The present disclosure provides air cooling systems and methods for propulsion systems (e.g., aviation or aerospace propulsion systems). More particularly, the present disclosure provides integrated air cooling systems and methods utilizing air cycle machine cooling for hybrid-electric aircraft or aerospace propulsion systems or the like. The present disclosure provides integrated air cycle machine cooling into the hybrid propulsion system (e.g., into the wing-mounted hybrid propulsion system). As such, the air cooling systems and methods of the present disclosure can minimize weight while improving electric motor/generator cooling.

A thermal management system for an aircraft comprises a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, a first thermal bus, and a first heat exchanger module. The first thermal bus comprises a first heat transfer fluid, with the first heat transfer fluid being in fluid communication, in a closed loop flow sequence, between the first gas turbine engine, the or each first electric machine, and the first heat exchanger. Waste heat energy generated by at least one of the first gas turbine engine, and the or each first electric machine, is transferred to the first heat transfer fluid. The first heat exchanger module comprises a first flow path and a second flow path. The first flow path is configured to direct a flow of the first heat transfer fluid either to a first heat dissipation portion in which a first proportion Q.sub.A of the waste heat energy from the first heat transfer fluid is transferred to a first dissipation medium, or additionally to a second heat dissipation portion in which a second proportion Q.sub.B of the waste heat energy from the first heat transfer fluid is transferred to a second dissipation medium, in dependence on a temperature of the first heat transfer fluid entering the first heat exchanger module, a temperature of the first heat dissipation medium, and a temperature of the second heat dissipation medium. The second flow path is configured to direct the flow of the first heat transfer fluid to a second heat dissipation portion in which a second proportion Q.sub.B of the waste heat energy from the first heat transfer fluid is transferred to a second dissipation medium, or additionally to the first heat dissipation portion in which a first proportion Q.sub.A of the waste heat energy from the first heat transfer fluid is transferred to a first dissipation medium, in dependence on a temperature of the first heat transfer fluid entering the first heat exchanger module, a temperature of the first heat dissipation medium, and a temperature of the second heat dissipation medium.

A power generation system for an aircraft includes a first power source, a second power source, and a power dispatch module communicatively coupled with the first and second power sources. The power dispatch module includes a controller having one or more processors configured to perform a plurality of operations, including but not limited to receiving a plurality of loading data associated with the power generation system, predicting a future power demand due to future load changes using the loading data, determining first and second power setpoints for the first and second power sources, respectively, based on the future power demand due to the future load changes, and controlling first and second power outputs of the first and second power sources based on the first and second power setpoints such that the future power demand of the power generation system is shared by the first and second power sources.

Provided are a hybrid engine system capable of protecting an engine from overspeeding when the load of a generator rapidly decreases, and a method of controlling the hybrid engine system.

The hybrid engine system includes: an engine; a generator driven by the engine to output electrical energy; a battery configured to store electrical energy produced by the generator or supply electrical energy together with the generator; and a controller configured to control the engine, wherein the controller includes a torque meter for measuring torque of an output shaft of the engine and a current meter for measuring output current of the generator, and is further configured to change a control mode of the engine when a reduction rate of at least one of the torque and the current is greater than a set value while the engine is operating in a fly mode.

An electric machine adapted for use in a gas turbine engine includes a shaft extending along a central axis, a magnetic rotor drum, and a non-magnetic rotor hub rotatably coupled with the shaft and the magnetic rotor drum. The magnetic rotor drum includes a rotor and a plurality of magnets arranged circumferentially about the central axis.

A propulsion assembly includes a first torque source coupled with a first shaft and a second torque source coupled with a second shaft. A coupler selectively couples the first and second torque sources. When the first and second torque sources are coupled via the coupler, in response to a command to decouple the first torque source, an unloading operation is performed to decrease the torque output provided by the first torque source to a threshold, and when reached, the first shaft is decoupled from the coupler. When the first torque source is coupled with the coupler but the second torque source is not, in response to a command to couple the second torque source, a speed matching operation is performed to increase the speed of the second shaft to match a speed of the first shaft, and when the speeds are matched, the second shaft is coupled to the coupler.

A propulsion assembly includes a first torque source coupled with a first shaft and a second torque source coupled with a second shaft. A coupler selectively couples the first and second torque sources. When the first and second torque sources are coupled via the coupler, in response to a command to decouple the first torque source, an unloading operation is performed to decrease the torque output provided by the first torque source to a threshold, and when reached, the first shaft is decoupled from the coupler. When the first torque source is coupled with the coupler but the second torque source is not, in response to a command to couple the second torque source, a speed matching operation is performed to increase the speed of the second shaft to match a speed of the first shaft, and when the speeds are matched, the second shaft is coupled to the coupler.

An electric aircraft includes a forward wing and an aft wing both located on a top side of a fuselage and both having an upward dihedral angle. The forward wing has a straight or slightly forward-swept leading edge. The aft wing has a swept leading edge. A plurality of propeller engines are located on a leading-edge side of the forward wing. A single propeller engine may be located on a top side of an aft end of the fuselage; alternatively, a plurality of propeller engines are located on the aft wing on either its leading or trailing edge. The propeller engines are each powered by an electric motor. An unobstructed cargo door is located on a side of the fuselage, aft of the forward wing.

An aircraft propulsion system having a variable airflow capture area is provided. The propulsion system includes a main propulsion source and an auxiliary propulsion source. In a first mode, the auxiliary propulsion source is stowed within an aerodynamic profile of the aircraft, and the main propulsion source provides all of the propulsion force for powering flight of the aircraft. In a second mode, the auxiliary propulsion source is deployed to augment the airflow capture area of the main propulsion source and increase an overall airflow capture area of the propulsion system. In the second mode, the auxiliary power source is operated by power extracted from the main propulsion source. The main propulsion source may include one or more low bypass ratio engines. The auxiliary power source may include one or more auxiliary thrust fans coupled at a plurality of locations on the aircraft.

An aircraft includes an engine and an electric power system. The electric power system includes a permanent magnet machine. The permanent magnet machine includes a stator and a rotor configured to rotate relative to the stator.