Abstract: The present disclosure relates generally to a gas turbine engine that includes a fan configured to generate a fanstream and a fanstream duct configured to receive the fanstream flowing therethrough. An engine electronic component is positioned in flow communication with the fanstream. A heating element is positioned in the fanstream upstream from the engine electronic component and is operative to heat at least a portion of the fanstream in flow communication with the engine electronic component. The position of the engine electronic component passively thermally conditions the engine electronic component and the heating element actively thermally conditions the engine electronic component.

Abstract: A compressor for use in a gas turbine engine comprises a compressor rotor including blades and a disc, with a bore defined radially inwardly of the disc. A radially outer housing surrounds an outer diameter of the blades. A lower pressure tap and a higher pressure tap tap air from two distinct locations within the compressor and radially outwardly through the outer housing. A valve selectively delivers at least one of the lower pressure tap and the higher pressure tap to the bore of the disc. A control for the valve is programmed to move the valve to a position delivering the higher pressure tap at a point prior to take-off when the compressor is mounted in a gas turbine engine on an aircraft. A gas turbine engine and a method of operating a gas turbine engine are also disclosed.

Abstract: A turbine section of a gas turbine engine according to an example of the present disclosure includes, among other things, a fan drive turbine section, and a second turbine section. The fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed. The second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is faster than the first speed.

Abstract: A rotor assembly of a gas turbine engine may be spoked and includes a rotor and a shell. The rotor has a rotor disk and a plurality of blades each having a platform attached to the rotor disk and with a first channel defined radially between the platforms and the rotor disk. The shell projects aft of the rotor and includes inner and outer walls with a passage defined therebetween. The passage is in fluid communication with the first channel and, together, form part of a secondary flowpath for cooling of adjacent components. The rotor assembly may further include a structure located radially inward of the rotor disk and shell. The structure defines a supply conduit for flowing air from the passage and into a rotor bore defined at least in part by adjacent rotor disks. The entering air, being pre-heated when flowing through the channel and passage, warms the bore and reduces thermal gradients, thus thermal fatigue, across the rotor disk.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan having one or more fan blades. A compressor section is in fluid communication with the fan. The compressor section includes a first compressor section and a second compressor section. A turbine section is in fluid communication with the compressor section. The turbine section includes a first turbine section and a second turbine section. The first turbine section has a first exit area at a first exit point and rotates at a first speed. The second turbine section has a second exit area at a second exit point and rotates at a second speed. A first performance quantity is defined as the product of the first speed squared and the first area. A second performance quantity is defined as the product of the second speed squared and the second area.

Abstract: A turbine section of a gas turbine engine according to an example of the present disclosure includes, among other things, a fan drive turbine section and a second turbine section. The fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed. The second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is faster than the first speed.

Abstract: A turbine section of a gas turbine engine according to an example of the present disclosure includes, among other things, a fan drive turbine section and a second turbine section. The fan drive turbine section has a first exit area at a first exit point and is configured to rotate at a first speed. The second turbine section has a second exit area at a second exit point and is configured to rotate at a second speed, which is faster than the first speed.

Abstract: A gas turbine engine is provided that includes a compressor section, a combustor section, a diffuser case module, a turbine section, and a manifold. The diffuser case module includes a multiple of struts within an annular flow path from said compressor section to said combustor section, wherein at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path. The manifold is in communication with said mid-span pre-diffuser inlet and said turbine section.

Abstract: A gas turbine engine is provided that includes a compressor section, a combustor section with a multiple of pre-swirlers, a diffuser case module, and a manifold. The diffuser case module includes a multiple of struts within an annular flow path from said compressor section to said combustor section, wherein at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path. The manifold is in communication with said mid-span pre-diffuser inlet and said multiple of pre-swirlers.

Abstract: A gas turbine engine that includes a compressor section, a combustor section, a diffuser case module, and a manifold. The diffuser case module includes a multiple of struts within an annular flow path from said compressor section to said combustor section, wherein at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path. The manifold is in communication with said mid-span pre-diffuser inlet and a bearing compartment.

Abstract: A gas turbine engine is provided that includes a compressor section, a combustor section, a diffuser case module, and a manifold. The diffuser case module includes a multiple of struts within an annular flow path from said compressor section to said combustor section, wherein at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path. The manifold in communication with said mid-span pre-diffuser inlet and said combustor section.

Abstract: A turbine section for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan drive turbine including a fan drive duct, the fan drive turbine being configured to drive a fan section through a geared architecture at a speed that is less than an input speed to the geared architecture. A fan drive turbine includes a fan drive duct, the fan drive turbine being configured to drive a fan section through a geared architecture at a speed that is less than an input speed to the geared architecture. At least one upstream turbine is configured to drive at least one compressor. The at least one upstream turbine includes a turbine duct defining a conical flow path having a conical inlet defined by a first diameter and a conical outlet defined by a second diameter greater than the first diameter. The conical outlet is in fluid communication with the fan drive duct downstream of the conical outlet.

Abstract: A diffuser for a gas turbine engine is provided that includes an outer shroud, an inner shroud, a multiple of struts between said outer shroud and said inner shroud to define an annular flow path, and a mid-span pre-diffuser inlet in communication with said annular flow path.

Abstract: A gas turbine engine is provided comprising a compressor section, a combustor section, a diffuser case module, and a manifold. The diffuser case module includes a multiple of struts within an annular flow path from said compressor section to said combustor section, wherein at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path. The manifold is in communication with said mid-span pre-diffuser inlet and said compressor section.

Abstract: An example turbomachine assembly includes, among other things, a nose cone of a turbomachine, and a pump at least partially within an interior of the nose cone. The pump is selectively rotated by a motor to communicate air to the interior.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a compressor section, a combustor in fluid communication with the compressor section and a turbine section in fluid communication with the combustor. A buffer system includes a first circuit that supplies a first buffer supply air and a second circuit that supplies a second buffer supply air. The first circuit includes a first bleed air supply and a second bleed air supply. The second circuit includes a third bleed air supply and a fourth bleed air supply and at least one of the first circuit and the second circuit includes a variable area ejector.

Abstract: A cooling system for providing a buffer cooled cooling air to a turbine section of a gas turbine engine is disclosed. The cooling system may comprise a first conduit configured to transmit a cooling air toward the turbine section, a heat exchanger configured to cool a bleed airflow diverted from the first conduit to provide a buffer air, and a bypass conduit configured to direct at least a portion of the buffer air through at least one passageway that bypasses a bearing compartment of the gas turbine engine. The cooling system may further comprise a manifold configured to allow the cooling air exiting the first conduit and the buffer air exiting the bypass conduit to mix and provide the buffer cooled cooling air, and a nozzle assembly configured to deliver the buffer cooled cooling air to the turbine section.

Abstract: An example turbomachine assembly includes, among other things, a nose cone of a turbomachine, and a pump at least partially within an interior of the nose cone. The pump is selectively rotated by a motor to communicate air to the interior.

Abstract: An example turbomachine assembly includes, among other things, a nose cone of a turbomachine, and a pump that is selectively driven by a motor or a shaft to communicate air to an interior of the nose cone. The pump shaft is driven by a dedicated geared architecture.

Abstract: A gas turbine engine includes a compressor section, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor, and a buffer system. The buffer system can include a first circuit that supplies a first buffer supply air and a second circuit that supplies a second buffer supply air. The first circuit can include a first bleed air supply a second bleed air supply and at least one of an ejector and a valve. The second circuit can include a third bleed air supply, a fourth bleed air supply and at least one of an ejector and a valve.