Abstract: A variable vane system includes a first variable vane stage with a multiple of first actuator gears, each of the multiple of first actuator gears mounted to a respective variable vane of the first variable vane stage, and a second variable vane stage with a multiple of second stage actuator gears. Each of the multiple of second stage actuator gears mounted to a respective variable vane of the second variable vane stage, an extended geared unison ring driven by the drive gear, wherein the extended geared unison ring spans at least a first variable vane stage and a second variable vane stage, each of the multiple of first and second stage actuator gears driven by the extended geared unison ring.

Abstract: A compressor intermediate case for a gas turbine engine includes a plurality of intermediate case struts joining the compressor intermediate case to an inner engine structure. Each strut of the plurality of intermediate case struts includes a leading edge. A turning scoop is disposed at the leading edge of each strut of the plurality of intermediate case struts. A plurality of diffusers extends radially outwardly from the compressor intermediate case so that each diffuser of the plurality of diffusers engages with a corresponding turning scoop. A substantially annular structural fire wall extends radially outwardly from the compressor intermediate case. An environmental control system manifold is disposed on the compressor intermediate case. The environmental control system manifold includes an exit port.

Abstract: In a featured embodiment, a gas turbine engine has a first compressor rotor driven by a first turbine rotor, and a second compressor rotor driven by a second turbine rotor. The second compressor rotor is upstream of the first compressor rotor and the first turbine rotor is upstream of the second turbine rotor. An air mixing system taps air from a location upstream of the first compressor rotor for delivery to an environmental control system. The air mixing system receives air from a first air source and a second air source. The first air source includes air at a first pressure upstream of the first compressor rotor. The second air source includes air at a lower second pressure. At least one valve controls a mixture of air from the first and second sources to achieve a predetermined pressure for the environmental control system.

Abstract: A gas turbine engine has a first source of air to be delivered into a core of the engine, and a second source of air, distinct from the first source of air and including separately controlled first and second fans, each delivering air into respective first and second conduits connected to distinct auxiliary applications.

Abstract: A gas turbine engine includes a harmonic drive driven by an actuator, a drive shaft driven by the harmonic drive, the drive shaft with a first drive gear and a second drive gear. A first variable vane stage with a first actuator gear to direct drive one of a multiple of variable vanes, the first actuator gear meshed with the first drive gear; and a second variable vane stage with a second actuator gear to direct drive one of a multiple of variable vanes, the second actuator gear meshed with the second drive gear.

Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream most end, and more upstream locations. A turbine section has a high pressure turbine. A first tap taps air from at least one of the more upstream locations in the main compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger. A second tap taps air from a location closer to the downstream most end than the location(s) of the first tap. The first and second tap mix together and are delivered into the high pressure turbine. An intercooling system for a gas turbine engine is also disclosed.

Abstract: A variable vane system includes a harmonic drive driven by an actuator, and a drive gear driven by the harmonic drive. A geared unison ring includes an actuator gear rack driven by the drive gear and a multiple of vane gears, each of the multiple of vane gears mounted to one of a multiple of variable vanes, the multiple of vane gears driven by a gear rack of the geared unison ring.

Abstract: A gas turbine engine comprises a gas generator rotating along a first axis of rotation, with at least one compressor rotor, at least one gas generator turbine rotor and a combustion section. A fan drive turbine rotates along a second axis of rotation, downstream of at least one gas generator turbine rotor. The fan drive turbine drives a pair of shaft portions extending in opposed directions beyond the axis of rotation of the gas generator.

Abstract: A variable vane system for a gas turbine engine, comprising an actuator, a harmonic drive comprising a strain wave gearing mechanism comprising a fixed circular spline, a flex spline attached to an output shaft, and a wave generator attached to an input shaft driven by the actuator, the flex spline driven by the wave generator with respect to the circular spline, a drive gear driven by the output shaft along a first axis, and an actuator gear mounted to one of a multiple of variable vanes along a second axis transverse to the first axis, the actuator gear driven by the drive gear.

Abstract: An aircraft engine includes a gas generator, a turbine fluidly connected to the gas generator, and a fan connected to the turbine via a shaft. The fan is positioned aft of the turbine, and the shaft is at least partially disposed in a fan inlet flowpath.

Abstract: An electromechanical component arrangement for a gas turbine engine includes a mechanical component located at a first side of a firewall of a gas turbine engine and an electronic module of the electromechanical component in communication with the mechanical component separated from the mechanical component by a firewall, the firewall comprising a first side and a second side, the second side having a lower operating temperature than the first side. A vibration isolation structure is located at the second side. The electronic module is connected thereto and includes at least one vibration isolator secured to the firewall to vibrationally isolate the electronic module from gas turbine engine vibrations.

Abstract: A gas turbine engine includes a core engine assembly including a compressor section communicating air to a combustor section where the air is mixed with fuel and ignited to generate a high-energy gas flow that is expanded through a turbine section. The turbine section is coupled to drive the compressor section. A propulsor section aft of the core engine is driven by the turbine section. An exhaust duct routing exhaust gases around the propulsor section. The exhaust duct includes an inlet forward of the propulsor section, an outlet aft of the turbine section and a passageway between the inlet and the outlet. An aircraft and exhaust duct are also disclosed.

Abstract: A gas turbine engine has a first source of air to be delivered into a core of the engine, and a second source of air, distinct from the first source of air and including separately controlled first and second fans, each delivering air into respective first and second conduits connected to distinct auxiliary applications.

Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream most end, and more upstream locations. A turbine section has a high pressure turbine. A first tap taps air from at least one of the more upstream locations in the main compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger. A second tap taps air from a location closer to the downstream most end than the location(s) of the first tap. The first and second tap mix together and are delivered into the high pressure turbine. An intercooling system for a gas turbine engine is also disclosed.

Abstract: A gas turbine engine has a nose cone, a fan for delivering air into a bypass duct as bypass flow, and into a core engine to be delivered to a compressor. The nose cone includes a vent to receive air and deliver the air across a heat exchanger, which receives a fluid to be cooled. The air from the vents is delivered to an outlet downstream of the heat exchanger, such that a majority of the air being delivered to the outlet becomes part of the bypass flow.

Abstract: A compressor intermediate case for a gas turbine engine includes a plurality of intermediate case struts joining the compressor intermediate case to an inner engine structure. Each strut of the plurality of intermediate case struts includes a leading edge. A turning scoop is disposed at the leading edge of each strut of the plurality of intermediate case struts. A plurality of diffusers extends radially outwardly from the compressor intermediate case so that each diffuser of the plurality of diffusers engages with a corresponding turning scoop. A substantially annular structural fire wall extends radially outwardly from the compressor intermediate case. An environmental control system manifold is disposed on the compressor intermediate case. The environmental control system manifold includes an exit port.

Abstract: A decoupled gas turbine engine includes a high spool assembly and a low spool assembly each having a rotational axis that are spaced from one-another. The engine further includes a combustor that may have a centerline spaced from the rotational axes of each spool assembly. Turning ducts of the engine are configured to re-direct airflow from one spool assembly to the next and/or between one spool assembly and the combustor.

Abstract: A gas turbine engine according to an example of the present disclosure includes a mount arrangement including a gas turbine engine, an adapter ring extending circumferentially around the gas turbine engine and mounted to both the gas turbine engine and to an aircraft beneath the gas turbine engine, and at least one thrust link including a first end mounted to the gas turbine engine fore of the adapter ring, and an opposite second end mounted to the adapter ring above the gas turbine engine.

Abstract: A gas turbine engine includes an actuator, a harmonic drive driven by the actuator, an engine case containing a multiple of variable vanes, a firewall transverse to the engine case, a first temperature on one side of the firewall and a second temperature on the other side of the firewall, the first temperature different than the second temperature and a drive shaft driven by the harmonic drive, the drive shaft extends through the firewall to operate the multiple of variable vanes.

Abstract: An active clearance control assembly for a gas turbine engine includes a firewall, a fluid intake and an active clearance control manifold. A conduit is configured to direct a fluid from the fluid intake on a first axial side of the firewall through the firewall to at least one active clearance control manifold on a second axial side of the firewall. A valve is located on the first axial side of the firewall and is configured to regulate the flow of the fluid through the conduit.