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 power lever angle setting indicative of a deceleration event, reduce fuel flow to the combustion system by the fuel metering unit, and to operate the first electric machine in a generator mode to reduce the HP spool rotational speed and engine core mass flow.

Methods and systems for reigniting an engine of an aircraft are described. An engine flameout event is detected during flight. Responsive to detecting the engine flameout event, an engine speed and a commanded engine operating state are monitored. The engine speed is compared to a predetermined threshold. A determination is made regarding whether the commanded engine operating state corresponds to an engine on state. When the commanded engine operating state corresponds to the engine on state, and when the engine speed is below the predetermined threshold, a predetermined ignition sequence for the engine is initiated.

A system for neural network compensated aero-thermodynamic gas turbine engine parameter/inlet condition synthesis. The system includes an aero-thermodynamic engine model configured to produce a real-time model-based estimate of engine parameters, a machine learning model configured to generate model correction errors indicating the difference between the real-time model-based estimate of engine parameters and sensed values of the engine parameters, and a comparator configured to produce residuals indicating a difference between the real-time model-based estimate of engine parameters and the sensed values of the engine parameters. The system also includes an inlet condition estimator configured to iteratively adjust an estimate of inlet conditions based on the residuals and adaptive control laws configured to produce engine control parameters for control of gas turbine engine actuators based on the inlet conditions.

A system and method for adaptively controlling a cooldown cycle of an auxiliary power unit (APU) that is operating and rotating at a rotational speed includes reducing the rotational speed of the APU to a predetermined cooldown speed magnitude that ensures combustor inlet temperature has reached a predetermined temperature value, determining, based on one or more of operational parameters of the APU, when a lean blowout of the APU is either imminent or has occurred, and when a lean blowout is imminent or has occurred, varying one or more parameters associated with the shutdown/cooldown cycle.

A fuel control system according to an exemplary aspect of the present disclosure includes, among other things, a fuel delivery valve selectively moveable to a closed position to shut off a flow of fuel to a downstream location. The system further includes a windmill bypass valve, and a shutoff pressure line between the windmill bypass valve and the fuel delivery valve. The windmill bypass valve is selectively operable to direct fuel to the shutoff pressure line to assist the movement of the fuel delivery valve to the closed position. A method is also disclosed.

A method of controlling at least part of a start-up or re-light process of a gas turbine engine, the method comprising: determining when a flame in a combustion chamber of a gas turbine engine is extinguished, during a start-up process or re-light process or during operation; purging the combustion chamber by controlling rotation of a low pressure compressor using a first electrical machine, and controlling rotation of a high pressure compressor using a second electrical machine, the combustion chamber downstream of the low pressure compressor and high pressure compressor; and controlling rotation of the low pressure compressor using the first electrical machine, and controlling rotation of the high pressure compressor using the second electrical machine to restart the start-up process or perform the re-light process.

Systems and methods for mitigating flameout risk in an engine of an aircraft are provided. A humidity value is obtained from a humidity sensor associated with the engine. A flameout risk for the engine is determined based on the humidity value. The flameout risk is compared to a predetermined risk threshold. When the flameout risk is above the predetermined risk threshold, a rate of fuel flow to the engine is increased.

Systems and methods for controlling fuel flow to an engine during start are provided. Fuel is caused to be injected into a combustor of the engine according to a first fuel schedule defining a minimum fuel flow limit required to achieve light-off of the engine, the minimum fuel flow limit set at an initial value. Following light-off of the engine, at least one operating parameter of the engine is monitored. Based on the at least one operating parameter, occurrence of flameout in the engine is detected. In response to detecting occurrence of flameout in the engine, the minimum fuel flow limit is increased from the initial value to a first value to obtain an adjusted fuel schedule, and fuel is caused to be injected into the combustor according to the adjusted fuel schedule.

A method of starting a gas turbine engine includes determining an abnormal shutdown condition during operation of the gas turbine engine and determining a first set of lightoff parameters for the gas turbine engine. The method also includes restarting the gas turbine engine using the first set of lightoff parameters. The method further includes iteratively determining subsequent first sets of lightoff parameters and restarting the gas turbine engine using a respective subsequent first set of the determined subsequent first sets of lightoff parameters until the gas turbine maintains a first set of operational parameters, where the first set of operational parameters is representative of a robust lightoff of the gas turbine engine.

A fuel injection system of a gas turbine includes a first pilot nozzle injecting fuel in a first flow rate range, a second pilot nozzle injecting fuel in a second flow rate range that is greater than the first flow rate range, a main nozzle injecting fuel in a third flow rate range that is greater than the second flow rate range, a first valve opening or closing a first supply pipe fueling the second pilot nozzle, a second valve opening or closing a second supply pipe fueling the main nozzle, and a controller selectively opening any one of the first supply pipe and the second supply pipe or opening or closing both of the first supply pipe and the second supply pipe by reflecting a change in altitude and thus applying control signals to the first valve and the second valve.