General Electric Company

Jorge Francisco Seda

Thomas Ory Moniz

Robert Joseph Orlando

Peter Nicholas Szucs

James Edward Johnson

Ambrose A Hauser

John Christopher Brauer

Lawrence Wayne Dunbar

Paul Daniel Kemper

Alan Roy Stuart

Philip Roger Gliebe

Justin P Stephenson

Gregory T Steinmetz

Lawrence Nmn Butler

Kenneth S Scheffel

Justin Paul Stephenson

An inner core cowl baffle assembly for a turbofan engine assembly is provided. The engine assembly includes a core gas turbine engine, a core cowl which circumscribes the core gas turbine engine, a nacelle positioned radially outward from the core cowl, and a fan nozzle duct defined between the core cowl and the nacelle. The inner core cowl baffle assembly includes an inner core cowl baffle, and an actuator assembly configured to vary the throat area of the fan nozzle duct by selectively repositioning the inner core cowl baffle with respect to the core cowl.

Method of operating a turbofan engine cowl assembly

A method for operating a turbofan engine includes providing a first fan disk configured to rotate in a first rotational direction, coupling a first fan blade to the first fan disk, the first fan blade including a dovetail, an airfoil coupled to the dovetail, and a plurality of channels extending through the dovetail, and channeling airflow through the at least one fan blade such that the airflow is discharged from the airfoil trailing edge to facilitate reducing the operational noise level of the turbofan engine and improve.

Gas turbine engine fan assembly and method for assembling to same

Turbofan engine cowl assembly and method of operating the same

A nozzle assembly for a gas turbine aircraft engine is provided. The nozzle assembly includes a nacelle and a core cowl positioned at least partially within the nacelle such that the core cowl and the nacelle are aligned substantially concentrically to each other defining an annular fan bypass duct. A first member couples the nacelle to the core cowl and includes a first flap hingedly coupled to each sidewall of the first member. An opposing second member couples the nacelle to the core cowl and includes a second flap hingedly coupled to each sidewall of the second member. The first flaps and the second flaps are selectively positionable between a first operational position and a second operational position to vary a throat area of said fan bypass duct. A turbofan engine assembly and a method for operating the same are also provided.

Turbofan engine nozzle assembly and method of operating same

A turbofan engine assembly is provided. The turbofan engine assembly includes a core gas turbine engine, a core cowl which circumscribes the core gas turbine engine, a nacelle positioned radially outward from the core cowl, a fan nozzle duct defined between the core cowl and the nacelle, and an inner core cowl baffle assembly positioned within the fan nozzle duct. The inner core cowl baffle assembly includes a first core cowl baffle coupled to a portion of the core cowl within the fan nozzle duct, a second core cowl baffle coupled to a portion of the core cowl and positioned a distance radially inward from the first core cowl baffle, and an actuator assembly configured to vary the throat area of the fan nozzle duct by selectively repositioning the second core cowl baffle with respect to the first core cowl baffle. A method for operating the turbofan engine assembly is also provided.

A method for operating a turbofan engine assembly is provided. The turbofan engine assembly includes a core cowl which circumscribes the core gas turbine engine, a nacelle positioned radially outward from the core cowl, a fan nozzle duct having an area defined between the core cowl and a portion of the nacelle, a first cowl and a second cowl that define a portion of the nacelle wherein the second cowl is repositionable with respect to the first cowl, and a thrust reverser assembly wherein the second cowl surrounds the thrust reverser assembly. The method includes varying an operating speed of the fan assembly from a first operating speed to a second operating speed. The method further includes selectively positioning the second cowl between a first operational position and a second operational position to vary the area of the fan nozzle duct to facilitate improving engine efficiency at the second operating speed. The method further includes selectively positioning the second cowl between the second operational position and a third operational position to vary an amount of air flowing through the fan nozzle duct and the thrust reverser assembly, wherein the second operational position substantially prevents airflow from flowing through the thrust reverser assembly to improve the efficiency of the turbofan engine, and wherein the third operational position directs airflow through the thrust reverser assembly.

Turbofan engine cowl assembly

A method for operating a turbofan engine assembly including a core gas turbine engine is provided. The method includes varying an operating speed of the turbofan engine assembly from a first operating speed to a second operating speed. The method also includes selectively positioning a first arcuate portion and a second arcuate portion of a split cowl assembly to vary a throat area of a fan nozzle duct defined downstream from the core gas turbine engine to facilitate improving engine efficiency at the second operating speed. The split cowl assembly is downstream from the core gas turbine engine and inside the fan nozzle duct. A turbofan engine assembly and nozzle assembly are also provided.

Turbofan engine nozzle assembly and method for operating the same

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.

Lubrication system for a counter-rotating turbine engine and method of assembling same

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.

Counter-rotating turbine engine and method of assembling same

An aircraft engine turbine frame includes a first structural ring, a second structural ring disposed co-axially with and radially spaced inwardly of the first structural ring about a centerline axis. A plurality of circumferentially spaced apart struts extend between the first and second structural rings. Forward and aft sump members having forward and aft central bores are fixedly joined to forward and aft portions of the turbine frame respectively. A frame connecting means for connecting the engine to an aircraft is disposed on the first structural ring. The frame connecting means may include a U-shaped clevis. The frame may be an inter-turbine frame axially located between first and second turbines of first and second rotors of a gas turbine engine assembly. An axial center of gravity of the second turbine passes though or very near a second turbine frame bearing supported by the aft sump member.

Aircraft engine with inter-turbine engine frame

An aircraft gas turbine engine includes high and low pressure turbines having respective counter rotatable low pressure inner and outer rotors with low pressure inner and outer shafts in part rotatably disposed co-axially within a high pressure rotor and drivingly connected to first and second fan blade rows and first and second boosters respectively. A bypass duct radially bounded by a fan casing and an annular radially inner bypass duct wall surrounds the boosters axially located between the first and second fan blade rows. The engine has a high pressure compressor operable to produce an overall pressure ratio in a range of about 40-65 and a fan inlet hub to tip radius ratio in a range between 0.20 and 0.35, a bypass ratio in a range of 5-15, an operational fan pressure ratio in a range of 1.4-2.5, and a sum of operational fan tip speeds in a range of 1000 to 2500 feet per second.

Counter rotating aircraft gas turbine engine with high overall pressure ratio compressor

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.

Aircraft engine with inter-turbine engine frame supported counter rotating low pressure turbine rotors

A recirculation passageway for a turbine engine provides stall protection in a booster by directing high pressure airflow from a flow path of the booster to the passageway. The high pressure airflow loses energy and decreases in pressure while traveling through the passageway until re-entry into the booster flow path. The airflow recirculates in the passageway until the airflow is discharged through a high pressure compressor.

Booster recirculation passageway and methods for recirculating air

A turbine engine including an oil flow path through the low pressure system shaft which, as compared to at least one known oil flow path, facilitates increasing the rated torque and life of the low pressure system shaft is described. In one embodiment, the lubrication opening in the low pressure shaft is oriented so that with respect to a plane along which an outer surface of the low pressure shaft substantially extends, the center axis of the lubrication opening is substantially normal to the plane. As a result of this configuration, the stress concentration in the low pressure shaft at the location of the lubrication opening is substantially reduced. In addition, such a lubrication opening is relatively easy to form as compared to the known, angularly oriented lubrication shaft. Further, the lubrication opening in the present configuration can have a larger diameter than the known angularly oriented lubrication opening due to the reduced stress concentration, and such a larger lubrication opening facilitates more free flow of lubrication through the opening to the differential bearing.

Growing community of inventors

Cincinnati, OH, United States of America

Jorge Francisco Seda