An aircraft that includes a hybrid-electric propulsion system is provided. In one aspect, the hybrid-electric propulsion system includes at least one propulsor that includes a gas turbine engine and an electric machine mechanically coupled with a spool of the gas turbine engine. When idle operation is commanded, electrical power is provided to the electric machine to cause the electric machine to apply torque to the spool and fuel provided to the engine can be reduced. Thus, the electric machine is controlled to provide a power assist to maintain the engine at the commanded idle speed whilst reducing fuel consumption.

A gas turbine engine is provided. The gas turbine engine includes a turbomachine having a low speed spool and a high speed spool; a rotor assembly coupled to the low speed spool; an electric machine rotatable with the low speed spool for extracting power from the low speed spool, for adding power to the low speed spool, or both; and an inter-spool clutch positioned between the low speed spool and the high speed spool for selectively coupling the low speed spool to the high speed spool.

A system and method is disclosed for the start-up and control of pulsejet engines and this system includes an Electronic Fuel Injection (“EFI”) system that further includes one or more electrically controlled fuel injectors that can be selectively operated for start-up and control of such pulsejet engines. According to the system and method, the rate and/or pattern of fuel delivery to pulsejet engines can be varied not only by controlling the amount of time the fuel injectors are open versus closed to define a “duty cycle,” but also with the capability to selectively disable one or more fuel injectors in the programmed manner for start-up and control of such pulsejet engines.

A system and method is disclosed for the start-up and control of pulsejet engines and this system includes an Electronic Fuel Injection (“EFI”) system that further includes one or more electrically controlled fuel injectors that can be selectively operated for start-up and control of such pulsejet engines. According to the system and method, the rate and/or pattern of fuel delivery to pulsejet engines can be varied not only by controlling the amount of time the fuel injectors are open versus closed to define a “duty cycle,” but also with the capability to selectively disable one or more fuel injectors in the programmed manner for start-up and control of such pulsejet engines.

A method of controlling the oil flow in an engine is provided. In preferred embodiments, the method comprises: flowing oil to a first oil pump upstream or downstream of a fuel oil heat exchanger and flowing oil to a second oil pump upstream or downstream of an air oil heat exchanger. One of two control functions to control the oil mass flow rate through the first oil pump is selected wherein the first control function minimizes specific fuel consumption (“SFC”) by the engine and the second control function minimizes average oil temperature. Preferably, the oil pumps are electric and the total combined oil mass flow rate of the first and second oil pumps is maintained constant.

A method of controlling the oil flow in an engine is provided. In preferred embodiments, the method comprises: flowing oil to a first oil pump upstream or downstream of a fuel oil heat exchanger and flowing oil to a second oil pump upstream or downstream of an air oil heat exchanger. One of two control functions to control the oil mass flow rate through the first oil pump is selected wherein the first control function minimizes specific fuel consumption (“SFC”) by the engine and the second control function minimizes average oil temperature. Preferably, the oil pumps are electric and the total combined oil mass flow rate of the first and second oil pumps is maintained constant.

A gas turbine engine comprising a variable geometry combustor having pilot fuel injectors and main fuel injectors; a fuel metering system configured to control fuel flow to the pilot fuel injectors and the main fuel injectors; a variable geometry airflow arrangement for the variable geometry combustor, which is configured to vary the airflow through the pilot fuel injectors and/or the main fuel injectors; a control system configured to control the variable geometry airflow arrangement in dependence upon airflow delivered to the combustor, the fuel flow to the pilot fuel injectors and the main fuel injectors, and a target index of soot emissions, thereby controlling airflow through the pilot fuel injectors and/or the main fuel injectors and hence the quantity of soot produced by combustion.

A gas turbine engine for an aircraft comprises a high-pressure (HP) spool comprising an HP compressor and a first electric machine driven by an HP turbine; a low-pressure (LP) spool comprising an LP compressor and a second electric machine driven by an LP turbine; a combustion system comprising a fuel metering unit; and an engine controller configured to, in response to a change of a throttle lever angle setting indicative of an acceleration event, increase fuel flow to the combustion system by the fuel metering unit, and to operate the first electric machine in a motor mode to increase the HP spool rotational speed and engine core mass flow.

A gas turbine engine for an aircraft comprises a high-pressure (HP) spool comprising an HP compressor and a first electric machine driven by an HP turbine; a low-pressure (LP) spool comprising an LP compressor and a second electric machine driven by an LP turbine; a combustion system comprising a fuel metering unit; and an engine controller configured to, in response to a change of a throttle lever angle setting indicative of an acceleration event, increase fuel flow to the combustion system by the fuel metering unit, and to operate the first electric machine in a motor mode to increase the HP spool rotational speed and engine core mass flow.

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.