Abstract: An exhaust system for an aircraft engine, has: a turbine exhaust duct (TED) having an inner wall and an outer wall defining an inlet conduit and outlet conduits; and an exhaust case enclosing the TED and defining exit ports and extending from a forward end to a rearward end and having: a first section extending between the ends, the first section extending from a first edge to a second edge and having a first flange and a second flange at the first and second edges; and a second section extending between the ends and from a third edge to a fourth edge and having a third flange and a fourth flange at the third and fourth edges, wherein the first section is detachably secured to the second section via the first flange being detachably secured to the third flange and the second flange being detachably secured to the fourth flange.

Abstract: A resonator for a combustor is provided. The resonator includes an outer body that partially defines a first chamber and one or more outer apertures configured to convey a first fluid between an outer fluid plenum and the first chamber. The resonator further includes an inner body that partially defines a second chamber and one or more inner apertures configured to convey a second fluid between the second chamber and a combustion chamber. The resonator further includes a flexible membrane that is disposed between the inner body and the outer body and partitioning the first chamber and the second chamber.

Abstract: A rotating detonation engine (RDE) injector and method of creating the same. In some embodiments, the RDE injector is configured such that the fuel to oxidizer pressure ratio is in a range of 0.8775 to 1.4917. In some embodiments of the method of creating a RDE a particular combination of propellants is identified, and their gas specific constant and specific heat ratio are recorded. A plurality of initial flow conditions is established, and the compressible mass flow equation is used to determine an appropriate cross-sectional area of each injector to produce a fuel to oxidizer pressure ratio is in a range of 0.8775 to 1.4917.

Abstract: An aircraft comprising at least one dihydrogen tank, at least one dihydrogen consumer system, a supply line arranged between a consumer system and a tank, for each consumer system, a loop containing a heat transfer fluid, for each tank and each supply line, a supply heat exchanger which exchanges calories between the heat transfer fluid and the dihydrogen and, for each consumer system, a return heat exchanger which exchanges calories between the heat transfer fluid and the consumer system.

Abstract: A gas turbine engine generally defines an axial direction and includes a combustor assembly. The combustor assembly includes a plurality of liner tiles defining at least part of a combustion chamber that extends between a forward end and an aft end generally along the axial direction. A combustor liner support assembly is positioned on a cold side of the plurality of liner tiles and defines a rail extending radially toward the combustion chamber. A retainer assembly includes a bushing where at least one liner tile of the plurality of liner tiles is floatably coupled to the rail via the bushing.

Abstract: A shower device includes a shower device body. The shower device body includes a body water channel, a shower head with spray holes, a storage box for storing cleaning liquid, a control valve, and a one-way valve. The shower head is connected to the shower device body, and the spray holes are fluidly coupled to the body water channel. The storage box includes a liquid inlet and a liquid outlet and is installed in the shower device body. The liquid inlet and the liquid outlet are fluidly coupled to the body water channel respectively. The control valve is installed in the liquid inlet and is configured to control opening and closing of the liquid inlet. The one-way valve is installed in the liquid outlet.

Abstract: The laser beam machine includes: a laser emitting system for emitting a laser beam, wherein the laser emitting system has a protective glass capable of transmitting the laser beam; and a housing attached to the laser emitting system. The housing includes, a first end facing the laser emitting system, a second end opposite to the first end, and a first blowing portion located over an entire circumference of an interior of the first end. The first blowing portion blows a gas toward a central axis of the housing.

Abstract: A combustor assembly for a gas turbine engine includes a combustor defining a combustion chamber. The combustor assembly includes a combustor liner support shell, and a plurality of liner tiles defining at least part of the combustion chamber. Each liner tile of the plurality of liner tiles defines a protrusion and a recess opposite the protrusion where the recess is configured to at least partially receive the protrusion of an adjacent liner tile of the plurality of liner tiles. A fastener assembly is at least partially disposed in the recess for coupling a respective liner tile to the combustor liner support shell.

Abstract: A fuel cell electric power system includes a fuel cell stack, and a charge air system including an intake air compressor, an exhaust turbine, and a flow control valve assembly. The flow control valve assembly includes at least one flow control valve rotatable in a valve housing to fluidly connect an exhaust inlet, selectively, to a turbine feed outlet via a variable flow area to adjust a torque applied to the exhaust turbine by a flow of exhaust.

Abstract: A plasma thruster has a tunable electron source configured to provide electrons with controllable energy. An entry electrode and an exit electrode permit an air flow to pass from the entry electrode to the exit electrode. The entry electrode and exit electrode receive electrons from the tunable electron source. A controller selectively controls the entry and exit electrodes to accelerate positive and negative ions in the air.

Abstract: An aircraft propulsion system includes a cryogenic fuel system and a fuel flow path where a fuel flow is circulated from upstream in a direction downstream to a combustor of the core engine. A first heat exchanger inputs heat into the fuel flow downstream of the first heat exchanger is used to heat the fuel flow from the fuel storage tank and a second heat exchanger inputs thermal energy from a heat source into the fuel flow from the core engine downstream of the first heat exchanger. A turboexpander is configured to generate shaft power from expansion of the fuel flow exhausted downstream from at least one of the first heat exchanger and the second heat exchanger.

Abstract: The invention relates to an aircraft turbine engine (101) which comprises, downstream of the low pressure compressor of its low pressure body (104) and upstream of its combustion chamber (102), an electric motor (206) which is configured to rotate a rotor blade ring so as to generate a flow of air. Moreover, the rotational speed of the electric machine rotor is independent of the rotational speed of the compressor rotor.

Abstract: A gas turbine engine includes a compressor section, a combustion section, and a turbine section arranged in serial flow order and defining a working gas flowpath, a compressor of the compressor section comprising an aft-most compressor stage; a stage of stator vanes located downstream of the aft-most compressor stage; a stator case including a seal pad; and a spool drivingly coupled to the compressor, the spool and the stator case together defining a rotor cavity in fluid communication with the working gas flowpath, the spool comprising a seal tooth assembly, the seal tooth assembly including a seal support extension, a seal tooth extending from the seal support extension toward the seal pad, and a dampener operable with the seal support extension.

Abstract: A fuel system can include a first fuel circuit, a second fuel circuit, and an inert gas purge system operatively connected to both the first fuel circuit and the second fuel circuit to purge at least a portion of either or both of the first and/or second fuel circuit. The first fuel can be a liquid fuel and the second fuel can be a gaseous fuel. The first fuel circuit can include a first fuel manifold configured to fluidly communicate a first fuel supply with at least one dual fuel nozzles downstream of the first fuel manifold.

Abstract: A method for energy conversion for a vehicle is provided. The method including extracting a flow of compressed fluid from a compressor section of a propulsion system; flowing the flow of compressed fluid to a turbine operably coupled to a driveshaft, in which the driveshaft is operably coupled to a load device; expanding the flow of compressed fluid through the turbine to generate an output torque at the driveshaft to operate the load device; and flowing the expanded flow of compressed fluid from the turbine to thermal communication with a thermal load.

Abstract: A nose cone for debris deflection comprising a deflector formed in the nose cone, and an actuator in operative communication with the deflector, the actuator configured to engage the deflector configured to create a deflection air stream; the deflection air stream being configured to manipulate a fan inlet air flow such that any debris entrained in the fan inlet air flow has a trajectory line directed away from a core flow of a gas turbine engine and toward a bypass flow of the gas turbine engine.

Abstract: An assembly is provided for an aircraft. This aircraft assembly includes a first propulsor rotor, a geartrain, a rotating assembly and an auxiliary turbine. The rotating assembly is rotatable about an axis and includes a turbine rotor. The rotating assembly is coupled to and is configured to drive rotation of the first propulsor rotor through the geartrain. The auxiliary turbine is coupled to the first propulsor rotor independent of the geartrain.

Abstract: What is provided is a metal ion recovery device including: a raw solution tank that is configured to store a metal ion containing raw solution including metal ions; a recovery liquid tank that is configured to store a metal ion recovery liquid including metal ions recovered from the metal ion containing raw solution; a metal ion selective permeable membrane that separates the raw solution tank and the recovery liquid tank and selectively transmits the metal ions; an anode that is arranged on a side of the metal ion selective permeable membrane close to the raw solution tank; a cathode that is arranged on a side of the metal ion selective permeable membrane close to the recovery liquid tank; a porous current collector that is formed of a conductive material; a first spacer that is configured to maintain a gap between the metal ion selective permeable membrane and the anode; and a second spacer that is configured to maintain a gap between the metal ion selective permeable membrane and the cathode, in which the

Abstract: An inlet scroll for a combustor includes an inlet, an annular outlet, a volute duct, and a transition. The volute extends from the inlet to circumscribe an axis to form a monotonically decreasing cross-sectional flow area from the inlet. The transition fluidly connects a radially inner boundary of the volute duct to the annular outlet. The transition includes a radial portion at the volute duct and an axial portion at the annular outlet.

Abstract: A combustor system for an aerial vehicle includes an outer liner with a first end and a second end, an inner combustor case contained within the outer liner and defining a central combustor chamber, a lumen extending into the central combustor chamber, an air flow path extending between the outer liner and the inner combustor case, an integrated air scoop attached to the inner combustor case at the lumen and directing air from the air flow path into the central combustor chamber, and a fuel bolt, the fuel bolt including a head located outside of the outer liner and a body starting outside of the outer liner and extending perpendicular to the outer liner through the outer liner, through the integrated fuel scoop, and ending at the lumen, and fuel manifold with a fuel orifice within the fuel bolt.