An arcjet thruster system for a spacecraft is provided. The arcjet thruster system may include a power supply that includes a radio-frequency start power supply and a continuous direct-current power supply, each selectively coupled to electrodes of an arcjet for initiation and maintenance of an arc between the electrodes. A radio-frequency/direct-current control module may be provided for selectively coupling the radio-frequency start power supply and a continuous direct-current power supply. The radio-frequency start power supply may be used to initiate an arc that is then sustained by the continuous direct-current power supply.

Solar radiation is collected and converted to microwave energy by means maintained in outer space on one or more platforms in space. The microwave energy is then transmitted to earth and directed to and constantly focused on one or more segments of a generally circular and slowly rotating tropical depression, which is a precursor to a tropical storm and then a hurricane (alternatively called a cyclone or typhoon). The focused microwave energy interrupts the cycle of vertical upward and downward movement of air, water, & water vapor, and destabilizes the developing rotational motion of the tropical depression, causing it to break up and dissipate, and preventing it from ultimately developing into a hurricane. When not focusing microwave energy on a tropical depression, the microwave energy can alternatively be used to create earth based solar energy, which would help alleviate costs of operating the system solely for hurricane prevention.

A representative spacecraft system includes a launch vehicle elongated along a launch vehicle axis and having at least one stage carrying a corresponding rocket engine. The representative system further includes an annular support structure carried by the at least one stage and positioned to support a cargo spacecraft having a service module and a cargo module. The cargo module of the cargo spacecraft is positioned along the launch vehicle axis in a direction distal from the support structure, and at least a portion of the service module of the cargo spacecraft positioned within an annulus of the support structure.

A combination two vehicle system is provided that includes a land-based vehicle configured for traversal over ground and an aerial vehicle configured for travel through air, where the aerial vehicle is configured to detachably couple with the land-based vehicle. The aerial vehicle includes stabilizing legs to provide stability for take-off or landing. A method of operating the vehicle system includes steps of: providing the land-based and the aerial vehicle having the stabilizing legs, and aerially transporting the vehicle system including the land-based vehicle and the aerial vehicle, where the stabilizing legs provide stability for take-off or landing. In addition, the land-based vehicle may transport a removable cargo pod, which the aerial vehicle may independently retrieve from the land-based vehicle.

A satellite rescue system (SRS) (1) for rescue and recertification of dormant satellites, said SRS having a thruster end (13) with a primary propulsion nozzle (11) and maneuvering thrusters (12) and a satellite connection end (8) with a body (15) between both ends. The satellite connection end of the SRS has an interface ring (14) with clinch clamps (4) that securely attach to a ring (3) on the rescued satellite. An umbilical connector (7) on the satellite connecting end of the SRS provides power and data to the rescued satellite.

A cable (130) includes an electrical wire band (134). A parallel electrical wire (133) is formed by arranging a pair of a first electrical wire (131) and a second electrical wire (132) without being twisted together. The electrical wire band is formed by arranging a plurality of parallel electrical wires in a lateral row. In addition, between two adjacent parallel electrical wires, a second electrical wire of one parallel electrical wire is arranged adjacently to a first electrical wire of the other parallel electrical wire. Further, the electrical wire band is wound and accommodated in a spiral spring shape in a cable wrap mechanism (100).

A matrix thruster that may be used to reposition and/or stabilize a CubeSAT satellite. The matrix thruster includes a conductive plate with an opening, a plurality of wires within the opening, a power supply electrically connected to the conductive plate or each of the plurality of wires via an inductor, and an electrical switch. The electrical switch creates a current change that creates an electric potential spike across the inductor. The electric potential spike across the inductor initiates an arc discharge between one of the wires and the conductive plate, which forms plasma that ejects cathode particles from the matrix thruster. Using multiple wires (e.g., four titanium wires) extends the lifetime of the thruster, as each wire restores an inter-electrode film needed for the other wires to continue generating plasma.

This invention relates to a power supply module for nanosatellite systems which will find application in the field of space technology and satellite communications, and in particular for powering nanosatellites. The created power supply module consists of at least one battery pack and at least one control and energy distribution module and provides maximum efficiency at a given illumination by adjusting the operating output voltage of the input stages (1.1, 1.2 and 1.3) according to the illumination of the panels. All nodes in the module are duplicated, which achieves complete redundancy of the module, which is activated after the main node is defective, or when the load is greater than the load which this main node can withstand. The use of power busbars, on the other hand, leads to a reduced voltage drop on the respective line, as well as to lower temperature losses. The input channels for the solar panels are transferred to the battery pack and it is possible to connect them in parallel when there are more than one.

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 thermal management system includes a slurry generator, an injector pump coupled to the slurry generator, a heat exchanger reactor coupled to the injector pump, wherein the heat exchanger reactor is adapted to subject a thermally expendable heat absorption material to a temperature above 60° C. and a pressure below 3 kPa, and wherein the expendable heat absorption material endothermically decomposes into a gaseous by-product. A vapor cycle system is coupled to the heat exchanger reactor and is operatively connected to a thermal load. A thermal energy storage system may be coupled to the vapor cycle system and the thermal load. The thermal energy storage system may isolate the heat exchanger reactor from thermal load transients of the thermal load.