Abstract: An assembly is provided for an aircraft powerplant. This assembly includes a vane structure, a first electric device and a cooling system. The first electric device is disposed within the vane structure. The first electric device includes a first device housing and first electrical circuitry. The first electrical circuitry is housed within an interior of the first device housing. The cooling system is configured to cool the first electrical circuitry. The cooling system includes a first heat exchanger and an air circuit. The first heat exchanger is disposed within the vane structure and thermally coupled to the first electric device. The air circuit extends within the vane structure from an airflow inlet into the air circuit, through the first heat exchanger, to an airflow outlet from the air circuit. The airflow inlet fluidly couples the air circuit to a volume external to the vane structure.

Abstract: A hybrid-electric aircraft propulsion system includes an engine, a first electric machine assembly, and a second electric machine assembly. The engine includes a first rotational assembly and a second rotational assembly. The first electric machine assembly includes a first electric machine, a first control unit, first electrical cables, and a first cable conduit. The first electric machine is coupled with the first rotational assembly. The first electrical cables extend between and electrically connect the first electric machine and the first control unit through the first cable conduit. The second electric machine assembly includes a second electric machine, a second control unit, second electrical cables, and a second cable conduit. The second electric machine is coupled with the second rotational assembly. The second electrical cables extend between and electrically connect the second electric machine and the second control unit through the second cable conduit.

Abstract: A gas turbine engine is provided that includes compressor and turbine sections, and a turbine cooling air (TCA) system. The turbine section has upstream and downstream rotor stages, and turbine sub-sections. The TCA system includes a flow mixing device and a compressor bleed air flow valve, and is configured to receive first and second compressor bleed air flows. The compressor bleed air flow valve may be in an open or a closed configuration. The TCA system operates in a first or a second mode. In the first mode, the compressor bleed air flow valve is closed and a conditioned air flow is from the first compressor bleed air flow. In the second mode, the compressor bleed air flow valve is open and the conditioned air flow is a mixture of the first and second compressor bleed air flows. The conditioned air flow is directed to the downstream turbine sub-section.

Abstract: A gas turbine engine is provided that includes fan, compressor, combustion, and turbine sections, a fan bypass air path, and a turbine active clearance control system. The turbine section has a rotor stage and a turbine case clearance control segment. The turbine active clearance control system includes a bleed air mixing system, a fan bypass air inlet device configured to receive a fan bypass air flow, and a compressor air bleed structure configured to receive a compressor bleed air flow. The turbine active clearance control system is configurable in a turbine case heating mode in which the bleed air mixing system mixes the fan bypass air flow and the compressor bleed air flow to produce a conditioned air flow that is operable to heat the turbine case clearance control segment. The turbine active clearance control system directs the conditioned air flow to the turbine case clearance control segment.

Abstract: A gas turbine engine disposed within a nacelle is provided that includes an open rotor propulsion system, compressor, combustion, and turbine sections, and a turbine active clearance control system. The turbine section has a turbine case clearance control segment disposed radially outside of a rotor stage. The turbine active clearance control system includes a bleed air mixing system, a nacelle air inlet device, and a compressor air bleed structure. The turbine active clearance control system is configurable in a turbine case heating mode in which the bleed air mixing system is used to mix an exterior air flow and a compressor bleed air flow to produce a conditioned air flow that is operable to heat the turbine case clearance control segment, and the turbine active clearance control system is configured to direct the conditioned air flow to the turbine case clearance control segment.

Abstract: A gas turbine engine is provided that includes a fan, compressor, combustion, and turbine sections, a fan bypass air path, a lubrication system, and a turbine active clearance control (ACC) system. The turbine section has a rotor stage and a turbine case clearance control segment. The turbine ACC system includes first and second valves and a heat exchanger. The heat exchanger may receive a first fan bypass air flow and a lubricant flow and permits heat transfer therebetween to produce a flow of conditioned air. The first valve receives conditioned air from the heat exchanger. The second valve may receive conditioned air and a second fan bypass air flow. The turbine ACC system may operate in a heating mode in which conditioned air flow, or a combination of conditioned air flow and the second fan bypass air flow is passed to the turbine case clearance control segment.

Abstract: A propulsion system for an aircraft includes a gas turbine engine, an electrical assembly, a powertrain, and a controller. The gas turbine engine includes a rotational structure and a second rotational structure each including a turbine rotor. The electrical assembly includes an electric machine and a control unit. The control unit is electrically connected to the electric machine. The powertrain includes a differential geartrain configured to couple the rotational structure and the second rotational structure together to the electric machine. The controller is connected in communication with the control unit. The controller is configured to maintain a predetermined power transfer ratio between a spool mechanical power transfer of the second rotational structure and an electric machine mechanical power transfer of the electric machine along a rotation speed range of the second rotational structure by controlling an electrical loading of the electric machine with the control unit.

Abstract: An aircraft powerplant assembly includes a gearbox, a first electric machine, a first electric cable, a second electric machine and a second electric cable. The first electric machine is mounted to the gearbox at a first side of the gearbox. A first machine terminal is disposed at a first side of the first electric machine. The first electric cable is electrically coupled to the first machine stator through the first machine terminal. The second electric machine is mounted to the gearbox at the first side of the gearbox. A second machine terminal is disposed at a second side of the second electric machine that faces the first side of the first electric machine. The second electric cable is electrically coupled to the second machine stator through the second machine terminal.

Abstract: A system includes a nacelle, a gas turbine engine, and an electrical machine. The nacelle includes an upper bifurcation and an engine core compartment. The gas turbine engine includes a low pressure spool and is disposed within the nacelle. The electrical machine is disposed within the upper bifurcation and is configured to transfer mechanical power between the low pressure spool and the electrical machine.

Abstract: An assembly is provided for an aircraft powerplant. This assembly includes a gearbox, an engine accessory, an engine core, a motor accessory and an electric motor. The gearbox includes an engine power transfer system and a motor power transfer system independent of the engine power transfer system. The engine accessory includes an engine accessory rotor. The engine core includes a flowpath, a compressor section, a combustor section, a turbine section and an engine rotating assembly. The flowpath extends through the compressor section, the combustor section and the turbine section. The engine rotating assembly includes a turbine rotor in the turbine section. The engine rotating assembly is operatively coupled to the engine accessory rotor through the engine power transfer system. The motor accessory includes a motor accessory rotor. The electric motor includes an electric motor rotor operatively coupled to the motor accessory rotor through the motor power transfer system.

Abstract: An aircraft powerplant assembly includes a cooling plate, a first control valve and a first electric machine. The cooling plate includes a fluid cooling circuit, a cooling circuit inlet and a cooling circuit outlet. The fluid cooling circuit is disposed in the cooling plate and extends from the cooling circuit inlet to the cooling circuit outlet. The fluid cooling circuit includes a first circuit passage fluidly coupled between the cooling circuit inlet and the cooling circuit outlet. The first control valve is fluidly coupled between the cooling circuit inlet and the first circuit passage. A first controller housing is removably attached to the cooling plate and overlaps the first circuit passage. First controller circuitry is disposed within an interior of the first controller housing. The first controller circuitry is in thermal communication with the cooling plate through a wall of the first controller housing.

Abstract: An aircraft powerplant assembly includes a cooling plate, a first electric machine controller and a second electric machine controller. The cooling plate includes a fluid cooling circuit, a cooling circuit inlet and a cooling circuit outlet. The fluid cooling circuit extends within the cooling plate from the cooling circuit inlet to the cooling circuit outlet. A first controller housing is removably attached to the cooling plate and overlaps the fluid cooling circuit. First controller circuitry is in thermal communication with the cooling plate through a wall of the first controller housing. A second controller housing is removably attached to the cooling plate and overlaps the fluid cooling circuit. Second controller circuitry is in thermal communication with the cooling plate through a wall of the second controller housing.

Abstract: An assembly for an aircraft powerplant includes an engine case, a gearbox, a conduit, a first electric machine, a first controller and a first electric cable. The gearbox is offset from the engine case. The conduit extends longitudinally from a first end of the conduit to a second end of the conduit. The conduit is anchored to the engine case at the first end of the conduit by a first conduit mount. The conduit is anchored to the gearbox at the second end of the conduit by a second conduit mount. The conduit is self-supported longitudinally between the first conduit mount and the second conduit mount. The first electric machine is mounted to the gearbox. The first controller is mounted to the engine case. The first electric cable electrically couples the first electric machine to the first controller. The first electric cable extends longitudinally through a bore of the conduit.

Abstract: An assembly is provided for an aircraft powerplant. This assembly includes an electric machine, an electric machine controller, an electric cable, a conduit and an air circuit. The electric machine controller is configured to control operation of the electric machine. The electric cable electrically couples the electric machine controller to the electric machine. The conduit extends longitudinally from a first end of the conduit to a second end of the conduit. The electric cable extends longitudinally through a bore of the conduit. The air circuit is configured to cool the electric cable within the conduit using a flow of air. The air circuit includes the bore of the conduit.

Abstract: A propulsion system for an aircraft includes an open propulsor rotor, a turbine engine and a thermal management system. The turbine engine includes an engine core and an engine flowpath. The engine core includes a compressor section, a combustor section and a turbine section. The engine flowpath extends longitudinally though the compressor section, the combustor section and the turbine section from an airflow inlet into the engine flowpath to a combustion products exhaust from the engine flowpath. The thermal management system is configured to manage heat energy generated by the propulsion system during operation of the turbine engine. The thermal management system includes an electric boost compressor, a heat exchanger and a system flowpath disposed outside of the engine core. The system flowpath extends longitudinally through the electric boost compressor and the heat exchanger from an airflow inlet into the system flowpath to an airflow exhaust from the system flowpath.

Abstract: An aircraft powerplant assembly includes a cooling plate and an electric machine controller. The cooling plate includes a fluid cooling circuit internal to a body of the cooling plate. The fluid cooling circuit includes an inlet manifold, an outlet manifold and a plurality of heat exchange passages. Each of the heat exchange passages extends longitudinally along a longitudinal centerline from the inlet manifold to the outlet manifold. A first heat exchange passage is configured with a plurality of first cooling elements arranged longitudinally along its longitudinal centerline. The electric machine controller includes a controller housing and controller circuitry. The controller housing is removably attached to the cooling plate and overlaps the heat exchange passages. The controller circuitry is disposed within an interior of the controller housing. The controller circuitry is in thermal communication with the cooling plate through a wall of the controller housing.

Abstract: A propulsion system for an aircraft includes a propulsor rotor, a powerplant, a nacelle and a lockout system. The powerplant is coupled to and configured to drive rotation of the propulsor rotor. An electrical system for the powerplant is configured to receive electrical power from an electrical power source. The nacelle includes an access structure configured to move between a first position and a second position. The access structure at least partially covers the electrical system when in the first position. The access structure is arranged in the first position for propulsion system operation. The access structure is operable to be arranged in the second position for propulsion system maintenance and/or propulsion system inspection. The lockout system is configured to automatically trigger an operation to cut off the electrical power to the electrical system from the electrical power source when the access structure moves away from the first position.

Abstract: A component for a gas turbine engine includes a component body formed of ceramic matrix composite lamina and has at least one hook. The at least one hook has an attachment region radially inward of the at least one hook. The attachment region is radially thinner from a hook end of the at least one hook to a remote end, and then becomes radially thicker. A slot is formed through a radial thickness of the at least one hook from the hook end in a remote direction, such that there are two sections of the attachment region. A gas turbine engine is also disclosed.

Abstract: An assembly is provided for an aircraft powerplant. This assembly includes a vane structure and a heat exchanger. The vane structure extends longitudinally to a leading edge of the vane structure. The vane structure includes a translating body, a stationary body and an inlet passage. The translating body is configured to translate longitudinally between a first position and a second position. The translating body is configured to form the leading edge of the vane structure and close an inlet into the inlet passage when in the first position. The translating body is configured to open the inlet into the inlet passage when in the second position. The heat exchanger is disposed within the stationary body. The inlet passage projects into the stationary body from the inlet into the inlet passage to the heat exchanger.

Abstract: A system is provided for an aircraft. This aircraft system includes a turbine engine and an air system. The turbine engine includes a flowpath, a compressor section, a combustor section and a turbine section. The flowpath extends through the compressor section, the combustor section and the turbine section from an airflow inlet into the flowpath to a combustion products exhaust from the flowpath. The air system includes an air circuit, a first boost compressor, a second boost compressor and a heat exchanger. A circuit inlet into the air circuit is fluidly coupled to the flowpath upstream of the combustor section. The air circuit projects out from the circuit inlet and extends through the first boost compressor, the second boost compressor and the heat exchanger.