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.

Abstract: An aircraft gas turbine engine includes a fan that in operation moves air through both a core airflow path and a bypass airflow path of the gas turbine engine. The core airflow path and the bypass airflow path converge at a jet nozzle of the gas turbine engine. A method of controlling the gas turbine engine includes detecting, at a controller of the gas turbine engine, a cruise operating condition of the gas turbine engine, and in response to detecting the cruise operating condition, operating a mechanism of the gas turbine engine to decrease an effective area of the jet nozzle.

Abstract: An aircraft gas turbine engine includes a fan that in operation moves air through both a core airflow path and a bypass airflow path of the gas turbine engine. The core airflow path and the bypass airflow path converge at a jet nozzle of the gas turbine engine. A method of controlling the gas turbine engine includes detecting, at a controller of the gas turbine engine, a cruise operating condition of the gas turbine engine, and in response to detecting the cruise operating condition, repositioning a plurality of overlap jointed split duct panels such that an effective area of the jet nozzle is reduced.

Abstract: A gas turbine engine for an aircraft includes an outer bypass section wall, and a jet nozzle including at least one inflatable diaphragm. The at least one inflatable diaphragm is disposed along the outer bypass section wall. The gas turbine engine also includes a fluid pressure sensor configured to measure a fluid pressure within the at least one inflatable diaphragm, an inlet valve configured to control a pressurized flow of a fluid into the at least one inflatable diaphragm in response to a command from a controller, and a release valve configured to control a release of the fluid from within the at least one inflatable diaphragm in response to a command from the controller.

Abstract: An aircraft gas turbine engine includes an exhaust nozzle having a nozzle flowpath, and a nacelle of the gas turbine engine at least partially defining an exhaust nozzle flow area of the exhaust nozzle. The nacelle extends axially along an engine central longitudinal axis of the gas turbine engine and circumferentially around the engine central longitudinal axis. A shutter vector apparatus is attached to the nacelle and is operable between a stowed position and a deployed position. The shutter vector apparatus is configured to modulate the exhaust nozzle flow area when the shutter vector apparatus is moved between the stowed position and the deployed position. The shutter vector apparatus includes a plurality of shutter elements, each shutter element having at least one aerodynamic surface extending into the nozzle flowpath of the exhaust nozzle when the shutter vector apparatus is in the deployed position.

Abstract: A seal assembly of a shaft bearing system of an aircraft propeller system includes an inner seal element fixed to a shaft and rotatable therewith about a shaft axis and an outer seal assembly located radially outboard of the inner seal element and rotationally fixed relative to the shaft axis. The outer seal assembly includes an outer seal interactive with the inner seal element at a plurality of knife edges of the inner seal element, and a lip seal having a seal lip interactive with the inner seal element. The outer seal assembly includes a first circumferential seal segment and a second circumferential seal segment abutting the first circumferential seal segment at a seal joint. The lip seal is positioned axially between the outer seal and a bearing of the shaft bearing system.

Abstract: An aircraft propulsion system includes an engine, an electric machine, and a gearbox. The gearbox includes an output shaft with an axis, an electric machine drive gear mounted on the output shaft and driven by the electric machine, and an engine drive gear couplable with the output shaft by a clutch assembly. The clutch assembly includes a first coupling assembly, a second coupling assembly, and a magnetic alignment system. The first coupling assembly includes a first body mounted on the engine drive gear. The second coupling assembly includes a sliding coupling housing mounted on the output shaft and movable between first and second positions. A second body is mounted on the housing. The second body is disengaged from the first in the first position and engaged in the second. The magnetic alignment system includes magnets on the bodies that couple magnetically in the second position.