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.

One embodiment is an electric drive system for an aircraft comprising a plurality of redundant motors, wherein power generated by the plurality of motors is used to drive a rotor system comprising a rotor shaft having a plurality of rotor blades connected thereto; a gear box associated with the plurality of redundant motors; a collective actuator for controlling a collective pitch of the rotor blades connected to the rotor shaft; and at least one structural element for retaining the redundant motors, the gear box, and the collective actuator together as a single integrated unit.

A hybrid multirotor propulsion system for an aircraft includes a plurality of propulsion units, each propulsion unit having a propeller, an electromotor and a peripheral differential gearbox; a plurality of driving elements, each of which is coupled to a respective one of the plurality of propulsion units; a mechanical power source; a main distributor gearbox; at least one electric machine; and a power management unit. The power management unit is configured according to a predetermined operating mode, which causes the mechanical power source to output first and second mechanical power components; and distributing the first mechanical power component to provide each driving element with a direct mechanical propeller power; and causes the electric machine to convert the second mechanical power component into electric power, part of which provides each electromotor with an electric propeller power. The direct mechanical propeller power causes each electromotor to convert the electric propeller power into an indirect mechanical propeller power, outputted to the peripheral differential gearbox; and causes the peripheral differential gearbox of each propulsion unit to aggregate the direct mechanical propeller power and the indirect mechanical propeller power to a total mechanical propeller power which drives the propeller of each propulsion unit.

A hybrid multirotor propulsion system for an aircraft includes a plurality of propulsion units, each propulsion unit having a propeller, an electromotor and a peripheral differential gearbox; a plurality of driving elements, each of which is coupled to a respective one of the plurality of propulsion units; a mechanical power source; a main distributor gearbox; at least one electric machine; and a power management unit. The power management unit is configured according to a predetermined operating mode, which causes the mechanical power source to output first and second mechanical power components; and distributing the first mechanical power component to provide each driving element with a direct mechanical propeller power; and causes the electric machine to convert the second mechanical power component into electric power, part of which provides each electromotor with an electric propeller power. The direct mechanical propeller power causes each electromotor to convert the electric propeller power into an indirect mechanical propeller power, outputted to the peripheral differential gearbox; and causes the peripheral differential gearbox of each propulsion unit to aggregate the direct mechanical propeller power and the indirect mechanical propeller power to a total mechanical propeller power which drives the propeller of each propulsion unit.

A twin-engine system includes a gas turbine engine comprising a core and a first output shaft drivable by the core. An electric engine has an electric motor configured to drive a second output shaft. A reduction gear box (RGB) has an RGB input drivingly engaged to both the first output shaft and the second output shaft. The RGB has an RGB output to provide rotational output to a rotatable load.

A twin-engine system includes a gas turbine engine comprising a core and a first output shaft drivable by the core. An electric engine has an electric motor configured to drive a second output shaft. A reduction gear box (RGB) has an RGB input drivingly engaged to both the first output shaft and the second output shaft. The RGB has an RGB output to provide rotational output to a rotatable load.

A hybrid propulsion system includes a heat engine configured to drive a heat engine shaft. An electric motor is configured to drive a motor shaft. A transmission system includes at least one gear box. The transmission system is configured to receive rotational input power from each of the heat engine shaft and the motor shaft and to convert the rotation input power to output power. The motor shaft includes a disconnect mechanism to allow the heat engine to rotate with the electric motor stopped. The heat engine shaft includes a disconnect mechanism to allow the electric motor to rotate with the heat engine stopped.

A hybrid propulsion system includes a heat engine configured to drive a heat engine shaft. An electric motor is configured to drive a motor shaft. A transmission system includes at least one gear box. The transmission system is configured to receive rotational input power from each of the heat engine shaft and the motor shaft and to convert the rotation input power to output power. The motor shaft includes a disconnect mechanism to allow the heat engine to rotate with the electric motor stopped. The heat engine shaft includes a disconnect mechanism to allow the electric motor to rotate with the heat engine stopped.

A method includes providing thrust to an aircraft through a power train from a heat engine connected to the power train. The method includes controlling an electric motor connected to the power train to counter torque ripple in the power train from the heat engine. A system includes a power train for providing thrust to an aircraft. A heat engine is connected to the power train. An electric motor is operatively connected to the power train. A controller is operatively connected to control the electric motor. The controller includes machine readable instructions configured to cause the controller to control the electric motor to counter torque ripple in the power train from the heat engine.

A method includes providing thrust to an aircraft through a power train from a heat engine connected to the power train. The method includes controlling an electric motor connected to the power train to counter torque ripple in the power train from the heat engine. A system includes a power train for providing thrust to an aircraft. A heat engine is connected to the power train. An electric motor is operatively connected to the power train. A controller is operatively connected to control the electric motor. The controller includes machine readable instructions configured to cause the controller to control the electric motor to counter torque ripple in the power train from the heat engine.