General Electric Company

Other

Gary Michael Itzel

Sandip Dutta

Benjamin Paul Lacy

Steven Sebastian Burdgick

Matthew Troy Hafner

Yufeng Phillip Yu

John McConnell Delvaux

Xiuzhang James Zhang

James Zhang

Ronald Scott Bunker

Srikanth Chandrudu Kottilingam

Ibrahim Sezer

James William Vehr

Robert Henry Devine, Ii

Adebukola Oluwaseun Benson

Doyle Clyde Lewis

Ariel Caesar-Prepena Jacala

Nicholas William Rathay

James Albert Tallman

Jaime Javier Maldonado

John Alan Eastman

Sergio Daniel Marques Amaral

Waylon Willard Webbon

Aaron Ezekiel Smith

David Wayne Weber

David Edward Schick

Victor John Morgan

David Vincent Bucci

Kevin Richard Kirtley

Ganjiang Feng

Joseph Anthony Weber

Iain Robertson Kellock

Scott Edmond Ellis

Jeffrey Clarence Jones

Christopher Donald Porter

Kathleen Blanche Morey

Dipankar Pal

Ariel Caesar Prepena Jacala

Michael Adelbert Sullivan

Sanjay Chopra

Robert Carl Meyer

Sarah Jane Osgood

Radhakrishna Bagepalli

Curtis Alan Johnson

Bin Wei

Zachary John Snider

Gregory Thomas Foster

Andrew Philip Shapiro

Neelesh Nandkumar Sarawate

Christian Lee Vandervort

Nesim Abuaf

Douglas Scott Byrd

Brendon James Leary

Fred Thomas Willett, Jr

Brian Peter Arness

Jacob John Kittleson

Stephen Paul Wassynger

Larry W Plemmons

Abdul-Azeez Mohammed-Fakir

Kevin Leon Bruce

Dean Thomas Lenahan

Ahmad Safi

Alan Majors Walker

Jason Dean Fuller

David Richard Johns

Robert Walter Coign

Philip Lynn Andrew

Jason Edward Albert

Victor H Correia

Carlos Miguel Miranda

Bodhayan Dev

Charles Lewis Davis, Iii

Satoshi Atsuchi

Gustavo Adolfo Ledezma

Bryan David Lewis

Kevin Thomas McGovern

Calvin Levy Sims

Christopher Michael Penny

Felipe Roman-Morales

Camilo Andres Sampayo

Nicholas Alvin Hogberg

Mitchell Allan Merrill

Charles Alan Bulgrin

Evan Andrew Sewall

Anshuman Singh

Markus Feigl

Subodh Diwakar Deodhar

Ralph Chris Bruner

Dimitrios Stathopoulos

Yang Liu

Ariel C Jacala

Victor H S Correia

Charles Andrew Malinowski

Onika Misasha Kerber

Matthew James Findlay

Andrew P Elovic

Nagendra Karthik Depuru Mohan

Joshua R Kornau

Thomas Nelson Toornman

Arun Meenakshinatha Iyer

The present disclosure is directed to a turbomachine that includes a hot gas path component having an inner surface and defining a hot gas path component cavity. An impingement insert is positioned within the hot gas path component cavity. The impingement insert includes an inner surface and an outer surface and defines an impingement insert cavity and a plurality of impingement apertures fluidly coupling the impingement insert cavity and the hot gas path component cavity. A plurality of pins extends from the outer surface of the impingement insert to the inner surface of the hot gas path component.

Impingement insert for a gas turbine engine

A component includes an outer wall that includes an exterior surface, and at least one plenum defined interiorly to the outer wall and configured to receive a cooling fluid therein. The component also includes a coating system disposed on the exterior surface. The coating system has a thickness. The component further includes a plurality of adaptive cooling openings defined in the outer wall. Each of the adaptive cooling openings extends from a first end inflow communication with the at least one plenum, outward through the exterior surface and to a second end covered underneath at least a portion of the thickness of the coating system.

Coated components having adaptive cooling openings and methods of making the same

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.

An airfoil having a wall structure including a plurality of spaced walls for improved cooling and lifetime is disclosed. The airfoil and walls are made by additive manufacturing. The airfoil includes an exterior wall, an intermediate wall, and an interior wall each separated from adjacent walls by a plurality of standoff members; a plurality of outer cooling chambers defined between the exterior and intermediate walls, the chambers partitioned by an outer partition; a plurality of intermediate cooling chambers defined between the intermediate and interior walls, the chambers partitioned by an intermediate partition; a thermal barrier coating on each of the exterior wall and the intermediate wall; a first plurality of impingement openings through the intermediate wall; a second plurality of impingement openings through the interior wall; and a plurality of cooling passages through the exterior wall.

Turbine airfoil with multiple walls and internal thermal barrier coating

An airport having an exterior wall including a plurality of spaced layers for improved cooling and lifetime is disclosed. The airfoil and exterior wall are made by additive manufacturing. The exterior wall includes an exterior layer, an intermediate layer, and an interior layer each separated from adjacent layers by a plurality of standoff members; a plurality of first cooling chambers between the exterior and intermediate layers, the chambers partitioned by a first partitioning wall; a plurality of second cooling chambers between in the intermediate and interior layers, the chambers partitioned by a second partitioning wall; a thermal barrier coating on the exterior layer; a plurality of impingement openings in the intermediate layer and a second plurality of impingement openings in the interior layer; and a plurality of cooling passages in the exterior layer. The exterior layer may also include plateaus on an exterior face through which the cooling passages extend.

Turbine airfoil multilayer exterior wall

Turbine components are disclosed including a component wall defining a constrained portion, a manifold having an impingement wall, and a post-impingement cavity disposed between the manifold and the component wall. The impingement wall includes a wall thickness and defines a plenum and a tapered portion. The tapered portion tapers toward the constrained portion and includes a plurality of impingement apertures and a wall inflection. The wall inflection is disposed proximal to the constrained portion, and the tapered portion is integrally formed as a single, continuous object. The wall inflection may include an inflection radius of less than about 3 times the wall thickness of the impingement wall, or the tapered portion may include a consolidated portion with the impingement wall extending across the plenum. A method for forming the turbine component is also disclosed, including forming the tapered portion as a single, continuous tapered portion by an additive manufacturing technique.

Turbine components and method for forming turbine components

The present disclosure is directed to a rotor blade for a turbomachine. The rotor blade includes an airfoil defining a cooling passage extending between a root and a tip of the airfoil. The airfoil further defines a pocket positioned between the cooling passage and an exterior surface of the airfoil and spaced apart from the cooling passage and the exterior surface. The pocket provides a thermal impediment between the cooling passage and the exterior surface of the airfoil.

Turbomachine rotor blade pocket

A cooling assembly includes a cooling chamber disposed inside of a turbine assembly. The cooling chamber directs cooling air inside an airfoil of the turbine assembly. The cooling assembly includes a metered channel fluidly coupled with the cooling chamber. The metered channel directs at least some of the cooling air out of the cooling chamber outside of a rail surface of the airfoil. The metered channel is elongated along and encompasses an axis. The metered channel has an interior surface with a distance between opposing first portions of the interior surface. The distance between the opposing first portions decreases at increasing distances along the axis from the cooling chamber toward the rail surface.

Cooling assembly for a turbine assembly

A turbine component includes an airfoil having an airfoil chamber disposed within the airfoil, the airfoil chamber configured to supply a coolant through the airfoil. The tip of the airfoil includes a rail extending radially from the tip plate, the rail including an inner rail surface defining a tip pocket therein, an outer rail surface and a radially outward facing rail surface between the inner rail surface and the outer rail surface. A tip rail cavity is within and partially circumscribes the rail, the tip rail cavity receiving a coolant flow. A tip rail cooling passage includes an inlet fluidly coupled to the tip rail cavity, a passage length fluidly coupled to the inlet and partially circumscribing the rail, a metering element fluidly coupled to the passage length, and an outlet fluidly coupled to the metering element and extending through the radially outward facing rail surface.

Turbine component with tip rail cooling passage

An apparatus is disclosed including a first and second article, a first interface volume disposed between and enclosed by the first article and second article, a cooling fluid supply, and at least one cooling fluid channel in fluid communication with the cooling fluid supply and the first interface volume. The first article includes a first material composition. The second article includes a second material composition. The at least one cooling fluid channel includes a heat exchange portion disposed in at least one of the first and second article downstream of the cooling fluid supply and upstream of the first interface volume. A turbine shroud is disclosed wherein the first and second articles are an outer and inner shroud. A turbine nozzle is disclosed wherein the first and second articles are an endwall and fairing.

Apparatus, turbine nozzle and turbine shroud

A component configured for impingement cooling includes an inner wall defining a plurality of apertures. Each aperture of the plurality of apertures is configured to emit a cooling fluid therethrough. The component also includes an outer wall spaced from the inner wall. The outer wall and the inner wall extend along a longitudinal axis of the component. The component further includes a plurality of angled walls extending between the inner wall and the outer wall. The plurality of angled walls define a plurality of angled channels in fluid communication with the plurality of apertures. Each angled wall of the plurality of angled walls extends at an acute angle relative to the longitudinal axis.

Components having channels for impingement cooling

A seal in a gas turbine for sealing a radial gap defined between rotating and stationary structure. The rotating structure may include a row of rotor blades. The seal may include a pad attached to the stationary structure. The pad may include an abradable structure. The pad may further include a liner attached to and covering an outer surface of the abradable structure. The stationary structure to which the pad is attached may define an axial section of the outer radial boundary of the annular flowpath, the axial section coinciding axially with an axial position of the row of rotor blades.

Gas turbine seals

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Simpsonville, SC, United States of America

Gary Michael Itzel