Abstract: A flight vehicle includes a fuselage and a gas turbine engine. The gas turbine engine is coupled to the fuselage to provide thrust when air surrounding the flight vehicle is admitted to the gas turbine engine and combusted with fuel. The flight vehicle further includes a generator coupled to the gas turbine engine to provide power to equipment included in the vehicle.

Abstract: A turbomachine includes a vane, a rotation support coupled to an end of the vane, and a spindle coupled to the rotation support. The spindle, the vane, and the rotation support are rotationally aligned. An annular sleeve defines the spindle. The annular sleeve contacts the rotation support at a radially inward extent and contacts a turbine casing at a radially outward extent. A first rolling element engages the annular sleeve substantially near the radially outward extent. The first rolling element is coupled to the turbine casing. A second rolling element engages the annular sleeve substantially near the radially inward extent. The second rolling element is coupled to an outer endwall ring. A center of mass of the annular sleeve is positioned between the first and second rolling elements.

Abstract: The present disclosure relates to a hot section gas turbine engine component assembly and a method for forming such.

Abstract: A turbomachine includes a variably positioned vane, an outer spindle integral with a vane outer button, where the vane is coupled to the vane outer button, an annular sleeve defining the spindle, where the annular sleeve contacts the vane outer button at a radially inward extent and contacts a turbine casing at a radially outward extent. The annular sleeve includes a wall portion having an aperture and the spindle extends through the aperture. A nut engages spindle threads, where the nut applies force to the wall portion toward the radially inward extent. An end of the spindle extends through the turbine casing, and a cantilever rotation actuator is coupled to the end of the spindle. A first bearing and second bearing engage the annular sleeve. A vane inner button is coupled to the vane, and a third bearing engages the vane inner button.

Abstract: A combustion system for gas turbine engine is disclosed. The combustion system includes a wave rotor drum and a drive coupled to the wave rotor drum. The wave rotor drum is mounted for rotation about an axis and is formed to include a plurality of combustion channels that extend along the axis to conduct transient waves along the axis during operation of the combustion system. The drive is coupled to the wave rotor drum and is adapted to drive rotation of the wave rotor drum about the axis.

Abstract: In one embodiment, a gas turbine engine may include an accessory system integrated in a frame between a bypass flow path and a core flow path. A shaft may be provided between the accessory system and a core shaft of the gas turbine engine. The integrated accessory system may include power devices such as motors and generators, among other types of devices. The gas turbine engine may have more than one device spaced at various circumferential locations. In one non-limiting example, two power devices may be spaced 180 degrees apart. In some embodiments, a shaft and gearing may be provided to a location such as a frame in preparation to receive a power device at a later time.

Abstract: A gas turbine engine having a ramburner is disclosed. The ramburner is disposed downstream of a gas turbine engine combustor and receives an engine exhaust flow from the gas turbine engine combustor. The ramburner also accepts a bypass air. Fuel is injected into the ramburner and a combustion reaction is auto-initiated based upon local gas temperatures. No mechanical flame holders need be used. A slidable valve may be used to vary the amount of bypass air into the ramburner. A movable cowl and a plug nozzle form an exit flow path of the gas turbine engine. The movable cowl can be positioned to vary a throat area and exit area of the gas turbine engine based upon the operation of the ramburner, which may be influenced by the amount of bypass air entering the ramburner.

Abstract: An exemplary gas turbine engine may include a compressor, a turbine, and a combustor disposed between the compressor and the turbine. The combustor generally may have an inner casing. The exemplary gas turbine may further include a pre-swirl nozzle configured to receive a cooling stream. The cooling stream may be supplied from a cooled cooling air stream. The pre-swirl nozzle may further be configured to direct at least a portion of the cooling stream to the turbine. The pre-swirl nozzle may be flexibly mounted to the inner casing of the combustor.

Abstract: A turbofan gas turbine engine is provided having a unique power off-take shaft and gear system. Other gas turbine engine types are also contemplated herein. Two power off-takes are provided, one each for the low pressure spool and high pressure spool. The power off-takes extend across a core flow path of the turbofan engine between the low and high pressure shafts to a fan frame of the turbofan. A drive gear is provided near the front end of the high pressure shaft, and another drive gear is provided on the low pressure shaft near the drive gear for the high pressure shaft. Both gears are located in a sump of the gas turbine engine. The power off-take shafts are coupled to the drive gears. Two power devices are coupled to the power off-take shafts and are located in the fan frame. The power devices can be electric generators or motors.

Abstract: A low leakage, integral fixed vane system for a gas turbine engine compression system which can accommodate the latest 3-D aerodynamic airfoil geometries and provide endwall ovalization control. A unitized construction wherein the vanes are integral with the case wall in the compression system.

Abstract: An adaptive cycle gas turbine engine is disclosed having a number of features. A fan arrangement is provided having counter-rotating fan stages, one fan stage is operable to be clutched and decoupled from the other stage. A high pressure compressor bypass is also provided. A clutch is provided to at least partially drive an intermediate pressure compressor with a high pressure turbine when the high pressure compressor is bypassed. A partial bypass of the high pressure turbine may be provided.

Abstract: One embodiment of the present invention is an integrated aero-engine flowpath structure. Another embodiment is a method of manufacturing integrated aero-engine flowpath structure. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for aero-engine flowpath structures. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Abstract: A flight vehicle includes a fuselage and a gas turbine engine. The gas turbine engine is coupled to the fuselage to provide thrust when air surrounding the flight vehicle is admitted to the gas turbine engine and combusted with fuel. The flight vehicle further includes a generator coupled to the gas turbine engine to provide power to equipment included in the vehicle.

Abstract: A gas turbine engine includes a compressor assembly that is rotationally coupled to a shaft. The engine includes a radial diffuser assembly coupled to a shroud of the compressor assembly and positioned to receive compressed air from a centrifugal impeller. The radial diffuser assembly includes a first arcuate wall and a second wall, and a near axial diffuser coupled to the radial diffuser assembly and positioned to receive the compressed air from the radial diffuser assembly. The gas turbine includes a gas seal coupled between the second wall and a wall of the near axial diffuser, the gas seal configured to prevent the compressed air from passing through the seal while allowing relative motion between the radial diffuser assembly and the near axial diffuser.

Abstract: A gas turbine engine includes a compressor assembly that is rotationally coupled to a shaft, the compressor assembly having a centrifugal impeller and a shroud covering a bladed portion of the centrifugal impeller. The compressor assembly includes a diffuser that is attached to the shroud via a pair of flanges, the diffuser including a strut that is mounted through an aft-extending leg to a base of an intercase. A sealing assembly is attached to the diffuser and is attachable to a transition duct that is positioned to receive air from the diffuser. The sealing assembly is configured to prevent air from passing through the sealing assembly while allowing relative motion to occur between the transition duct and the diffuser.

Abstract: One embodiment of the present invention is an integrated aero-engine flowpath structure. Another embodiment is a method of manufacturing integrated aero-engine flowpath structure. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for aero-engine flowpath structures. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Abstract: One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique gas turbine engine main engine rotor. Still another embodiment is a unique method for assembling a gas turbine engine main engine rotor. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and gas turbine rotors. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Abstract: One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique variable cycle gas turbine engine. Another embodiment is a unique adaptive fan system for a variable cycle turbofan engine having at least one turbine. Another embodiment is a unique method for operating a variable cycle gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and related systems.

Abstract: A turbomachine includes a variably positioned vane, an outer spindle integral with a vane outer button, where the vane is coupled to the vane outer button, an annular sleeve defining the spindle, where the annular sleeve contacts the vane outer button at a radially inward extent and contacts a turbine casing at a radially outward extent. The annular sleeve includes a wall portion having an aperture and the spindle extends through the aperture. A nut engages spindle threads, where the nut applies force to the wall portion toward the radially inward extent. An end of the spindle extends through the turbine casing, and a cantilever rotation actuator is coupled to the end of the spindle. A first bearing and second bearing engage the annular sleeve. A vane inner button is coupled to the vane, and a third bearing engages the vane inner button.

Abstract: In one embodiment a centrifugal impeller has an axis of rotation, a plurality of blades, and a plurality of pitches defined between the blades. A cavity may be formed in a back side of the impeller, opposite with at least one of the plurality of pitches. In one form the cavity may not extend below a radius determined according to a stress criteria or stress limitation.