General Electric Company

Ge Infrastructure Technology, LLC

Ge Vernova Infrastructure Technology LLC

General Electric Technology Gmbh

Matthew Troy Hafner

John McConnell Delvaux

Zachary John Snider

Glenn Curtis Taxacher

Gary Michael Itzel

James Zhang

Sandip Dutta

Jason Robert Parolini

Srikanth Chandrudu Kottilingam

Canan Uslu Hardwicke

Victor John Morgan

Christopher Donald Porter

Frederic Woodrow Roberts, Jr

Peter De Diego

James Murray

Brad Wilson Vantassel

Jeffrey Clarence Jones

John Ellington Greene

Markus Feigl

John M Matthews

Jonathan Dwight Berry

Neelesh Nandkumar Sarawate

Stanley Kevin Widener

Joshua Lee Margolies

Thomas Raymond Farrell

Lyndsay Marie Kibler

Rajendra Kumar Bordia

Quan Li

Benjamin Paul Lacy

Michael John Hughes

Richard Martin DiCintio

Jacob John Kittleson

Russell Pierson DeForest

Joseph Vincent Citeno

Michael Anthony Wondrasek

David John Humanchuk

Bodhayan Dev

Brad Wilson VanTassel

Adam John Fredmonski

James Joseph Murray, Iii

Scott Francis Johnson

James Bernard Fehlberg

David Randall Hobart

Christopher West

Brad Wilson Van Tassel

John Mcconnell Delvaux

James Joseph Murray

A stator vane includes a platform that defines an opening. The stator vane further includes an airfoil that has a leading edge, a trailing edge, a suction side wall, and a pressure side wall. The airfoil extends radially between a base and a tip. At least one of the base or the tip includes a protrusion. The protrusion extends into the opening of the platform such that the platform surrounds the protrusion of the airfoil. The stator vane further includes a braze joint disposed between and fixedly coupling the platform and the protrusion of the airfoil. The stator vane further includes a cooling circuit defined in at least one of the protrusion or the platform to cool the braze joint.

Cooling circuit for a stator vane braze joint

A stator vane assembly includes a platform that defines an opening. The stator vane assembly further includes an airfoil segment that includes an airfoil and a boss. The airfoil includes a leading edge, a trailing edge, a suction side wall, and a pressure side wall. The airfoil extends radially between a base and a tip. The boss extends from at least one of the base or the tip of the airfoil. The boss is disposed in the opening of the platform. The stator vane assembly further includes a mechanical retention device that couples the platform to the boss of the airfoil segment.

Stator vane assembly having mechanical retention device

An airfoil includes a leading edge, a trailing edge, a base, and a tip. The airfoil further includes a pressure side wall and a suction side wall that extend between the leading edge, the trailing edge, the base, and the tip. The airfoil further includes a plurality of passages that are defined within the airfoil and extend from an inlet at one of the base or the tip. Each passage of the plurality of passages is defined at least partially by a primary impingement wall and a solid side wall. The primary impingement wall is spaced apart from one of the pressure side wall or the suction side wall such that a primary impingement gap is defined therebetween. The primary impingement wall defines a plurality of impingement apertures that direct air in discrete jets across the impingement gap to impinge upon an interior surface of the airfoil.

Turbomachine airfoil having impingement cooling passages

A ceramic matrix composite (CMC) component includes at least one seal surface, the at least one seal surface disposed adjacent an interfacing surface for providing a seal therebetween; and a coating disposed on the seal surface. The coating includes an aluminum oxide and/or a silicon dioxide.

Coating for improved surface finish

An additively manufactured (AM) object may include a body including an opening in an exterior surface thereof, the opening having a shape and a first area at the exterior surface of the body. A mask may be positioned over the opening. The mask has the shape of the opening and a second area that is larger than the first area so as to overhang the exterior surface of the body about the opening. A plurality of support ligaments couple to the mask and the exterior surface of the body at a location adjacent to the opening to support a portion of the mask. A coating can be applied to the object, and the mask removed. The final AM object includes a plurality of ligament elements extending from the exterior surface of the body and through the coating adjacent the opening, each ligament element at least partially surrounded by the coating.

Additively manufactured object using mask over opening for coating

A coolant delivery system for a component of a gas turbine system includes: a plurality of independent circuits of cooling channels embedded within an exterior wall of the component, each independent circuit of cooling channels including a plurality of headers and a plurality of feed tubes fluidly coupling the plurality of headers to a supply of cooling fluid; and an impingement plate connected to the exterior wall of the component by the plurality of feed tubes of the independent circuits of cooling channels, wherein, in each of the plurality of independent circuits of cooling channels, the cooling fluid flows through the plurality of feed tubes and the plurality of headers into the circuit of cooling channels only in response to a formation of a breach in the exterior wall of the component that exposes at least one of the cooling channels of the circuit of cooling channels.

Coolant delivery via an independent cooling circuit

A method may include: in a used metal turbomachine blade including a root including a shank, a platform coupled to the shank and an airfoil coupled to the platform, removing the airfoil, leaving a remaining base including the platform, the shank and the root. The method may also form a radially extending opening through the platform into the shank, and insert a ceramic shank nub extending from a ceramic airfoil into the radially extending opening of the remaining base. The ceramic airfoil is fixedly attached to the remaining base. The method allows reuse of the metal shank while providing the lower cooling requirements of a ceramic airfoil.

Method for replacing metal airfoil with ceramic airfoil, and related turbomachine blade

A turbine nozzle assembly system includes a plurality of nozzle sets, where each nozzle set forms an annulus. The nozzles in each set include an inner endwall and an outer endwall that include joint openings to receive the respective endwall mount ends of an airfoil. The airfoils across the plurality of nozzle sets have an inner endwall mount end and an outer endwall mount end that are identical amongst the plurality of nozzle sets. A wing portion of the airfoil has a selected wing shape that is identical within the respective nozzle set but different amongst the plurality of nozzle sets. In this manner, the endwalls can be removed from an airfoil and replaced with an airfoil having a different wing shape that provides a different pairwise throat area. The system allows changing of a pairwise throat area for a nozzle set without replacing the entirety of each nozzle.

Turbine nozzle assembly system with nozzle sets having different throat areas

An embodiment of an independent cooling circuit for selectively delivering cooling fluid to a component of a gas turbine system includes: a plurality of independent circuits of cooling channels embedded within an exterior wall of the component, wherein the plurality of circuits of cooling channels are interwoven together; an impingement plate; and a plurality of feed tubes connecting the impingement plate to the exterior wall of the component and fluidly coupling each of the plurality of circuits of cooling channels to at least one supply of cooling fluid, wherein, in each of the plurality of circuits of cooling channels, the cooling fluid flows through the plurality of feed tubes into the circuit of cooling channels only in response to a formation of a breach in the exterior wall of the component that exposes at least one of the cooling channels of the circuit of cooling channels.

An embodiment of an independent cooling circuit for selectively delivering cooling fluid to a component of a gas turbine system includes: at least one coolant feed channel fluidly coupled to a supply of cooling fluid; and an interconnected circuit of cooling channels, including: an interconnected circuit of cooling channels embedded within an exterior wall of the component; an impingement plate; and a plurality of feed tubes connecting the impingement plate to the exterior wall of the component and fluidly coupling a supply of cooling fluid to the interconnected circuit of cooling channels; wherein the cooling fluid flows through the plurality of feed tubes into the interconnected circuit of cooling channels only in response to a formation of a breach in the exterior wall of the component that exposes at least one of the cooling channels.

A turbine component includes a root and an airfoil extending from the root to a tip opposite the root. The airfoil forms a leading edge and a trailing edge portion extending to a trailing edge. Radial cooling channels in the trailing edge portion of the airfoil permit radial flow of a cooling fluid through the trailing edge portion. Each radial cooling channel has a first end at a lower surface at a root edge of the trailing edge portion or at an upper surface at a tip edge of the trailing edge portion and a second end opposite the first end at the lower surface or the upper surface. A method of making a turbine component and a method of cooling a turbine component are also disclosed.

Turbine component and methods of making and cooling a turbine component

A turbine component includes a root and an airfoil extending from the root to a tip opposite the root. The airfoil forms a leading edge and a trailing edge portion extending to a trailing edge. A plurality of axial cooling channels in the trailing edge portion of the airfoil are arranged to permit axial flow of a cooling fluid from an interior of the turbine component at the trailing edge portion to an exterior of the turbine component at the trailing edge portion. A method of making a turbine component includes forming an airfoil having a trailing edge portion with axial cooling channels. The axial cooling channels are arranged to permit axial flow of a cooling fluid from an interior to an exterior of the turbine component at the trailing edge portion. A method of cooling a turbine component is also disclosed.

A flow regulating system for increasing a flow of cooling fluid supplied to a cooling system of a component of a gas turbine system is provided. The flow regulating system includes: a pneumatic circuit embedded within a section of the component, the pneumatic circuit including a set of interconnected pneumatic passages; and a pressure-actuated switch fluidly coupled to the pneumatic circuit. The pressure-actuated switch is activated in response to a formation of a breach in the section of the component and an exposure of at least one of the pneumatic passages of the pneumatic circuit embedded in the section of the component. The activation of the pressure-actuated switch increases the flow of cooling fluid supplied to the cooling system of the component.

Oxidation activated cooling flow

An apparatus is disclosed, including a first article, a second article, at least one interface structure, and a thermal break directly adjacent to the at least one interface structure. The first article includes a first material composition having a first thermal tolerance. The second article includes a second material composition having a second thermal tolerance greater than the first thermal tolerance. The first article and the second article are in contact with one another through the interface structure. The thermal break interrupts a thermal conduction path from the second article to the first article.

Apparatus with thermal break

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Honea Path, SC, United States of America

Matthew Troy Hafner