A fuel flow system for a gas turbine engine includes a fuel inlet to admit a flow of fuel into the fuel flow system and a fuel outlet to direct the flow of fuel to the gas turbine engine from the fuel flow system. One or more pumps are positioned along a fuel flow path connecting the fuel inlet to the fuel outlet. A metering valve is in flow communication with the one or more pumps to meter the flow of fuel. A minimum pressure shut-off valve is fluidly connected to the one or more pumps, and a shutoff apparatus configured to permit selective energizing to stop the flow of fuel from the fuel outlet without operation of the metering valve.

A pumping system includes a pump array of multiple pump-engine assemblies. Each pump-engine assembly comprises a pump and a gas turbine engine driving the pump. A manifold is coupled to the pumps. A master controller is coupled to each of the pump-engine assemblies either directly or via one or more intermediate controllers. The master controller and any intermediate controllers are collectively programmed to respond to user input including a desired hydraulic output at the manifold by automatically calculating and applying inputs to the individual pump-engine assemblies to provide the desired hydraulic output.

A gas turbine engine for an aircraft. The engine including: an engine core with a turbine, a compressor, a fan located upstream of the compressor and including a plurality of fan blades, and a core shaft connecting the turbine to the compressor; a gearbox which receives an input from the core shaft and outputs drive, via a driveshaft, to the fan so as to drive the fan at a lower rotational speed than the turbine, the drive shaft and core shaft forming a shaft system.

An aircraft is provide including an airframe, an extending tail, and a counter rotating, coaxial main rotor assembly including an upper rotor assembly and a lower rotor assembly. A translational thrust system positioned at the extending tail, the translational thrust system providing translational thrust to the airframe. At least one flight control computer configured to independently control the upper rotor assembly and the lower rotor assembly through a fly-by-wire control system. A plurality of sensors to detect sensor data of at least one environmental condition and at least one aircraft state data, wherein the sensors provide the sensor data to the flight control computer.

A system and method can include a rotational component, a fluid pump, and a shaft transferring rotational power toward the fluid pump. A fluid circuit can include a valve and the fluid pump wherein pump is configured to motivate fluid toward the valve disposed downstream of the pump. The system can be configured to raise fluid pressure at the pump outlet by closing the valve to thereby effect an increased braking load on the shaft. The rotational component can be an electric machine mechanically coupled to a gas turbine engine. The fluid circuit can include a heat exchanger configured to transfer heat between the rotational component and the fluid. The system can include a second heat exchanger configured to transfer heat from the fluid to a heat sink. A processing system can be configured to receive a command to increase the hydraulic braking load to the rotational component by closing the valve to raise fluid pressure at the pump outlet based on a braking command.

There are described methods and systems for operating an aircraft turboprop engine. The method comprises controlling a propeller of the turboprop engine based on a selected one of a reference propeller rotational speed and a minimum propeller blade angle while the turboprop engine is running; detecting an inflight restart of the turboprop engine; and controlling the propeller during the inflight restart in accordance with at least one of a modified reference propeller rotational speed and a modified minimum propeller blade angle to maintain an actual propeller blade angle above an aerodynamic disking angle during the inflight restart.

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.

The invention regards a gas turbine engine control system and a method for limiting turbine overspeed in case of a shaft failure. The control system includes: an overspeed protection system that activates an activation member in case a shaft failure is detected; a fuel limiting mechanism coupled with the activation member, wherein the fuel limiting mechanism is configured to limit the fuel supply to the gas turbine engine combustor if the activation member is activated; a variable stator vane mechanism which is configured to adjust variable stator vanes of a compressor of the gas turbine engine in their rotational position, the variable stator vanes having a closed position which blocks air flow through the compressor. A connecting fuel line connecting the fuel limiting mechanism and the variable stator vane mechanism is provided, wherein upon activation of the activation member the fuel limiting mechanism pressurizes the connecting fuel line, thereby activating the variable stator vane mechanism to move at least one row of the variable stator vanes into the closed position.

A system for piloting an aircraft includes a throttle, a flight parameter sensor system, and an engine calculator controlling thrust parameters of an engine of the aircraft. The piloting system includes a flight control unit connected to the engine calculator, the sensors and the throttle. The flight control unit generates a thrust control vector from a flight control law and sends a digital signal including the generated thrust control vector to the engine calculator. The engine calculator controlling said thrust parameters of the engine depending on the generated thrust control vector received.

An aircraft propulsion system includes a gas turbine engine; a generator; a storage battery; a motor which drives a rotor, using at least one of the electric power which is output from the generator and the electric power which is output from the storage battery; a detection unit which detects the number of revolutions of the engine shaft; an engine control unit which controls at least a fuel flow rate of the gas turbine engine; and a generator control unit which controls the operation of the generator. When the number of revolutions satisfies a predetermined condition, at least the generator control unit executes a control for reducing a sudden change in the number of revolutions.