Abstract: A hybrid-electric propulsion (HEP) system for powering an aircraft is provided. The system includes a gearbox, a gas turbine engine, and an electric motor. The gearbox includes an input shaft, an output shaft, a drive gear, a first layshaft assembly, a second layshaft assembly, and a first gearbox accessory mount. The gas turbine engine is in drive communication with the input shaft of the gearbox. The electric motor is engaged with the first gearbox accessory mount. The first layshaft (FL) assembly includes a first layshaft, a first FL gear, and a second FL gear. The first and second FL gears are attached to the first layshaft. The first FL gear is engaged with the input shaft and the second FL gear is engaged with the drive gear. The first gearbox accessory mount is in geared communication with the drive gear independent of the first layshaft assembly.

Abstract: An aircraft propulsion system includes an engine, an electric machine, a gearbox, and a propulsor. The gearbox includes an output shaft with a rotational axis, an electric machine drive gear mounted on the output shaft and driven by the electric machine, and an engine drive gear selectively couplable with the output shaft by a clutch assembly. The clutch assembly includes a coupling drive shaft mounted to the output shaft, a clutch housing mounted to the engine drive gear, and sliding clutch members moveably mounted to the clutch housing. The sliding clutch members are radially movable between inner and outer positions. In the outer position, the sliding clutch members engage the coupling drive shaft to couple the engine and output shaft. In the inner position, the sliding clutch members disengage from the coupling drive shaft to decouple the engine from the output shaft. The propulsor is driven by the output shaft.

Abstract: A system is provided for an aircraft. This aircraft system includes an engine, a fuel system and an air system. The fuel system includes a fuel circuit, a fuel source, a heat exchanger and a turboexpander. The fuel circuit is configured to direct fuel from the fuel source, through the heat exchanger and the turboexpander, to the engine for combustion. The heat exchanger is configured to transfer heat energy from combustion products generated by the combustion of the fuel in the engine into the fuel in the fuel circuit. The air system is discrete from the engine. The air system includes an air compressor and the turboexpander. The air system is configured to expand the fuel directed across the turboexpander by the fuel circuit to power operation of the air compressor.

Abstract: The hybrid aircraft power plant can have a load, a thermal engine coupled to drive the load via a clutch, and coupled to drive a first lubricant pump, an electric motor coupled to drive the load and a second lubricant pump. In a first mode, the clutch is in an engaged configuration, and the thermal engine drives the load via the clutch and drives the first lubricant pump. In a second mode, in which the clutch is in the disengaged configuration and in which the thermal engine may be inoperative, the electric motor is operated to drive the load and the second lubricant pump.

Abstract: A propulsion system for an aircraft is provided. The propulsion system includes an engine including a first input shaft and a drivetrain coupled to the first input shaft. The drivetrain includes an output shaft, a gear assembly, an electric motor, and a second input shaft. The gear assembly includes a main gear mounted to the output shaft, and a bevel ring gear integrated on a face of the main gear. The main gear, the bevel ring gear, and the output shaft are rotatable about a first rotational axis. The electric motor includes a rotor rotatable about a second rotational axis. The second rotational axis is perpendicular to the first rotational axis. The second input shaft couples the rotor to the bevel ring gear. The second input shaft is rotatable about the second rotational axis.

Abstract: A heat exchanger assembly for an aircraft includes a housing, a heat exchanger, and a nozzle. The housing includes an inner flow path wall and an outer flow path wall. The inner flow path wall and the outer flow path wall form an air inlet, an air outlet, and a venturi channel. The heat exchanger includes a heat exchanger body, a heat transfer plenum, a plenum inlet, and a plenum outlet. The heat exchanger body is disposed downstream of the venturi channel within the air flow path. The heat exchanger is configured to direct a fluid from through the heat transfer plenum from the plenum inlet to the plenum outlet. The nozzle includes a nozzle housing disposed within the air flow path. The nozzle housing forms a nozzle inlet and a nozzle outlet. The nozzle outlet is disposed within the venturi channel.

Abstract: A gas turbine engine is provided that includes an axial centerline, a compressor section, a combustor, a turbine section, and engine shaft, and a shaft retention system. The engine shaft is engaged with the compressor and turbine sections. The engine shaft has an aft end and axially extends along the engine axial centerline. The shaft retention system includes a shaft retainer cap and a retainer piston. The shaft retainer cap is attached to the aft end of the engine shaft. The retainer piston is mounted for axial translation. The shaft retention system is disposable in an engaged configuration wherein the retainer piston is coupled with the shaft retainer cap, and in a disengaged configuration wherein the retainer piston is disengaged with the shaft retainer cap.

Abstract: A fuel system is configured to deliver a first fuel and/or a second fuel to a combustor section. The fuel system includes a first fuel source, a second fuel source, a first control valve, a second control valve, a metering valve and a fuel injector. The first fuel source is configured to crack ammonia into hydrogen gas and nitrogen gas and subsequently output the first fuel. The second fuel source is configured to output the second fuel. The metering valve is configured to meter a flow of the first fuel received from the first fuel source through the first control valve to the fuel injector during a first operating mode. The metering valve is configured to meter a flow of the second fuel received from the second fuel source through the second control valve to the fuel injector during a second operating mode.

Abstract: A fuel system includes a reservoir, a heat exchanger, a heating device, a cracker and a fuel injector. The heat exchanger is configured to vaporize a first flow of liquid ammonia received from the reservoir into ammonia gas using heat energy. The heat exchanger is configured to receive the heat energy from the heating device during a first operating mode. The heat exchanger is configured to receive the heat energy from a turbine engine a second operating mode. The cracker configured to at least partially crack the ammonia gas into hydrogen gas and nitrogen gas. The fuel system is configured to deliver the hydrogen gas or a combination of the hydrogen gas and the nitrogen gas to a fuel injector for injection into a combustion chamber.

Abstract: A hybrid-electric propulsion system for an aircraft is provided that includes a propulsion unit, a propulsion unit gear box, a gas turbine engine, and an electric motor drive system. The gas turbine engine has compressor, combustion, and turbine sections, and is configured to provide motive force to the propulsion unit gear box. The electric motor drive system is disposed to provide motive force to the propulsion gear box or to the gas turbine engine. The electric motor drive system (EMDS) is adapted to interchangeably use a plurality of EMDS packages, each EMDS package including an electric motor and a battery. The electric motor drive system has a set of performance parameters that vary depending upon the particular EMDS package that is used with the electric motor drive system. The electric motor drive system includes interfaces that allow each EMDS package to be interchangeably used.

Abstract: An aircraft system is provided that includes a compressor section, a combustor section, a turbine section, an exhaust duct and a flowpath. The exhaust duct includes an upstream duct section, a plurality of intermediate duct sections and a downstream duct section. A first of the intermediate duct sections is fluidly discrete from a second of the intermediate duct sections. The intermediate duct sections are fluidly coupled in parallel between the upstream duct section and the downstream duct section. The flowpath extends sequentially longitudinally through the compressor section, the combustor section, the turbine section, the upstream duct section, the intermediate duct sections and the downstream duct section from an inlet into the flowpath to an outlet from the flowpath.

Abstract: An aircraft system is provided that includes a compressor section, a combustor section, a turbine section, an exhaust duct and a flowpath. The exhaust duct includes an upstream duct section, a plurality of intermediate duct sections and a downstream duct section. A first of the intermediate duct sections is fluidly discrete from a second of the intermediate duct sections. The intermediate duct sections are fluidly coupled in parallel between the upstream duct section and the downstream duct section. The flowpath extends sequentially longitudinally through the compressor section, the combustor section, the turbine section, the upstream duct section, the intermediate duct sections and the downstream duct section from an inlet into the flowpath to an outlet from the flowpath.

Abstract: A hybrid aircraft power plant includes an air mover for propelling an aircraft, a gas turbine engine operable to drive the air mover, and an electric machine operable as a motor to drive the air mover. The gas turbine engine includes a turbine disk for extracting energy from a stream of combustion gas and an exhaust duct defining: an exhaust passage for conveying the stream of combustion gas from the turbine disk; and a cavity disposed radially inward of the exhaust passage. A shaft drivingly connects the turbine disk with the electric machine and extends inside the cavity defined by the exhaust duct. A fan is drivingly connected to the shaft and operable to circulate cooling air inside the cavity of the exhaust duct.

Abstract: A propulsion system is provided for an aircraft. This propulsion system includes a geartrain, a propulsor rotor, a turbine engine core, an electric machine and an accessory gearbox system. The propulsor rotor is disposed to a first side of the geartrain. The turbine engine core is disposed to a second side of the geartrain. The turbine engine core is operatively coupled to the propulsor rotor through the geartrain. The electric machine is disposed to the second side of the geartrain. The electric machine is operatively coupled to the propulsor rotor and/or the turbine engine core through the geartrain. The accessory gearbox system is operatively coupled to the turbine engine core. The accessory gearbox system is circumferentially offset from the electric machine about a centerline axis of the turbine engine core. The accessory gearbox system axially overlaps the electric machine along the centerline axis of the turbine engine core.

Abstract: A gas turbine engine is provided that includes a compressor section, a combustor, a turbine section, an engine shaft, a shaft retention system, and an axial centerline. The engine shaft is engaged with the compressor section and the turbine section. The engine shaft has an aft end and axially extends along the axial centerline of the engine. The shaft retention system includes a shaft nut and a shaft retainer. The shaft nut is attached to the aft end of the engine shaft. The shaft retainer is mounted for axial translation normal to the axial centerline. The shaft retainer is disposable in an engaged configuration wherein the shaft retainer is coupled with the shaft nut, and in a disengaged configuration wherein the shaft retainer is disengaged with the shaft nut.

Abstract: An assembly for a propulsion system of an aircraft includes an electric motor, a first gearbox module, a second gearbox module, and a propeller. The electric motor includes a rotor. The rotor includes a first axial end and a second axial end. The first gearbox module includes a first gear assembly. The first gear assembly is coupled to the first axial end. The second gearbox module includes a second gear assembly. The second gear assembly is coupled to the second axial end. The propeller is coupled to the first gear assembly. The first gear assembly is configured to drive rotation of the propeller in response to rotation of the rotor.

Abstract: An assembly for an aircraft powerplant includes a compressor rotor and an air system. The compressor rotor includes a plurality of compressor blades. The compressor blades are arranged circumferentially around the axis in a compressor blade array and are disposed in a flowpath. The air system includes an air source, an air injector and an air circuit fluidly coupling the air source to the air injector. The air system is configured to direct compressed air from the air source to the air injector through the air circuit. The air injector is configured to direct a stream of the compressed air into the flowpath and against one or more of the compressor blades to apply a rotational driving force onto the compressor rotor about an axis. The flowpath has a radial height at an injector location. The air injector projects a radial distance into the flowpath at the injector location.

Abstract: A method is provided for operating a powerplant. During this method, a hydrocarbon fuel is delivered to a turbine engine for combustion within a combustion chamber of the turbine engine during initial startup operation of the turbine engine. Ammonia is at least partially cracked into hydrogen gas and nitrogen gas. A non-hydrocarbon fuel is delivered to the turbine engine for combustion within the combustion chamber of the turbine engine during post-startup operation of the turbine engine. The non-hydrocarbon fuel is or otherwise includes the hydrogen gas or a combination of the hydrogen gas and the nitrogen gas.

Abstract: An assembly for an aircraft powerplant includes a compressor rotor and an air system. The compressor rotor includes a plurality of compressor blades. The compressor blades are arranged circumferentially around the axis in a compressor blade array and are disposed in a flowpath. The air system includes an air source, an air injector and an air circuit fluidly coupling the air source to the air injector. The air system is configured to direct compressed air from the air source to the air injector through the air circuit. The air injector is configured to direct a stream of the compressed air into the flowpath and against one or more of the compressor blades to apply a rotational driving force onto the compressor rotor about an axis. The flowpath has a radial height at an injector location. The air injector projects a radial distance into the flowpath at the injector location.

Abstract: A gas turbine engine for an aircraft includes a jet nozzle. The jet nozzle includes an upper split duct panel and a lower split duct panel coupled with the upper split duct panel via a first overlap joint and a second overlap joint. The jet nozzle also includes at least one actuator configured to receive a command signal from a controller, and based on the command signal, reposition the upper split duct panel and lower split duct panel such that an effective area of the jet nozzle is adjusted.