Abstract: A ducted fan engine is deiced using a ground support deicing apparatus having a support structure, a plurality of sonic wave transmitters, an imaging device, and a controller that controls the sonic wave transmitters to emit sonic waves at varying frequencies. A deicing program causes the controller to control (a) providing imaging signals to obtain image data from imaging sensors, (b) receiving image data provided by each of the imaging sensors, and generates images of at least one component part of the engine, (c) detecting a presence or an absence of ice on the at least one component part of the engine, and (d) controlling the plurality of sonic wave transmitters to emit sonic waves in a given frequency range so as to remove the ice from the component part of the engine.

Abstract: A gas turbine engine includes a compressor section comprising a compressor, a combustion section, and a turbine section arranged in serial flow order and defining a working gas flowpath, the compressor comprising an aft-most compressor stage; a spool drivingly coupled to the compressor; a stage of stator vanes located downstream of the aft-most compressor stage; and a stator case, the spool and the stator case together defining a rotor cavity in fluid communication with the working gas flowpath, the stage of stator vanes including a first stator vane defining a fluid passage, and the stator case defining a plenum and a supplemental airflow passage, the plenum in fluid communication with the fluid passage in the first stator vane, the supplemental airflow passage in fluid communication with the plenum and the rotor cavity for proving an airflow to the rotor cavity

Abstract: A gas turbine engine includes a compressor section having rotor shaft assembly and a stator shroud assembly including a stator shroud casing surrounding the compressor rotor shaft assembly, a compressor flow passage being defined between the compressor rotor shaft assembly and the stator shroud casing. A compressor bleed air system includes a compressor bleed air duct that includes a sloped inlet portion, such as a NACA submerged inlet portion, in the stator shroud casing. Alternatively, stator vanes at a compressor bleed air passage may have a NACA inlet scoop and an airflow passage through the stator vane providing a bleed airflow to a cavity, which has airflow passages providing bleed air to an upstream side of the bleed air passage.

Abstract: A gas turbine engine includes a compressor with an inner casing and an outer casing. The inner casing defines a primary flow path for airflow through the compressor, the inner casing and the outer casing define a bleed air cavity therebetween. The inner casing at least partially defines a first bleed air channel between the primary flow path and the first bleed air cavity. The first bleed air channel is defined between a first wall and a second wall, wherein the first wall is upstream from the second wall. The first wall includes a plurality of holes in fluid communication with the primary flow path and fluidly coupled to a suction manifold. The plurality of holes is configured to energize a fluid boundary layer along the first wall.

Abstract: A cooling system for a turbine engine. The turbine engine includes a compressor, a turbine, and a shaft that drivingly couples the compressor and the turbine. The cooling system includes one or more cavities of the compressor. The cooling system includes a shaft flowpath defined in the shaft. The shaft includes one or more shaft apertures that provide fluid communication between the shaft flowpath and the one or more cavities. Air passes through the one or more shaft apertures and the shaft flowpath during a shutdown of the turbine engine to cool the one or more cavities.

Abstract: A gas turbine engine that defines a radial direction and axial direction. The gas turbine engine includes a compressor section that includes a high pressure compressor. The high pressure compressor includes an aft-most compressor stage. The aft-most compressor stage includes a rotor, a low pressure shaft, a high pressure shaft drivingly coupled to the high pressure compressor. The high pressure shaft includes a central spool member and a forward spool member. The forward spool member extending between the central spool member and the aft-most compressor stage. The aft-most compressor stage further includes an air guide coupled to the aft-most compressor stage and the high pressure shaft. The air guide and a radially inward side of the forward spool member together defines an air flow passage for providing a cooling air flow.

Abstract: A gas turbine engine includes a compressor section having rotor shaft assembly and a stator shroud assembly including a stator shroud casing surrounding the compressor rotor shaft assembly, a compressor flow passage being defined between the compressor rotor shaft assembly and the stator shroud casing. A compressor bleed air system includes a compressor bleed air duct that includes a sloped inlet portion, such as a NACA submerged inlet portion, in the stator shroud casing. Alternatively, stator vanes at a compressor bleed air passage may have a NACA inlet scoop and an airflow passage through the stator vane providing a bleed airflow to a cavity, which has airflow passages providing bleed air to an upstream side of the bleed air passage.

Abstract: A gas turbine engine includes a compressor with an inner casing and an outer casing. The inner casing defines a primary flow path for airflow through the compressor, the inner casing and the outer casing define a bleed air cavity therebetween. The inner casing at least partially defines a first bleed air channel between the primary flow path and the first bleed air cavity. The first bleed air channel is defined between a first wall and a second wall, wherein the first wall is upstream from the second wall. The first wall includes a plurality of holes in fluid communication with the primary flow path and fluidly coupled to a suction manifold. The plurality of holes is configured to energize a fluid boundary layer along the first wall.

Abstract: A gas turbine engine that defines a radial direction and axial direction. The gas turbine engine includes a compressor section that includes a high pressure compressor. The high pressure compressor includes an aft-most compressor stage. The aft-most compressor stage includes a rotor, a low pressure shaft, a high pressure shaft drivingly coupled to the high pressure compressor. The high pressure shaft includes a central spool member and a forward spool member. The forward spool member extending between the central spool member and the aft-most compressor stage. The aft-most compressor stage further includes an air guide coupled to the aft-most compressor stage and the high pressure shaft. The air guide and a radially inward side of the forward spool member together defines an air flow passage for providing a cooling air flow.

Abstract: A turbofan engine is provided. The turbofan engine includes a fan; a turbomachine operably coupled to the fan for driving the fan, wherein the turbomachine, the fan, or both include an engine component; a heat source; and a heat transfer system configured to reduce ice buildup or ice formation in the engine component, the heat transfer system in communication with the heat source, the heat transfer system comprising: a first heat transfer component in communication with the heat source; and a second heat transfer component that extends from the first heat transfer component to or through the engine component, wherein the first heat transfer component comprises one of a heat pipe or a graphene rod, and wherein the second heat transfer component comprises the other of the heat pipe or the graphene rod.

Abstract: A ducted fan engine is deiced using a ground support deicing apparatus having a support structure, a plurality of sonic wave transmitters, an imaging device, and a controller that controls the sonic wave transmitters to emit sonic waves at varying frequencies. A deicing program causes the controller to control (a) providing imaging signals to obtain image data from imaging sensors, (b) receiving image data provided by each of the imaging sensors, and generates images of at least one component part of the engine, (c) detecting a presence or an absence of ice on the at least one component part of the engine, and (d) controlling the plurality of sonic wave transmitters to emit sonic waves in a given frequency range so as to remove the ice from the component part of the engine.

Abstract: A cooling system for a turbine engine. The turbine engine includes a compressor, a turbine, and a shaft that drivingly couples the compressor and the turbine. The cooling system includes one or more cavities of the compressor. The cooling system includes a shaft flowpath defined in the shaft. The shaft includes one or more shaft apertures that provide fluid communication between the shaft flowpath and the one or more cavities. Air passes through the one or more shaft apertures and the shaft flowpath during a shutdown of the turbine engine to cool the one or more cavities.

Abstract: In one aspect, a device for preventing bird strikes to an engine. The device includes a projector mounted on a component of the engine. The projector is positioned to project an image outside of the engine. In another aspect, an engine for an aircraft. The engine includes an engine component and a projector mounted on the engine component. The projector is positioned to project an image outside of the engine.

Abstract: In one aspect, the present disclosure is directed a gas turbine engine including a compressor, a combustor, and a turbine in a serial flow arrangement. The compressor includes a casing having an inner wall and an outer wall. The inner wall defines a passageway for airflow through the compressor. The casing defines a bleed cavity between the inner wall and the outer wall. The inner wall has an opening. The casing defining a channel between the opening and the bleed cavity to direct airflow into the bleed cavity. The channel is defined by a first wall and a second wall of the casing. The second wall being downstream of the first wall. A surface of the first wall has a pattern to reduce flow separation in the channel.

Abstract: A turbofan engine is provided. The turbofan engine includes a fan; a turbomachine operably coupled to the fan for driving the fan, wherein the turbomachine, the fan, or both include an engine component; a heat source; and a heat transfer system configured to reduce ice buildup or ice formation in the engine component, the heat transfer system in communication with the heat source, the heat transfer system comprising: a first heat transfer component in communication with the heat source; and a second heat transfer component that extends from the first heat transfer component to or through the engine component, wherein the first heat transfer component comprises one of a heat pipe or a graphene rod, and wherein the second heat transfer component comprises the other of the heat pipe or the graphene rod.

Abstract: In one aspect, the present disclosure is directed a gas turbine engine including a compressor, a combustor, and a turbine in a serial flow arrangement. The compressor includes a casing having an inner wall and an outer wall. The inner wall defines a passageway for airflow through the compressor. The casing defines a bleed cavity between the inner wall and the outer wall. The inner wall has an opening. The casing defining a channel between the opening and the bleed cavity to direct airflow into the bleed cavity. The channel is defined by a first wall and a second wall of the casing. The second wall being downstream of the first wall. A surface of the first wall has a pattern to reduce flow separation in the channel.

Abstract: An example gas turbine engine includes a compressor including a casing defining a passageway and a shaft extending through the passageway. The shaft drivingly coupling the compressor and a turbine of the gas turbine engine. The shaft has an opening to receive airflow from the passageway. The gas turbine engine also includes an inner shroud, stator vanes coupled to and extending radially between the casing and the inner shroud, and an offtake scoop disposed on a downstream side of the inner shroud. The offtake scoop has a channel to direct the airflow radially inward toward the opening in the shaft.

Abstract: An apparatus and method of cooling a hot portion of a gas turbine engine, such as a multi-stage compressor of a gas turbine engine, by reducing an operating air temperature in a space between a seal and a blade post of adjacent stages by routing cooling air through an inlet in the vane, passing the routed cooling air through the vane, and emitting the routed cooling air into the space.

Abstract: According to exemplary embodiments, a rotor off-take assembly is provided by positioning an angled hole or aperture in a stator assembly. This angled hole provides improved pressure recovery and utilizes higher dynamic pressure to drive the bleed air flow into the off-take cavity.

Abstract: An apparatus and method of cooling a hot portion of a gas turbine engine, such as a multi-stage compressor of a gas turbine engine, by reducing an operating air temperature in a space between a seal and a blade post of adjacent stages by routing cooling air through an inlet in the vane, passing the routed cooling air through the vane, and emitting the routed cooling air into the space.