An ion thruster (1) and a method for providing trust is disclosed. The ion thruster comprises a sputtering magnetron (2), a target (3) arranged at the sputtering magnetron, and a second electrode (4). During a first pulse, the target is at a negative potential (U.sub.HiP) with respect to a second electrode and a plasma is sustained whereby atoms are sputtered from the target and at least a portion thereof become ionised by the plasma. During a second pulse, a reversed potential (U.sub.rev) is applied between the target and the second electrode. This increases the potential of a volume of the plasma adjacent to the target, which in turn accelerates ions in a direction away from the target. Thereby, thrust is provided.

The disclosure further relates to a computer program and a computer readable medium, as well as a spacecraft comprising the ion thruster.

An atmosphere-breathing pulsed plasma thruster includes an inlet, a discharge section with an anode in fluid communication with the inlet, and a nozzle in fluid communication with the discharge section. Electrode assemblies extend radially through the discharge section and include a second electrode in the discharge section and an elongate portion extending outwardly. An electrode control mechanism moves the plurality of electrode assemblies between an inner position nearer to the anode and an outer position farther from the anode. At least one igniter extends between the anode and a cathode. An ignition circuit connects the anode and the cathodes to a first source of electric energy, and connects the igniter to a second source of electric energy through a controllable switch. A processor controls the position of the second electrodes, for example, in response to changes in atmospheric pressure.

A propulsion system for a spacecraft includes at least one chemical propulsion engine, at least one electrical propulsion engine, and a propellant feed system that includes a propellant supply system including at least one tank to store a common propellant. The propellant supply system is configured to deliver a controlled flow of the common propellant to both the chemical propulsion engine and the electrical propulsion engine. Also disclosed is the use of a common propellant in a spacecraft and a method of providing a common propellant to an electrical and a chemical propulsion engine.

A propulsion system for a spacecraft includes at least one chemical propulsion engine, at least one electrical propulsion engine, and a propellant feed system that includes a propellant supply system including at least one tank to store a common propellant. The propellant supply system is configured to deliver a controlled flow of the common propellant to both the chemical propulsion engine and the electrical propulsion engine. Also disclosed is the use of a common propellant in a spacecraft and a method of providing a common propellant to an electrical and a chemical propulsion engine.

A Hall-effect thruster assembly includes a plurality of magnetic sources for creating a magnetic circuit. The plurality of magnetic sources are positioned between a first end and a second, opposite end of the Hall-effect thruster. The plurality of magnetic sources define a longitudinal axis extending through the first end and the second end. The first end is configured as a discharge end. A mount assembly is coupled to the second end. The mount assembly is configured to secure the plurality of magnetic sources to a spacecraft. A magnetic element is supported by the mount assembly. The magnetic element is positioned relative to the plurality of magnetic sources by the mount assembly.

A time average peak value of low frequency magnetic noise or low frequency conductive noise generated from a power supply device which drives a Hall thruster is suppressed without mass of a satellite significantly increased. A pulse width control circuit (22) of a Hall thruster power supply device (10) outputs a spread signal (58) obtained by performing spread spectrum on a pulse signal based on a control signal (54). A voltage output circuit (21) outputs output voltage (52) to a Hall thruster (50) in accordance with the spread signal (58) output by the pulse width control circuit (22).

The disclosed subject matter relates to an ion thruster for thrust vectored propulsion of a spacecraft, comprising a reservoir for a propellant, an emitter having a base and, on one side of the base, at least one outlet for emitting ions of the propellant, wherein the base is connected to the reservoir for providing flow of propellant from the reservoir to said at least one outlet, and an extractor facing said one side of the emitter for extracting and accelerating the ions from the emitter, wherein the extractor is split into sectors about an axis which orthogonally runs through said one side of the emitter, wherein said sectors are electrically insulated from one another.

The invention relates to an ion thruster, comprising: a chamber, a reservoir, comprising a solid propellant (PS), housed in the chamber and comprising a conductive jacket provided with an orifice; means for forming an ion-electron plasma in the chamber, which means are able to sublime the solid propellant in the reservoir, then to generate said plasma in the chamber from the sublimed propellant coming from the reservoir through the orifice; a means for extracting and accelerating the ions and electrons of the plasma out of the chamber, which means comprises at least two grids at one end (E) of the chamber; a radiofrequency AC voltage source for generating a radiofrequency signal comprised between the plasma frequencies of the ions and of the electrons, arranged in series with a capacitor and connected, by one of its outputs and via this capacitor, to one of the grids, with the other grid being connected to the other output of said voltage source;
said means for extracting and accelerating and said voltage source making it possible to form, at the output of the chamber, an ion-electron beam.

A thruster for a micro-satellite is disclosed. The thruster includes a cathode composed of a propellant material and an anode composed of ablative material. The thruster includes a housing having a proximate end and an opposite distal end having a thrust channel. The housing holds the anode and the cathode. A pulsed voltage source is coupled between the cathode and the anode causing current sufficient to create ablation of the anode and a plasma jet including ablated particles from the anode to be emitted from the thrust channel.

The invention provides a plasma production and control device that may be used in propulsion (e.g., satellite propulsion) and/or industrial applications. The plasma production system comprises a unidirectional magnetic field.