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: A turbine engine can include a compressor section, a combustion section, and a turbine section in serial flow arrangement. The combustion section can have a combustor with a combustor liner at least partially defining a combustion chamber, as well as a fuel-air mixer including at least one set of mixing tubes.

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: A combustion chamber, in particular for a gas turbine, includes a support structure, a plurality of retaining elements fastened to the support structure, and a plurality of heat shield elements which jointly form a heat shield and which each have a hot gas side, a cold gas side and end faces which interconnect the hot gas side and the cold gas side, the retaining elements interlockingly engaging in recesses in the heat shield elements. The retaining elements each have at least two engagement portions for interlockingly engaging in the recesses in a heat shield element, which engagement portions are interconnected in a tensionally rigid manner and are tensionally rigid themselves. Spring elements extend between the support structure and the heat shield elements, which spring elements are designed in particular as leaf springs and effect a frictional connection between the engagement portions of the retaining elements and the heat shield elements.

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: 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: A gas turbine engine includes a fan shaft diving a fan having fan blades. A gear system includes a sun gear surrounded by a plurality of intermediate gears. A carrier at least partially supports the plurality of intermediate gears. A ring gear surrounds the plurality of intermediate gears. The sun gear is driven by a turbine. At least one fan shaft support bearing is located forward of the gear system. A coupling fixes the ring gear from rotation relative to an engine static structure. A lubrication system lubricates components across a rotation gap. The lubrication system includes a lubricant input. A stationary first bearing receives lubricant from the lubricant input and has a first race in which lubricant flows. A second bearing rotates within the first bearing. The second bearing has a first opening in registration with the first race such that lubricant may flow from the first race through the first opening into a first conduit.

Abstract: A fuel nozzle for a combustor may include a primary fuel passage having a primary fuel outlet. A secondary fuel passage having a secondary fuel outlet. A valve located within the secondary fuel passage. The valve may include a stationary hub having a plurality of first angled surfaces. The valve may include a movable sleeve having a plurality of second angled surfaces. Each of the plurality of first angled surfaces aligns axially with a respective one of the plurality of second angled surfaces to define a slot therebetween. The valve is movable between a plurality of positions to open, to close, or to partially open the slots.

Abstract: A gas turbine engine includes a propulsor with a power turbine, a power turbine shaft extending forward therefrom defining a centerline axis, and a fan driven by the power turbine shaft. The fan is aligned with the centerline axis forward of the power turbine and is operatively connected to be driven by the power turbine through the power turbine shaft. A gas generator operatively connected to the propulsor is included downstream from the fan and forward of the power turbine, wherein the gas generator defines a generator axis offset from the centerline axis. The gas generator is operatively connected to the power turbine to supply combustion products for driving the power turbine.

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: 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 fluid supercharging device (3, 5), comprising: a rotating shaft (7); a vane disc (308) coaxially fixed to the rotating shaft (7); a plurality of fan blades (301) fixed around a perimeter of the vane disc (308); the back side of the fan blades 301 being provided with at least one fluid guiding inlet (305), an end of the back side distal from the vane disc (308) is provided with a fluid guiding outlet (306, 307), a fluid channel (304) communicating the fluid guiding inlet (305) with the fluid guiding outlet (306, 307) is provided along a lengthwise direction inside the fan blades; the fan blades (301) rotate to generate a centrifugal force such that a fluid flows into the fluid channel via the fluid guiding inlet on the back side, and flows out of the fluid guiding outlet along the lengthwise direction of the fan blades.

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.