A turboshaft engine for a rotorcraft includes a first spool and an independently rotatable second spool. The first spool includes a low pressure compressor, a first set of variable guide vanes disposed at an entry of the low pressure compressor, and a low pressure turbine drivingly engaged to the low pressure compressor. The second spool includes a high pressure compressor, a second set of variable guide vanes disposed at an entry of the high pressure compressor, the second set of variable guide vanes independently operable relative to the first set of variable guide vanes, and a high pressure turbine drivingly engaged to the high pressure compressor. One or both of the low pressure compressor and the high pressure compressor includes a mixed flow rotor.

A turboshaft engine for a rotorcraft includes a first spool and an independently rotatable second spool. The first spool includes a low pressure compressor, a first set of variable guide vanes disposed at an entry of the low pressure compressor, and a low pressure turbine drivingly engaged to the low pressure compressor. The second spool includes a high pressure compressor, a second set of variable guide vanes disposed at an entry of the high pressure compressor, the second set of variable guide vanes independently operable relative to the first set of variable guide vanes, and a high pressure turbine drivingly engaged to the high pressure compressor. One or both of the low pressure compressor and the high pressure compressor includes a mixed flow rotor.

A reverse-flow gas turbine engine includes a core of the gas turbine engine comprising multiple spools rotatable about a center axis of the gas turbine engine. Each spool is configured to pressurize air and to extract energy from combustion gases. The air and combustion gases are configured to flow through the core in a forward direction from an air inlet at an aft end of the core to an outlet at a forward end of the core. A propeller is disposed forward of the outlet. A reduction gearbox (RGB) is drivingly engaged to the core. An electric motor is drivingly engaged to the propeller and disposed axially between the RGB and the propeller.

A reverse-flow gas turbine engine includes a core of the gas turbine engine comprising multiple spools rotatable about a center axis of the gas turbine engine. Each spool is configured to pressurize air and to extract energy from combustion gases. The air and combustion gases are configured to flow through the core in a forward direction from an air inlet at an aft end of the core to an outlet at a forward end of the core. A propeller is disposed forward of the outlet. A reduction gearbox (RGB) is drivingly engaged to the core. An electric motor is drivingly engaged to the propeller and disposed axially between the RGB and the propeller.

The torque transfer assembly can include a sleeve having an elongated internal opening, a shaft having an elongated body extending in the internal opening of the sleeve, a first coupler at a first end of the elongated body coupled to a gas turbine engine rotor, and a second coupler at a second end of the elongated body coupled to a generator, the second coupler opposite the first coupler relative to a length of the elongated body, the shaft being made of a metal, and a bushing extending around the elongated body of the shaft, trapped between the shaft and the sleeve, the bushing made of plastic, the bushing having a coefficient of thermal expansion greater than a coefficient of thermal expansion of the sleeve.

A method for monitoring an aircraft turboshaft engine including a vibration sensor capable of outputting a vibration signal and a spark plug. The method includes, based on the vibration signal output by the vibration sensor, a step of determining a level of vibration of the turboshaft engine as well as a step of determining an indicator of spark plug wear.

A method for monitoring an aircraft turboshaft engine including a vibration sensor capable of outputting a vibration signal and a spark plug. The method includes, based on the vibration signal output by the vibration sensor, a step of determining a level of vibration of the turboshaft engine as well as a step of determining an indicator of spark plug wear.

A power management system for a multi engine rotorcraft having a main rotor system with a main rotor speed. The power management system includes a first engine that provides a first power input to the main rotor system. A second engine selectively provides a second power input to the main rotor system. The second engine has at least a zero power input state and a positive power input state. A power anticipation system is configured to provide the first engine with a power adjustment signal in anticipation of a power input state change of the second engine during flight. The power adjustment signal causes the first engine to adjust the first power input to maintain the main rotor speed within a predetermined rotor speed threshold range during the power input state change of the second engine.

Methods and systems for operating a rotorcraft comprising a plurality of engines are provided. A request to enter into an asymmetric operating regime (AOR), in which at least one active engine of the plurality of engines is operated in an active mode to provide motive power to the rotorcraft and at least one standby engine of the plurality of engines is operated in a standby mode to provide substantially no motive power, is obtained. Engine usage data for the plurality of engines, including at least one first engine and at least one second engine, is determined. Based on the engine usage data, one of the at least one first and second engines is operated as the at least one active engine for the AOR, and the other one of the at least one first and second engines is operated as the at least one standby engine for the AOR.

Methods and systems for operating a rotorcraft comprising a plurality of engines are provided. A request to enter into an asymmetric operating regime (AOR), in which at least one active engine of the plurality of engines is operated in an active mode to provide motive power to the rotorcraft and at least one standby engine of the plurality of engines is operated in a standby mode to provide substantially no motive power, is obtained. Engine usage data for the plurality of engines, including at least one first engine and at least one second engine, is determined. Based on the engine usage data, one of the at least one first and second engines is operated as the at least one active engine for the AOR, and the other one of the at least one first and second engines is operated as the at least one standby engine for the AOR.