An aircraft propulsion system comprises a gas turbine engine arranged to provide propulsive thrust and a fuel cell system having an air input port, the aircraft propulsion system being configured such that air from a compressor of the gas turbine engine is provided to the air input port during operation of the aircraft propulsion system. The fuel cell system is able to provide appreciable electrical power at altitude without the need for a dedicated compressor.

A power supply device includes a power generator, a drive source, a plurality of power supply lines, a plurality of batteries, a diode, a difference calculating unit 11, 12, 13, or 14, and a difference summing unit 16. The difference calculating unit 11, 12, 13, or 14 is configured to calculate a difference between demanded electric power P1, P2, P3, or P4 in the corresponding power supply line and a charge state of the corresponding battery. The difference summing unit 16 is configured to sum the differences in electric power in the power supply lines calculated by the difference calculating units 11, 12, 13, and 14. In the power supply device, the drive source is controlled such that electric power equal to or higher than the electric power calculated by the difference summing unit 16 is generated by the power generator.

The disclosure relates to an assembly for an aircraft turbomachine, having a speed reduction gear comprising an input pinion connected to a power shaft of the turbomachine and an output pinion connected to a propeller shaft of the turbomachine. According to the disclosure, the assembly includes two electric machines which are each configured to provide electrical power to the propeller shaft or to draw mechanical power from the propeller shaft, each electric machine comprising a rotor and a stator, the stator configured to be connected to a casing of the turbomachine. The speed reduction gear can include two substantially parallel intermediate transmission lines configured to transmit the torque from the input pinion to the output pinion, each rotor being driven in rotation respectively by a pinion of an intermediate line.

A power supply device includes a power generator, a drive source, a plurality of power supply lines, a plurality of batteries, a difference calculating unit 11, a difference summing unit 12, and an electric power summing unit 13. The difference calculating unit 11 is configured to calculate differences D1, D2, D3, and D4 between a target charge state set for each battery and an estimated charge state. The difference summing unit 12 is configured to sum the differences D1, D2, D3, and D4 calculated by the difference calculating unit 11. The electric power summing unit 13 is configured to sum the charge state calculated by the difference summing unit 12 and electric power used for the electric loads. A control unit 9 controls the drive source such that electric power calculated by the electric power summing unit 13 is generated by the power generator.

A system is provided for an aircraft. This system includes a propulsor rotor and a powerplant configured to drive rotation of the propulsor rotor. The powerplant includes a turbo-compounded intermittent internal combustion engine and a dual rotor electric machine. The turbo-compounded intermittent internal combustion engine includes a turbine rotor operatively coupled to the propulsor rotor through the dual rotor electric machine. The dual rotor electric machine includes a first rotor, a second rotor, a first stator and a second stator. The first stator and the second stator are arranged radially between and axially aligned with the first rotor and the second rotor.

A system includes a propeller assembly comprising a propeller, a drive shaft coupled to the propeller assembly, wherein the drive shaft when in operation generates rotation of the propeller, a reduction gearbox coupled to the drive shaft, wherein the reduction gearbox when in operation alters a rotation speed between a power turbine shaft and the drive shaft, and an electric motor disposed between the propeller assembly and the reduction gearbox, wherein the electric motor is coupled to the drive shaft, wherein the electric motor when in operation imparts torque to at least partially rotate the drive shaft.

A hybrid electric aircraft equipped with gyroscopic stabilization control is provided. In one aspect, a hybrid electric aircraft includes a turbo-generator having a gas turbine engine and an electric generator operatively coupled thereto for generating electrical power. The turbo-generator defines a rotation axis. The aircraft also includes one or more electrically-driven propulsors for producing thrust for the aircraft. In addition, the aircraft includes a pivot mount operatively coupled with the turbo-generator. To provide gyroscopic stabilization control of the aircraft, the pivot mount is controlled to adjust the rotation axis of the turbo-generator relative to a prime stability axis of the aircraft. Additionally or alternatively, a rotational speed of the turbo-generator can be changed to provide gyroscopic stabilization control of the aircraft.

A fuel cell assembly is provided, including a fuel cell stack having a fuel cell, the fuel cell having a cathode and an anode; and a multi-gas sensor configured to sense gas composition data of a flow of output products from the cathode, gas composition data of a flow of output products from the anode, gas composition data of a fluid surrounding the fuel cell, or a combination thereof to determine fuel cell leakage diagnostic information.

A fuel cell assembly is provided, including a fuel cell stack having a fuel cell, the fuel cell having a cathode and an anode; and a multi-gas sensor configured to sense gas composition data of a flow of output products from the cathode, gas composition data of a flow of output products from the anode, gas composition data of a fluid surrounding the fuel cell, or a combination thereof to determine fuel cell leakage diagnostic information.

Disclosed herein are high-power hybrid electric jet propulsion systems and methods for creating thrust using electric propulsion and chemical reaction. Embodiments include powering a Stage1 (6) electric machines (6b) through motor controllers for spinning a shaft (5), which spins a fan (2) to draw in air through the engine cowl (1) into Stage1 (6) where it gets compressed and heated. The air is then passed into a cyclonic combustion chamber (7) for mixing with a fuel to create a medium (X1). The medium (X1) is then passed into the turbine blades (8a) in Stage2 (8) to the spin shaft (5). The shaft (5) is configured to produce the high bypass thrust (10) and the jet exhaust thrust (9).