General Electric Company

The United States of America as Represented by the Administrator of the

Dean Thomas Lenahan

Robert J Corsmeier

Larry W Plemmons

William Andrew Bailey

Carl Francis Holdren

Andrew P Elovic

Chi-Ling Michael Tsou

Christopher Charles Glynn

John Howard Starkweather

Iain Robertson Kellock

Robert Joseph Albers

Alan Majors Walker

Gulcharan Singh Brainch

John Christopher Brauer

David A Di Salle

Edward P Brill

Charles H Gay, Jr

Gary Michael Itzel

Jeffrey Donald Clements

Robert Proctor

Thomas G Wakeman

Nesim Abuaf

David Alan Frey

Monty Lee Shelton

Harvey M Maclin

Sean Robert Keith

Daniel Ross Predmore

Steven Alan Ross

Victor H Correia

Curtis Walton Stover

Rollin G Giffin, Iii

Leslie Eugene Leeke, Jr

Kevin Joseph Barb

Thomas Tracy Wallace

Daniel Edward Demers

Jeffrey Glover

Martin C Hemsworth

David M Kercher

Sanjay Chopra

Douglas Arthur Lupe

Richard H Andersen

Peter Andrew Simeone

Frederick W Tegarden

Frederick Martin Miller

Ronald Eugene Mc Rae, Jr

Daniel John Harris

Poul D Pedersen

James R Reigel

Charles F Riedmiller

Jeremy Stephen Wigon

L D Shotts

Albert P Sterman

James P Rauf

John J Bourneuf

Dennis Centeno Iglesias

Albert P Sterman, Deceased

James H Bertke

James Patrick McGovern

Tuley, Eugene N

Barry John Kalb

Bandadi S Shetty

Alan Patrick St Hilaire

Sankar Nellian

Jeremy Stephan Wigon

Nesim Abauf

A baffle seal for a thrust reverser of a gas turbine engine includes a seal body of a resilient material. The seal body has a relatively flat central portion with opposed front and back sides. A raised sealing rim extends at least partially around the central portion. An outer periphery of the sealing rim extends beyond an outer periphery of the central portion so as to define an exposed sealing surface.

Baffle seal for gas turbine engine thrust reverser

A transcowl for a gas turbine engine thrust reverser includes an arcuate outer wall; an arcuate inner wall; and an arcuate baffle positioned between the inner and outer walls at a forward end of the transcowl. The baffle has an arcuate cross-sectional shape which defines a forward-facing interior area. An arcuate forward seal is carried at a forward end of the inner wall. The forward seal includes a plurality of axially-extending, spaced-apart arcuate seal teeth which collectively define a labyrinth seal.

Sealing arrangements for gas turbine engine thrust reverser

A method facilitates assembling a gas turbine engine. The method comprises providing a rotor assembly including a rotor shaft and a rotor disk that includes a radially outer rim, a radially inner hub, and an integral web extending therebetween, wherein the rotor assembly is rotatable about an axis of rotation extending through the rotor shaft, and coupling a disk retainer including at least one discharge tube to the rotor disk wherein the discharge tube extends outwardly from the disk retainer for pumping the air to a higher pressure before discharging cooling fluid therefrom in a direction that is substantially perpendicular with respect to the axis of rotation.

Methods and apparatus for cooling gas turbine engine rotor assemblies

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.

Methods and apparatus for assembling gas turbine nozzles

A gas turbine engine blade includes pressure and suction sides extending between leading and trailing edges and root to tip. The pressure side includes a tip rib recessed therein to define a tip shelf terminating in an inclined ramp. The ramp may be aligned with streamlines of combustion gas flow for preventing interruption thereof and the increase of heat transfer therefrom.

Ramped tip shelf blade

A turbine bucket includes an airfoil extending from a platform, having high and low pressure sides; a wheel mounting portion; a hollow shank portion located radially between the platform and the wheel mounting portion, the platform having an under surface. An impingement cooling plate is located in the hollow shank portion, spaced from the under surface, and the impingement plate is formed with a plurality of impingement cooling holes therein.

Bucket platform cooling scheme and related method

A gas turbine engine rotor assembly includes a plurality of aerodynamic devices to direct airflow radially inward. The gas turbine engine rotor assembly includes a rotor shaft that includes a plurality of openings. The aerodynamic devices include a pair of vane segments and a pair of sidewalls. A contoured outer surface includes an opening and permits the aerodynamic device to be positioned against an inner surface of the rotor shaft, and a flange ring defines a pocket. The aerodynamic device fits within the pocket to concentrically align the openings.

Methods and apparatus for supplying cooling airflow in turbine engines

A blocker and swirl inducer hole configuration for use in connection with a high pressure turbine is described. In one embodiment, the blocker holes are oriented to a 45-degree tangential angle with respect to the direction of rotation of the seal, which results in pre-swirling the air before being injected into the swirl cavity. In addition, the number of blocker holes is reduced by as much as 50% of the number of blocker holes used in the known CFM56 turbine. Further, rather than injecting the air into the first swirl cavity as is known, the air is injected into a second swirl cavity. The combined effect of orienting the holes to the 45-degree tangential angle with respect to the direction of rotation of the seal, locating the holes to open into the second swirl cavity, and reducing the flow area by about 50%, results in an increase in blocker hole pressure ratio. Increasing the blocker hole pressure ratio results in a higher hole exit velocity which maximizes the cavity inlet swirl. The blocker holes therefore not only provide back-pressure, but also function as swirl-inducers. By inducing swirl into the air injected into the second swirl cavity, better turbine disk rim cooling effectiveness is provided. This result facilitates maintaining reasonable metal temperatures at increasingly severe cycle conditions without the normally expected engine performance penalties.

Reduced windage high pressure turbine forward outer seal

A seal joint including a sealing member for sealing segmented components of a gas turbine engine is spaced from the gas path to provide more effective cooling. The segments include radial extension flanges spacing the seal joint from the gas path. Various cooling arrangements are adaptable to the modified segments to provide for open or closed circuit impingement or convection cooling plus film cooling. Excessive thermal gradients are avoided as the entire corner of the segment is evenly cooled. Moreover, air usage at the joint is bounded by seal leakage with no extra dedicated cooling air.

Cooling joint connection for abutting segments in a gas turbine engine

A turbine rotor includes a rotor disk having a plurality of dovetail slots defining therebetween respective disk posts. Each post has a top land and a plurality of tangs spaced radially therebelow inside adjacent ones of the slots. A plurality of turbine blades each having a dovetail are mounted in respective ones of the dovetail slots for radially retaining the blades. A platform is joined integrally with the dovetail and extends circumferentially to adjoin adjacent ones of the platforms to define a radially inner flowpath. An airfoil extends integrally from the platform for extracting energy from combustion gases flowable thereover. A thermal barrier coating is fixedly bonded to each of the disk post top lands for providing thermal insulation between the platforms and the disk posts for reducing disk post temperature.

Shielded turbine rotor

In a gas turbine including a compressor having a bore and a rotor comprised of multiple stages extending between a first stage at a forward end of the compressor and a last stage at an aft end of the compressor, each stage including a rotor disk having a peripheral rim and multiple blades secured to the peripheral rim, a combustion system comprising a plurality of combustors utilizing discharge air from the compressor for combustion, and multiple turbine stages driven by combustion gases from the combustion system, the improvement comprising means for supplying cooling air at least to a peripheral rim of the last stage of the compressor.

Compressor rotor cooling system for a gas turbine

A turbine cooling supply circuit for a gas turbine engine is disclosed in which the flow of coolant through the engine is directed to minimize temperature rise prior to discharge into the turbine. Bleed air from an impeller stage of a compressor is channeled along a backwall thereof into a cavity disposed radially inwardly of a combustor casing. The cavity is divided by a lightweight tubular member having a circumferentially uniform contour into a non-flow zone proximate the casing and a flow zone proximate a rotor. Maintenance of a high tangential flow component in the bleed flow in the flow zone facilitates discharge of the flow through apertures in a rotor into a bore portion thereof, minimizing work input and temperature rise of the coolant. Further, the coolant is isolated from the hot casing wall and nonuniform structures disposed thereon, thereby reducing heat flux and temperature rise associated with aerodynamic drag caused by increased surface area, nonuniform interruptions in the stationary surface or any combination thereof. Features are incorporated in the tubular member to provide margin against buckling.

Gas turbine engine cooling supply circuit

Improved operation can be achieved from an enhanced gas turbine engine having a segmented annular shroud which radially expands and contracts to match the expansion and contraction of engine rotor blades. The segmented annular shroud is supported by a structure which includes an annular ring having two radially outwardly extending flanges, and forward and aft annular segmented brackets which attach the segmented shroud to the forward and aft side of the ring respectively. In a preferred embodiment of the invention, two circumferentially extending separate and distinct air impingement manifolds surrounds each of the outwardly extending flanges. Each manifold is provided with a valve for controlling the amount and temperature of the airflow entering each manifold. The air from each manifold then impinges upon each of the outwardly extending flanges, thereby controlling the radial movement of the corresponding shroud segments and the associated clearances with the rotor blade tips. The use of separate and distinct manifolds and the corresponding values which regulate the amount and temperature of airflow to each manifold allows individual shroud portions to be separately expanded or contracted.

Clearance control system for separately expanding or contracting

The invention comprises the use of a thermal barrier coating on a member of a gas turbine engine component with one surface of the member exposed to high temperature but not exposed to high velocity gases, as an insulating material to slow the heat transfer to the substrate of the component member and allow thermal gradients within acceptable material limits between the thermal barrier coated member and the remainder of the component. In a particular application, the use of a thermal barrier coating enables making a high pressure turbine nozzle integral with a discourager seal lip, eliminating fasteners protruding into the rotor/stator cavity and split lines between segments of the prior art that allowed high temperature gas ingestion further into the cavity and caused windage losses.

Integral turbine nozzle support and discourager seal

Growing community of inventors

Cincinnati, OH, United States of America

Dean Thomas Lenahan