Fluid flow control valve comprising a tubular body extending in a longitudinal direction with a fluid inlet and a fluid outlet situated respectively at the two longitudinal ends of the body, the valve comprising a nozzle and a piston connected to the body, the piston being housed in the body, the nozzle being made up of a part provided with a fluid passage having a calibrated dimension, the passage emerging at one end of the nozzle and forming a seat, said seat being situated against a terminal end of the piston forming a shutter preventing the flow of fluid in the closed position of the valve, the piston comprising a body defining a passage for the fluid in the body for the flow of the fluid between the inlet and the outlet, the body of the valve consisting of a material having a different expansion coefficient from the piston or the nozzle, the valve comprising a heating member which, depending on the heating power delivered, makes it possible to separate the end of the nozzle and the piston by differential expansion, to allow the flow of fluid between the inlet and the outlet in an open position of the valve, characterized in that the terminal end of the piston comprises a ball that is crimped into the body of the piston.

Aspects disclosed herein include graphite and hexagonal boron nitride bimaterials, methods of making these bimaterials, and electric propulsion devices or thrusters with these bimaterials. Aspects disclosed herein include electric propulsion devices comprising: at least one portion comprising or formed of a monolithic bimaterial; wherein the monolithic bimaterial comprises a graphite material and a hexagonal boron nitride material; and wherein the graphite material and hexagonal boron nitride material are monolithically integrated in the bimaterial.

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.

Spacecraft propulsion systems and methods featuring a first storage tank containing a metallic propellant and a second storage tank containing a non-metallic propellant are provided. A selected one of the metallic propellant and the non-metallic propellant is supplied to an electric propulsion thruster, depending on an operational mode of the spacecraft. The metallic propellant is stored at a relatively high density, while the non-metallic propellant is stored at a lower density than the metallic propellant. Moreover, the non-metallic propellant is preferably utilized to produce thrust through the electric propulsion thruster during operational maneuvers, while the metallic propellant is reserved for producing thrust through the electric propulsion thruster during end-of-life, such as deorbiting, maneuvers.

An apparatus generates energetic particles and generates a plasma of a vaporized solid material and gaseous precursors for the application of coatings to surfaces of a substrate by way of condensation of plasma and for electric propulsion applications.

A dual mode engine for propelling a spacecraft, including a combustion chamber having a flange end, an open nozzle end, and an enclosed chamber portion extending therebetween, a propellant tank in fluidic communication with the combustion chamber, an electronic controller, a power source operationally connected to the electronic controller, and a fluid flow motivator operationally connected to the electronic controller and connected in fluidic communication with the propellant tank. The engine further includes a chemical propulsion portion further including a propellant inlet port operationally connected to the combustion chamber and disposed adjacent the flange end, an ignition trigger electrode positioned in the combustion chamber adjacent the propellant inlet port and operationally connected to the electronic controller and operationally connected to the power source, wherein the propellant inlet port is fluidically connected to the propellant tank. The engine also includes an electric propulsion portion, further including at least two spaced electrodes for ionizing the propellant positioned in the combustion chamber adjacent the nozzle end, a plurality of attitude control thrusters operationally connected to the electronic controller and in fluidic communication with the propellant tank, and a plurality of respective valves, each respective valve fluidically connected between a respective attitude control thruster and the propellant tank, wherein the propellant inlet port is fluidically connected to the propellant tank.

A method of operating a spacecraft propulsion system comprises injecting electrons into the plasma surrounding the spacecraft prior to creating the stream of ions, and after commencing creation of the ion stream, continuing the injection of electrons in an amount sufficient to maintain the spacecraft at a positive potential. This method may be implemented in a single thruster. In spacecraft with multiple thrusters the same method may be implemented in each thruster.

Where the propulsion system comprises a plurality of thrusters, the method may comprise: operating at least one of the thrusters as a drive thruster, and operating at least one of the thrusters as an auxiliary or reserve thruster. The electron source of the at least one auxiliary thruster may be operated before creation of the ion stream to inject the electrons into the plasma surrounding the spacecraft.

An electric propulsion simulator console (EPSC) which electronically simulates an electric propulsion assembly of a spacecraft as well as propulsion fuel control components and positioning components of the spacecraft. The EPSC simulates a spacecraft thruster electrical interface can test four thruster interfaces simultaneously and continuously. The simulator additionally facilitates the testing of spacecraft fault detection, isolation, and recovery by simulating failed magnet circuits, open anode paths, and flameout conditions. The EPSC includes an electrical propulsion unit load simulator adapted to receive propulsion unit control signals from a spacecraft under test and a spacecraft propulsion unit positioner simulator the simulator adapted to display a simulated state of three axes of movement for at least one propulsion unit positioner responsive to positioning signals received from the spacecraft under test. A propulsion unit fuel valve simulator is also provided and can display a simulated state of propulsion unit fuel valves responsive to control signals received from the spacecraft under test.

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.

An orbiting satellite can be maintained in a virtual orbit, having an orbital period equal to the natural orbit of a satellite at a different altitude, by equipping the satellite with at least one radial thruster. Radial thrusters on the anti-nadir-facing side of the satellite allow for virtual orbits higher than the natural altitude, while radial thrusters on the nadir-facing side of the satellite allow for virtual orbits lower than the natural altitude. This allows a satellite to evade threats, such as orbital debris and/or hostile spacecraft, without losing its relative position within a satellite constellation or experiencing the diminished services often attendant such maneuvers. Similar techniques can also be used for surveillance of orbiting satellites.