A propulsion unit, notably for an aircraft, including a turbomachine; a propeller that can be selectively coupled to the turbomachine; a rotary electrical machine able to drive at least the turbomachine, and selective coupling of the rotary electric machine to the propeller and/or the turbomachine that is able to couple only the propeller to the rotary electrical machine during a defined stage of operation of the propulsion unit and that is able to couple or not couple the rotary electric machine to a gas generator and/or a free turbine of the turbomachine.

A hybrid gas-electric turbine engine for turboprop or turboshaft applications is disclosed together with associated methods. In various embodiments disclosed herein, the turbine engine comprises a turbine configured to be driven by a flow of combustion gas; a turbine shaft configured to be driven by the turbine and transfer power to a load coupled to the turbine engine and an electric motor configured to transfer power to the load coupled to the turbine engine. The rotor may have a rotor axis of rotation that is radially offset from a shaft axis of rotation of the turbine shaft. In some embodiments, the electric motor may be a multi-rotor electric motor.

Air inlet duct of a turbine engine, in particular an aircraft turbine engine comprising a gas generator, which extends axially between the air inlet and the gas generator and has a first axial wall part and a second wall part which is angularly offset with respect to the first part, which duct is capable of causing, in a shedding region, shedding of the boundary layer formed by an air flow along the wall of the duct; and a device for controlling said shedding of the boundary layer, characterised in that the control device comprises an air-blowing pipe which opens via at least one air-injection opening which is directly upstream of the shedding region, the blowing pipe being connected to an air intake positioned upstream of said air-injection opening or in the shedding region and comprising an air compressor means between the air intake and the air-injection opening.

In one aspect, there is provided a planetary gearbox, comprising a sun gear, a plurality of planet gear assemblies, each planet gear assembly having a main gear meshed with the sun gear, a fore lateral gear and an aft lateral gear disposed on opposite sides of the main gear and rotating therewith, a diameter of the main gear being different than a diameter of the fore and aft lateral gears, a planet carrier rotatably supporting at least some of the planet gear assemblies, and at least one fore ring gear meshed with the fore lateral gears, at least one aft ring gear meshed with the aft lateral gears, wherein one of the sun gear, the planet carrier, and the ring gears is configured to be operatively connected to an input, one is configured to be operatively connected to an output, and rotation of a remaining one is limited.

A gas turbine engine has a first spool having a low pressure compressor section disposed forward of an air inlet along a direction of travel of the engine, and a low pressure turbine section disposed forward of the low pressure compressor section and drivingly engaged thereto. A second spool has a high pressure compressor section disposed forward of the low pressure compressor section, and a high pressure turbine section disposed forward of the high pressure compressor section and drivingly engaged thereto. The high pressure turbine section is disposed aft of the low pressure turbine section. An output drive shaft drivingly engages the low pressure turbine section and extends forwardly therefrom to drive a rotatable load. A method of operating a gas turbine engine is also discussed.

An hybrid aircraft power plant, has: a gas turbine engine having a high-pressure spool including a high-pressure compressor, a high-pressure turbine, and a high-pressure shaft drivingly engaging the high-pressure turbine to the high-pressure compressor, a low-pressure spool including a low-pressure compressor, a low-pressure turbine, and a low-pressure shaft drivingly engaging the low-pressure turbine to the low-pressure compressor; an electric motor drivingly engaged to the low-pressure shaft; and a torque-transmitting device operatively connected to the HP-shaft and having an engaged configuration in which the torque-transmitting device drivingly engages the electric motor to the high-pressure shaft and a disengaged configuration in which the torque-transmitting device disconnects the electric motor from the high-pressure shaft.

A reverse flow gas turbine engine has a low pressure (LP) spool and a high pressure (HP) spool arranged sequentially in an axial direction. The LP spool comprises an LP compressor disposed forward of an LP turbine and drivingly connected thereto via an LP compressor gear train. The HP spool comprises an HP compressor in flow communication with the LP compressor, and an HP turbine disposed forward of the HP compressor and drivingly connected thereto via an HP shaft.

Systems and methods include providing an aircraft with a fuselage and a convertible engine disposed within the fuselage. The convertible engine is operable as a turbofan engine in a thrust mode and a turboshaft engine in a shaft power mode. The convertible engine includes a housing, an engine core having a low pressure turbine shaft, and a bypass fan system. The bypass fan system includes a bypass fan having a fan clutch. The fan clutch selectively couples at least a portion of the bypass fan to the low pressure turbine shaft when the convertible engine is operated in the thrust mode and decouples at least a portion of the bypass fan from the low pressure turbine shaft when the convertible engine is operated in the shaft power mode.

Turbine engine systems and aircraft having such systems are described. The turbine engine systems include a combustor arranged along a core flow path of the turbine engine, a drive shaft having at least a compressor section and a turbine section coupled thereto, a cryogenic fuel tank configured to supply a fuel to the combustor, and an expansion turbine mechanically coupled to the drive shaft, the expansion turbine configured to receive fuel from the cryogenic fuel tank and expand said fuel, wherein expansion of said fuel by the expansion turbine drives rotation of the expansion turbine to provide power input to the drive shaft.

There is provided a blade angle feedback system for an aircraft-bladed rotor rotatable about a longitudinal axis and having an adjustable blade pitch angle. A feedback device is coupled to rotate with the rotor and to move along the axis with adjustment of the blade pitch angle. The feedback device comprises a body having position marker(s) embedded therein, the body made of a first material having a first magnetic permeability and the position marker(s) comprising a second material having a second magnetic permeability greater than the first. Sensor(s) are positioned adjacent the feedback device and configured for producing, as the feedback device rotates about the axis, sensor signal(s) in response to detecting passage of the position marker(s). A control unit is communicatively coupled to the sensor(s) and configured to generate a feedback signal indicative of the blade pitch angle in response to the sensor signal(s) received from the sensor(s).