General Electric Company

Mark Willard Marusko

Joshua Brian Jamison

Mark Eugene Noe

Aaron Michael Dziech

Michael James Verrilli

Joseph Charles Albers

Paul Izon

James Herbert Deines

Christopher Donald Johnson

Barry Allan Wilson

Toby George Darkins, Jr

Eric Alan Estill

John Peter Heyward

Michael Alan Stieg

James Thomas Hawkins

Brian Christopher Towle

Anthony Reid Paige

Toby George Darkins

Gregory Terrence Garay

Xi Yang

Douglas Gerard Konitzer

Suresh Subramanian

Brian David Przeslawski

Michael John McCarren

Brian Patrick Peterson

Michael John Mccarren

Steven Robert Brassfield

James Thompson

Donald Michael Corsmeier

Richard William Albrecht, Jr

Craig Douglas Young

Kevin Samuel Klasing

David Edward Bulman

Donald Lee Gardner

Mark Joseph Mielke

Daniel Lee Durstock

Daniel Dale Brown

Mark Edward Stegemiller

James Edwin Rhoda

David Bruce Patterson

Donald Scott Yeager

Richard William Jendrix

Glenn Edward Wiehe

Brandon Flowers Powell

Michael Verrilli

Thomas Ryan Wallace

Robert Paul Tyler

Andrew John Breslin

Christopher Ryan Johnson

Michael D Carroll

Anthony Joseph Venzon

Cory Michael Williams

Michael Joseph Zerbe

Christopher Hasse

Steve Mark Molter

Joshua B Jamision

Renee Solomon Wenstrup

Robert Tyler

Ii Paige Anthony Reid

Toby George Darkins Jr

Anthony Reid Paige Ii

A combustor assembly for a gas turbine engine including an outer casing, the gas turbine engine defining an axial direction, the combustor assembly includes a liner at least partially defining a combustion chamber extending between an aft end and a forward end generally along the axial direction within the outer casing, the liner including an inner surface facing the combustion chamber and an outer surface facing away from the combustion chamber, and a damper assembly extending between the outer casing and the outer surface of the liner, the damper assembly including a selectively separable support and a damper spring, the damper spring being disposed between the support and the liner.

Combustor assembly for a turbine engine

A combustor assembly for a gas turbine engine includes an outer liner that at least partially defines a combustion chamber. The outer liner extends between an aft end and a forward end generally along an axial direction within an outer casing of the gas turbine engine. The outer liner includes an inner surface facing the combustion chamber and an outer surface facing away from the combustion chamber. The combustor assembly includes a radial damper assembly disposed against the outer liner at a plurality of circumferential points.

Partial integrated core-shell investment casting molds that can be assembled into complete molds are provided herein. Each section of the partial mold may contain both a portion of a core and portion of a shell. Each section can then be assembled into a mold for casting of a metal part. The partial integrated core-shell investment casting molds and the complete molds may be provided with filament structures corresponding to cooling hole patterns on the surface of the turbine blade or the stator vane, which provides a leaching pathway for the core portion after metal casting. Core filaments that can be used to supplement the leaching pathway, for example in a core tip portion of the mold are also provided herein.

Method for making an integrated core-shell structure

Multi-piece integrated core-shell structure for making cast component

A gas turbine engine and combustor assembly is generally provided. The gas turbine engine may include an outer casing while the combustor assembly may include a liner and a damper assembly. The liner may at least partially define a combustion chamber extending between an aft end and a forward end generally along an axial direction within the outer casing. The liner may include an inner surface facing the combustion chamber and an outer surface facing away from the combustion chamber. The damper assembly may extend between the outer casing and the outer surface of the liner. The damper assembly may include a selectively separable support and damper spring. The damper spring may be disposed between the support and the liner.

A turbine blade sealing structure for a ceramic matrix composite component is provided. The turbine sealing structure includes at least one top ply abutting a top portion of a shank. The at least one top ply extends out over the top portion of the shank. The turbine sealing structure includes at least one side ply abutting a side portion of the shank. The at least one side ply extends along a side portion of the shank. The at least one top ply and at least one side ply form an angel wing and a skirt with on the shank of the ceramic matrix composite component.

Turbine blade sealing structure

A rotor blade assembly for a rotor of a gas turbine engine having an axis of rotation includes a shank portion formed from a ceramic matrix composite (CMC) material. The rotor blade assembly also includes a platform portion formed from a substantially similar CMC material as that of the shank portion. The platform portion is coupled to the shank portion. The platform portion and the shank portion cooperate to at least partially define two opposing side portions of the rotor blade assembly. The opposing side portions are angularly separated with respect to the axis of rotation. The rotor blade assembly further includes a damper retention apparatus. The damper retention apparatus is coupled to the shank portion. The damper retention apparatus includes at least one angled bracket apparatus extending toward a circumferentially adjacent rotor blade assembly.

Turbine rotor blade assembly and method of assembling same

Methods for fabricating a component of a gas turbine engine are provided. In one embodiment, the method includes molding a CMC material to form a first portion of the gas turbine engine component, processing the first portion to form a first assembly, preparing the first assembly and a second portion of the gas turbine engine component for processing, and processing the first assembly and second portion to form a second assembly. In another embodiment, the method includes processing a first plurality of CMC plies to form a first assembly; positioning the first assembly and a second plurality of CMC plies on a tool for processing, the first assembly defining a first plane, the second plurality of plies defining a second plane, wherein the second plane is perpendicular to the first plane; and processing the first assembly and the second plurality of plies to form a second assembly.

Fabrication of gas turbine engine components using multiple processing steps

A clearance control apparatus for a gas turbine engine including an annular turbine case having opposed inner and outer surfaces; an annular manifold surrounding a portion of the turbine case, the manifold including: an inlet port in fluid communication with the manifold and the outer surface of the turbine case, and an exit port; and a bypass pipe having an upstream end coupled to the exit port, a downstream end coupled to a low-pressure sink, and a valve disposed between upstream and downstream ends, the valve selectively moveable between a first position which blocks flow between the upstream and downstream ends, and a second position which permits flow between the upstream and downstream ends.

Suction-based active clearance control system

A process for producing airfoil components containing ceramic-based materials and having a tip cap. The process entails forming an airfoil portion of the component from an airfoil portion material that contains a precursor of a ceramic-based material. The airfoil portion material defines concave and convex walls of the airfoil portion, and the concave and convex walls define a tip region of the airfoil portion and at least one cavity within the airfoil portion. At least a first ply is formed that contains a precursor of a ceramic-based material, and the first ply at least partially closes the cavity at the tip region of the airfoil portion. The airfoil portion material and the first ply are then cured so that the first ply forms a tip cap that closes the cavity and the precursors of the airfoil portion material and first ply are converted to the ceramic-based materials thereof.

Airfoil components containing ceramic-based materials and processes therefor

Hybrid turbine airfoil components containing a ceramic material, in which detailed features of the components are formed of materials other than ceramic materials. The components include a first component formed of a ceramic-based material and a second component formed of a metallic material. The first component comprises an airfoil portion and a nub, and the second component is separately formed and attached to the first component by casting the metallic material around the nub of the first component. The second component includes a platform portion between the airfoil portion and the nub of the first component and a dovetail portion on the nub of the first component. Each of the platform and dovetail portions has at least one off-axis geometric feature that results in the second component having a more complex geometry than the first component.

Turbine airfoil components containing ceramic-based materials and processes therefor

An airfoil for a gas turbine engine includes a first sidewall and a second sidewall coupled together at a leading edge and a trailing edge, such that a cavity is defined therebetween. A plurality of cooling circuits are defined within the cavity. Each cooling circuit channels cooling fluid through at least one cooling chamber to facilitate cooling the airfoil. More specifically, a cascade impingement circuit, a down pass circuit, a flag tip circuit, and a trailing edge circuit are provided. The cascade impingement circuit includes a central chamber and a plurality of impingement chambers.

Method and apparatus for cooling gas turbine rotor blades

A turbine flowpath apparatus for a gas turbine engine includes: a flowpath component exposed at least partially to a primary combustion gas flowpath of the engine, the flowpath component comprising low-ductility material; a metallic annular stationary structure surrounding the flowpath component, including a bearing surface which bears against the flowpath component, so as to restrain the flowpath component from axial movement in a first direction; and a spring element disposed between the flowpath component and the stationary structure which resilient urges the flowpath component in the first direction against the bearing surface.

Resilient mounting apparatus for low-ductility turbine shroud

A method of assembling a combustion system for a gas turbine engine includes providing a combustion chamber frame, an inner casing structure, and an outer casing structure. The method also includes mounting the combustion chamber frame between the inner casing structure and the outer casing structure such that the combustion chamber frame is coupled to the inner casing structure and the outer casing structure.

Combustion system and method of assembling the same

A turbine shroud apparatus for a gas turbine engine having a central axis includes: an arcuate shroud segment comprising low-ductility material and having a cross-sectional shape defined by opposed forward and aft walls, and opposed inner and outer walls, the walls extending between opposed first and second end faces and collectively defining a shroud cavity; and an annular stationary structure surrounding the shroud segment, where the shroud segment is mechanically coupled to the stationary structure. The stationary structure includes at least one axially-facing bearing surface which is in direct contact with the shroud segment, and the shroud segment is disposed so as to absorb at least one axially-aligned force and transfer the axially-aligned force to the bearing surface.

Structural low-ductility turbine shroud apparatus

Growing community of inventors

Springboro, OH, United States of America

Mark Willard Marusko