General Electric Company

Richard William Jendrix

Steven Robert Brassfield

Roger Lee Doughty

Michael Alan Hile

James Hamilton Grooms, Iii

Cory Michael Williams

Ching-Pang Lee

Bangalore Aswatha Nagaraj

Christopher Charles Glynn

Scott Alan Schimmels

Robert Proctor

David John Wortman

Larry W Plemmons

Mark Willard Marusko

Travis J Packer

Mark Edward Stegemiller

John David Bibler

Julius John Montgomery

Glen William Royal

Thomas Ryan Wallace

Nathan Gerard Cormier

William B Connor

Bradford Alan Tracey

Charles Thomas McMillan

Brian Jared Agnew

Samuel Joseph Stoughton

Steve Mark Molter

Renee Solomon Wenstrup

Christopher Thomas Schuppe

David Fasig

Christopher John Van Derven

James Hamilton Grooms

Cory Williams

A clearance control assembly for a gas turbine engine that defines an axial direction and a radial direction and includes a stage of rotor blades and a shroud hanger. The assembly includes a case configured to be positioned outward along the radial direction from the stage of rotor blades when installed in the gas turbine engine. The case is further configured to be engaged with the shroud hanger at a first location when installed in the gas turbine engine. The assembly also includes a baffle positioned outward along the radial direction from the case to define a chamber therebetween. The baffle has a forward end and an aft end. The forward end of the baffle is engaged with the case to form a first seal and the aft end of the baffle is engaged with the case to form a second seal. The baffle, the case, or both define an inlet to allow a fluid to enter the chamber and the case defines an outlet to allow the fluid to exit the chamber.

Clearance control assembly

Gas turbine engine turbine blade assembly includes a hollow airfoil joined to blade root, dovetail slot heat shield bonded or attached to a bottom surface of the root, and a shield outlet from heat shield open to inlet apertures extending radially through a radially inner root end of the root. Heat shield may have body with legs extending upwardly from heat shield bottom, slanted open upstream end, and free ends of the legs longer than the heat shield bottom. Flanges may be located along free ends and bonded to bottom surface. Body, heat shield bottom and/or the legs may be rounded. Disk includes a plurality of dovetail slots formed in a rim, complimentary plurality of turbine blades removably retained in dovetail slots by the roots, slot bottoms of the dovetail slots extending circumferentially between disk posts in rim. Heat shield bottoms may be radially spaced apart the slot bottoms.

Turbine dovetail slot heat shield

A turbine disk and blade assembly comprises a turbine disk including a plurality of disk posts extending radially therefrom and a plurality of disk slots between adjacent disk posts. The turbine disk is of a first material having a first coefficient of thermal expansion. The assembly further comprises a plurality of turbine blades. One of the turbine blades is received in each disk slot. The turbine blades are of a second material having a second coefficient of thermal expansion. The assembly further comprises a plurality of seal plates, wherein one of the seal plates is positioned in each disk slot radially inward of the turbine blade. The seal plates are of a material having a coefficient of thermal expansion substantially similar to that of either the disk posts or the turbine blade.

Device, system and method for preventing leakage in a turbine

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

An airfoil is disclosed having at least first and second cast, axially-stacked internal airflow cooling circuits. Each circuit defines multiple air flow passages positioned laterally between a pressure sidewall side and a suction sidewall side of respective ones of the circuits. Each of the circuits is formed by a separate casting core. Methods of forming a axially-stacked core and an airfoil are also disclosed.

Airfoil cooling circuits and method

A method of manufacturing a gas turbine engine includes providing a turbine mid-frame, coupling a plurality of rotor blades to a rotor disk, the rotor disk is coupled axially aft from the turbine mid-frame such that a cavity is defined between the rotor disk and the turbine mid-frame, and forming at least one opening extending through the turbine mid-frame to facilitate channeling cooling air into the gap, the opening configured to impart a high relative tangential velocity into the cooling air discharged from the opening.

Methods and apparatuses for cooling gas turbine engine rotor assemblies

A turbine rotor for a gas turbine engine includes a rotor disk having a plurality of circumferentially alternating dovetail slots and disk posts and a plurality of blades disposed in the dovetail slots. A cooling air plenum is defined in the portion of each dovetail slot located radially inward of the corresponding blade. Air from the plenums is directed through slots formed in axial surfaces of the blades or disk posts into an adjacent wheel cavity, thereby purging the wheel cavity.

Method and apparatus for purging turbine wheel cavities

A gas turbine engine hollow airfoil includes an airfoil outer wall having width wise spaced aparted pressure and suction side walls joined together at chordally spaced apart leading and trailing edges of the airfoil and extending longitudinally from a root to a tip. At least one internal aft flowing serpentine cooling circuit inside the airfoil has an exit that is positioned aft of the entrance so as to have a chordal flow direction afterwards from the leading edge to the trailing edge within the serpentine circuit. At least one longitudinally extending first side wall impingement chamber is in downstream fluid communication with the serpentine cooling circuit within the airfoil and positioned between one of the side walls, preferably the pressure side wall, and a first inner wall bounding the serpentine cooling circuit. Impingement cooling apertures are disposed in the first inner wall between one of the serpentine channels and the side wall impingement chamber. A first plurality of side wall film cooling holes may extend out from the first side wall impingement chamber through the pressure side wall. At least one tip cooling hole extending out of at least one of the impingement chambers may be disposed through a longitudinally outer tip wall of the tip of the airfoil.

AFT flowing serpentine airfoil cooling circuit with side wall

A coated article comprises a base article having a substrate made of a material selected from the group consisting of a nickel-base alloy and a cobalt-base alloy, an intermediate metallic coating structure, and a thermal barrier coating. The intermediate metallic coating structure may be a protective layer of platinum, rhodium, or palladium on the base article, with no bond coat. The intermediate metallic coating may also be a bond coat, either oxidized or unoxidized, with a protective layer of platinum, rhodium, or palladium thereon. The protective layer typically partially interdiffuses with the underlying metallic structure.

Platinum, rhodium, or palladium protective coatings in thermal barrier

Growing community of inventors

Liberty Township, OH, United States of America

Richard William Jendrix