A circuit for igniting and sustaining an electron discharge includes an ignitor circuit. The igniter circuit includes a high voltage transformer and a switch connected in series between a primary of the transformer and a DC source return. The switch is configured to receive a driving signal. A reset circuit is connected in parallel to the primary of the high voltage transformer. A first rectifier is connected in series between a secondary of the high voltage transformer and a keeper. A terminal of the secondary of transformer is connected to a cathode. The circuit for igniting and sustaining the electron discharge also includes a sustaining circuit having a current source with a return connected to a cathode and a second rectifier connected in series between the current source and the keeper.

A micro-cathode arc propulsion system. By replacing an inductive circuit in a traditional micro-cathode arc propulsion system with a capacitor circuit, the stability of the operation of a micro-cathode arc thruster can be improved due to the stable discharging mode of the capacitor, and as the internal resistance of the capacitor is small during operation, the additional power consumption of the circuit is reduced, and the efficiency of the system is improved. In addition, as a pulse power supply is used to power in a pulse manner, the average power inputted into the micro-cathode arc thruster is greatly reduced.

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.

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 system operating in a spacecraft includes a tank storing a propellant, a thermal receiver configured to change the propellant from a first phase to a second phase by providing heat to the propellant, and an energy conversion device fluidically coupled to the tank and configured to generate electric energy using the propellant in the second phase.

A spacecraft propulsion system comprises two thrusters, each operating in accordance to a corresponding propulsion technique. A controller is configured to direct collected solar energy to heat a propellant for consumption in one of the two thrusters, or to generate electric energy for the other one of the two thrusters.

A propulsion system for a spacecraft includes at least one electrical propulsion engine comprising at least one neutraliser; and a pressurant gas system comprising a pressurant gas which is fed directly into the at least one neutraliser.

A space-based solar power station, a power generating satellite module and/or a method for collecting solar radiation and transmitting power generated using electrical current produced therefrom is provided. Power transmitters can be coordinated as a phased array and the power generated by the phased array is transmitted to one or more power receivers to achieve remote wireless power generation and delivery. In many embodiments, a reference signal is distributed within the space-based solar power station to coordinate the phased array. In several embodiments, determinations of the relative locations of the antennas in the array are utilized to evaluate the phase shift and/or amplitude modulation to apply the reference signal at each power transmitter.

In various embodiments, an electrothermal propulsion system for imparting velocity to a spacecraft features a joint-free monolithic integrated thruster component including an inlet section, a resistively heated heat-exchanger section, and a nozzle section.

A modular and scalable system to transfer space articles between space orbits. In one embodiment, the system employs a rendezvous vehicle which docks with a space article in an initial orbit, the connected stack then docking with a locomotive vehicle which maneuvers to a targeted orbit where the space article is detached. In one feature, the rendezvous vehicle and locomotive vehicle use a common propellant and the space article is a satellite.