A method of pumping fuel in a fuel system of a composite aircraft having a gas turbine engine includes feeding un-heated fuel from a fuel tank to the gas turbine engine, by using an ejector pump to draw the fuel from the fuel tank and feeding the fuel through an engine fuel pump within a main fuel line upstream of the gas turbine engine. Fuel from the main fuel line is bled and directed to the ejector pump via a motive flow pump assembly, the motive flow pump assembly including a motive flow pump generating a motive flow for the ejector pump. The bleed fuel flow is then passed through a hydrophobic fuel screen located upstream of the motive flow pump.

A fluid delivery system can include a main pump disposed in a primary fluid line configured to receive a low pressure fluid on a low pressure side and output a high pressure fluid on a high pressure side to provide flow from the fluid source to a fluid destination via the primary fluid line. An interconnected valve system can be disposed in the primary fluid line configured to control flow through the primary fluid line.

A fuel ecology system includes a fuel ecology reservoir having a reservoir volume, includes a moveable barrier disposed within the reservoir volume, the movable barrier defining a first volume and a second volume within the reservoir volume, a fuel inlet port in fluid communication with the first volume, a fuel outlet port in fluid communication with the first volume, and a vent port in fluid communication with the second volume.

A fuel ecology system includes a fuel ecology reservoir having a reservoir volume, includes a moveable barrier disposed within the reservoir volume, the movable barrier defining a first volume and a second volume within the reservoir volume, a fuel inlet port in fluid communication with the first volume, a fuel outlet port in fluid communication with the first volume, and a vent port in fluid communication with the second volume.

A system and medium for controlling a fuel gas compressor of a gas turbine system that compresses a gaseous fuel for consumption in a high-pressure combustor. Moreover, the compressor is configured to generate a discharge pressure for the combustor based at least in part on a load demand for the gas turbine system.

A combustion staging system includes a splitting unit which receives a metered fuel flow and controllably splits the received fuel flow into pilot and mains flows. Pilot and mains fuel manifolds distribute fuel from the splitting unit to the pilot and mains stages. The splitting unit selects and deselects pilot-only operation. Both pilot and mains manifolds are selectable for pilot and mains operation. A cooling flow recirculation line has a delivery section arranged to provide a cooling flow of fuel to the mains manifold when it is deselected during pilot-only operation. Cooling flow enters the delivery section from a high pressure fuel zone of the engine and exits the return section to a low pressure fuel zone of the engine. A controller adjusts the splitting unit during pilot-only operation to partially select the mains manifold thereby increasing the pressure in the mains manifold to meet a target fuel pressure therein.

A combustion staging system includes a splitting unit which receives a metered fuel flow and controllably splits the received fuel flow into pilot and mains flows. Pilot and mains fuel manifolds distribute fuel from the splitting unit to the pilot and mains stages. The splitting unit selects and deselects pilot-only operation. Both pilot and mains manifolds are selectable for pilot and mains operation. A cooling flow recirculation line has a delivery section arranged to provide a cooling flow of fuel to the mains manifold when it is deselected during pilot-only operation. Cooling flow enters the delivery section from a high pressure fuel zone of the engine and exits the return section to a low pressure fuel zone of the engine. A controller adjusts the splitting unit during pilot-only operation to partially select the mains manifold thereby increasing the pressure in the mains manifold to meet a target fuel pressure therein.

A fuel supply system for a turbine engine that provides a modulated thrust control malfunction accommodation (TCMA). The fuel supply system can include a fuel line that fluidly connects a fuel tank and the turbine engine. A fuel pump and a fuel metering valve can be fluidly connected to the fuel line. A bypass line can fluidly connect to the fuel line. Flow through the bypass line can be controlled using a bypass valve and a balancing pressure valve. The TCMA can then modulate the fuel flow using the valves.

A fuel supply system for a turbine engine that provides a modulated thrust control malfunction accommodation (TCMA). The fuel supply system can include a fuel line that fluidly connects a fuel tank and the turbine engine. A fuel pump and a fuel metering valve can be fluidly connected to the fuel line. A bypass line can fluidly connect to the fuel line. Flow through the bypass line can be controlled using a bypass valve and a balancing pressure valve. The TCMA can then modulate the fuel flow using the valves.

A fuel supply circuit for an aeronautical cryogenic turbomachine including: at least one cryogenic reservoir containing a liquid fuel topped with a boil-off gas and including a high-pressure liquid pump to supply at least one main propulsion device of the turbomachine with liquid fuel, an auxiliary turbomachine including an electric generator, a gas compressor to supply the auxiliary turbomachine with gaseous fuel, and a buffer gas reservoir connecting the gas compressor to the cryogenic reservoir to reinject gas into the cryogenic reservoir in order to maintain the pressure in the cryogenic reservoir above a predefined value, the electric generator supplying the high-pressure liquid pump and the gas compressor.