Abstract: A gas generator has at least one compressor rotor, at least one gas generator turbine rotor and a combustion section. A fan drive turbine is positioned downstream of a path of the products of combustion having passed over the at least one gas generator turbine rotor. The fan drive turbine drives a shaft and the shaft engages gears to drive at least three fan rotors.

Abstract: A gas turbine engine includes a main compressor. A tap is fluidly connected downstream of the main compressor. A heat exchanger is fluidly connected downstream of the tap. An auxiliary compressor unit is fluidly connected downstream of the heat exchanger. The auxiliary compressor unit is configured to compress air cooled by the heat exchanger with an overall auxiliary compressor unit pressure ratio between 1.1 and 6.0. An intercooling system for a gas turbine engine is also disclosed.

Abstract: A boundary layer ingestion engine includes a gas generator and a turbine fluidly connected to the gas generator. A fan is mechanically linked to the turbine via a shaft such that rotation of the turbine is translated to the fan. A boundary layer ingestion inlet is aligned with an expected boundary layer, such that the boundary layer ingestion inlet is configured to ingest fluid from a boundary layer during operation of the boundary layer ingestion engine.

Abstract: A gas turbine engine includes a core engine including a central engine axis and a nacelle surrounding the core engine. At least a portion of the nacelle is axially movable relative to the core engine between open and fully closed positions. A ceramic-based liner is located at an aft portion of the core engine. The ceramic-based component mechanically interfaces with the movable portion of the nacelle when the nacelle is in the fully closed position. A turbine section and a method of accommodating thermally-induced dimensional change of engine components 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 boundary layer ingestion engine includes a gas generator and a turbine fluidly connected to the gas generator. A fan is mechanically linked to the turbine via a shaft such that rotation of the turbine is translated to the fan. A boundary layer ingestion inlet is aligned with an expected boundary layer, such that the boundary layer ingestion inlet is configured to ingest fluid from a boundary layer during operation of the boundary layer ingestion engine.

Abstract: A gas turbine engine includes a plurality of rotating components housed within a compressor section and a turbine section. A first tap is connected to the compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap and configured to deliver air to an aircraft fuselage. A cooling compressor is connected downstream of the heat exchanger. A high pressure feed is configured to deliver air at a second pressure which is higher than the first pressure. The cooling compressor is configured to deliver air to at least one of the plurality of rotating components. A valve assembly that can select whether air from the first tap or air from the high pressure feed is delivered to the aircraft pneumatic system.

Abstract: A boundary layer ingestion engine includes a gas generator, a turbine fluidly connected to the gas generator, and a fan mechanically linked to the turbine via at least one shaft. The linkage is configured such that rotation of the turbine is translated to the fan. The boundary layer ingestion engine further includes an exhaust duct fluidly connected to an outlet of the turbine. The exhaust duct is positioned radially inward of the fan.

Abstract: A gas turbine engine comprises a main compressor section having a downstream most end, 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 cooling air downstream of the heat exchanger, and delivers air into the high pressure turbine. The heat exchanger has at least two passes, with one of the passes passing air radially outwardly, and a second of the passes returning the air radially inwardly to the compressor. An intercooling system for a gas turbine engine is also disclosed.

Abstract: A gas turbine engine comprises a housing having an inlet leading to a fan rotor. A bypass door is mounted upstream of the inlet to the fan rotor, and is moveable away from a non-bypass position to a bypass position to selectively bypass boundary layer air vertically beneath the engine. An aircraft is also disclosed.

Abstract: A turbine engine such as a pusher fan engine is provided. This turbine engine includes a nacelle with a bypass flowpath. A fan rotor is configured to propel air out of the bypass flowpath. A plurality of guide vanes are configured to direct the air to the fan rotor.

Abstract: A gas turbine engine includes a main engine compressor section. A booster compressor changing a pressure of airflow received from the main engine compressor section to a pressure desired for a pneumatic system. The booster compressor operates at airflow conditions greater than a demand by the pneumatic system. A speed change system driving the booster compressor at speeds corresponding to a 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 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: 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.

Abstract: A propulsion system for an aircraft including at least two main gas turbine engines and a plurality of dedicated boundary layer ingestion fans. The propulsion system is arranged such that a combined thrust produced by the boundary layer ingestion fans is less than 20 percent of a total thrust of the main engines and the boundary layer ingestion fans. The boundary layer ingestion fans are controllable and selectively turned off at lower speeds.

Abstract: A gas turbine engine includes a main compressor section. A booster compressor includes an inlet and an outlet. The inlet receives airflow from the main compressor section and the outlet provides airflow to a pneumatic system. A recirculation passage is between the inlet and the outlet. A flow splitter valve controls airflow between the outlet and the inlet through the recirculation passage for controlling airflow to the pneumatic system based on airflow output from the booster compressor. A bleed air system for a gas turbine engine and a method of controlling engine bleed airflow are also disclosed.

Abstract: A gas turbine engine comprises an engine having a compressor section, and a turbine section. A firewall and accessory pumps are mounted on a downstream side of the firewall. The accessory pumps are driven by electric motors mounted on the firewall on an upstream side of the firewall.

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 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 electromechanical system for a gas turbine engine includes a mechanical component located at a first side of a firewall of a gas turbine engine, and an electrical motor located at a second side of the firewall and configured to drive the mechanical component. The electrical motor mechanically connected to the mechanical component through a firewall opening in the firewall, the first side having a higher operating temperature than the second side. An electrical connection extends between the mechanical component and the electrical motor via the same firewall opening.