Abstract: A gas turbine engine includes a main engine compressor section. A booster compressor changes a pressure of airflow received from the main engine compressor section to a pressure desired for a pneumatic system. The booster compressor is configured to operate at airflow conditions greater than a demand of the pneumatic system. An exhaust valve controls airflow between an exhaust outlet and an outlet passage to the pneumatic system. The exhaust valve is operable to exhaust airflow from the booster compressor in excess of the demand of the pneumatic system. A bleed air system for a gas turbine engine and a method of controlling engine bleed airflow are also disclosed.

Abstract: A propulsion system for an aircraft includes a furcated nozzle that has at least a first duct that extends from a first propulsor to a first trailing end duct opening and a second duct that extends from a second propulsor to a second trailing end duct opening.

Abstract: A hybrid electric gas turbine engine includes a fan section, a turbine section, a first gas generating core, and a generator assembly. The fan section has a fan. The turbine section has a turbine drivably connected to the fan through a main shaft that extends along a central longitudinal axis. The first gas generating core extends along a first axis that is radially offset from the central longitudinal axis. The generator assembly is drivably connected to the main shaft through a gear assembly.

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, passing the tapped air through a first heat exchanger and then to a cooling compressor. A second tap taps air from a location closer to the downstream most end than the location of the first tap, and the first and second taps mix together and are delivered into the high pressure turbine. The cooling compressor is positioned downstream of the first heat exchanger, and upstream of a location where air from the first and second taps mix together.

Abstract: A cross flow fan to be incorporated into an aircraft fuselage comprises an ingestion fan rotor to be positioned in a tail section of an aircraft fuselage to reduce boundary layer air from a top surface of the fuselage and to drive the air away from the top surface, and a drive arrangement for the ingestion fan rotor. An aircraft is also disclosed.

Abstract: A cable drive system for variable vane operation comprising an actuator, a harmonic drive driven by the actuator, a cable drive system driven by the harmonic drive; and a unison ring driven by the cable drive system.

Abstract: A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream discharge, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the 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, and delivers air into the high pressure turbine. An intercooling system for a gas turbine engine is also disclosed.

Abstract: A variable vane system includes a unison ring driven by a harmonic drive. A multiple of pins extend from the unison ring. A multiple of variable vanes are driven by the unison ring, each of the multiple of variable vanes connected to the unison ring through a respective drive arm with a slot that receives one of the multiple of pins to accommodate axial motion a support that extends from an engine case to support the unison ring and a multiple of rollers to support the unison ring.

Abstract: A gas turbine engine comprises a core engine housing. A nacelle is positioned radially outwardly of the core engine housing. An outer bypass housing is positioned outwardly of the nacelle. There is at least one accessory to be cooled positioned in a chamber radially between the core engine housing and the nacelle. A manifold delivers cooling air into the chamber, and extends ng circumferentially about a central axis of the core engine. The nacelle has an asymmetric flow cross-section across a circumferential extent.

Abstract: A variable vane system including an actuator; a harmonic drive driven by the actuator, a multi-planar drive gear driven by the harmonic drive, a first actuator gear mounted to a first variable vane of a first variable vane stage, the first actuator gear driven by the multi-planar drive gear, the first actuator gear comprises a drive arm, and a second actuator gear mounted to a second variable vane of a second variable vane stage, the second actuator gear driven by the multi-planar drive gear.

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: A propulsion system for an aircraft has at least two fans with each fan having a fan drive shaft. A turboshaft gas turbine engine drives each of the at least two fans. The turboshaft engine drives an output shaft which drives a gear to, in turn, engage to drive a gear on a first intermediate shaft extending from the turboshaft gas turbine engine in a rearward direction toward an intermediate fan drive shaft. The intermediate fan drive shaft drives the fan drive shaft, and the at least two first intermediate drive shafts extending over a distance that is greater in an axial dimension defined between the turboshaft gas turbine engine and the fan than it is in a width dimension defined between the at least two fans.

Abstract: A drive arrangement for an aircraft comprises a pair of propulsor units each having a fan and a fan shaft for driving the fan. A core engine has a turbine driving a core engine shaft. A mechanical connection connects the core engine shaft to drive the fan shafts for each of the propulsor units. An aircraft also has such an arrangement.

Abstract: An aircraft is provided that includes a wing and a pusher fan engine. The pusher fan engine is configured in the wing.

Abstract: An actuator system including a harmonic drive operable to drive a variable vane system of a gas turbine engine. A drive arm is driven by the harmonic drive to drive a unison ring. A multiple of variable vanes driven by the unison ring.

Abstract: A propulsor and mount arrangement comprises a propulsor rotor and a fan casing surrounding the propulsor rotor. The fan casing receives two side mounts and a thrust link pivotally attached to the fan casing at a location that will be within 10° of a vertically lowermost location when the propulsor is mounted on an aircraft, and the side mounts being at circumferentially opposed positions, and within a lower 270° when the propulsor is mounted on an aircraft. At least a portion of both the side mounts, and a pivot point connect the thrust link to the fan casing in a common plane defined perpendicular to a rotational axis of the propulsor rotor. An aircraft is also disclosed.

Abstract: A gas turbine engine includes a fluid intake. A turbine section includes a turbine case. A firewall is located upstream of the turbine section. A conduit is configured to direct a fluid from the fluid intake to the turbine case. A valve is located on a first side of the firewall opposite from the turbine section and is configured to regulate a flow of the fluid through the conduit.

Abstract: An aircraft engine includes a gas powered turbine core. A first fan is connected to the turbine core via a shaft. The fan is positioned aft of the turbine. A second fan is connected to the first fan via a geared connection.

Abstract: A compressor section for use in a gas turbine engine comprises a compressor rotor having a hub and a plurality of blades extending radially outwardly from the hub and an outer housing surrounding an outer periphery of the blades. A tap taps air at a radially outer first location, passing the tapped air through a heat exchanger, and returning the tapped air to an outlet at a second location which is radially inward of the first location, to provide cooling air adjacent to the hub. A gas turbine engine is also disclosed.

Abstract: A variable vane system for a gas turbine engine includes a geared unison ring driven by a drive gear, the geared unison ring axially slidable parallel to an engine axis with respect to the drive gear as the rotation of the unison ring results in axial motion that is accommodated by the axial sliding between a gear mesh of the geared unison ring and the drive gear.