Abstract: A turbine engine that includes a stationary assembly, and a rotor assembly configured to rotate relative to the stationary assembly. The rotor assembly includes a plurality of unitary turbine and fan blades. Each unitary turbine and fan blade includes a single turbine airfoil, a single fan airfoil positioned radially outward from the single turbine airfoil, and a midspan shroud segment defined between the single turbine airfoil and the single fan airfoil.

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: A turbine engine that includes a stationary assembly, and a rotor assembly configured to rotate relative to the stationary assembly. The rotor assembly includes a plurality of unitary turbine and fan blades. Each unitary turbine and fan blade includes a single turbine airfoil, a single fan airfoil positioned radially outward from the single turbine airfoil, and a midspan shroud segment defined between the single turbine airfoil and the single fan airfoil.

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 turbomachine includes a compressor, an inner annular casing, and an outer annular casing. The inner annular casing and the outer annular casing define at least one cavity therebetween. The clearance control system includes a manifold system including at least one conduit disposed within the cavities and configured to channel a flow of cooling fluid between the cavities. The clearance control system also includes an impingement system including a header and at least one plenum configured to channel the flow of cooling fluid to the inner annular casing. The conduits configured to channel the flow of cooling fluid to the impingement system. The clearance control system further includes a channel system including at least one channels configured to channel the flow of cooling fluid to the turbomachine. The channels are configured to control the flow of cooling fluid to the manifold system.

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 gas turbine engine includes a core engine, a low pressure turbine, a fan assembly, a gearbox, and a lubrication scavenge pump. The core engine includes a high pressure compressor, a combustor, and a high pressure turbine configured in a serial flow arrangement. The low pressure turbine is positioned axially aft of the core engine. The fan assembly is positioned axially forward of the core engine. The gearbox is positioned axially forward of the fan assembly. The lubrication scavenge pump is positioned forward of the gearbox.

Abstract: A counter-rotating shaft assembly of a gas turbine engine includes an outer shaft rotatable in a first direction about a virtual rotational axis, an inner shaft counter-rotatable about the virtual rotational axis in a second direction that is opposite to the first direction, a differential bearing rotatably connecting the two shafts, a centering spring connecting the inner shaft to the differential bearing, and a squeeze film damper between the differential bearing and the inner shaft.

Abstract: A turbine assembly includes a rotor assembly including a shaft coupled to a plurality of rotor stages including a plurality of turbine blades. The shaft and the plurality of turbine blades define a wheelspace therein. The turbine assembly further includes a plurality of seals in series, at least one seal of the plurality of seals is coupled between a static support member and a respective rotor stage such that a plurality of turbine cavities in series are defined within the wheelspace. Each turbine cavity of the plurality of turbine cavities defined by the plurality of seals receives a pressurized fluid flow that applies an axially aft force to the respective rotor stage of the plurality of rotor stages that at least partially reduces net rotor thrust generated by the rotor assembly during operation, the pressurized fluid flow further provides turbine purge within the wheel space.

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: 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 fluid transfer assembly for use in a gas turbine engine includes a rotor shaft, a stationary assembly circumscribing the rotor shaft, and a rotating component coupled to the rotor shaft and positioned radially inward of the stationary assembly. The rotating assembly includes a hub coupled to the rotor shaft, a plurality of rotor blades coupled to the hub, and a shroud coupled to the plurality of rotor blades.

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: An actuator includes a pump including a first cavity and a diaphragm coupled in flow communication with the first cavity. The diaphragm is configured to pressurize a fluid contained in the first cavity. The pump further includes a first valve coupled in flow communication with the first cavity. The first valve is configured to release fluid from the first cavity when the first cavity is pressurized. The actuator also includes a piston assembly operatively coupled to the pump.

Abstract: A variable core cowl vent nozzle system is described herein, the system including a core casing at least partially surrounding a core engine of a gas turbine engine, and a core cowl extending aftward from the core casing. The core cowl defines a core cowl vent area between the core cowl and a primary nozzle. At least one of the core cowl and the primary nozzle is movable to vary the core cowl vent area. A method for varying the core cowl vent area by moving one or more of the core cowl and the primary nozzle is also described herein.

Abstract: A system for mounting an engine to an aircraft includes an engine forward mount angled toward the forward end of the engine at a first angle. At least two thrust links extend between an engine aft mount to a link support connection at a second angle. The engine aft mount is angled toward the forward end of the engine at a third angle. A projection of a load vector of the engine forward mount onto a vertical plane extending through the axis of rotation of the engine and a projection of a load vector of each of the at least two thrust links onto the vertical plane intersect the axis of rotation of the engine within a first vertical plane segment extending between a forward end of a nose of a fan assembly and forward of a forward mount interface.

Abstract: The present disclosure is directed to a gas turbine engine defining a longitudinal direction, an axial centerline extended along the longitudinal direction, an upstream end and a downstream end opposite of the upstream end along the longitudinal direction, a radial direction, and a circumferential direction. The gas turbine engine includes a high speed turbine rotor coupled to a high pressure (HP) shaft and HP compressor, a low speed turbine rotor comprising an axially extended hub, and a first turbine bearing disposed radially between the low speed turbine rotor and the high speed turbine rotor. The high speed turbine rotor defines a turbine cooling conduit through the high speed turbine rotor. The low speed turbine rotor includes a rotating nozzle adjacent to the turbine cooling conduit. The first turbine bearing defines an outer air bearing and an inner air bearing. The first turbine bearing defines a stationary nozzle adjacent to the rotating nozzle of the first turbine rotor.