Abstract: An assembly of a gas turbine engine includes a low pressure compressor, a high pressure compressor, an intermediate case between the low pressure compressor and the high pressure compressor, and a cleaning apparatus having an annular arrangement about the intermediate case. A core flow path is defined through the low pressure compressor, the intermediate case, and the high pressure compressor. The annular arrangement includes a scroll portion configured to extract debris from an airstream at an outer diameter wall of the core flow path and return the airstream to the core flow path.

Abstract: An assembly of a gas turbine engine includes a low pressure compressor, a high pressure compressor, an intermediate case between the low pressure compressor and the high pressure compressor, and a cleaning apparatus having an annular arrangement about the intermediate case. A core flow path is defined through the low pressure compressor, the intermediate case, and the high pressure compressor. The annular arrangement includes a scroll portion configured to extract debris from an airstream at an outer diameter wall of the core flow path and return the airstream to the core flow path.

Abstract: A cooling system for a plurality of conductive cables in a gas turbine engine includes a cooling source and an electric motor disposed in a tail cone. The cooling source may comprise an electric fan or an oil pump. The cooling source may be configured for active cooling of the plurality of conductive cables. The electric fan may be in fluid communication with ambient air during operation of the gas turbine engine.

Abstract: A fan section includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor is a four-stage low pressure compressor. The low pressure turbine is a three-stage low pressure turbine. A high spool including a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine-stage high pressure compressor. The high pressure turbine is a two-stage high pressure turbine. An exhaust gas exit temperature of greater than 900 degrees Fahrenheit and less than 1000 degrees Fahrenheit at maximum take-off.

Abstract: An aircraft propulsion system includes a fan section that includes a fan shaft that is rotatable about a fan axis. The fan shaft includes a fan gear. The aircraft propulsion system also includes a boost turbine engine that includes a first output shaft that includes a first gear that is coupled to the fan gear. The boost turbine engine has a first maximum power capacity. The aircraft propulsion system further includes a cruise gas turbine engine that includes a second output shaft that includes a second gear that is coupled to the fan gear. The cruise turbine engine has a second maximum power capacity that is less than the first maximum power capacity of the boost turbine engine. The fan section produces a thrust that corresponds to power input through the fan gear from the boost turbine engine and the cruise turbine engine.

Abstract: A hybrid-electric aircraft system is provided and includes first and second hybrid-electric engines, first and second ducting systems fluidly communicative with each other and with the first and second hybrid-electric engines, respectively, and a control system. The control system is operably coupled to each of the first and second hybrid-electric engines and to each of the first and second ducting systems. The control system is configured to run the first hybrid-electric engine normally, to run the second hybrid-electric engine in a lower power mode and to control each of the first and second ducting systems to direct bleed air from the first hybrid-electric engine to the second hybrid-electric engine.

Abstract: Aircraft propulsion systems and aircraft are described. The aircraft propulsion systems include aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and an expander, wherein the main engine core comprises a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the expander, prior to being injected into the burner for combustion.

Abstract: A fan section includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor is a four-stage low pressure compressor. The low pressure turbine is a three-stage low pressure turbine. A high spool including a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine-stage high pressure compressor. The high pressure turbine is a two-stage high pressure turbine. An exhaust gas exit temperature of greater than 900 degrees Fahrenheit and less than 1000 degrees Fahrenheit at maximum take-off.

Abstract: A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction. An inner core engine has an inner core engine housing surrounding a compressor section, including a low pressure compressor. A rigid connection between a fan case and the inner core engine includes A-frames rigidly connected at a connection point to the fan case. Fan exit guide vanes rigidly connect to the fan case, and to the inner core engine. A fan intermediate case is positioned forward of a first rotor stage in the low pressure compressor. A rigid structure is connected to the inner core engine and to the fan exit guide vanes. The rigid structure defines a structure moment stiffness. The fan intermediate case defines an intermediate case moment stiffness. A ratio of the structure moment stiffness to the intermediate case moment stiffness is between 5 and 15.

Abstract: A turbine engine has: a compressor; a combustor; a turbine; a gaspath passing downstream from the compressor through the combustor and then through the turbine; a fuel source; a fuel flowpath from the fuel source to the combustor; and a heat exchanger for transferring heat from the gaspath to the fuel flowpath The heat exchanger has: an inner wall in heat transfer relation with the gaspath; an outer wall; tubes between the inner wall and the outer wall bounding respective segments of the fuel flowpath; and a heat transfer fluid between the inner wall and the outer wall and in heat transfer relation with the tubes and the inner wall.

Abstract: A gas turbine engine includes a fan section that includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor includes a greater number of stages than the low pressure turbine. A high spool includes a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine stage high pressure compressor and includes a pressure ratio per stage of greater than or equal to 1.20 and less than or equal to 1.33.

Abstract: An engine system of an aircraft includes a gas turbine engine comprising at least one spool and at least one electric machine operably coupled with the at least one spool. A controller is configured to detect if the at least one spool of the gas turbine engine is in or is approaching an overspeed condition and apply a load to the at least one spool via the at least one electric machine.

Abstract: Aircraft propulsion systems and methods of operation thereof, include aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and a turbo expander assembly. The main engine core includes a compressor section, a combustor section having a burner, and a turbine section. Fuel is supplied from the at least one fuel tank through a fuel flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and selectively through the turbo expander assembly, prior to being injected into the burner for combustion. The turbo expander assembly is operably coupled to at least two load sources through a selective coupler and configured to selectively drive operation of the at least two load sources.

Abstract: A cooling system for a plurality of conductive cables in a gas turbine engine includes a cooling source and an electric motor disposed in a tail cone. The cooling source may comprise an electric fan or an oil pump. The cooling source may be configured for active cooling of the plurality of conductive cables. The electric fan may be in fluid communication with ambient air during operation of the gas turbine engine.

Abstract: A gas turbine engine includes a fan section that includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor includes a greater number of stages than the low pressure turbine. A high spool includes a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine stage high pressure compressor and includes a pressure ratio per stage of greater than or equal to 1.20 and less than or equal to 1.33.

Abstract: A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction. An inner core engine has an inner core engine housing surrounding a compressor section, including a low pressure compressor. A rigid connection between a fan case and the inner core engine includes A-frames rigidly connected at a connection point to the fan case. Fan exit guide vanes rigidly connect to the fan case, and to the inner core engine. A fan intermediate case is positioned forward of a first rotor stage in the low pressure compressor. A rigid structure is connected to the inner core engine and to the fan exit guide vanes. The rigid structure defines a structure moment stiffness. The fan intermediate case defines an intermediate case moment stiffness. A ratio of the structure moment stiffness to the intermediate case moment stiffness is between 5 and 15.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a rotating structure, a stationary structure, a bearing and a gearbox. The rotating structure is configured to rotate about a rotational axis. The rotating structure includes a shaft and a bladed rotor connected to the shaft. The stationary structure circumscribes the rotating structure. The bearing rotatably mounts the shaft to the stationary structure. The gearbox is disposed radially inboard of the bearing.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a stationary structure, a rotating structure and an electric machine. The rotating structure is rotatably mounted to the stationary structure by a first bearing and a second bearing. The electric machine is between the first bearing and the second bearing. The electric machine includes a rotor and a stator circumscribing the rotor. The rotor is connected to the rotating structure. The stator is connected to the stationary structure.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a first rotating structure, a turbine engine apparatus, a rotating coupler and a seal assembly. The first rotating structure is configured to rotate about a rotational axis. The turbine engine apparatus includes an electric machine and a second rotating structure. The electric machine includes an electric machine rotor and an electric machine stator. The second rotating structure is configured to rotate about the rotational axis and is coupled to the electric machine rotor. The rotating coupler is coupled to the first rotating structure by a first connection. The rotating coupler is coupled to the second rotating structure by a second connection. The seal assembly includes a rotating seal land and a stationary seal element. The rotating seal land is mounted onto the rotating coupler. The stationary seal element sealingly engages the rotating seal land.

Abstract: A gear reduction reduces a speed of a fan rotor relative to a speed of a fan drive turbine. A rigid connection between a fan case and an inner core housing includes a plurality of A-frames connected at a connection point to the fan case. Legs in the A-frames extend away from the connection point in opposed circumferential directions to be connected to a compressor wall of the inner core housing. The rigid connection also includes a plurality of fan exit guide vanes rigidly connected to the fan case. A lateral stiffness ratio of the lateral stiffness of the plurality of fan exit guide vanes and a lateral stiffness of a combination of the plurality of A-frame, the compressor wall, and a fan intermediate case which is forward of the low pressure compressor being greater than or equal to 0.6 and less than or equal to 2.0.