General Electric Company

Gulcharan Singh Brainch

Michael Joseph Danowski

Robert David Briggs

Jonathan Philip Clarke

Robert Proctor

Curtis Walton Stover

Ronald Scott Bunker

Roger Dale Wustman

Dean Thomas Lenahan

Robert Joseph Albers

John Christopher Brauer

Ching-Pang Lee

Ambarish Jayant Kulkarni

Byron Andrew Pritchard

Larry W Plemmons

Kevin Robert Feldmann

David B Stewart

Sean Robert Keith

Steven Alan Ross

Jason Randolph Allen

Shawn Michael Pearson

Todd Stephen Heffron

Satoshi Atsuchi

John R Hess

Anne M Isburgh

David A Di Salle

Ronald Eugene McRae, Jr

William Ross Miller

Jonathan Michael Rausch

Andrew David Kemp

Edward P Brill

Frederick Alan Buck

Richard W Albrecht

Michael Alan Meade

Jonathan Phillip Clarke

Ronald Eugene Mcrae Jr

An engine component for a gas turbine engine includes a film-cooled substrate having a hot surface facing hot combustion gas and a cooling surface facing a cooling fluid flow. The substrate includes one or more film holes that have a metering section defining a metering diameter and a diffusing section that defines a hooded length.

Engine component with film hole

Gas turbine engine components are provided which utilize an insert to provide cooling air along a cooled surface of an engine component. The insert provides cooling holes or apertures which face the cool side surface of the engine component and direct cooling air onto that cool side surface. The apertures may be formed in arrays and directed at an oblique or a non-orthogonal angle to the surface of the insert and may be at an angle to the surface of the engine component being cooled. An engine component assembly is provided with counterflow impingement cooling, comprising an engine component cooling surface having a cooling fluid flow path on one side and a second component adjacent to the first component. The second component may have a plurality of openings forming an array wherein the openings extend through the second component at a non-orthogonal angle to the surface of the second component.

Engine components with impingement cooling features

An apparatus for a gas turbine engine can include an airfoil having an interior. The interior can be separated into one or more cooling air channels extending in a span-wise direction. An accelerator insert can be placed in one or more cooling air channels to define a reduced cross-sectional area within the cooling air channel to accelerate an airflow passing through the cooling air channel.

Accelerator insert for a gas turbine engine airfoil

A structure for disrupting the flow of a fluid comprises and a second lateral wall spaced apart from one another, yet joined, by a floor and a ceiling; and, (b) a vorticor pin extending in a direction parallel to an X-axis, the vorticor pin concurrently rising above and extending away from the floor to a height, in a direction parallel to a Y-axis, the vorticor pin comprising: (i) a front surface extending incompletely between the first lateral wall and the second lateral wall, the front surface extending above the floor and having an arcuate portion that is transverse with respect to a Z-axis, which is perpendicular to the X-axis and the Y-axis, and (ii) a rear surface extending between the first lateral wall and the second lateral wall, the rear surface extending between the front surface and the floor, the rear surface having an inclining section that tapers in height, taken parallel to the Y-axis, in a direction parallel to the Z-axis.

Fastback vorticor pin

A shroud assembly for a turbine section of a turbine engine includes a shroud plate in thermal communication with a hot combustion gas flow and a baffle overlying the shroud plate to define a region. One or more shaped cooling features are located along the region such that a cooling fluid flow passing through the region encounters the shaped cooling features to increase the turbulence of the cooling fluid flow.

Shroud assembly for a turbine engine

An airfoil core die includes a tab for forming a continuous and smooth contour and extension of a first trailing edge slot recessed all to a bottom of the airfoil. The contour is formed on each airfoil core at injection and is maintained through to finish casting of the airfoil.

Apparatus and methods for cooling slot step elimination

A method facilitates fabricating a rotor assembly for a gas turbine engine. The method comprises providing a plurality of rotor blades that each include an airfoil, a dovetail, a shank, and a platform, wherein the platform extends between the shank and the airfoil, and wherein the dovetail extends outwardly from the shank, and forming a cooling circuit within a portion of the shank to supply cooling air to the rotor blade for supplying cooling air to the rotor blade for impingement cooling a portion of the rotor blade and for supplying cooling air to the rotor blade for purging a cavity defined downstream from the rotor blade.

Methods and apparatus for cooling gas turbine engine rotor assemblies

A method facilitates assembling a turbine nozzle for a gas turbine engine. The method includes providing a turbine nozzle including a plurality of airfoil vanes that extend between an inner band and an outer band, and forming a compound radii fillet that extends between a first of the airfoil vanes and the outer band, such that at least a second of the airfoil vanes is coupled to the outer band only by a single radii fillet.

Methods and apparatus for assembling gas turbine nozzles

A method enables a turbine nozzle for a gas turbine engine to be assembled. The method includes providing a turbine nozzle including a plurality of airfoil vanes extending between an inner band and an outer band, wherein the outer band includes at least one mounting system that extends radially outwardly therefrom and includes a rail and at least one hook, and coupling the turbine nozzle into the gas turbine engine using the mounting system such that the turbine nozzle is at least partially supported by at least one hook. The method also includes positioning a seal assembly between at least one hook and the outer band to facilitate reducing radial leakage through the turbine nozzle.

A cooled turbine element including an airfoil and a flowpath boundary member extending laterally from either an inboard end or an outboard end of the airfoil. The member has a flowpath face and an outside face which is cooler than said flowpath face creating a tendency for the member to deflect in a direction away from the flowpath face and causing a thermally induced tensile radial stress in a region of the trailing edge of the airfoil. The element has an interior cooling passage and at least one cooling hole extending from the interior cooling passage to an opening located in an area upstream from the stressed region of the trailing edge to cool the area so the airfoil thermally deflects to a shape corresponding to that of the boundary member thereby lowering the thermally induced tensile radial stress in the airfoil at the trailing edge thereof.

Method for reducing cooled turbine element stress and element made thereby

Overall performance in enhanced in gas turbine engines by providing a rotating seal including a rotating member arranged to rotate about an axis and having at least one annular projection extending radially outwardly therefrom, and a stator element having a first surface arranged to contact the projection. The stator element includes at least one slot formed in the first surface, the slot axially traversing the projection so as to allow a flow of purge air to pass. More than one such slot can be used, and each slot is preferably angled circumferentially in the direction of rotation of the rotating member.

Rotating seal

An article comprising an airfoil surface is provided with a thermal insulating outer layer on at least one region, but not all, of the airfoil surface. The region is selected from the airfoil surface which has been observed to experience more strenuous environmental operating conditions during service operation than the airfoil surface outside of the region. In one form, the thermal insulating outer layer is applied to at least one region of each of a plurality of airfoil surfaces. Each airfoil surface is disposed about an axis of rotation with the region of each airfoil surface facing generally outwardly from the axis of rotation and from each other. The airfoil surfaces are rotated about the axis of rotation while a thermal insulating material contacts at least the regions.

Partially coated airfoil and method for making

An article having a design external aerodynamic shape includes a body with a body external surface and an outer portion external surface of an outer portion bonded with a part of but less than substantially all of the body surface. The external surfaces of the body and outer portion each define less than the design external aerodynamic shape. However, together they define such shape. The article is made by selecting the design external aerodynamic shape from an existing article having a first body and a first outer portion. A replacement outer portion is selected in an amount less than the first outer portion and bonded with a region less than substantially all of a replacement body external surface. The replacement body has a replacement body external surface and replacement body external shape different from the first body substantially by that amount. The replacement body is made with the replacement body external shape. Then the replacement outer portion is bonded with the region of the replacement body external surface. The result is that the external surfaces of the replacement body and replacement outer portion together define the design external aerodynamic shape.

Aerodynamic article with partial outer portion and method for making

A turbine blade includes a platform and an airfoil having a plurality of trailing edge slots separated by land areas larger than the slots. The slots have an exit diffusion half angle of about two degrees. Cooling air flows through the slots and over a trailing edge of the airfoil. The platform includes a plurality of openings extending through the platform at an angle. The openings are positioned between a suction side of the blade and a second end of the platform. The openings transport disk post cooling air to a surface of the platform and provide cooling for the platform.

Growing community of inventors

West Chester, OH, United States of America

Gulcharan Singh Brainch