Abstract: An aircraft propulsion system includes a gas turbine engine, an electric machine, and a drive train. The gas turbine engine includes a first spool and a second spool. The electric machine includes a first rotor and a second rotor. The first rotor and the second rotor are concentric about a first axis. The drive train couples the first spool with the first rotor. The drive train is configured to couple the second spool with the second rotor. The drive train includes a tower shaft, a drive shaft, and an angled gear box. The tower shaft is coupled with the first spool. The tower shaft rotatable about a second axis. The drive shaft is operably coupled with the first rotor. The drive shaft is rotatable about a third axis transverse to the second axis. The angled gear box couples the tower shaft with the drive shaft.

Abstract: An 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 combustor section includes a combustor. 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. The air circuit projects out from the circuit inlet and extends through the first boost compressor, the second boost compressor and the heat exchanger. The first boost compressor and the second boost compressor are configurable upstream of the heat exchanger along the air circuit and independently operable.

Abstract: An assembly is provided for an aircraft propulsion system. This assembly includes a compressor section, a combustor section, a turbine section, a flowpath, a tail cone structure, an electric machine and an air system. The flowpath includes a flowpath inlet and a flowpath exhaust. The flowpath extends through the compressor section, the combustor section and the turbine section from the flowpath inlet to the flowpath exhaust. The tail cone structure forms an inner peripheral boundary of the flowpath at the flowpath exhaust. The tail cone structure includes an internal volume. The electric machine is disposed within the internal volume. The air system is configured to direct air, bled from the flowpath upstream of the compressor section, into the tail cone structure to cool the electric machine.

Abstract: An assembly is provided for an aircraft propulsion system. This assembly includes a compressor section, a combustor section, a turbine section, a flowpath, a tail cone structure, an electric machine and an electric fan. The flowpath includes a flowpath inlet and a flowpath exhaust. The flowpath extends through the compressor section, the combustor section and the turbine section from the flowpath inlet to the flowpath exhaust. The tail cone structure forms an inner peripheral boundary of the flowpath at the flowpath exhaust. The tail cone structure includes an internal volume and an outlet fluidly coupled with the internal volume. The electric machine is disposed within the internal volume. The electric fan is configured to flow air within the internal volume such that the air cools the electric machine and is exhausted from the tail cone structure through the outlet.

Abstract: An aircraft propulsion system assembly includes a vane structure, a tail cone structure, an electric machine and an air system. The vane structure includes an inner platform, an outer platform and a plurality of vanes. Each of the vanes extends radially across the flowpath from the inner platform to the outer platform. The vanes include a first vane and a second vane. The tail cone structure forms an inner peripheral boundary of the flowpath downstream of the vane structure. The tail cone structure includes an internal volume. The electric machine is disposed within the internal volume. The air system is configured to direct air from an air source radially outboard of the flowpath, radially inward across the flowpath through the first vane and the second vane, into the tail cone structure to cool the electric machine.

Abstract: An aircraft system includes an environmental control system and a propulsion system. The propulsion system includes a compressor section, a combustor section, a turbine section, a bypass flowpath, an air circuit, a working fluid system and a heat exchanger. The bypass flowpath bypasses the compressor section, the combustor section and the turbine section. The air circuit includes a circuit inlet fluidly coupled to the bypass flowpath. The air circuit is configured to direct air received from the bypass flowpath through the circuit inlet to the environmental control system. The working fluid system includes a fluid circuit. The working fluid system is configured to service one or more components of the propulsion system using a liquid working fluid directed through the fluid circuit. The heat exchanger is configured to transfer heat energy between the air in the air circuit and the liquid working fluid in the fluid circuit.

Abstract: An aircraft propulsion system includes an open propulsor rotor, a turbine engine and an air system. The turbine engine is configured to drive rotation of the open propulsor rotor about an axis. The turbine engine includes an engine rotor and a shroud. The engine rotor includes a plurality of rotor blades arranged circumferentially around the rotor base in an array. The shroud is next to and circumscribes the array of the rotor blades. The air system includes an electric boost compressor, an air circuit and a clearance control device. The air circuit extends longitudinally through the electric boost compressor from an air source to the clearance control device. The clearance control device is configured to control a clearance between the shroud and the rotor blades using air received from the air source through the air circuit.

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: A gas turbine engine having an axial centerline is provided that includes compressor, combustor, and turbine sections, a tangential on board injector (TOBI) system, and an HPC leakage guide structure. The compressor section has a high pressure compressor (HPC) that includes an HPC aft hub. The TOBI system extends circumferentially around the engine axial centerline, and has a plurality of nozzles and an inner radial flange. The HPC leakage guide structure has forward and aft ends. The forward end is disposed and configured to receive a leakage flow from the HPC and the aft end is engaged with the TOBI inner radial flange. The HPC leakage guide structure and the HPC aft hub define an HPC aft hub cavity. The HPC aft hub cavity extends between the forward and aft ends and has a flow area that is non-decreasing in a direction from the forward end to the aft end.

Abstract: A gas turbine engine disposed within a nacelle is provided that includes an open rotor propulsion system, compressor, combustion, and turbine sections, a lubrication system, and a turbine active clearance control system. The turbine active clearance control system includes first and second valves, a heat exchanger, and a nacelle air inlet device. The heat exchanger receives air from the air inlet device, accepts a lubricant flow therethrough, and permits heat transfer between the air flow and the lubricant to produce a conditioned air flow. The first valve receives conditioned air flow from the heat exchanger and may pass the conditioned air flow to the second valve. The second valve receives a second air flow from the air inlet device. The turbine active clearance control system may operate in a heating mode in which conditioned air flow, or a combined air flow is provided 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, a compressor section, an airflow inlet, and an inlet panel structure actuation system. The airflow inlet is in fluid communication with the compressor section. The airflow inlet is defined by a plurality of inlet panel structures. Each inlet panel structure includes an inlet side segment and an outer panel segment. The inlet side segments collectively define an outer radial surface of the airflow inlet. The inlet side segment and the outer panel segment of each inlet panel structure intersect at an inlet panel leading edge. The inlet panel structure actuation system is configured to selectively actuate the inlet panel leading edge of each inlet panel structure in a radial inward direction or in a radial outward direction.

Abstract: An aircraft propulsion system includes a gas turbine engine and a power transfer assembly. The gas turbine engine includes a high-pressure (HP) spool and a low-pressure (LP) spool. The power transfer assembly includes a first motor-generator, a second motor-generator, a motor control assembly, and a common assembly housing. The first motor-generator is operably coupled with the HP spool. The second motor-generator is operably coupled with the LP spool. The first motor-generator is electrically connected to the second motor-generator. The motor control assembly is electrically connected to the first motor-generator and the second motor-generator. The motor control assembly is configured to operate the first motor-generator to apply a first rotational force to the HP spool or to operate the second motor-generator to apply a second rotational force to the LP spool. The common assembly housing houses the first motor-generator and the second motor-generator.

Abstract: An aircraft propulsion system includes a gas turbine engine, a power transfer assembly, and a controller. The gas turbine engine includes a first spool and a second spool. The power transfer assembly includes a first motor-generator, a second motor-generator, and a motor control assembly. The first motor-generator is operably coupled with the first spool. The second motor-generator is operably coupled with the second spool. The motor control assembly is electrically connected to the first motor-generator and the second motor-generator. The motor control assembly is configured to control an operating voltage of the power transfer assembly. The controller is connected in signal communication with the motor control assembly. The is controller is configured to identify a target operating voltage of the power transfer assembly selected in response to an aircraft altitude input to the controller and control the motor control assembly to control the operating voltage at the target operating voltage.

Abstract: A gas turbine engine is provided that includes a fan section, a fan bypass air path, a compressor section, a combustion section, a turbine section, an annular compressor inlet, and an inlet panel structure actuation system. The annular compressor inlet is defined by an inner radial surface and a plurality of inlet panel structures. Each inlet panel structure includes an inlet side segment and a bypass side segment. The inlet side segments collectively define an outer radial surface of the compressor inlet. The inlet side segment and the bypass side segment of each inlet panel structure intersect at an inlet panel leading edge. The inlet panel structures are disposable in a radial inner most configuration and in a radial outer most configuration. The inlet panel structure actuation system is configured to selectively actuate the inlet panel structures between the radial inner most configuration and the radial outer most configuration.

Abstract: An apparatus is provided for an aircraft. This apparatus includes an electric machine and an air cooling circuit. The electric machine includes an electric machine rotor and an electric machine stator. The electric machine is configured to generate an electromagnetic field with the electric machine rotor and the electric machine stator. The electric machine rotor is configured to rotate about an axis. The air cooling circuit includes a cooling boot. The air cooling circuit is configured to direct air from an air source into the cooling boot. The cooling boot forms an air plenum with an exterior surface of the electric machine. The cooling boot includes a plurality of air outlets. The cooling boot is configured to direct the air through the air outlets and into the air plenum to air cool the electric machine.

Abstract: An assembly is provided for an aircraft. This assembly includes an engine core, an engine case, a nacelle wall and an air circuit. The engine core includes a compressor section, a combustor section and a turbine section. The engine case houses the engine core. The nacelle wall is radially outboard of and covers the engine case. A compartment is formed by and extends radially between the engine case and the nacelle wall. The air circuit includes a manifold disposed within the compartment. The air circuit is configured to direct air from an air source into the manifold. The manifold includes a plurality of outlets. The manifold is configured to direct the air into the compartment through the outlets.

Abstract: An apparatus is provided for an aircraft. This apparatus includes a turbine engine, a first electric machine, a first controller, a second electric machine, a second controller and a first fluid circuit. The first electric machine is operatively coupled to the turbine engine. The first controller is configured to control operation of the first electric machine. The second electric machine is operatively coupled to the turbine engine. The second controller is configured to control operation of the second electric machine. The first fluid circuit is configured to circulate a first liquid to cool and/or lubricate the first electric machine, the first controller, the second electric machine and the second controller.

Abstract: An aircraft apparatus includes a turbine engine, a first electric machine, a first controller, a first fluid circuit, a second electric machine, a second controller and a second fluid circuit. The turbine engine includes a first rotating structure and a second rotating structure. The first electric machine is operatively coupled to the first rotating structure. The first controller is configured to control operation of the first electric machine. The first fluid circuit is configured to circulate a first liquid and service the first electric machine and the first controller. The second electric machine is operatively coupled to the second rotating structure. The second controller is configured to control operation of the second electric machine. The second fluid circuit is configured to circulate a second liquid and service the second electric machine and the second controller. The second fluid circuit is fluidly discrete from the first fluid circuit.

Abstract: An aircraft assembly includes a turbine engine, a first electric machine, a first controller, an electric machine fluid circuit and a controller fluid circuit. The turbine engine includes a compressor section, a combustor section, a turbine section, a flowpath and a first rotating structure. The first rotating structure includes a first bladed rotor disposed in one of the compressor section or the turbine section. The first electric machine includes a first electric machine rotor. The first electric machine rotor is operatively coupled to the first rotating structure. The first controller is configured to electrically couple the first electric machine to an electrical system. The electric machine fluid circuit is configured to circulate a first liquid and service the first electric machine. The controller fluid circuit is configured to circulate a second liquid and service the first controller. The controller fluid circuit is fluidly discrete from the electric machine fluid circuit.

Abstract: An apparatus is provided for an aircraft that includes an open rotor propulsion system. The open rotor propulsion system includes a turbine engine, a first electric machine, a second electric machine, a first controller, a second controller and a first fluid circuit. The first controller is configured to control operation of the first electric machine. The second controller is configured to control operation of the second electric machine. The first fluid circuit is configured to circulate a first liquid to cool and/or lubricate the first electric machine, the first controller, the second electric machine and the second controller.