Methods and apparatuses for emitting electrons from a hollow cathode are provided. The cathode includes a plasma holding region configured to hold a plasma, a gas supply source configured to supply gas to the plasma holding region, and an orifice plate disposed on a periphery of the plasma holding region. The orifice plate comprises a plurality of openings constructed to receive electrons from the plasma. The plurality of openings decouple gas conductance and electrical conductance across the orifice plate. The diameters of the plurality of openings are within a range of 20%-60%, inclusive, of a diameter of a circular opening with an area equal to a sum of the areas of the plurality of openings.

A Hall thruster includes an annular discharge region and an annular cathode concentric to the annular discharge region.

An ion propulsion device including emission modules in an emission plane, each module having an insulating support, an emission electrode on the support, and a conductive liquid with a microfluidic channel depositing conductive liquid on the electrode; an extraction electrode common to the emission modules and facing the modules; and a control unit, in which each module is configured to emit an ion beam when an electric field is applied to the liquid; each control unit controls an ion emission current emitted by applying a potential difference between each emission electrode and the extraction electrode; the emission electrodes are spaced apart by a linear distance that is greater than a distance between two adjacent emission electrodes separated by an empty space; and a length of the insulating support between the electrodes is greater than a propagation distance of an electric leakage current by charge jumping along the support between the electrodes.

A heaterless hollow cathode provides electron emission current in an electric space propulsion system. A mechanical, thermal, and electromagnetic design of the cathode apparatus is presented, and a method of operation for rapid ignition and stabilization of the cathode is provided. The keeper of the cathode apparatus has a thickness change which reduces the flow of heat away from the cathode's emitter assembly. The method for heating the emitter assembly includes controlling applied voltages so that the current flowing from the emitter assembly to the keeper is maintained at a predetermined fixed value. By this method, damage to the electron emitting surfaces of the emitter assembly by electric arcing and/or by depletion of dopant materials is avoided.

The present invention relates to a neutralizer (4) for an ion thruster (1) of a spacecraft (S), comprising: a cathode (5) for emission of electrons (6), a support (7) with an opening (8) inside which the cathode (5) is supported in a radially spaced manner, and an electrically conductive shielding (9) which surrounds said opening (8) and is electrically insulated from the support (7), wherein a ring (11) is mounted between the shielding (9) and the cathode (5) and is electrically insulated from the shielding (9) and radially spaced from the cathode (5).

A neutralizer suitable for use in an ion engine comprises a halogen gas source and an electrode tube comprising an inlet opening connected to the halogen gas source for supplying a halogen gas provided by the halogen gas source into the electrode tube, a discharge space for generating a plasma from the halogen gas supplied into the electrode tube, and an outlet opening for discharging the plasma generated in the discharge space and free electrons from the electrode tube. An electron emitter is arranged in the discharge space of the electrode tube, which is at least partially made of tungsten, a tungsten alloy or a tungsten composite material containing at least one of iridium, rhenium, ruthenium, rhodium and osmium.

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 field emission neutralizer is provided. The field emission neutralizer includes a bottom plate and a field emission cathode unit located on the bottom plate. The field emission cathode unit includes a substrate, a shell located on the substrate, a cathode emitter located inside the shell, a mesh grid insulated from the cathode emitter, and a shielding layer insulated from the mesh grid. The cathode emitter includes a cathode substrate and a graphitized carbon nanotube array. The graphitized carbon nanotube array is in electrical contact with the cathode substrate. The graphitized carbon nanotube array is fixed on a surface of the substrate body, and the carbon nanotubes of the graphitized carbon nanotube array are substantially perpendicular to the cathode substrate.

A field emission electron source and a method of manufacturing the same. A field emission electron source comprises an emitting electrode and an extractor gate electrode. The emitting electrode comprising a plurality of particles with nanosharp protrusions. The extractor gate electrode comprises a metal. The extractor gate electrode is formed in a same plane as the emitting electrode. The extractor gate electrode is formed surrounding the emitting electrode. A method of manufacturing a field emission electron source comprises forming an emitting electrode comprising a plurality of particles with nanosharp protrusions using a direct ink writing (DIW) printer. The method comprises forming an extractor gate electrode comprising a metal using the DIW printer.

A structure for preventing an electric discharge between devices without any need for a dedicated device when mounted on an artificial satellite is provided. A charging mitigation device for reducing a potential difference between two or more physical objects includes an electric thruster (100) configured to generate plasma (PL), wherein the two or more physical objects are covered with the plasma (PL) generated from the electric thruster (100).