Abstract: An assembly for an aircraft propulsion system includes a housing structure, a guide vane structure, and a front mount assembly. The housing structure includes an engine case assembly. The engine case assembly includes a front center body. The guide vane structure includes a guide vane structure base. The front mount assembly is mounted on the front center body. The front mount assembly is disposed axially coincident with the guide vane structure. The front mount assembly includes at least one pylon attachment linkage.

Abstract: An assembly for an aircraft propulsion system includes an open propulsor rotor, a housing structure, and a turbine engine core. The open propulsor rotor includes a plurality of open propulsor blades arranged circumferentially around an axis. Each of the plurality of open propulsor blades projects radially out into an environment external to the aircraft propulsion system. The housing structure includes an engine case assembly and a nacelle. The nacelle forms at least a portion of an exterior flow boundary for the aircraft propulsion system. The nacelle forms a guide vane structure receptacle at the exterior flow boundary. The nacelle further includes a nacelle interior skeletal structure disposed at the guide vane structure receptacle. The turbine engine core is housed within the engine case assembly. The turbine engine core is configured to drive rotation of the open propulsor rotor about the axis.

Abstract: An assembly for an aircraft propulsion system includes a housing structure. The housing structure includes an engine case assembly and a nacelle. The nacelle is disposed radially outboard of the engine case assembly. The nacelle forms at least a portion of an exterior flow boundary for the aircraft propulsion system that borders an environment external to the aircraft propulsion system. The nacelle includes a nacelle interior skeletal structure mounted to the engine case assembly at a forward mount and an aft mount.

Abstract: An assembly for an aircraft propulsion system includes a housing structure, an open guide vane structure, and a flowpath. The housing structure includes an upstream housing end and a pylon structure. The pylon structure includes a main pylon body and a front center body attached to the main pylon body. The front center body is disposed at the upstream housing end. The front center body includes an outer radial case structure and an inner radial case structure extending circumferentially about an axis of the aircraft propulsion system. The open guide vane structure includes a guide vane structure base disposed at the outer radial case structure. The flowpath extends through the aircraft propulsion system from an inlet into the flowpath to an exhaust from the flowpath. The inlet is formed by the front center body between the outer radial case structure and the inner radial case structure.

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 first core turbine rotor of a first propulsion system is disposed in a first core turbine section and is configured to rotate in a first rotational direction. A first power turbine rotor is disposed in a first power turbine section and is configured to rotate in the first rotational direction. The first power turbine rotor is operatively coupled to a first open propulsor rotor. A second core turbine rotor of a second propulsion system is disposed in a second core turbine section and is configured to rotate in the first rotational direction. A second power turbine rotor is disposed in a second power turbine section and is configured to rotate in a second rotational direction that is opposite the first rotational direction. The second power turbine rotor is operatively coupled to a second open propulsor rotor.

Abstract: A gas turbine engine includes an engine core with a compressor section, a combustion section downstream of the compressor section relative to a core flow path, and a turbine section downstream of the combustion section relative to the core flow path. A core casing extends around the compressor section, the combustion section, and the turbine section. A second casing extends around the core casing and a core compartment is between the core casing and the second casing. An electric compressor is in the core compartment. A first port is formed in the second casing and fluidically is connected to the electric compressor. A second port is formed in the core casing in the combustion section and is fluidically connected to the electric compressor.

Abstract: An assembly includes first vanes, second vanes, a rotor and an actuation system. The actuation system includes a first ring and a second ring. The first ring is coupled to each first vane. The second ring is coupled to each second vane. During a first mode, the actuation system is configured to rotate the first ring in a first rotational direction about the axis, rotate the second ring in a second rotational direction about the axis, and pivot each of the first vanes and each of the second vanes in a common first pivot direction. During a second mode, the actuation system is configured to rotate the first ring in the second rotational direction about the axis, rotate the second ring in the first rotational direction about the axis, and pivot each of the first vanes and each of the second vanes in a common second pivot direction.

Abstract: An assembly is provided for an aircraft propulsion system. This assembly includes a propulsion module, a gas generator and an anti-icing system. The propulsion module includes an open propulsor rotor. The gas generator is configured to drive rotation of the open propulsor rotor about an axis. The gas generator includes a flowpath, a compressor section, a combustor section, a turbine section and an exhaust section. The flowpath extends through the compressor section, the combustor section, the turbine section and the exhaust section. The turbine section is axially between the combustor section and the open propulsor rotor along the axis. The anti-icing system is configured to bleed gas from the flowpath in the exhaust section to provide heated gas. The anti-icing system is configured to direct the heated gas to the propulsion module to reduce or prevent ice accumulation on an exterior surface of the propulsion module.

Abstract: A buffer air assembly for an aircraft engine includes a first pressurized air header, at least one first bearing compartment, a first bleed air source, and at least one electric buffer compressor. The at least one first bearing compartment is connected in fluid communication with the first pressurized air header. The first bleed air source is connected in fluid communication with the first pressurized air header by a first bleed check valve. The first bleed air source is configured to direct first pressurized bleed air to the first pressurized air header through the first bleed check valve. The at least one electric buffer compressor is connected in fluid communication with the first pressurized air header by a first buffer compressor check valve.

Abstract: An aircraft assembly includes first and second propulsion systems. Each propulsion system includes an open propulsor rotor and a turbine engine configured to drive rotation of the open propulsor rotor. The turbine engine includes a compressor section, a combustor section, a turbine section, a first rotating structure, a second rotating structure and a flowpath. The flowpath extends through the compressor section, the combustor section and the turbine section with the first bladed rotor disposed between the second bladed rotor and the combustor section along the flowpath. The open propulsor rotor of the first propulsion system is configured to rotate a first rotational direction. The open propulsor rotor of the second propulsion system is configured to rotate a second rotational direction that is opposite the first rotational direction.

Abstract: A gas turbine engine is provided that includes compressor and turbine sections, an engine compartment, an air turbine starter (ATS), and a buffer air augmentation system. The compressor and turbine sections are engaged by a shaft. The ATS is disposed to drive the compressor and turbine sections. The air turbine starter is configured to receive a flow of compressed air from a compressed air source. The air turbine starter has an exhaust port configured to pass a flow of ATS exhaust gas flow. The buffer air augmentation system includes flow ducting configured to selectively provide fluid communication of the flow of ATS exhaust gas between the exhaust port of the air turbine starter and the engine compartment. The system is configured to provide the flow of ATS exhaust gas to the engine compartment during a start-up segment of an engine operational cycle.

Abstract: An aircraft system includes a turbine engine case and an attachment. The turbine engine case includes a sidewall, a first mount and a second mount. The first mount and the second mount are radially outboard of and fixed to the sidewall. The second mount is circumferentially offset from the first mount about the axis. The attachment is configured to mount the turbine engine case to a pylon structure and transfer one or more loads between the turbine engine case and the pylon structure. The attachment comprises a bracket. The bracket is mounted to the first mount through a first mount-bracket pin connection. The first mount-bracket pin connection includes a first mount-bracket connection pin seated in a first mount-bracket connection aperture having an elongated cross-sectional geometry in a reference plane perpendicular to the axis. The bracket is mounted to the second mount through a second mount-bracket pin connection.

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 propulsion system is mounted to an aircraft airframe. An open propulsor rotor of the propulsion system is configured to rotate about a rotational axis in a rotational direction within a plane of rotation. The propulsor rotor includes a propulsor blade with a blade tip and a blade span. The propulsor blade is configured to rotate about the rotational axis through a blade reference position in a reference direction towards a reference point on a fuselage. The propulsor blade is oriented vertically at the blade reference position. The blade tip is vertically offset from the reference point by an offset distance when the propulsor blade is at the blade reference position. The offset distance is equal to or less than fifty percent of the blade span. The reference point is coincident with the plane of rotation at a vertical periphery of the fuselage.

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: An assembly for an aircraft propulsion system includes a first open propulsor rotor, a turbine engine, a pylon structure, a support structure, a second open propulsor rotor and a drive system. The turbine engine includes a flowpath, a compressor section, a combustor section, a turbine section and a rotating structure. The rotating structure includes a turbine rotor in the turbine section. The rotating structure is configured to drive rotation of the first open propulsor rotor. The pylon structure is mounted to the turbine engine. The support structure is arranged with the pylon structure independent of the turbine engine. The second open propulsor rotor is rotationally supported by the support structure. The drive system is configured to drive rotation of the second open propulsor rotor.

Abstract: An aircraft propulsion system, a propulsor section and a core engine section configured to generate an exhaust gas flow that is expanded through a turbine to drive a shaft for driving the fan. A propulsor mount assembly is attached to the propulsor section and supports the propulsor section on an aircraft structure and isolates loads generated in the propulsor section from the core engine section.

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.