A power management system for an aircraft with a hybrid power source comprises a rechargeable electricity source and an electricity generating source, a detector determining status data of the elements of the aircraft power consumption electrical circuit, and power data relating to the instantaneous electrical power demanded by the aircraft and/or the charging status of the rechargeable electricity sources, an automaton receiving the power data from the detector and determining a control status for the power sources, an adapter determining a backup electrical configuration when the status data indicate a failure, a controller determining an electrical control for the rechargeable electrical source and the electrical generating source based on the instantaneous electrical power demanded, and a switch emitting commands to the switches of the aircraft power consumption electrical circuit to implement a nominal electrical configuration, or, a backup electrical configuration for the adapter in case of receipt thereof.

A power management system for an aircraft with a hybrid power source comprises a rechargeable electricity source and an electricity generating source, a detector determining status data of the elements of the aircraft power consumption electrical circuit, and power data relating to the instantaneous electrical power demanded by the aircraft and/or the charging status of the rechargeable electricity sources, an automaton receiving the power data from the detector and determining a control status for the power sources, an adapter determining a backup electrical configuration when the status data indicate a failure, a controller determining an electrical control for the rechargeable electrical source and the electrical generating source based on the instantaneous electrical power demanded, and a switch emitting commands to the switches of the aircraft power consumption electrical circuit to implement a nominal electrical configuration, or, a backup electrical configuration for the adapter in case of receipt thereof.

A hybrid electric system for a rotorcraft can include a first thermal engine, a second thermal engine, and an electrical machine. The first thermal engine can be sized to produce a maximum first thermal engine power that is below a one-or-more-engine-inoperative (OEI) requirement power and the second thermal engine can be sized to produce a maximum second thermal engine power that is below the OEI requirement power. The electrical machine can be sized to provide at least a remaining power needed to reach the OEI requirement power in an OEI state.

A hybrid electric propulsion system includes a gas turbine engine and an electric machine coupled to the gas turbine engine. A method for operating the propulsion system includes determining, by one or more computing devices, a baseline power output for the gas turbine engine; operating, by the one or more computing devices, the gas turbine engine to provide the baseline power output; determining, by the one or more computing devices, a desired power output greater than or less than the baseline power output; and providing, by the one or more computing devices, power to, or extracting, by the one or more computing devices, power from, the gas turbine engine using the electric machine such that an effective power output of the gas turbine engine matches the determined desired power output.

A vehicle includes a gas turbine engine having at least two spools and an associated power system. The power system includes two independent power subsystems, including a first power subsystem for managing power transfer between spools and a second power subsystem for supplying a base power load to the vehicle. The first power subsystem has a first electric machine mechanically coupled with a first spool of the gas turbine engine and a second electric machine mechanically coupled with a second spool. The second electric machine is electrically coupled with the first electric machine such that electrical power is transmittable therebetween. The second power subsystem has a third electric machine mechanically coupled with one of the spools. The third electric machine is electrically coupled with a load positioned offboard the gas turbine engine. The first power subsystem and the second power subsystem are electrically decoupled from one another.

A power management system for an aircraft with a hybrid power source comprises a rechargeable electricity source and an electricity generating source, a detector determining status data of the elements of the aircraft power consumption electrical circuit, and power data relating to the instantaneous electrical power demanded by the aircraft and/or the charging status of the rechargeable electricity sources, an automaton receiving the power data from the detector and determining a control status for the power sources, an adapter determining a backup electrical configuration when the status data indicate a failure, a controller determining an electrical control for the rechargeable electrical source and the electrical generating source based on the instantaneous electrical power demanded, and a switch emitting commands to the switches of the aircraft power consumption electrical circuit to implement a nominal electrical configuration, or, a backup electrical configuration for the adapter in case of receipt thereof.

A power management system for an aircraft with a hybrid power source comprises a rechargeable electricity source and an electricity generating source, a detector determining status data of the elements of the aircraft power consumption electrical circuit, and power data relating to the instantaneous electrical power demanded by the aircraft and/or the charging status of the rechargeable electricity sources, an automaton receiving the power data from the detector and determining a control status for the power sources, an adapter determining a backup electrical configuration when the status data indicate a failure, a controller determining an electrical control for the rechargeable electrical source and the electrical generating source based on the instantaneous electrical power demanded, and a switch emitting commands to the switches of the aircraft power consumption electrical circuit to implement a nominal electrical configuration, or, a backup electrical configuration for the adapter in case of receipt thereof.

A computer-implemented method for optimally operating a hybrid-electric propulsion system by control of equipment dynamics. Prior to start of a mission, an original energy management plan is generated which is calculated to minimize estimated life-cycle operating costs for the vehicle during the mission. During an initial portion of the mission, operations of first and second power sources, a power distribution system, and a propulsion system are controlled such that a power mixture is supplied to the propulsion system from the first and second power sources in accordance with the original energy management plan. During the initial portion of the mission, a modified energy management plan is generated which is calculated to minimize estimated life-cycle operating costs for the vehicle. During a subsequent portion of the mission, operations of the first and second power sources, power distribution system, and propulsion system are controlled such that a power mixture is supplied to the propulsion system from the first and second power sources in accordance with the modified energy management plan.

An aircraft hybrid electrical propulsion (HEP) system includes an electrical system configured to deliver power to a plurality of electrical loads, a propulsion system configured to generate thrust in response to an input power, and an HEP controller in signal communication with the electrical system and the propulsion system. The HEP controller is configured to monitor a load demand of at least one electrical load among the plurality of electrical loads and to actively control the input power to actively control the thrust in response to changes in the load demand.

A hybrid gas turbine engine for use on an aircraft includes a motor/generator and gas turbine engine placed in parallel power communication with a rotating bladed component, such as an aircraft propeller, through a combining gear box. Power can be modulated with the propeller using the motor/generator. An aircraft having the aircraft propeller can also include several aircraft systems such as an air data computer, automatic flight control system (AFCS), a guidance and navigation system, a full authority digital engine controller/flight control computer (FADEC/FCC), and a fault detection and mitigation controller (FDMC). Data from each of these respective systems can be communicated over an aircraft data bus. In one form data from the AFCS and guidance and navigation system can be provided over the aircraft bus to the FDMC to modulate power to the propeller and in some forms act as a backup to the FADEC/FCC.