Abstract: A method for assembling a gas turbine engine includes coupling a gearbox to a low-pressure turbine, the gearbox includes a plurality of planetary gears intermeshed with the sun gear, each of the planetary gears includes a first gear portion having a first diameter and a second gear portion having a second diameter that is different than the first diameter.

Abstract: A method for assembling a gas turbine engine including providing a core gas turbine engine at least partially defined by a frame and having a drive shaft rotatable about a longitudinal axis of the core gas turbine engine. A low-pressure turbine is coupled to a core turbine engine. A counter-rotating fan assembly is coupled to the low-pressure turbine. The counter-rotating fan assembly includes a first fan assembly and a second fan assembly. A booster compressor is coupled to the second fan assembly. A gearbox is securely coupled to the frame so that the gearbox is circumferentially positioned about the drive shaft. The first fan assembly is rotatably coupled to an input of the gearbox such that the first fan assembly rotates in a first direction. The second fan assembly is rotatably coupled to an output of the gearbox such that the second fan assembly rotates in a second direction opposite the first direction.

Abstract: A turbofan engine includes a fan, compressor, combustor, high pressure turbine, and low pressure turbine joined in serial flow communication. The high pressure turbine includes two stages of rotor blades to effect corresponding exit swirl in the combustion gases discharged therefrom. A transition duct includes fairings extending between platforms for channeling the combustion gases to the low pressure turbine with corresponding swirl. First stage rotor blades in the low pressure turbine are oriented oppositely to the rotor blades in the high pressure turbine for counterrotation.

Abstract: A turbine nozzle includes a row of vanes joined to radially outwardly inclined outer and inner bands. Each vane has camber and an acute twist angle for importing swirl in combustion gases discharged at trailing edges thereof. The trailing edges are tilted forwardly from the inner band to the outer band for increasing aerodynamic efficiency.

Abstract: A turbofan engine includes a fan, compressor, combustor, single-stage high pressure turbine, and low pressure turbine joined in serial flow communication. First stage rotor blades in the low pressure turbine are oriented oppositely to the rotor blades in the high pressure turbine for counterrotation. First stage stator vanes in the low pressure turbine have camber and twist for carrying swirl directly between the rotor blades of the high and low pressure turbines.

Abstract: A method for assembling a gas turbine engine includes providing a first fan assembly including a plurality of rotor blades that are configured to rotate in a first rotational direction at a first rotational speed, rotatably coupling a second fan assembly axially aft of the first fan assembly, wherein the second fan assembly includes a plurality of rotor blades that are configured to rotate in a second rotational direction, and coupling a gearbox to the second fan assembly that is configured to rotate the second fan assembly at a second rotational speed that is different than the first rotational speed.

Abstract: A method for assembling a gas turbine engine includes providing a high pressure turbine, providing a low-pressure turbine inner rotor that includes a first plurality of turbine blade rows configured to rotate in a first direction, providing a low-pressure turbine outer rotor that includes a second plurality of turbine blade rows configured to rotate in a second direction that is opposite the first direction, and coupling a plurality of bearings between the outer rotor and a turbine mid-frame between the high pressure turbine and at least one of the inner and outer rotors such that the plurality of bearings support a forward end of the outer rotor.

Abstract: A method for assembling a gas turbine engine that includes providing a low-pressure turbine inner rotor that includes a first plurality of turbine blade rows configured to rotate in a first direction, providing a low-pressure turbine outer rotor that includes a second plurality of turbine blade rows configured to rotate in a second direction that is opposite the first direction, coupling a turbine mid-frame assembly including a plurality of spokes within the engine such that the spokes are spaced axially forward of the inner rotor, coupling a bearing between the turbine mid-frame assembly and the inner rotor such that the inner rotor is rotatably coupled to the turbine mid-frame, and adjusting the plurality of spokes to align the bearing in a radial direction.

Abstract: A method for assembling a gas turbine engine includes providing a first fan assembly configured to rotate in a first rotational direction, rotatably coupling a second fan assembly to the first fan assembly, wherein the second fan assembly is configured to rotate in a second rotational direction that is opposite the first rotational direction, and coupling a lubrication system to the gas turbine engine such that a lubrication fluid is channeled through at least a portion of the second fan assembly.

Abstract: A method for assembling a gas turbine engine that includes providing a first fan assembly configured to rotate in a first rotational direction, rotatably coupling a second fan assembly to the first fan assembly, wherein the second fan assembly is configured to rotate in a second rotational direction that is opposite the first rotational direction, coupling a first shaft to the first fan assembly and to a first turbine rotor that is configured to rotate in a first rotational direction, coupling a second shaft coupled to the second fan assembly and to a second turbine rotor that is configured to rotate in a second rotational direction that is opposite the first rotational direction, and coupling a lubrication system to the gas turbine engine such that a lubrication fluid is channeled through the first shaft to lubricate at least one of the first and second fan assemblies.

Abstract: A method for assembling a gas turbine engine that includes providing a low-pressure turbine inner rotor that includes a first plurality of turbine blade rows configured to rotate in a first direction, and rotatably coupling a low-pressure turbine outer rotor to the inner rotor, wherein the outer rotor includes a second plurality of turbine blade rows that are configured to rotate in a second direction that is opposite the first rotational direction of the inner rotor and such that at least one of the second plurality of turbine blade rows is coupled axially forward of the first plurality of turbine blade rows.

Abstract: A method for assembling a gas turbine engine that includes providing a low-pressure turbine inner rotor that includes a first plurality of rows of turbine blades configured to rotate in a first rotational direction, providing a low-pressure turbine outer rotor that includes a second plurality of rows of turbine blades configured to rotate in a rotational direction that is opposite the first rotational direction, and coupling a support structure between the outer rotor and a turbine mid-frame such that the support structure supports a forward end of the outer rotor, and wherein the support structure includes a first portion that has a first coefficient of thermal expansion and a second portion that has a second coefficient of thermal expansion that is different than the first coefficient of thermal expansion.

Abstract: A method for assembling a gas turbine engine that includes providing a low-pressure turbine inner rotor configured to rotate in a first direction, providing a low-pressure turbine outer rotor configured to rotate in a second direction that is opposite the first rotational direction, and coupling at least one foil bearing to at least one of the inner and outer rotors to facilitate improving clearance control between a first rotating component and at least one of a second rotating component and a non-rotating component.

Abstract: A gas turbine engine having a longitudinal centerline axis therethrough, including: a fan section at a forward end of the gas turbine engine including at least a first fan blade row connected to a first drive shaft; a booster compressor positioned downstream of and in at least partial flow communication with the fan section including a plurality of stages, each stage including a stationary compressor blade row and a rotating compressor blade row connected to a drive shaft and interdigitated with the stationary compressor blade row; a core system positioned downstream of the compressor, where the core system further includes an intermediate compressor positioned downstream of and in flow communication with the booster compressor, the intermediate compressor being connected to a second drive shaft, and a combustion system for producing pulses of gas having increased pressure and temperature from a fluid flow provided to an inlet thereof so as to produce a working fluid at an outlet; and, a low pressure turbine

Abstract: A gas turbine engine includes a compressor powered by a turbine. The turbine includes a nozzle having vanes extending between outer and inner bands. Each vane includes an internal cooling plenum and a bypass tube extending through the bands. First and second manifolds surround the outer band and are disposed in flow communication with the plenums and bypass tubes, respectively. A bleed circuit joins the compressor to the manifolds for providing pressurized air thereto. A control valve modulates airflow to the first manifold and in turn through the cooling plenums of the vanes.

Abstract: A method facilitates assembling a gas turbine engine including a compressor and a rotor assembly coupled in axial flow communication downstream from the compressor. The method comprises coupling a bypass system in flow communication with the compressor to channel a portion of flow discharged from the compressor towards the rotor assembly is channeled through the bypass system, and coupling a downstream end of the bypass system within the gas turbine engine such that the flow entering the bypass system flows past the rotor assembly and is discharged downstream from the rotor assembly.

Abstract: An aircraft gas turbine engine assembly includes an inter-turbine frame axially located between high and low pressure turbines. Low pressure turbine has counter rotating low pressure inner and outer rotors with low pressure inner and outer shafts which are at least in part rotatably disposed co-axially within a high pressure rotor. Inter-turbine frame includes radially spaced apart radially outer first and inner second structural rings disposed co-axially about a centerline and connected by a plurality of circumferentially spaced apart struts. Forward and aft sump members having forward and aft central bores are fixedly joined to axially spaced apart forward and aft portions of the inter-turbine frame. Low pressure inner and outer rotors are rotatably supported by a second turbine frame bearing mounted in aft central bore of aft sump member. A mount for connecting the engine to an aircraft is located on first structural ring.

Abstract: An aircraft gas turbine engine assembly includes an inter-turbine frame axially located between high and low pressure turbines. Low pressure turbine has counter rotating low pressure inner and outer rotors with low pressure inner and outer shafts which are at least in part rotatably disposed co-axially within a high pressure rotor. Inter-turbine frame includes radially spaced apart radially outer first and inner second structural rings disposed co-axially about a centerline and connected by a plurality of circumferentially spaced apart struts. Forward and aft sump members having forward and aft central bores are fixedly joined to axially spaced apart forward and aft portions of the inter-turbine frame. Low pressure inner and outer rotors are rotatably supported by a second turbine frame bearing mounted in aft central bore of aft sump member. A mount for connecting the engine to an aircraft is located on first structural ring.

Abstract: A gas turbine engine including a temperature sensor assembly that continuously monitors an operating temperature within a multi-stage rotor assembly of the turbine engine is described. A plurality of cavities are defined within the multi-stage rotor assembly. The temperature sensor assembly includes a plurality of temperature sensor assemblies that monitor the temperature within each of the rotor assembly cavities. Each temperature sensor assembly includes a plurality of support guide tubes attached to the engine and extending from an outer casing of the engine to each cavity. A temperature sensor is inserted through the guide tubes and positioned within each cavity.

Abstract: An actuation system which includes apparatus for transferring control signals for varying the blade pitch from a non-rotating member to a rotating member of a gas turbine engine is described. In one embodiment, the engine includes a main disk containing a row of blades, and the disk is coupled to a main shaft which rotates about the engine axis. An actuating disk, or cone, rotates with the main disk, and the torque required to rotate the actuating cone is transmitted through ball splines attached to the main shaft. The ball splines allow transmission of torque between the disk and the cone and at the same time enable the disk and the cone to move axially relative to each other without binding. The actuating cone is supported at one end by the ball splines and at its other end by ball thrust bearings. The ball thrust bearings are pre-loaded to enable the bearings to withstand axial loads in opposite directions.