General Electric Company

Charles Stanley Orkiszewski

Praveen Sharma

Gert Johannes Van Der Merwe

Shivam Mittal

Brian Drummond

Nitesh Jain

Daniel Edward Mollmann

Richard Alan Wesling

Kudum Shinde

Weize Kang

Tod Robert Steen

Steven Alan Ross

Milind Chandrakant Dhabade

Shashank Suresh Puranik

Nagamohan Govinahalli Prabhakar

Thomas P Joseph

Anthony Michael Metz

Michael Riehle

Carl Lawrence MacMahon

Thomas Edward Agin

Steven Andrew Strang

Hahn M Spofford

Subramanya Shankar

Jorg Arrien Tiemon

Laurence Seelen

Andrew Bryan Cannella

A lubricant supply system for a squeeze film damper. The lubricant supply system includes a fluid passage configured to supply a lubricant to the squeeze film damper of a bearing assembly. The fluid passage includes a first flow path for lubricant into the lubricant supply system and into the squeeze film damper of the bearing assembly and a second flow path for lubricant into the squeeze film damper of the bearing assembly. The lubricant supply system includes a valve assembly configured to allow the fluid passage to move from the first flow path to the second flow path in response to an interruption of the first flow path.

Lubricant supply system

The present invention is directed to a turbomachinery blade and a method of fabricating the turbomachinery blade. The turbomachinery blade comprises a plurality of blade segments, and the plurality of the blade segments are separate from one another at least partially in an axial direction forming radial seams between adjacent blade segments. The method comprises forming a plurality of blade segments which are separate from one another at least partially in an axial direction; and forming radial seams between adjacent blade segments.

Turbomachinery blade and method of fabricating

A turbine engine assembly includes: a fan assembly; a turbine coupled to the fan assembly through a gearbox; a stationary component; and an assembly extending between the gearbox and the stationary component to couple the gearbox to the stationary component, wherein the assembly includes at least one vibration-reducing mechanism configured to isolate a vibratory response of the gearbox from the stationary component.

Turbine engine assembly and method of manufacturing thereof

Embodiments of systems and methods for operating a gas turbine engine defining a bowed rotor condition are generally provided. The systems and methods include rotating a rotor assembly defining a bowed rotor condition from approximately zero revolutions per minute (RPM) to within a bowed rotor mitigation speed range, in which the bowed rotor mitigation speed range is defined by a lower speed limit greater than zero RPM and an upper speed limit less than or equal to an idle speed condition of the gas turbine engine; applying a load at the rotor assembly via an energy storage device; adjusting the load to limit rotational speed or acceleration of the rotor assembly to within the bowed rotor mitigation speed range for a period of time; and removing the load to enable rotation of the rotor assembly to the idle speed condition following the period of time.

Method and system for mitigating bowed rotor operation of gas turbine engine

Embodiments are generally provided of a gas turbine engine including a rotor assembly comprising a shaft extended along a longitudinal direction, in which a compressor rotor and a turbine rotor are each coupled to the shaft; a casing surrounding the rotor assembly, in which the casing defines a first opening radially outward of the compressor rotor, the turbine rotor, or both, and a second opening radially outward of the compressor rotor, the turbine rotor, or both; a first manifold assembly coupled to the casing at the first opening; a second manifold assembly coupled to the casing at the second opening, in which the first manifold, the casing, and the second manifold together define a thermal circuit in thermal communication with the rotor assembly; and a fluid flow device in fluid communication with the first manifold assembly, in which the fluid flow device provides a flow of fluid to the first manifold assembly and through the thermal circuit, and further wherein the flow of fluid egresses the thermal circuit at the second manifold assembly.

Thermal gradient attenuation structure to mitigate rotor bow in turbine engine

The turbine assembly includes a shaft, a plurality of rotor disks, a plurality of blades, and at least one cutting mechanisms. The rotor disks are spaced axially along, and coupled to, the shaft and include a radially outer portion. The blades are spaced circumferentially around the radially outer portion. The cutting mechanisms are configured to cut the blades from the rotor disk during a disk over speed condition.

System and method for HPT disk over speed prevention

The present disclosure is directed to a gas turbine engine structure and method for reducing or mitigating bowed rotor. The method includes coupling a rotor assembly to a mechanical energy storage device via a clutch mechanism when the rotor assembly is at or below a speed limit below an idle speed condition; storing mechanical energy at the mechanical energy storage device via rotation of the rotor assembly at or below the speed limit; releasing mechanical energy from the mechanical energy storage device to rotate the rotor assembly following shutdown of the gas turbine engine; and rotating the rotor assembly via the mechanical energy from the mechanical energy storage device.

System and method for removing rotor bow in a gas turbine engine using mechanical energy storage device

A damper assembly includes a bearing assembly including a radially outer surface. A housing surrounds the bearing assembly and includes a radially inner surface facing the radially outer surface. The radially outer and inner surfaces define a plurality of annuli therebetween. The damper assembly further includes a plurality of fluid supplies coupled in flow communication with the plurality of annuli and configured to deliver a fluid to each annulus of the plurality of annuli. Each fluid supply of the plurality of fluid supplies independently controls the fluid within the respective annulus.

System and method for a variable squeeze film damper

A turbine engine assembly is provided. The assembly includes a stationary component, a drive shaft, and a gearbox coupled along the drive shaft, and coupled to the stationary component. The assembly also includes a vibration-reducing mechanism coupled between the stationary component and the gearbox. The vibration-reducing mechanism is configured to isolate a vibratory response of the gearbox from the stationary component.

Dampers for bearing assemblies of gas turbine engines are provided. For example, a damper comprises a damper housing, a lubricant passage defined in the damper housing, and a valve positioned within the lubricant passage. The lubricant passage includes an inlet and an outlet. In one embodiment, the valve is configured to restrict a flow of lubricant from the inlet into the damper when a reference temperature is below a threshold temperature. In another embodiment, the valve is configured to increase leakage of a lubricant through the outlet when a reference temperature is below a threshold temperature. In other embodiments, the valve is passively actuated and is configured to move within the lubricant passage to control a flow of lubricant from the inlet to the outlet based on a reference temperature.

Bearing damper for cold engine ground starts or in-flight re-starts

A method and apparatus for reducing dynamic loading of a gas turbine engine is provided. The gas turbine engine includes a rotor shaft assembly including a rotor shaft, a bearing assembly, a mounting race, and a support frame, the mounting race including a spherical surface. The method includes supporting the rotor shaft on the gas turbine engine support frame with the bearing assembly, the rotor shaft including a yield portion configured to permit bending of the rotor shaft during an imbalance operation.

Method and apparatus for supporting rotor assemblies during unbalances

An aircraft engine mount includes an annular casing fixedly disposed in the engine, a mounting platform spaced radially outwardly from the casing and fixedly joinable to an aircraft pylon, and a pair of circumferentially spaced apart links pivotally joined to the platform and the casing at respective radially outer and inner joints. The inner joints are circumferentially spaced closer together than the outer joints so that the intersection point of the longitudinal axes of the two links is disposed radially below the two links themselves for reducing lateral deflection of the casing relative to the engine mount.

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Cincinnati, OH, United States of America

Charles Stanley Orkiszewski