Abstract: A buffer air assembly for an aircraft engine includes a low-pressure header, a high-pressure header, a low-pressure bleed air source, a high-pressure bleed air source, and an electric buffer compressor. The low-pressure header is connected to at least one low-pressure bearing compartment. The high-pressure header is connected to at least one high-pressure bearing compartment. The low-pressure bleed air source is connected to the low-pressure header. The low-pressure bleed air source is configured to direct a low-pressure buffering air to the at least one low-pressure bearing compartment through the low-pressure header. The high-pressure bleed air source is configured to direct a high-pressure buffering air to the at least one high-pressure bearing compartment through the high-pressure header. The electric buffer compressor is connected to the low-pressure header and the high-pressure header.

Abstract: A tail cone ventilation system including a tail cone case defining a tail cone interior and a tail cone exterior, the tail cone having a forward portion and an aft portion separated axially along a tail cone axis; a distribution manifold located within the tail cone interior proximate the forward portion, wherein the distribution manifold comprises nozzles arranged radially around the axis, the nozzles configured to direct a cooling air over at least one electronic component within the tail cone interior and along an inner surface of the tail cone case; an air inlet fluidly coupled with the distribution manifold through ducting, the air inlet located externally from the tail cone interior; and a tail cone discharge located proximate the tail cone aft portion, the tail cone discharge being fluidly coupled with the distribution manifold.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, propulsor means, first compression means, second compression means, reduction means for reducing a rotational speed of an output that drives the propulsor means relative to an input, first expansion means, and second expansion means. The reduction means includes an epicyclic gear system with a gear reduction. The gear system is straddle-mounted by first and second bearings on opposite sides of the gear reduction relative to an engine longitudinal axis.

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 and a ventilation system. The flowpath extends through the compressor section, the combustor section and the turbine section from an inlet into the flowpath to an exhaust from the flowpath. The tail cone structure is arranged at the exhaust from the flowpath. The ventilation system is configured to bleed compressed air from the flowpath along the compressor section to provide ventilation air. The ventilation system is configured to direct the ventilation air into an internal volume in the tail cone structure.

Abstract: An assembly for a gas turbine engine includes at least one rotational assembly, an engine static structure, a compressor, one or more compressed air loads, and a particulate separator assembly. The at least one rotational assembly includes a shaft, a bladed compressor rotor, and a bladed turbine rotor. The engine static structure includes an engine case assembly. The engine case assembly surrounds the at least one rotational assembly. The compressor includes the bladed compressor rotor. The compressor is configured to form a compressed air flow. The one or more compressed air loads are disposed within the engine case assembly. The particulate separator assembly includes a plurality of particulate separators. The plurality of particulate separators are disposed outside of the engine case assembly. The plurality of particulate separators are configured to separate particulate from the compressed air flow and direct the compressed air flow to the one or more compressed air loads.

Abstract: A gas turbine engine includes a compressor section including a first compressor, a turbine section including a first turbine and a second turbine, a first shaft and a second shaft, the first shaft interconnecting the first turbine and the second compressor, and a geared architecture. The first shaft is supported on a first bearing in an overhung manner. A performance ratio is between 0.5 and 1.5.

Abstract: A gas turbine engine is provided that includes a high pressure compressor (HPC), a combustor section, a high pressure turbine (HPT), and a bypass tangential on board injector (TOBI) system. The combustor section has a combustor. A core gas path extends through the HPC, the combustor section, and the HPT. The bypass TOBI system extends circumferentially around the engine axial centerline, and has a plurality of nozzles, inner and outer radial sides, a plurality of first type and second type radial passages configured to allow the gas from the HPC to pass from the inner radial side of the bypass TOBI system to the outer radial side of the bypass TOBI system, wherein the first type radial passages are differently configured from the second type radial passages.

Abstract: A gas turbine engine includes an open rotor configured to rotate with a nose cone. A heat exchanger is positioned within the nose cone and a system for using a working fluid. An inlet from the system is connected to the heat exchanger and an outlet from the heat exchanger connected back to the system. A central opening is in a central portion of the nose cone to deliver cooling air across the heat exchanger, and a duct downstream of the heat exchanger to direct the cooling air radially to at least one static opening such that the cooling air can move radially outwardly through static structure and be directed into a propulsion airflow path. A method and a heat exchange system are also disclosed.

Abstract: A gas turbine engine includes a rotor having a plurality of blades without an outer housing, and rotating with a nose cone. The rotor, plurality of blades and nose cone provide a rotating structure. A static heat exchanger is positioned within the nose cone and a system for using a working fluid. An inlet from the system is connected to the heat exchanger and an outlet from the heat exchanger is connected back to the system. A central opening is in a central portion of the nose cone to deliver cooling air across the heat exchanger. A duct downstream of the heat exchanger directs the cooling air such that the cooling air can move radially outwardly through the rotating structure such that the cooling air can be directed into a propulsion airflow path. A method and a heat exchange system are also disclosed.

Abstract: An assembly for a gas turbine engine includes at least one rotational assembly, an engine static structure, a compressor, one or more compressed air loads, and a particulate separator assembly. The at least one rotational assembly includes a shaft, a bladed compressor rotor, and a bladed turbine rotor. The engine static structure includes an engine case assembly. The engine case assembly surrounds the at least one rotational assembly. The compressor includes the bladed compressor rotor. The compressor is configured to form a compressed air flow. The one or more compressed air loads are disposed within the engine case assembly. The particulate separator assembly includes a plurality of particulate separators. The plurality of particulate separators are disposed outside of the engine case assembly. The plurality of particulate separators are configured to separate particulate from the compressed air flow and direct the compressed air flow to the one or more compressed air loads.

Abstract: A particle separator configured to be disposed within a turbine engine diffuser flow path is provided. The particle separator includes a housing that extends between a forward end and an aft end, first and second swirl vane stages, and first and second air flow exits. The first swirl vane stage is configured to cause an air flow passing through the first swirl vane stage to swirl within the housing as the air flow travels axially in a direction from the forward end toward the aft end of the housing. The second swirl vane stage is configured to cause a first portion of the air flow passing through the second swirl vane stage to redirect into a substantially axially directed flow that enters the second air flow exit.

Abstract: A particle separator configured to be disposed within a turbine engine diffuser flow path is provided. The particle separator includes a housing, a hollow cavity member, and a helical member. The housing has a forward end inlet, an exterior wall, an aft wall, an interior cavity, and an aft exhaust passage extending outwardly from the exterior wall adjacent the aft wall. The hollow cavity member has a first portion, an interior region, and an open aft end. The first portion is disposed within the housing interior cavity and includes perforations. The helical member is disposed within the housing interior cavity. The helical member forms a helical passage between the housing exterior wall and the first portion of the hollow cavity member. The helical passage has a passage inlet disposed adjacent to the housing forward end inlet and a terminal end in communication with the aft exhaust passage.

Abstract: A tie-shaft with thermal barrier coating system including a tie-shaft including an inside diameter and an outside diameter opposite the inside diameter, an exterior surface proximate the outside diameter; a piston-ring seal in operative communication with the tie-shaft proximate the outside diameter; and the thermal coating system disposed on the exterior surface of the tie-shaft proximate the piston-ring seal.

Abstract: A buffer air assembly for an aircraft engine includes a low-pressure header, a high-pressure header, a low-pressure bleed air source, a high-pressure bleed air source, and an electric buffer compressor. The low-pressure header is connected to at least one low-pressure bearing compartment. The high-pressure header is connected to at least one high-pressure bearing compartment. The low-pressure bleed air source is connected to the low-pressure header. The low-pressure bleed air source is configured to direct a low-pressure buffering air to the at least one low-pressure bearing compartment through the low-pressure header. The high-pressure bleed air source is configured to direct a high-pressure buffering air to the at least one high-pressure bearing compartment through the high-pressure header. The electric buffer compressor is connected to the low-pressure header and the high-pressure header.

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: 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 according to an example of the present disclosure includes, among other things, propulsor means, first compression means, second compression means, reduction means for reducing a rotational speed of an output that drives the propulsor means relative to an input, first expansion means, and second expansion means. The reduction means includes an epicyclic gear system with a gear reduction. The gear system is straddle-mounted by first and second bearings on opposite sides of the gear reduction relative to an engine longitudinal axis.

Abstract: A bowed-rotor-resistant low shaft of a gas turbine engine is provided. The bowed-rotor-resistant low shaft includes a low shaft and a thermal barrier coating applied to at least a lengthwise portion of an outside diameter of the low shaft, wherein the thermal barrier coating reduces bowing of the low shaft due to latent heating effects from the gas turbine engine.

Abstract: A gas turbine engine is provided having an axial centerline, a compressor section, a turbine section, and a combustor section. The turbine section has a turbine first vane assembly that includes an annular first vane inner radial support. The combustor section has an outer casing, an annular combustor, and a unitary inner diffuser structure. The annular combustor has an inner radial flange. The unitary inner diffuser structure includes a compressor discharge, an inner diffuser case, and a tangential onboard injector (TOBI) inseparably attached to one another. The unitary inner diffuser structure further includes an outer radial flange and a TOBI connection flange. The annular FV inner radial support, the combustor inner radial flange, and the TOBI connection flange are secured to one another.

Abstract: A gas turbine engine is provided that includes compressor and combustor sections, inner and outer casings, an annular diffuser, an inner diffuser casing, a heat exchanger, and an HPT stator vane stage. An annular combustor is disposed radially inward of the outer casing and has inner and outer radial wall structures. The outer casing and the combustor outer radial wall structure define a diffuser OD flow path. The annular diffuser directs diffuser gas towards the combustor section. The inner diffuser casing is disposed radially inward of the annular combustor and spaced apart from the combustor inner radial wall structure. The inner casing is disposed radially inward of and spaced apart from the inner diffuser casing. The inner diffuser casing and the inner casing define an ICF passage. The heat exchanger is configured to produce intercooler gas. Intercooler gas is directed through the ICF passage and into the HPT stator vanes.