformer. An Electric Propulsion System (EPS) includes the PPU, HET and PMA.