Abstract: Aircraft hydrogen fuel systems and methods and systems of starting such systems are described. The aircraft hydrogen fuel systems include a hydrogen burning main engine, a main tank configured to contain liquid hydrogen to be supplied to the main engine during a normal operation, and a starter tank configured to contain gaseous hydrogen to be used during a startup operation of the main engine. Methods and processes for starting and/or restarting such systems are described.

Abstract: Aircraft propulsion systems include a fan shaft connected to a fan, the fan shaft defining a centerline axis of the aircraft propulsion system, one or more offset cores arranged at an angle to the centerline axis, the one or more offset cores each comprising a hydrogen burning combustor, a centerline cavity defined along the centerline axis, and a heat exchanger arranged within the centerline cavity. In operation, a portion of air is directed from the fan into the centerline cavity to provide a first working fluid to the heat exchanger within the centerline cavity.

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: An assembly is provided for a turbine engine. This turbine engine assembly includes an electric machine, a turbine engine component and a lubrication system. The electric machine includes a rotor and a stator. The turbine engine component is discrete from the electric machine. The lubrication system includes a first lubricant circuit, a second lubricant circuit and a lubricant source fluidly coupled in parallel with the first lubricant circuit and the second lubricant circuit. The first lubricant circuit includes a machine volume for the electric machine. The second lubricant circuit includes a component volume for the turbine engine component.

Abstract: A gas turbine engine includes a fan section including 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. The gear reduction is a planetary gear system. 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. A high spool includes a high pressure turbine that drives a high pressure compressor. A first low spool support bearing is located axially between the low pressure compressor and the gear reduction. A second low spool support bearing is located axially between the low pressure compressor and the high pressure compressor.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a rotating structure, a stationary structure and an electric machine. The rotating structure is configured to rotate about a rotational axis. The stationary structure circumscribes the rotating structure. The electric machine includes a rotor and a stator. The rotor circumscribes the rotating structure and is coupled to the rotating structure through a spline connection. The stator is connected to the stationary structure.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a turbine engine shaft, a gearbox and a coupler. The turbine engine shaft is configured to rotate about a rotational axis. The gearbox includes a gear configured to rotate about the rotational axis. The coupler is coupled to the turbine engine shaft by a coupler-shaft connection. The coupler-shaft connection is configured as or otherwise includes a crowned spline connection. The coupler is coupled to the gear by a coupler-gear connection. The coupler-gear connection is configured as or otherwise includes a crowned spline connection.

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 gas turbine engine has a low speed input shaft drives a first plurality of accessories. A high speed input shaft drives a second plurality of accessories. The first plurality of accessories rotating about a first set of rotational axes perpendicular to a first plane. The second plurality of accessories rotating about a second set rotational axes perpendicular to a second plane. The first and second planes extending in opposed directions away from a drive input axis. Compressed air is tapped and passes through a heat exchanger, then to a boost compressor, and then to at least one rotatable components in a main compressor section and a main turbine section. The boost compressor driven on a boost axis, which is non-parallel to the first set of rotational axes and the second set of rotational axes.

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.

Abstract: An ejector assembly for a cooling system of a gas turbine engine may comprise: a tail cone having a tail cone outlet in fluid communication with a cooling air flow of the cooling system; an ejector body defining a mixing section, a constant area section, and a diffuser section; and a nozzle section in fluid communication with an exhaust air flow of the gas turbine engine, the ejector assembly configured to entrain the cooling air flow via the exhaust air flow.

Abstract: The cooling system may comprise: an electric machine; a first conduit including a cable housing and an inlet; a plurality of conductive cables extending from the electric machine, the plurality of conductive cables disposed at least partially in the cable housing; and an electric fan disposed in the first conduit, the cooling system configured to passively flow air through the first conduit to cool the plurality of conductive cables during operation of the gas turbine engine, and the electric fan configured to actively cool the plurality of conductive cables after an engine shutdown 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 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 is a five-stage low pressure compressor. The low pressure turbine is four-stage 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. The high pressure turbine is a two-stage high pressure turbine.

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 turbine engine system includes aircraft systems including at least one hydrogen fuel tank, engine systems comprising a compressor section, a combustor section having a burner, and a turbine section, and a hydrogen fuel flow supply line configured to supply hydrogen fuel from the at least one hydrogen fuel tank into the burner for combustion. The turbine engine system has a bypass ratio between 5 to 20.

Abstract: 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.

Abstract: A hybrid electric engine including a gas turbine engine including a low speed spool, a high speed spool a fan section, a compressor section, a combustor section, and a turbine section. The hybrid electric engine further includes an electric generator configured to convert rotational power of the high or low speed spool to electricity and a variable area turbine control system electrically connected to the electric generator. The variable area turbine control system being configured to adjust a cross-sectional area of a core flow path of the hybrid electric engine. The variable area turbine control system including a plurality of variable turbine vanes located in the turbine section and a variable area turbine actuator configured to rotate each of the plurality of variable turbine vanes to adjust the cross-sectional area of the core flow path of the hybrid electric engine. The variable area turbine actuator is an electromechanical actuator.

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: A hybrid electric propulsion system including: a gas turbine engine comprising a low speed spool and a high speed spool, the low speed spool comprising a low pressure compressor and a low pressure turbine, and the high speed spool comprising a high pressure compressor and a high pressure turbine; an electric motor configured to augment rotational power of the high speed spool or the low speed spool; at least one blade outer air seal positioned between an outer case of the high pressure turbine and a plurality of blades of the high pressure turbine; a clearance control system operably coupled to the at least one blade outer air seal, the clearance control system configured to vary a position of the at least one blade outer air seal with respect to the plurality of blades of the high pressure turbine; and a controller operably coupled to the electric motor and the clearance control system, wherein the controller is configured to operate the clearance control system based upon an operational state of the electri