Abstract: An assembly for a propulsion system of an aircraft includes a gearbox module, an electric motor assembly, a first accessory load assembly, and a propulsor. The gearbox module includes a gear assembly and an output shaft. The gear assembly includes a main gear and a plurality of offset gears. The electric motor assembly includes an electric motor. The electric motor includes a rotor. The rotor is mounted to a first offset gear of the plurality of offset gears. The rotor is configured for rotation about a rotational axis to drive rotation of the gear assembly and the output shaft. The first accessory load assembly includes at least one first accessory load. The at least one first accessory load is mounted to a second offset gear of the plurality of offset gears. The propulsor is coupled to the output shaft.

Abstract: An assembly for a propulsion system of an aircraft includes a gearbox, at least one accessory load assembly, a propulsor, and an electric motor. The gearbox module includes a gear assembly and an output shaft. The gear assembly is connected to the output shaft. The at least one accessory load assembly includes at least one accessory load coupled to the output shaft. The propulsor is coupled to the output shaft. The electric motor includes a rotor. The rotor is coupled to the gear assembly to drive rotation of the output shaft about a rotational axis. The rotation of the output shaft drives rotation of the propulsor and the at least one accessory load for each accessory load assembly of the at least one accessory load assembly.

Abstract: An aircraft system is provided that includes a thrust rotor and a powerplant coupled to and configured to drive rotation of the thrust rotor. The powerplant includes a gas turbine engine, a drivetrain and a fluid motor. The gas turbine engine includes a rotating assembly, a compressor section, a combustor section, a turbine section and a flowpath extending through the compressor section, the combustor section and the turbine section. The rotating assembly includes a turbine rotor in the turbine section. The turbine rotor is configured to convert fluid power of combustion products flowing through the flowpath within the turbine section into mechanical power for driving rotation of the rotating assembly during a mode of engine operation. The drivetrain is configured to rotatably couple the fluid motor to the rotating assembly. The fluid motor is configured to boost the mechanical power for driving the rotation of the rotating assembly during the mode of engine operation.

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 aircraft system is provided that includes a thrust rotor and a powerplant coupled to and configured to drive rotation of the thrust rotor. The powerplant includes a gas turbine engine, a drivetrain and a fluid motor. The gas turbine engine includes a rotating assembly, a compressor section, a combustor section, a turbine section and a flowpath extending through the compressor section, the combustor section and the turbine section. The rotating assembly includes a turbine rotor in the turbine section. The turbine rotor is configured to convert fluid power of combustion products flowing through the flowpath within the turbine section into mechanical power for driving rotation of the rotating assembly during a mode of engine operation. The drivetrain is configured to rotatably couple the fluid motor to the rotating assembly. The fluid motor is configured to boost the mechanical power for driving the rotation of the rotating assembly during the mode of engine operation.

Abstract: An aircraft system is provided that includes an open propulsor rotor, a gas turbine engine, an electric machine and a drivetrain. The gas turbine engine includes a first rotating assembly, a compressor section, a combustor section, a turbine section and a flowpath extending through the compressor section, the combustor section and the turbine section. The first rotating assembly is coupled to and is configured to drive rotation of the open propulsor rotor. The first rotating assembly includes a first turbine rotor in the turbine section. The electric machine is outside of the gas turbine engine. The electric machine includes an electric machine rotor. The drivetrain couples the electric machine rotor to the first rotating assembly.

Abstract: An aircraft system is provided that includes a thrust rotor and a powerplant coupled to and configured to drive rotation of the thrust rotor. The powerplant includes a gas turbine engine, a drivetrain and a fluid motor. The gas turbine engine includes a rotating assembly, a compressor section, a combustor section, a turbine section and a flowpath extending through the compressor section, the combustor section and the turbine section. The rotating assembly includes a turbine rotor in the turbine section. The turbine rotor is configured to convert fluid power of combustion products flowing through the flowpath within the turbine section into mechanical power for driving rotation of the rotating assembly during a mode of engine operation. The drivetrain is configured to rotatably couple the fluid motor to the rotating assembly. The fluid motor is configured to boost the mechanical power for driving the rotation of the rotating assembly during the mode of engine operation.

Abstract: An exhaust nozzle assembly for a propulsion system include a primary nozzle, an outer shroud, an ejector nozzle, and an actuator. The primary nozzle extends along an exhaust centerline. The primary nozzle includes a downstream axial end. The outer shroud surrounds the primary nozzle. The ejector nozzle extends axially between a first axial end and a second axial end. The second axial end forms a nozzle exit plane for the exhaust nozzle assembly. The ejector nozzle converges in a direction from the first axial end to the second axial end. The ejector nozzle forms a mixing cross-sectional area between the primary nozzle and the ejector nozzle at the downstream axial end. The actuator is mounted on the ejector nozzle. The actuator is configured to move the ejector nozzle between a first position and a second position, relative to the outer shroud, to control an area of the mixing cross-sectional area.

Abstract: The aircraft can include a rotary airfoil device having an output shaft rotatable around an airfoil rotation axis; an electric engine having a source shaft drivingly coupled to the output shaft; a controller in communication with the electric engine; a low voltage electric system operable to power the controller; a high voltage electric system having a high voltage battery operable to power the electric engine, the high voltage battery further operable to power the low voltage electric system via a voltage converter; and an electromagnetic generator having a generator shaft coupled to the output shaft, the electromagnetic generator operable to power the low voltage electric system.

Abstract: An exhaust nozzle assembly for a propulsion system include a primary nozzle, an outer shroud, an ejector nozzle, and an actuator. The primary nozzle extends along an exhaust centerline. The primary nozzle includes a downstream axial end. The outer shroud surrounds the primary nozzle. The ejector nozzle extends axially between a first axial end and a second axial end. The second axial end forms a nozzle exit plane for the exhaust nozzle assembly. The ejector nozzle converges in a direction from the first axial end to the second axial end. The ejector nozzle forms a mixing cross-sectional area between the primary nozzle and the ejector nozzle at the downstream axial end. The actuator is mounted on the ejector nozzle. The actuator is configured to move the ejector nozzle between a first position and a second position, relative to the outer shroud, to control an area of the mixing cross-sectional area.

Abstract: There is provided an exhaust mixer arrangement for a turbofan engine having a bypass passage for channelling a bypass flow and a core passage for channelling a core flow around a central axis. The exhaust mixer arrangement comprises a mixer body mounted for rotation about the central axis. The mixer body has an annular wall extending around the central axis. The annular wall defines a plurality of circumferentially distributed alternating inner and outer lobes, with each inner lobe protruding into the core passage, and each outer lobe protruding into the annular bypass passage. A driving unit is operatively connected to the mixer body for selectively driving the mixer body in rotation about the central axis. A controller is operatively connected to the driving unit for controlling a rotational speed of the mixer body as a function of a flight operating condition.

Abstract: A method and a system for synthesizing output power provided by an engine are provided. The engine comprising a compressor section, a combustor, and a turbine section in serial fluid flow communication. The engine is operated and, during the operating of the engine, a pressure of fluid at an exit of the compressor section, a temperature upstream of the exit of the compressor section, and a fuel flow rate to the engine are determined. A synthesized value of output power provided by the engine is determined based on a product of at least a first factor, a second factor, and a third factor, the first factor being a function of the pressure, the second factor being a function of the temperature, and the third factor being a function of the fuel flow rate. The synthesized value of output power provided by the engine is output.

Abstract: A method and a system for synthesizing output power provided by an engine are provided. The engine comprising a compressor section, a combustor, and a turbine section in serial fluid flow communication. The engine is operated and, during the operating of the engine, a pressure of fluid at an exit of the compressor section, a temperature upstream of the exit of the compressor section, and a fuel flow rate to the engine are determined. A synthesized value of output power provided by the engine is determined based on a product of at least a first factor, a second factor, and a third factor, the first factor being a function of the pressure, the second factor being a function of the temperature, and the third factor being a function of the fuel flow rate. The synthesized value of output power provided by the engine is output.

Abstract: A method and a system for synthesizing thrust from a turbofan engine are provided. The turbofan engine comprising a compressor section, a combustor, and a turbine section in serial fluid flow communication. The engine is operated and, during the operating of the turbofan engine, a pressure of fluid at an exit of the compressor section and a temperature of fluid at a location upstream of the exit of the compressor section are determined. A synthesized value of thrust from the turbofan engine is determined based on a product of at least a first factor and a second factor, the first factor being a function of the pressure and the second factor being a function of the temperature. The synthesized value of thrust from the turbofan engine is output.

Abstract: A variable geometry inlet system of an aircraft engine includes an inlet duct. The inlet duct includes at least first and second sections moveable between extended and retracted positions such that the inlet duct defines a variable axial length of an inlet passage for selective flight conditions. The inclusion of acoustic treatment may assist in controlling noise.

Abstract: An aircraft includes one or more engines configured to generate high-pressure bleed air and low-pressure bleed air. An aircraft environmental control system (ECS) is in fluid communication with one or more of the engines to receive the high-pressure bleed air and the low-pressure bleed air. The ECS calculates a synthesized low-pressure value associated with the low-pressure bleed air while still providing bleed air using the high pressure bleed air. The ECS further switches from the high pressure bleed air to the low-bleed pressure air through the ECS while blocking flow of the high-pressure bleed air through the ECS based on the synthesized low-pressure value.

Abstract: A variable geometry inlet system of an aircraft engine includes an inlet duct. The inlet duct includes at least first and second sections moveable between extended and retracted positions such that the inlet duct defines a variable axial length of an inlet passage for selective flight conditions. The inclusion of acoustic treatment may assist in controlling noise.

Abstract: The present disclosure provides a method and a system for controlling operation of an engine of an aircraft. The method comprises, at a controller of the engine, obtaining an actual engine output power for the engine, the actual engine output power based on a propeller rotation speed and a propeller blade pitch angle; converting the actual engine output power to a predicted thrust value; determining an actual engine power limit associated with a thrust limit of a propeller coupled to the engine from the predicted thrust value; and setting a maximum engine power limit of the engine using the actual engine power limit associated with the thrust limit of the propeller.

Abstract: The present disclosure provides a method and a system for controlling operation of an engine of an aircraft. A first engine power limit associated with a thrust limit for a propeller coupled to the engine is obtained. The first engine power limit is compared to a second engine power limit associated with a mechanical limit for the engine and to a third engine power limit associated with a thermal limit for the engine. A maximum engine power limit is set at a lowest value of the first, second, and third engine power limits.

Abstract: An aircraft includes one or more engines configured to generate high-pressure bleed air and low-pressure bleed air. An aircraft environmental control system (ECS) is in fluid communication with one or more of the engines to receive the high-pressure bleed air and the low-pressure bleed air. The ECS calculates a synthesized low-pressure value associated with the low-pressure bleed air while still providing bleed air using the high pressure bleed air. The ECS further switches from the high pressure bleed air to the low-bleed pressure air through the ECS while blocking flow of the high-pressure bleed air through the ECS based on the synthesized low-pressure value.