A satellite having a cooling system to remove heat from an electric rocket engine using a working fluid. The cooling system can include a pump that circulates working fluid along a cooling loop between the rocket engine and a radiator. The cooling system can also utilize thermoacoustic, Stirling refrigeration, and/or heat pipe techniques. One or more reservoirs can be provided to store the working fluid, and in some forms a secondary reservoir can be provided to aid in management of a center of mass of the satellite. A fluid reaction loop can be provided in which working fluid is accelerated to impart a torque on the satellite. In some forms, the working fluid can be utilized as both a coolant and a propellant for the rocket engine. The electric rocket thruster can also include one or more internal pathways for the conveyance of working fluid.

This application describes creating, modifying, and bending electromagnetic solitons at large scales for the various applications. An electromagnetic soliton generator system controls the magnetic soliton such that the orientation, rotation rate, pitch angle, and magnetic field strength of the solitons are modified to provide the described standing waves and generate a magnetic flux differential.

An electrospray thruster including an emitter, an extractor, a propellant storage vessel for a primary liquid propellant, a propellant delivery pathway from the vessel to the emitter, and an ionic liquid. The ionic liquid is configured to have a solid phase at temperatures less than a predetermined temperature and a liquid phase at temperatures greater than the predetermined temperature, and the ionic liquid is configured to create a propellant isolation barrier in the solid phase to prevent absorption by the primary liquid propellant. The electrospray thruster also includes a heater associated with the vessel and configured to heat the ionic liquid to above the predetermined temperature for mixing with the primary liquid propellant.

A satellite thruster increases satellite efficiency. The Linear Actuated μCAT has a stepper motor to move the ablative electrode forward. A LabVIEW program and Arduino microcontroller are used to analyze the Linear Actuated μCAT to determine how many steps are required for re-ignition, arc current, and the validity of the feed system. Results from testing show that micro-stepping the stepper motor is an effective way to replenish the cannibalized electrode for propellant.

Embodiments relate to a hallow cathode with integral layers of radiation shielding. The hollow cathode includes an inner cathode tube that forms a gas feed to direct gas toward a downstream end, where the directed gas forms plasma. A heater element is positioned at the downstream end of the inner cathode tube, the heater element to heat the plasma. The hollow cathode further includes an outer cathode tube with a keeper electrode to sustain a bias voltage across a gap at a downstream end of the outer cathode tube for igniting the plasma. The integral layers of radiation shielding are connected by offset radial supports and are incorporated as a single element with either the inner or outer cathode tube, where the integral layers are nested with torturous conductive paths to reduce radiation and conduction losses from the downstream end of the inner cathode tube.

A transfer system for supplying a receiving tank of a receiving spacecraft with a supply gas from a supply spacecraft. A transfer tank is disposed on the supply spacecraft and configured to retain a supply gas. A transfer line is coupled to the transfer tank, and one end thereof maybe coupled to the receiving tank. A transfer valve is operatively coupled to the transfer line. A heating system is thermally coupled to the transfer tank. A control system is operatively coupled to the transfer valve and the heating system. The control system is operable to cause a transfer quantity of the supply gas to be heated, and to open the transfer valve, such that a difference between the increased pressure of the supply gas in the transfer tank and a pressure in the receiving tank causes the transfer quantity of the supply gas to flow to the receiving tank.

A control assembly for a spacecraft includes a propellant management assembly configured to adjust a supply of propellant from a storage unit to a thrust generator. The control assembly further includes a controller having a processor configured to receive an input from the spacecraft, and receive at least one input from the propellant management assembly or from the thrust generator. The processor is further configured to, based on the inputs, determine a desired operating mode of the thrust generator, and based on the determination, either 1) send an output to the propellant management assembly to operate in a first mode in which the thrust generator uses propellant to electrostatically generate thrust or 2) send an output to the propellant management assembly to operate in a second mode in which the thrust generator uses propellant to gas-dynamically generate thrust.

According to some aspects of the subject disclosure, a spacecraft comprises first and second pluralities of thrusters. The pluralities of thrusters are attached to a spacecraft body by booms configured to move the first plurality of thrusters between stowed and deployed positions. The deployed position of the first plurality of thrusters is farther north than is the stowed position of the first plurality of thrusters. The deployed position of the second plurality of thrusters is farther south than is the stowed position of the second plurality of thrusters. The first plurality of thrusters comprises a first thruster and a second thruster separated from each other in an east-west direction. The second plurality of thrusters comprises a third thruster and a fourth thruster separated from each other in the east-west direction.

A plasma thruster comprises a cylindrical discharge channel (1), an injector (4), a RF antenna surrounding the discharge channel (1) and a device (3) for generating an axial static magnetic field in the discharge channel (1). The RF antenna is a cylindrical birdcage antenna (2) formed of several electrically conductive parallel legs (10) connected by two end rings (11) including capacitors (12) between adjacent legs (10) in each case. The two end rings (11) with the capacitors (12) are formed on two printed circuit boards (14) to which the legs (10) are attached, said printed circuit boards (14) having a through opening for the discharge channel (1). The antenna maximizes electrical coupling efficiency and provides resulting electromagnetic fields for quasi-neutral plasma acceleration along with the magnetic field effect provided by the externally applied magnetic field. This plasma thruster allows an easy upscaling or downscaling due to the printed circuit boards and is particularly suitable for low power applications like propulsion for smaller spacecrafts or satellites.

Described herein is a thrust system for a vehicle that includes at least three electrical power controllers, at least four electrical switches each configured to receive electrical power from at least one of the at least three electrical power controllers, and at least three thrusters each configured to receive electrical power from at least one of the at least three electrical switches. The at least four electrical switches are operable to switch a supply of electrical power from any of the at least three electrical power controllers to any one of the at least three thrusters.