A rocket engine system comprising a rocket engine, coolant and a coolant source, propellant and a propellant source, a turbopump, and a heat source. The coolant is pressurized and then heated through a heat source to a supercritical state for augmented heat transfer. The heat source may be a heat exchanger with returning coolant, or a preburner. The rocket engine system may further comprise at least one additional rocket engine with a pump to provide pressure for multiple engine. The rocket engine system may further comprise multiple turbopump shafts for independent control of propellants.

A planetary gearbox includes a planet carrier, with a planetary gear rotatably disposed on the planet carrier, and a gear that meshes with the planetary gear, as well as a gas turbine engine having such a planetary gearbox. The planet carrier has an oil supply installation including a supply line for oil to an opening. The oil is directed out of the opening in the direction of the planetary gear and/or the gear. The oil supply installation in relation to a primary rotating direction of the planetary gear and/or the gear in front of the opening includes a shielding region which protrudes from an external side of the oil supply installation and which by way of the external side of the oil supply installation on a side that faces the primary rotating direction of the planetary gear and/or the gear, delimits an oil collection groove.

Flow diverters for installation in mid-turbine frame systems at a conduit outlet of gas turbine engines are described. The flow diverters include a diverter body having a connector portion defining a diverter inlet, a diverter extension at least partially defining a diverter outlet, and a curved portion arranged between the connector portion and the diverter extension, the curved portion configured to change a direction of flow from a first direction to a second direction that is about 90° from the first direction as the flow passes from the diverter inlet to the diverter outlet.

A turbine engine system includes at least one hydrogen fuel tank, a core flow path heat exchanger in a core flow path; and engine systems located in the core flow path. The engine system including at least a compressor section, a combustor section having a burner, and a turbine section. The core flow path heat exchanger is arranged in the core flow path downstream of the combustor section. The hydrogen fuel is supplied from the at least one hydrogen fuel tank through a hydrogen fuel supply line, passing through the core flow path heat exchanger and then supplied into the burner for combustion.

A system and method of increasing the temperature of oil that is used to anti-ice a gas turbine propulsion engine includes supplying oil from an oil supply system to an anti-ice oil circulation system that is disposed within the front frame of the gas turbine propulsion engine, and selectively injecting compressed air from the gas turbine propulsion engine into the oil supply system.

The present application provides an exhaust frame differential cooling system of a gas turbine engine to mitigate a temperature differential along a compressor and/or a turbine to minimize centerline eccentricity of a shaft. The exhaust frame differential cooling system may include a number of compressor temperature sensors positioned about the compressor and/or a number of turbine temperature sensors positioned about the turbine, an exhaust frame including an inner barrel with a bearing tunnel for the shaft, an outer barrel, and a number of struts extending from the inner barrel to the outer barrel, a blower, and a cooling air metering system that provides cooling air from the blower to the bearing tunnel and through the inner barrel, the struts, and the outer barrel in response to the temperature differential being determined along the compressor and/or the turbine.

A gas turbine engine includes a compressor section, a combustor, and a turbine section. A bleed tap taps air from the compressor section through a bleed valve. The bleed valve is selectively opened by a control to dump air from the compressor section to a dump outlet. A heat exchanger duct includes a duct air inlet to cool a fluid in a heat exchanger and a duct air outlet. The dump outlet is within the heat exchanger duct.

An aircraft turbine engine includes a gas generator and a fan arranged upstream from the gas generator and configured to generate a main gas flow, one portion of which flows in a flow path of the gas generator to form a primary flow, and another portion of which flows in a flow path around the gas generator to form a secondary flow. The gas generator includes a low-pressure compressor that includes a rotor driving the fan. The turbine engine further includes an electric machine. The electric machine includes a rotor rotated by the rotor of the low-pressure compressor, and a stator extending around the rotor of the electric machine and configured to be cooled by the primary flow.

A compact cooling and boosting gas turbine system provides combined cooling, heating, and electrical power with high energy efficiency. The system has a pressure booster and a turbo-compressor. The pressure booster includes a fuel inlet, a fuel outlet, and a piston, and is in fluid communication with a gas turbine engine. The pressure booster also includes a coolant inlet, a coolant chamber, and a coolant outlet, and is in fluid communication with a closed pressurized coolant flow. The turbo-compressor includes a compressor and a turbine, and is in fluid communication with a water input flow and with the closed pressurized coolant flow. A coolant flow control valve controls the closed pressurized coolant flow. The system is configured to provide a cold water flow for a first position of the flow control valve and to provide a hot water flow for a second position of the flow control valve.

A combustor component according to at least one embodiment of the present invention includes a cylindrical body which internally includes a combustion chamber, and includes a weld part where a plurality of through holes opening to the combustion chamber are formed, and a housing which is disposed on an outer circumferential side of the cylindrical body to cover a part of the weld part, and defines an acoustic damping space communicating with the combustion chamber via at least one of the through holes. The plurality of through holes in the weld part has a formation density which is higher in a first region of the weld part covered with the housing than in a second region of the weld part positioned outside the housing.