Abstract: Some embodiments of the invention include a thruster system comprising a thruster and a pulsing power supply. The thruster may include a gas inlet port; a plasma jet outlet; and a first electrode. In some embodiments, the pulsing power supply may provide an electrical potential to the first electrode with a pulse repetition frequency greater than 10 kHz, a voltage greater than 5 kilovolts. In some embodiments, the pressure downstream from the thruster can be less than 10 Torr. In some embodiments, when a plasma is produced within the thruster by energizing a gas flowing into the thruster through the gas inlet port, the plasma is expelled from the thruster through the plasma jet outlet.

Abstract: The present invention discloses an electric field propulsion system for spacecraft. The system includes a capacitor stack comprising an array of supercapacitors. Solid-state electronic circuits generate modulated currents and electric fields in pulse coils. The pulse coils direct the electric fields onto separated electric charges stored in the capacitor stack. The resulting unidirectional Lorentz Forces thereby generate thrust without reaction mass. Reaction momentum is carried away by Poynting Vector fields in conformity with the currently understood principles of electrodynamics. The design is scalable down to micro-chip sized thrusters.

Abstract: 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.

Abstract: An inductive plasma acceleration apparatus, comprising a pulse laser assembly, a pulsed discharge assembly, an exciting coil assembly, a solid-state working medium, and a control assembly; the exciting coil assembly is electrically connected to the pulsed discharge assembly such that a strong pulse current is produced in the exciting coil assembly during the discharge process of the pulse discharge assembly, and an inductive pulse electromagnetic field is excited around the exciting coil assembly; the solid-state working medium is positioned on the optical path of a pulse laser emitted by the pulse laser assembly such that the solid-state working medium produces a pulse gas under the ablation action of the pulse laser, and the inductive pulse electromagnetic field is positioned on the circulation gas path of the pulse gas such that the pulse gas can enter the inductive pulse electromagnetic field.

Abstract: A method for quickly stopping the propulsion rotor of a helicopter after landing, comprising, following a request for quickly stopping the engine by a helicopter pilot, the following steps managed by the control unit of the turbomachine: Detecting the absence of the thermal stabilization phase of the gas generator of at least one turbomachine, controlling an extinction of the combustion chamber of the gas generator of at least one turbomachine, maintaining the rotation of the gas generator of which the combustion chamber is extinguished by means of said at least one electrical machine to ventilate the gas generator and stopping the main rotor of the helicopter by means of a mechanical brake.

Abstract: A rectifying device includes an air flow generator. The air flow generator is disposed at an exterior member of a vehicle. The exterior member is adjacent to a detector of a sensor that is disposed such that at least a portion of a detection range of the detector includes a rear region behind a plane in a traveling direction of the vehicle. The plane is parallel to a width direction and a vertical direction of the vehicle. The air flow generator is configured to generate an air flow that separates, from the detector of the sensor, travelling wind that accompanies travel of the vehicle. The air flow generator includes a plasma actuator that includes at least a pair of electrodes and a power source that is configured to apply an alternating current voltage to the electrodes.

Abstract: 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.

Abstract: According to certain aspects, an electric-propulsion thruster is used as part of a base or platform which also includes a power converter, having a plurality of inductors and other electrical components, and a printed circuit board (PCB). The PCB includes a layer at which the other electrical components and printed circuit inductor traces, for the plurality of inductors, are secured. The electric-propulsion thruster includes a housing (e.g., as part of the base or platform) providing a cavity and having at least one structurally-rigid side wall along the cavity, where the PCB is integrated with the electric-propulsion thruster for a compact arrangement which can be used to propel the apparatus. Such a compact design might be used as an important part of thruster spacecraft architecture such as micro-satellites (e.g., CubeSats).

Abstract: The present disclosure provides methods and systems for determining a synthesized engine parameter of a gas turbine engine. An initial model parameter is obtained from an onboard model associated with the gas turbine engine. A correction factor for the onboard model is determined by modifying a difference between the onboard model and an aero-thermal model of the gas turbine engine using first and second engine parameters and first and second operating conditions, wherein the first and second engine parameters are independent from one another over an operating envelope of the gas turbine engine. The initial model parameter is scaled by applying the correction factor thereto to obtain a corrected model parameter. The corrected model parameter is output as the synthesized engine parameter.

Abstract: 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.

Abstract: The present disclosure relates to a system for supplying fuel to a turbomachine. In some embodiments, the system includes a pump, a hydromechanical group, a drive device positioned to drive the pump, and a branch comprising actuators for controlling variable geometry. In some embodiments, the system regulates a fuel flow rate based at least in part on a flow rate set point value. In some embodiments, the system may include a flow rate sensor. A flow rate loop may be arranged to determine a pressure set point value at the outlet of the pump according to the flow rate set point value and a measurement supplied by the flow rate sensor. In some embodiments, a pressure sensor may be positioned in the fuel circuit at the outlet of the pump. In further embodiments, a pressure loop may control the speed of the drive device based at least on a difference between a pressure measurement supplied by the pressure sensor and the pressure set point value.

Abstract: A field emission propulsion system for a spacecraft includes a control unit, a propulsion assembly, and a plurality of extractor electrode voltage sources. The propulsion assembly comprises a plurality of field emission propulsion units having an ion source with a plurality of ion emitters and extractor electrodes associated with the ion emitters and disposed in a field arrangement. The plurality of extractor electrode voltage sources, each associated with the extractor electrodes to control the same, are controlled by the control unit using an individual extractor electrode voltage.

Abstract: The present disclosure provides a cooling structure of heat pipe for superconducting magneto plasma dynamic thruster having a cylindrical structure and includes a cathode, an intermediate connector and an anode. The cathode is arranged inside the intermediate connector, the anode is arranged outside the intermediate connector; the cathode is provided with a cathode cooling mechanism, and the anode is provided with an anode cooling mechanism. The cathode cooling mechanism includes a cathode heat pipe and a cathode heat dissipation fin. The anode heat pipe cooling mechanism includes an anode heat pipe and an anode heat dissipation fin.

Abstract: Disclosed is a closed drift, narrow channel Hall thruster configured to operate at powers <30 W. The thruster includes a thruster body and a neutralizing cathode. The thruster body includes a magnetic circuit including a magnetic source and two magnetic poles, a metallic, annular thruster channel formed by the magnetic poles with a downstream channel width smaller than about 3 mm and an upstream channel width greater than the downstream channel width, an anode positioned at the channel's entry, and a gas distributor configured to release a propellant gas into the thruster channel. The magnetic circuit is configured to generate a magnetic field in the thruster channel for trapping electrons therein. The channel walls (the magnetic poles) are under bias potential. The anode and the cathode are configured to generate a substantially axial electric field in the thruster channel.

Abstract: A controller may use energy packets to control a prime mover of a machine. The controller may include an energy packet measurement control to calculate energy packets, perform post-processing actions on the energy packets to generate processed energy packets, and convert the processed energy packets into a fuel valve reference. Post-processing may include a calibration correction to remove measurement artifacts.

Abstract: A gas turbine includes: multiple fuel systems; a combustor that combusts fuels from the multiple fuel systems in compressed air to generate combustion gas; and a turbine that is driven by the combustion gas. A flow volume ratio calculation device which calculates the flow volume ratio of the fuels flowing in the multiple fuel systems includes calculators that receive values of a first parameter and a second parameter capable of expressing the combustion state in the combustor, and that calculate the flow volume ratio relative to the received values of the two parameters from a predetermined relationship between the two parameters and the flow volume ratio.

Abstract: Systems and method for filling a fuel manifold comprising at least a primary and a second manifold of a gas turbine engine are described. The method comprises providing fuel flow to the secondary manifold of the gas turbine engine, the secondary manifold being partly or completely empty; monitoring at least one engine operational parameter of the gas turbine engine as fuel fills the secondary manifold; and accelerating the engine when a transition threshold is reached, the transition threshold being associated with the engine operational parameter and indicative that fuel has reached the combustor.

Abstract: A control system of a gas turbine engine is provided. The engine has a fuel flow metering valve which regulates a fuel flow to the engine, and one or more variable geometry components which are movable between different set points to vary an operating configuration of the engine. The control system has an engine fuel control sub-system which provides a fuel flow demand signal for controlling the fuel flow metering valve. The control system further has a variable geometry control sub-system which determines current set points to be adopted by the variable geometry components given the current engine operating condition in order to comply with one or more engine constraints. The control system further has an optimiser that receives the current set points and determines adjusted values of the set points which optimise, while complying with the engine constraints, an objective function modelling a performance characteristic of the engine, the objective function adapting to change in engine performance with time.

Abstract: Electrical propulsion systems and related methods are generally described. In some embodiments, an electrical propulsion system may include an electrically-actuated valve to selectively permit flow of propellant from a reservoir tank to a thruster. The valve may physically isolate the propellant from the thruster when inactivated, exhibiting a non-wetting surface which may inhibit propellant from passing through the valve towards the thrusters. In some embodiments, a valve may be activated through application of a voltage potential to the valve relative to the propellant, which may change the wettability of the valve, permitting propellant to wet and subsequently pass through the valve. The voltage potential may be adjusted to vary the wettability of the valve, resulting in the valve effectively regulating propellant flow rate. The valve may include a conductive layer, a dielectric or insulating layer, and a non-wetting layer to enhance the non-wetting behavior of the valve.

Abstract: A diffuser assembly includes a casing at a compressor aft end; an inner barrel member radially inward of the casing; and an array of radial flow splitters extending between the inner barrel member and the casing. Each radial flow splitter includes a leading edge facing into a flow of air, a trailing end wall opposite the leading edge, a pair of side walls extending between the leading edge and the trailing end wall, and an axis extending through the leading edge and the trailing end wall. A width of each radial flow splitter increases from the leading edge to the trailing end wall. The side walls diverge away from the axis in a downstream direction corresponding to the flow of air. Optionally, the side walls also diverge away from the axis in a radial direction between the inner barrel member and the casing.