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: An engine driven environmental control system (ECS) air circuit includes a gas turbine engine having a compressor section. The compressor section includes a plurality of compressor bleeds. A selection valve selectively connects each of said bleeds to an input of an intercooler. A second valve is configured to selectively connect an output of said intercooler to at least one auxiliary compressor. The output of each of the at least one auxiliary compressors is connected to an ECS air input.

Abstract: An exemplary gas turbine engine assembly includes a fan section including a fan, a first spool having a first turbine operatively mounted to a first turbine shaft, and a second spool having a second turbine operatively mounted to a second turbine shaft. The first and second towershafts are respectively coupled to the first and second turbine shafts. An accessory drive gearbox includes a set of gears. A compressor is driven by the first towershaft. A transmission couples a starter generator assembly to the set of gears. The transmission is transitionable between a first mode where the starter generator assembly is driven at a first speed relative to the second towershaft in response to rotation of the second towershaft, and a second mode where the starter generator assembly is driven at a different, second speed relative to the second towershaft in response to rotation of the second towershaft.

Abstract: A turbine exhaust assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a turbine exhaust case comprised of CMC material and attachable to a turbine case, a tail cone comprised of CMC material that has a leading edge and a trailing edge, and an exhaust mixer comprised of CMC material and coupled to the turbine exhaust case. The exhaust mixer has a plurality of lobes arranged about the tail cone to define an exhaust flow path. A plurality of struts extend from the tail cone to support the exhaust mixer at a location aft of the leading edge of the tail cone. A method of assembling a propulsion system is also disclosed.

Abstract: A boundary layer ingestion engine includes a fan section configured to extend into a boundary layer of a full annulus of an aft end of a fuselage of an aircraft. The fan section includes a first fan stage and a second fan stage. The boundary layer ingestion engine also includes a differential planetary gear system is operable to transform rotation of an input shaft into counter rotation of a first shaft coupled to the first fan stage and a second shaft coupled to the second fan stage. The boundary layer ingestion engine further includes a motor operable to drive rotation of the input shaft.

Abstract: An integrated propulsion system according to an example of the present disclosure includes, among other things, components that include a gas turbine engine, a nacelle assembly and a mounting assembly, the system designed by a process comprising identifying two or more of internal structural loading requirements, external structural mount loading requirements, aerodynamic requirements, and acoustic requirements for the system, and interdependently designing said components to meet said requirements. The nacelle assembly includes a fan nacelle and an aft nacelle, the fan nacelle arranged at least partially about a fan and the engine, and the fan nacelle arranged at least partially about a core cowling to define a bypass flow path.

Abstract: A gas turbine engine includes a shaft and a hub supported by the shaft. A housing includes an inlet and an intermediate case that respectively provide an inlet and an intermediate case flow path. A rotor is connected to the hub and supports a compressor section arranged axially between the inlet and the intermediate case flow paths. A compressor section inlet has a radially inner boundary that is spaced a second radial distance from the rotational axis different from the first radial distance. First and second bearings support the shaft relative to the intermediate case and the inlet case, respectively. An inner race of the first bearing and an inner race of the second bearing engage and rotate with the hub. A fan shaft is drivingly connected to a fan having fan blades. A gear system is connected to the fan shaft and driven through a flex shaft.

Abstract: A gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan rotor. Aspects of the gear system are provided with some flexibility. The fan drive turbine has a first exit area and rotates at a first speed. A second turbine section has a second exit area and rotates at a second speed, which is faster than said first speed. A performance quantity can be defined for both turbine sections as the products of the respective areas and respective speeds squared. A performance quantity ratio of the performance quantity for the fan drive turbine to the performance quantity for the second turbine section is relatively high.

Abstract: A disclosed gas turbine engine includes a fan section delivering air into a compressor section. An environmental control system includes a higher pressure tap at a higher pressure location in the compressor section, and a lower pressure tap at a lower pressure location. The lower pressure tap communicates to a first passage leading to a downstream outlet, and having a second passage leading into a compressor section of a turbocompressor. The higher pressure tap leads into a turbine section of the turbocompressor such that air in the higher pressure tap drives the turbine section to in turn drive the compressor section of the turbocompressor. A combined outlet of the compressor section and the turbine section of the turbocompressor intermixes and passes downstream to be delivered to an aircraft use.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan section including a fan rotatable about an engine axis with a plurality of fan blades rotatable about a fan blade axis. A geared architecture is in communication with the fan and driven by a turbine section. The fan rotates at a first speed and the turbine section rotates at a second speed different from the first speed and a fixed area fan nozzle in communication with the fan section.

Abstract: An aircraft propulsion system includes a fan section that includes a fan shaft that is rotatable about a fan axis. The fan shaft includes a fan gear. The aircraft propulsion system also includes a boost turbine engine that includes a first output shaft that includes a first gear that is coupled to the fan gear. The boost turbine engine has a first maximum power capacity. The aircraft propulsion system further includes a cruise gas turbine engine that includes a second output shaft that includes a second gear that is coupled to the fan gear. The cruise turbine engine has a second maximum power capacity that is less than the first maximum power capacity of the boost turbine engine. The fan section produces a thrust that corresponds to power input through the fan gear from the boost turbine engine and the cruise turbine engine.

Abstract: A gas turbine engine includes a gear system that provides a speed reduction between a fan drive turbine and a fan rotor. Aspects of the gear system are provided with defined flexibility. The fan drive turbine has a first exit area and rotates at a first speed. A second turbine section has a second exit area and rotates at a second speed, which is faster than said first speed. A performance quantity can be defined for both turbine sections as the products of the respective areas and respective speeds squared. A performance quantity ratio of the performance quantity for the fan drive turbine to the performance quantity for the second turbine section is between 0.5 and 1.5.

Abstract: A hybrid electric gas turbine engine includes a fan section having a fan, a turbine section having a turbine drivably connected to the fan through a main shaft that extends along a central longitudinal axis, a gas generating core extending along a first axis that is radially offset from the central longitudinal axis, and an electric motor drivably connected to the main shaft, wherein the electric motor is colinear with the main shaft.

Abstract: A gas turbine engine includes a main compressor section with a downstream most location. A turbine section has a high pressure turbine. A tap line is connected to tap air from a location upstream of the downstream most location in the main compressor section. The tapped air is connected to a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and is connected to deliver air into the high pressure turbine. A bypass valve is positioned downstream of the main compressor section, and upstream of the heat exchanger. The bypass valve selectively delivers air directly to the cooling compressor without passing through the heat exchanger under certain conditions.

Abstract: A gas turbine engine includes a main compressor section with a downstream most location. A turbine section has a high pressure turbine. A tap line is connected to tap air from a location upstream of the downstream most location in the main compressor section. The tapped air is connected to a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger, and is connected to deliver air into the high pressure turbine. A bypass valve is positioned downstream of the main compressor section, and upstream of the heat exchanger. The bypass valve selectively delivers air directly to the cooling compressor without passing through the heat exchanger under certain conditions.

Abstract: An exemplary gas turbine engine assembly includes a first spool having a first turbine operatively mounted to a first turbine shaft, and a second spool having a second turbine operatively mounted to a second turbine shaft. The first and second turbines are mounted for rotation about a common rotational axis within an engine static structure. The first and second turbine shafts are coaxial with one another. First and second towershafts are respectively coupled to the first and second turbine shafts. An accessory drive gearbox has a set of gears. A compressor is driven by the first towershaft. The engine assembly further includes a starter generator assembly, and a transmission coupling the starter generator assembly to the first set of gears.

Abstract: A turbine engine is disclosed that includes a fan case surrounding a fan rotatable about an axis. A core is supported relative to the fan case by support structure arranged downstream from the fan. The core includes a core housing having an inlet case arranged to receive airflow from the fan. A compressor case is arranged axially adjacent to the inlet case and surrounds a compressor stage having a rotor blade with a blade trailing edge. The support structure includes a support structure leading edge facing the fan and a support structure trailing edge on a side opposite the support structure leading edge. The support structure trailing edge is arranged axially forward of the blade trailing edge. In one example, a forward attachment extends from the support structure to the inlet case.

Abstract: An embodiment of a turbine engine outer case ring assembly includes a substantially circular first main body and a sealant layer covering at least a main gas-facing surface of the inner gas path side, the sealant layer having at least one coating sufficient to prevent infiltration of working gas into pores of the main body.

Abstract: A turbofan engine according to an example of the present disclosure includes, among other things, a fan including a circumferential array of fan blades, a low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, and a fan duct annulus area outboard of the low pressure compressor section inlet, and a fan drive turbine section. The fan drive turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is equal to or greater than 0.35, and is less than 0.55.

Abstract: A turbofan gas turbine engine includes, among other things, a fan section, a core engine, a bypass passage, and a bypass ratio defined as the volume of air passing into the bypass passage compared to the volume of air passing into the core engine, the bypass ratio being greater than or equal to 10. A gear arrangement drives the fan section. A compressor section includes a low pressure compressor section and a high pressure compressor section. A turbine section drives the gear arrangement. An overall pressure ratio is provided by the combination of a pressure ratio across the low pressure compressor section and a pressure ratio across the high pressure compressor section, and greater than 40. The pressure ratio across the high pressure compressor section is between 7 and 15, and the pressure ratio across the low pressure compressor section is between 4 and 8.