General Electric Company

James Edwin Rhoda

Mark Joseph Mielke

Seetharamaiah Mannava

David Edward Bulman

William D Cowie

Herbert Halila

Albert E McDaniel

Nathan Gerard Cormier

Andrew Breeze-Stringfellow

Stephen Joseph Ferrigno

Mark Willard Marusko

Andrew J Lammas

James Michael Forrester

Joseph T Stevenson

Wayne Ray Bowen

Mitchell Jay Headley

Michael Jason Bishop

Steven Mark Ballman

James Edward Gutknecht

David Scott Clark

Craig Patrick Burns

Gary Robert Peters

Paul Michael Smith

Michael D Carroll

Steven Roy Manwaring

Janine Elizabeth Bolt

Lawrence J Egan

Richard P Zylka

Daniel G Suffoletta

Thomas A Bradley

Daniel Gerard Suffoletta

Richard Patrick Zylka

A variable stator vane airfoil is mounted on a button centered about a rotational axis and includes circular leading and trailing edges circumscribed about the rotational axis at a button radius. Contoured pressure and suction sides extend from the circular leading edge to the circular trailing edge and are recessed inwardly from a perimeter circumscribed about the rotational axis at the button radius. One of upstream and downstream pressure side portions of the contoured pressure side is straight and another of the upstream and downstream pressure side portions is convexly curved. One of the upstream and downstream suction side portions is straight and another of the upstream and downstream suction side portions is convexly curved. One of the upstream pressure side portion and upstream suction side portion is straight and another of the upstream pressure side portion and upstream suction side portion is convexly curved.

Variable stator vane contoured button

A method facilitates assembling a variable vane assembly for a gas turbine engine including a casing and an inner shroud. The method comprises providing at least one variable vane including a radially inner spindle that includes a groove defined therein that has at least one machined face, and coupling the variable vane radially between the casing and the inner shroud such that at least a portion of the radially inner spindle is inserted at least partially through an opening extending radially through the inner shroud. The method also comprises securing the variable vane to the inner shroud by engaging the spindle machined face with a retainer coupled to the inner shroud.

Gas turbine engine variable vane assembly

A method of assembling a gas turbine engine includes providing a plurality of stator vane sectors that each include an equal number of stator vanes that are circumferentially-spaced such that a first circumferential spacing is defined between each pair of adjacent stator vanes within the sector, and coupling the plurality of stator vane sectors together to form a stator vane assembly such that a second circumferential spacing is defined between each pair of adjacent stator vanes coupled to adjacent sectors, wherein the second circumferential spacing is different from the first circumferential spacing.

Methods and apparatus for assembling gas turbine engines

A gas turbine engine rotor disk has a number of annular hubs circumscribed about a centerline and each of the hubs is connected to a disk rim by a web. A plurality of circumferentially spaced apart dovetail slots are disposed through the rim, extending circumferentially between disk posts, extending axially from a forward end to an aft end of the rim, and extending radially inwardly from a disk outer surface of the rim. Circumferentially extending annular burst slots extend radially through the rim into the dovetail slots between each adjacent pair of the webs. In the exemplary embodiment of the disk illustrated herein, the dovetail slots are circular arc dovetail slots. An axially forwardly extending overhang is located radially outwardly on each of the disk posts and a chamfer is on a radially outer corner of the each of the disk posts on a portion of the overhangs. A plurality of post holes extending axially aftwardly into the rim wherein each of the post hole extends axially aftwardly into a corresponding one of the disk posts.

Circular arc multi-bore fan disk

A gas turbine engine rotor assembly including a rotor having a radially outer rim with an outer surface shaped to reduce circumferential rim stress concentration between each blade and the rim. Additionally, the shape of the outer surface directs air flow away from an interface between a blade and the rim to reduce aerodynamic performance losses between the rim and blades. In an exemplary embodiment, the outer surface of the rim has a concave shape between adjacent blades with apexes located at interfaces between the blades and the rim.

Reduced-stress compressor blisk flowpath

A compressor casing is configured to surround blade tips in a compressor stage. The casing includes stall grooves with adjoining lands defining respective local gaps with the blade tips. At least one of the lands is offset to locally increase a corresponding one of the gaps larger than the nominal gap for the casing to reduce tip rubbing thereat.

Rub resistant compressor stage

A bushing assembly as a means to enable guided movement of a shaft in an opening of a housing comprises a stationary bushing support carried by a housing internal wall. The bushing support has a support internal wall defining a bushing support opening through the bushing support to receive a movable shaft. Carried by the shaft is a removable wear sleeve having a wear sleeve outer wall in juxtaposition with and spaced apart from the support internal wall to permit relative axial movement of the wear sleeve with the shaft within the bushing support opening. Disposed between and for contact with the support internal wall and with the wear sleeve outer wall is an anti-friction layer.

Bushing assembly with removable wear sleeve

Gas turbine engine blade with a metallic airfoil having a radially outer tip extending chordwise between a leading edge and a trailing edge and at least one laser shock peened surface extending radially along at least a portion of the tip and a region having deep compressive residual stresses imparted by laser shock peening (LSP) extending into the airfoil from the laser shock peened surface.

Laser shock peened gas turbine engine blade tip

Gas turbine engine fan blade with a metallic airfoil having a leading edge, a trailing edge, and a radially outer tip, and at least one laser shock peened surface located along the leading edge with a region having deep compressive residual stresses imparted by laser shock peening (LSP) extending into the airfoil from the laser shock peened surface.

Laser shock peened gas turbine engine fan blade edges

A blade support includes first and second abutting disks, with the first disk having a plurality of axial dovetail slots at one end thereof, and a circumferential dovetail first half-slot at an opposite end thereof, with the second disk having a circumferential dovetail second half-slot therein. A plurality of rotor blades have axial dovetails disposed in respective ones of the axial dovetail slots, and circumferential dovetails disposed in part in each of the first and second half-slots. The first and second disks are retained together so that the first and second half-slots collectively support the circumferential dovetails.

Split disk blade support

A bladed rotor having blades with blade shrouds includes a deformable protective coating applied to the blades to increase the damage tolerance of the blades during foreign object ingestion. The protective coating deforms upon impact between the shroud and airfoil of adjacent blades. The deformation absorbs impact energy and improves distribution of the impact load transferred between the blades.

Blade shroud deformable protective coating

Growing community of inventors

Mason, OH, United States of America

James Edwin Rhoda