Abstract:
A compressor section of a gas turbine engine includes a bleed port having a flow splitter therein so as to define a downstream bleed channel having a downstream inlet and an upstream bleed channel having an upstream inlet that is positioned radially outward from the downstream inlet.
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 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 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 pivot thrust reverser includes a first tandem pivot door subassembly comprising an inner panel and an outer panel. The inner panel and outer panel are connected by a first sliding rail. A second tandem pivot door subassembly is included comprising an inner panel and an outer panel. The inner panel and outer panel are connected by a second sliding rail.
Abstract:
A gas turbine engine includes a first annular portion that is stationary and adapted for partially surrounding an engine core. The first annular portion includes a fore pylon connecting portion. A rail is coupled to the fore pylon portion and extends in the aft direction from the stationary portion. A second annular portion is positioned aft of the first portion and coupled to the rail, the second portion being movable along an engine core centerline between a closed position and an open position, wherein the second annular portion is configured to engage with the first annular portion in the closed position to provide access to an engine core. The second annular portion includes a nozzle plug.
Abstract:
A gas turbine engine comprises a main fan that delivers air into a bypass duct and into a core engine. A heat exchanger is positioned within the bypass duct and receives a fluid to be cooled from a component associated with the gas turbine engine. A heat exchanger fan is positioned to draw air across the heat exchanger and a control for the heat exchanger fan. The control is programmed to stop operation of the fan during certain conditions, and to drive the heat exchanger fan under other conditions. A method of forming a heat exchanger is also disclosed.
Abstract:
A compressor section of a gas turbine engine includes a bleed port having a flow splitter therein so as to define a downstream bleed channel having a downstream inlet and an upstream bleed channel having an upstream inlet that is positioned radially outward from the downstream inlet.
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 mounting system for a gas turbine engine includes a compressor case portion, an inlet frame, an outlet frame, and a mounting structure. The compressor case portion houses rotatable compressor blades. The inlet frame connects to an inlet end of the compressor case. The outlet frame connects to an outlet end of the compressor case portion at an end opposite the compressor case inlet end. An axially fore mounting structure of the mounting structure connects to the inlet frame. An axially aft mounting structure of the mounting structure connects to the outlet frame. A bridging structure of the mounting structure is offset from the compressor case and connects the fore and aft mounting structures, thereby bridging engine loads across the inlet and outlet frames to reduce load induced distortion of the compressor case portion.