Abstract: A turbofan engine includes a propulsor means, an epicyclic gear system, a gear system input lateral stiffness is less than 5% of a gear mesh lateral stiffness, first compression means for compressing airflow from the propulsor means, the first compression means includes an entrance at a splitter, and at least one propulsor blade of the propulsor means is forward of the splitter relative to an engine longitudinal axis, second compression means for compressing airflow from the first compression means, first expansion means for driving the second compression means, the first expansion means includes an inlet and an outlet, second expansion means for driving the input of the gear system, and a performance quantity ratio of a first performance quantity to a second performance quantity is between 0.5 and 1.5.

Abstract: A gas turbine engine is provided that includes fan, compressor, combustion, and turbine sections, and an entrained debris removal system. The compressor section includes a compressor inlet disposed downstream of the fan section. A core gas path begins at the compressor inlet and extends through the compressor section. The bypass air path is disposed downstream of the fan section. The structural member extends radially across the bypass air path. The entrained debris removal system includes an inlet port engaged with the core gas path, a passage that extends through the structural member, and an exit port that is disposed in contact with ambient air. The entrained debris removal system is configured to selectively receive core gas from the core gas path and pass the core gas through the passage that extends through the structural member and out of the exit port.

Abstract: A turbofan engine includes a propulsor section that has a propulsor shaft in driving engagement with a propulsor. An epicyclic gear system has a gear mesh lateral stiffness and a gear mesh transverse stiffness. A gear system input defines a gear system input lateral stiffness and a gear system input transverse stiffness. The gear system input lateral stiffness is less than 5% of the gear mesh lateral stiffness. A first turbine section rotates at a first speed, and a second turbine rotates at a second speed that is faster than the first speed. A first performance quantity is defined as the product of the first speed squared and the first area of the first turbine, a second performance quantity is defined as the product of the second speed squared and the second area of the second turbine, and a performance quantity ratio is between 0.5 and 1.5.

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 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 compressor section having a downstream most location and more upstream locations. A compressor housing surrounds the compressor rotors. A combustor is downstream of the compressor and a turbine section is downstream of the combustor. At least one bleed inlet passage bleeds air compressed from the compressor section, and delivers the air for use on an associated aircraft. A baffle plate is intermediate the at least one bleed inlet passage and a bleed exit port from the gas turbine engine. The baffle plate has a plurality of holes for allowing the air to flow from the bleed inlet passage to the bleed exit port, with an area of the holes increasing from a location adjacent the at least one bleed inlet passage and in a circumferential direction away from the at least one bleed inlet passage. An aircraft air supply system and an aircraft are also disclosed.

Abstract: An engine system comprises a gas turbine engine that has a spool that includes a compressor rotor and a turbine, and a combustor connected to receive compressed air from the compressor rotor. There is an electric engine motor electrically connected to a battery and coupled to drive the spool. A lubrication circuit includes a bearing compartment, a scavenger line connected with the bearing compartment, and a mechanically-driven scavenger pump operable to suction the bearing compartment via the scavenger line. A bearing compartment pump is located in the scavenge line between the bearing compartment and the scavenger pump. The bearing compartment pump is electrically connected to the battery and is operable to reduce pressure in the bearing compartment independently of operation of the mechanically-driven scavenger pump.

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.

Abstract: An environmental control system bleed cooling system including bleed air from the environmental control system bleed air system; a heat exchanger fluidly coupled between the environmental control system bleed air system and a gas turbine engine; and ambient cooling air fluidly coupled with the heat exchanger, wherein the ambient cooling air is configured to remove thermal energy from the bleed air flowing through the heat exchanger.

Abstract: An environmental control system bleed cooling system including bleed air from the environmental control system bleed air system; a heat exchanger fluidly coupled between the environmental control system bleed air system and a gas turbine engine; and ambient cooling air fluidly coupled with the heat exchanger, wherein the ambient cooling air is configured to remove thermal energy from the bleed air flowing through the heat exchanger.

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 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 includes an open rotor propulsor for 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. An exit guide vane extends radially outward of the outer core panel. The exit guide vane has a pressure side and a suction side defined between a leading edge and a trailing edge. The exit guide vane includes at least one air scoop configured to capture the air from the open rotor propulsor and direct the captured air into the exit guide vane. A heat exchanger is positioned in the outer core pane. The captured air is passed over the heat exchanger and is configured to cool another fluid.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a propulsor, a compressor section, a turbine section and an epicyclic gear system. The turbine section includes a first turbine and a second turbine. The second turbine drives the propulsor through the epicyclic gear system. The epicyclic gear system includes a carrier. A drive shaft interconnects the carrier and the propulsor. A frame supports at least a portion of the drive shaft. A flexible support at least partially supports the gear system relative to an engine static structure.

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 bleed air cooling system for a gas turbine engine includes a bleed port located at an axial location of the gas turbine engine to divert a bleed airflow from a gas turbine engine flowpath, a bleed outlet located at a cooling location of the gas turbine engine and a bleed duct in fluid communication with the bleed port and the configured to convey the bleed airflow from the bleed port to the bleed outlet. A modulating valve is located at the bleed duct and is movable between a fully open position and a fully closed position to regulate the bleed airflow through the bleed duct based on one or more operating conditions of the gas turbine engine.

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 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.