The present technology is directed to plasma systems and associated methods, including propulsion systems for flight vehicles. A representative system includes a plurality of coils. The coils include a first coil positioned along a force axis, a second coil positioned along the force axis and spaced apart from the first coil, and a third coil that is magnetically shielded. A controller is operatively coupled to the coils and is configured to (a) increase energy to the first coil to generate a magnetic field in a portion of the plasma adjacent to the first coil, (b) decrease energy to the first coil and increase energy to the second coil to translate the resulting superposed magnetic field through the plasma to a position adjacent the second coil, and (c) transfer energy from the second coil to the third coil and decrease energy to the second coil to reduce the magnetic field in the plasma.

A method for controlling an ion thruster including an emission electrode, an extraction electrode and a conductive liquid which is deposited on the emission electrode, the ion thruster configured for emitting an ion beam when an electric field is applied to the conductive liquid, the ion beam providing thrust to the thruster, the thrust depending on an emission current I.sub.em and an ion emission speed, the method including the following steps: adjusting the emission current to a setpoint value I.sub.c by applying a threshold emission potential V.sub.thresh to the emission electrode by means of a current generator; and when the setpoint value I.sub.c of the emission current is reached, adjusting the emission speed by applying an extraction potential V.sub.ext to the extraction electrode by means of a voltage generator in order to bring the emission potential V.sub.em to a predetermined value V.sub.empr=V.sub.thresh+V.sub.ext.

A reconfigurable power processing unit for a spacecraft including a plurality of power modules. Each of the power modules includes a first power source and a second power source. The first power source and the second power source are configured to be in series in a first state and in parallel in a second state. A plurality of contactors connect each power module to at least one of another power module in the plurality of power modules and a power processing output and are configured to control the state of the power modules.

A plasma propulsion system with no internal electrodes is described. Gas is flowed into an insulated axisymmetric plasma liner. A radio frequency antenna generates an inductive or helicon plasma discharge within the liner. The plasma is accelerated through a converging/diverging magnetic field out of the liner, generating thrust.

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.

Spacecraft thruster systems are disclosed. In some instances, a spacecraft thruster system may include stacked ion thrusters and/or ion thruster layers. The ion thrusters and/or ion thruster layers may be sequentially activated and jettisoned from the thruster system after use.

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.

Scalable power processing units (PPUs) for Hall-effect thrusters (HETs) and terrestrial systems are disclosed. A technique for current estimation may be employed on each output of parallel isolated discharge supply modules (DSMs) to force proper current/load sharing between the DSMs. A flyback power supply may be used that performs the dual functions of a cathode keeper plasma ignitor and sustainer. The flyback power supply may be tuned for a high no-load direct current (DC) output voltage to achieve cathode keeper ignition rather than requiring a separate ignitor supply, which reduces circuit complexity. To address requirements for higher voltage DC ignition than are achievable with a flyback power supply alone, a supplemental DC ignitor may be placed in parallel with the flyback power supply of some embodiments. Such simplified PPU architectures may provide a high efficiency, low part count, scalable architecture suitable for more compact and lower cost system designs.

A pulsed metal plasma thruster (MPT) cube has a plurality of thrusters, each having a first cathode electrode and a trigger electrode separated from the first electrode by an insulator sufficient to support an initiation plasma, and a porous anode electrode positioned a separation distance from the face of all of the cathode electrodes. The cathode electrode can be either the inner electrode or the outer electrode. A power supply delivers a high voltage pulse to the trigger electrode with respect to the cathode electrode sufficient to initiate a plasma on the surface of the insulator. The plasma transfers between the anode electrode and cathode electrode of selected thrusters, thereby generating a pulse of thrust.