Abstract: The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction extended from an axial centerline, and a circumferential direction. The gas turbine engine includes a compressor section, a combustion section, and a turbine section in serial flow arrangement along the longitudinal direction. The gas turbine engine includes a low speed turbine rotor including a hub extended along the longitudinal direction and radially within the combustion section; a high speed turbine rotor including a high pressure (HP) shaft coupling the high speed turbine rotor to a HP compressor in the compressor section; and a first turbine bearing disposed radially between the hub of the low speed turbine rotor and the HP shaft. The HP shaft extends along the longitudinal direction and radially within the hub of the low speed turbine rotor. The high speed turbine rotor defines a turbine cooling conduit extended within the high speed turbine rotor.

Abstract: A turbomachinery apparatus (10) includes: a turbine (22), including: a turbine component (36) defining at least one arcuate flowpath surface (40); an array of axial-flow turbine airfoils (46) extending from the flowpath surface, the turbine airfoils defining spaces therebetween; and a plurality of splitter airfoils (146) extending from the at least one flowpath surface, in the spaces between the turbine airfoils, each splitter airfoil having opposed pressure and suction sides extending between a leading edge and a trailing edge, wherein the splitter airfoils have a thickness ratio which is less than a thickness ratio of the turbine airfoils.

Abstract: A gas turbine engine includes a fan section and a core turbine engine operable with the fan section. The gas turbine engine also includes a sump pressurization assembly including a valve, a first duct in airflow communication with the valve and a first pressurized air source, and a second duct in airflow communication with the valve and a second pressurized air source. Further, the sump pressurization assembly includes a supply duct selectively in airflow communication with the first duct and the second duct through the valve. The supply duct is located at least partially inward of a core air flowpath of the core turbine engine for pressurizing a sump of the gas turbine engine.

Abstract: The present disclosure is directed to a gas turbine engine defining an axial centerline, a longitudinal direction, a radial direction, and a circumferential direction. The gas turbine engine includes one or more frames in which the frame defines an inner ring and an outer ring generally concentric to the inner ring about the axial centerline. The frame defines a plurality of struts extended outward along the radial direction from the inner ring to the outer ring. One or more struts define one or more service passages extended at least partially along the radial direction within the strut, and wherein the inner ring, the outer ring, and the struts together define an integral structure.

Abstract: The present disclosure is directed to a gas turbine engine defining a radial direction, a circumferential direction, an axial centerline along a longitudinal direction, and wherein the gas turbine engine defines an upstream end and a downstream end long the longitudinal direction. The gas turbine engine includes a turbine section including a low speed turbine rotor, a high speed turbine rotor, and an intermediate speed turbine rotor. The low speed turbine rotor includes an inner shroud and an outer shroud outward of the inner shroud in the radial direction. The outer shroud defines a plurality of outer shroud airfoils extended inward of the outer shroud along the radial direction. The low speed turbine rotor further includes at least one connecting airfoil coupling the inner shroud to the outer shroud. The high speed turbine rotor is disposed upstream of the one or more connecting airfoils of the low speed turbine rotor along the longitudinal direction.

Abstract: The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction, and a circumferential direction, and wherein the gas turbine engine defines an upstream end and a downstream end along the longitudinal direction. The gas turbine engine includes a turbine section that includes a first rotating component and a second rotating component. The first rotating component includes an inner shroud and an outer shroud outward of the inner shroud in the radial direction. The outer shroud defines a plurality of outer shroud airfoils extended inward of the outer shroud along the radial direction. The first rotating component further includes at least one connecting airfoil coupling the inner shroud and the outer shroud. The second rotating component is upstream of the one or more connecting airfoils of the first rotating component along the longitudinal direction. The second rotating component includes a plurality of second airfoils extended outward in the radial direction.

Abstract: The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction, and a circumferential direction, and an upstream end and a downstream end along the longitudinal direction. The gas turbine engine includes a turbine section, a gearbox proximate to the turbine section, and a driveshaft. The turbine section includes a first rotating component interdigitated with a second rotating component along the longitudinal direction. The first rotating component includes an outer shroud defining a plurality of outer shroud airfoils extended inward of the outer shroud along the radial direction and one or more connecting airfoils coupling the outer shroud to a radially extended rotor. The second rotating component includes an inner shroud defining a plurality of inner shroud airfoils extended outward of the inner shroud along the radial direction. The second rotating component is coupled to an input shaft connected to an input gear of the gearbox.

Abstract: The present disclosure is directed to a gas turbine engine defining a radial direction, a circumferential direction, an axial centerline along a longitudinal direction. The gas turbine engine defines an upstream end and a downstream end along the longitudinal direction and includes a turbine frame defined around the axial centerline. The turbine frame includes a first bearing surface, a second bearing surface, and a third bearing surface. The first bearing surface corresponds to a first turbine rotor, the second bearing surface corresponds to a second turbine rotor, and the third bearing surface corresponds to a third turbine rotor, and each turbine rotor is independently rotatable.

Abstract: The present disclosure is directed to a gas turbine engine defining a longitudinal direction, a radial direction, and a circumferential direction, and an upstream end and a downstream end along the longitudinal direction. The gas turbine engine includes a turbine section, a gearbox proximate to the turbine section, and a driveshaft. The turbine section includes a first rotating component interdigitated with a second rotating component along the longitudinal direction. The first rotating component includes an outer shroud defining a plurality of outer shroud airfoils extended inward of the outer shroud along the radial direction and one or more connecting airfoils coupling the outer shroud to a radially extended rotor. The second rotating component includes an inner shroud defining a plurality of inner shroud airfoils extended outward of the inner shroud along the radial direction. The second rotating component is coupled to an input shaft connected to an input gear of the gearbox.

Abstract: A turbine engine including a core engine cowl including a compartment, a cooling airflow source positioned within the compartment, and a full authority digital engine control (FADEC) system coupled in communication with the cooling airflow source. The FADEC system is configured to determine a flight status of the turbine engine, and actuate the cooling airflow source when the turbine engine is not in flight, and before the turbine engine has been shut down, such that heat is exhausted from the compartment.

Abstract: The present disclosure is directed to a gas turbine engine defining a radial direction, a circumferential direction, an axial centerline along a longitudinal direction, and wherein the gas turbine engine defines an upstream end and a downstream end along the longitudinal direction, and wherein the gas turbine engine defines a core flowpath extended generally along the longitudinal direction. The gas turbine engine includes a turbine frame defined around the axial centerline, the turbine frame comprising a first bearing surface disposed inward along the radial direction. The gas turbine engine further includes a turbine rotor assembly including a bearing assembly coupled to the first bearing surface of the turbine frame and the turbine rotor assembly. The turbine rotor assembly further includes a first turbine rotor disposed upstream of the turbine frame and a second turbine rotor disposed downstream of the turbine frame.

Abstract: The present disclosure is directed to a gas turbine engine including a torque frame. The torque frame includes an inner shroud defined circumferentially around the axial centerline, an outer shroud surrounding the inner shroud and defined circumferentially around the axial centerline, and a structural member extended along the radial direction and coupled to the inner shroud and the outer shroud. The torque frame is configured to rotate around the axial centerline.

Abstract: The present disclosure is directed to a gas turbine engine including a turbine section including a first rotating component interdigitated along a longitudinal direction with a second rotating component. The first rotating component and the second rotating component are each coupled to a speed reduction assembly in counter-rotating arrangement. The first rotating component comprising an outer shroud and a plurality of outer shroud airfoils extended inward along a radial direction from the outer shroud. A connecting member couples the outer shroud to a radially extended first rotor. The second rotating component comprising an inner shroud and a plurality of inner shroud airfoils extended outward along the radial direction from the inner shroud, the plurality of inner shroud airfoils in alternating arrangement along the longitudinal direction with the plurality of outer shroud airfoils.

Abstract: A gas turbine engine includes a first compressor, a casing surrounding the first compressor, and a liner extending forward from the first compressor. The gas turbine engine also includes a core turbine frame assembly extending between the liner and the casing and a bleed air assembly. The bleed air assembly includes a bleed valve positioned in the liner, and a duct in airflow communication with the bleed valve and defining in outlet. The duct is positioned within the core turbine frame assembly and extends to the casing.

Abstract: A gas turbine engine includes a casing surrounding a first compressor, as well as a liner extending forward from the first compressor. Additionally, the gas turbine engine includes a bleed air assembly having a plurality of bleed valves positioned in the liner and spaced along a circumferential direction of the gas turbine engine. The bleed air assembly also includes a duct in airflow communication with the plurality of bleed valves and defining an outlet at the casing. The duct provides a flow bleed air from the plurality bleed valves to the outlet during operation.

Abstract: The present disclosure is directed to a turbine engine defining a longitudinal direction, a radial direction, and a circumferential direction. The turbine engine includes a power turbine including a first turbine rotor assembly interdigitated with a second turbine rotor assembly along the longitudinal direction; a gear assembly coupled to the first turbine rotor assembly and the second turbine rotor assembly, wherein the gear assembly includes a first input interface coupled to the first turbine rotor assembly, a second input interface coupled to the second turbine rotor assembly, and one or more third gears coupled to the first input interface and the second input interface therebetween; and a first output shaft and a second output shaft, wherein each of the first output shaft and the second output shaft are configured to couple to an electrical load device.

Abstract: A gas turbine engine includes a turbomachine and a fan rotatable by the turbomachine. The fan includes a plurality of fan blades. The gas turbine engine also includes an outer nacelle surrounding the plurality of fan blades and defining an nacelle inlet, the outer nacelle including an inner wall defining a plurality of openings located aft of the nacelle inlet and forward of the plurality of fan blades of the fan along an axial direction for providing a swirl airflow upstream of the plurality of fan blades of the fan at a swirl angle greater than zero relative to a radial direction.

Abstract: A gas turbine engine includes a turbomachine and a fan rotatable by the turbomachine. The fan includes a plurality of fan blades, each of the plurality of fan blades defining a fan blade span along a radial direction. The gas turbine engine further includes an outer nacelle surrounding the plurality of fan blades and a plurality of part-span inlet guide varies attached to the outer nacelle at a location forward of the plurality of fan blades along an axial direction. Each of the plurality of inlet guide vanes defines an IGV span along the radial direction, the IGV span being at least about five percent of the fan blade span and up to about fifty-five percent of the fan blade span.

Abstract: A gas turbine engine includes a turbomachine and a fan rotatable by the turbomachine. The fan includes a plurality of fan blades. The gas turbine engine also includes an outer nacelle surrounding the plurality of fan blades and including an inner wall, the inner wall of the outer nacelle including a plurality of pre-swirl contours positioned forward of the fan blades of the fan along an axial direction and extending inwardly along a radial direction.

Abstract: A gas turbine engine includes a turbomachine and a fan rotatable by the turbomachine. The fan includes a plurality of fan blades. The gas turbine engine also includes an outer nacelle surrounding the plurality of fan blades and a plurality of part-span inlet guide vanes cantilevered from the outer nacelle at a location forward of the plurality of fan blades along the axial direction. Each of the plurality of inlet guide vanes defines an inner end along the radial direction and it is unconnected with an adjacent part-span inlet guide vane at the inner end.