Abstract: A propulsion system for an aircraft includes a gas generating core engine generates an exhaust gas flow that is expanded through a turbine section, a power turbine driven by the exhaust gas flow, a propulsor coupled to the power turbine, a hydrogen fuel system configured to supply hydrogen fuel to the combustor through a fuel flow path, a condenser arranged along the core flow path and configured to extract water from the exhaust gas flow, and an evaporator arranged along the core flow path receiving a portion of the water extracted by the condenser to generate a steam flow that is injected into the core flow path upstream of the turbine section.

Abstract: A gas turbine engine and engine mount structure includes a core engine including a compressor section, a combustor section and a turbine section mounted within a core engine housing. The fan, the compressor section and the turbine section rotate about an axis of rotation. An outer nacelle surrounds the fan, and is spaced from the core engine housing to define a bypass duct. The fan delivers air into the bypass duct and into the core engine housing. The nacelle is formed with camber so as to be curved in a first plane away from the axis of rotation in a first lateral direction. An engine mount structure extends from the nacelle at an angle that is non-parallel and non-perpendicular to the first plane, and has a component in a lateral direction that is opposed to the first lateral direction. An aircraft is also disclosed.

Abstract: A gas turbine engine includes a primary flow path fluidly connecting a compressor section, a combustor section and a turbine section. A cooling air flowpath is disposed radially outward of the primary flowpath. A first seal spans from an inner diameter of the cooling air flowpath to an outer diameter of the cooling air flowpath. The first seal includes at least one axial convolution and a plurality of pass through features defining a purge airflow. A heat shield is positioned immediately downstream of the first seal and is configured in relation to the first seal such that the purge airflow enters a mixing plenum defined between the heat shield and the first seal.

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 primary flow path fluidly connecting a compressor section, a combustor section and a turbine section. A cooling air flowpath is disposed radially outward of the primary flowpath. A first seal spans from an inner diameter of the cooling air flowpath to an outer diameter of the cooling air flowpath. The first seal includes at least one axial convolution and a plurality of pass through features defining a purge airflow. A heat shield is positioned immediately downstream of the first seal and is configured in relation to the first seal such that the purge airflow enters a mixing plenum defined between the heat shield and the first seal.

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: A gas turbine engine includes a propulsor delivering air into a core engine inward of an outer core panel. The core engine has a compressor section, a combustor and a turbine section. A compressor bleed system includes a bleed valve on a conduit communicating with compressed air from the compressor section. The bleed valve is configured to deliver air into a chamber where it can flow outwardly of the core engine at an exit location radially inward of the outer core panel.

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: A heat exchanger for heat exchange between a first fluid and a second fluid has a plurality of tube sections, each comprising; an interior for passing the first fluid; an exterior for exposure to the second fluid; a first leg; a second leg; a turn joining the first leg to the second leg; and a first face and a second face. A support has at least one carbon member engaging the plurality of tube sections.

Abstract: Aircraft propulsion systems have aircraft systems including at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems includes at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and a turbo expander. The main engine core includes a compressor section, a combustor section having a burner, and a turbine section arranged along an engine shaft. 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 turbo expander, prior to being injected into the burner for combustion. The turbo expander is configured to impart power to the engine shaft.

Abstract: 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 a turbo expander. The main engine core includes 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 turbo expander, prior to being injected into the burner for combustion. The turbo expander includes a rotor separated into a first expander portion and a second expander portion arranged about an output shaft and the output shaft is operably connected to a generator configured to generate electrical power.

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 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: 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: A heat exchanger for heat exchange between a first fluid and a second fluid has a plurality of tube sections, each having; an interior for passing the first fluid; an exterior for exposure to the second fluid; a first leg; a second leg; and a turn joining the first leg to the second leg. A has: fiber members passing between legs of the tube sections; and an end plate.

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 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

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.