General Electric Company

United States of America As Represented by the Secretary of the Air Force

Michael Philip Hagle

James Neil Cooper

Junhaur Jih

Paul R Bien

Gilbert Farmer

Wayne Ray Grady

Brian Alan Norton

Paul Stuart Wilson

Robert B Foster

Jeffrey J Reverman

James Walter Caddell

Timothy Lee Siebert

Gregory Matthew Ford

Reed Roy Oliver

John Philip McKay

Max Joseph Freidauer

Dana Leigh Ferguson

Jeffrey John Reverman

A method is provided for making a repaired turbine engine stationary vane assembly including at least one airfoil bonded between spaced apart first and second platforms. The method comprises providing a first vane assembly member including at least a portion of the first platform, the airfoil, and a first segment of the second platform bonded with the airfoil. Also provided is a second vane assembly member including a second segment of the second platform with a recess of a shape and size sufficient to receive the first segment of the second platform. The first segment is placed in the recess after which the first and second segments are bonded, for example by brazing or welding. Provided is a more structurally sound repaired turbine engine stationary vane assembly including the first and second vane assembly members bonded together.

Method for making a repaired turbine engine stationary vane assembly and repaired assembly

A method is provided for repairing a turbine nozzle segment having at least one vane disposed between outer and inner bands. The method includes separating the inner band from the nozzle segment, and joining the inner band to a newly manufactured replacement casting having an outer band and a vane. The replacement casting includes an airfoil stub. The airfoil stub is received in a recess formed in the inner band. Joining is completed by joining the airfoil stub to the inner band. The replacement casting may be modified from a newly manufactured singlet casting.

Turbine nozzle segment and method of repairing same

An air-cooled combustor liner segment, including a plurality of spaced apart radially outwardly extending integral heat exchange protuberances or pins is repaired by providing a replacement edge member for a service damaged edge portion including such pins. The damaged edge portion is removed to provide a segment body with a body repair surface, of a first shape, from which the damaged edge portion was removed. A replacement edge member is provided including a plurality of spaced apart radially outwardly extending integral pins to replace pins in the removed edge portion. The replacement edge member includes a replacement edge member repair surface of a second shape matched with the first shape of the body repair surface. The body repair surface and the replacement edge member repair surface are disposed in juxtaposition at a matched junction therebetween so that their respective radially outer surfaces substantially are coextensive. Then the replacement edge member and the segment body are joined by bonding at the matched junction substantially without change in the cooling air flow control about the protuberances.

Method for repairing an air cooled combustor liner segment edge portion and repaired segment

An assembly of parts is prepared by providing a female part having a recess and a male part having an inserted surface received within the recess. A relief pattern such as knurling is formed on the inserted surface. The inserted surface of the male part is inserted into the recess of the female part with a contacting-channeled fit between the inserted surface and the recess surface, so that the knurl serves to align the inserted surface within the recess. The inserted surface of the male part is brazed or otherwise joined to the recess of the female part.

Self-fixtured joint assembly and its preparation

A combustor dome assembly includes a spectacle plate having an opening formed therein that defines an inner circumferential edge, at least one swirler assembly defining an inner cylindrical surface, and at least one deflector having an outer cylindrical surface. These components are arranged such that the outer cylindrical surface is joined to the inner circumferential edge and to the inner cylindrical surface. One preferred method of manufacturing the combustor dome assembly includes placing a first ring of joining material over the outer cylindrical surface. Then, the outer cylindrical surface is inserted into the spectacle plate opening so that the first ring is sandwiched between a first surface of the spectacle plate and a shoulder formed on the deflector. A second ring of joining material is next placed over the outer cylindrical surface, and a third ring of joining material is inserted into an annular groove formed in the inner cylindrical surface. The next step is to place the swirler assembly over the outer cylindrical surface so that the second ring is sandwiched between a second surface of the spectacle plate and an end of the swirler assembly and the third ring surrounds the outer cylindrical surface. The first, second and third rings are then heated to a temperature greater than their melting points so that they become molten joining material. The molten joining material is cooled to complete the joint.

Combustor dome assembly and method of assembling the same

A combustor liner has a stepped combustor liner surface and an overhang portion forming an air cooling slot. A contoured rear facing edge of the overhang portion reduces turbulence of combustion gas flow and reduces a combustor liner surface area exposed to combustion gases. A thermal barrier coating is also applied to the contoured rear facing edge, reducing heat flow into the overhang portion and hence reducing the operating temperature of the combustor liner. Thus, the amount of cooling air is reduced, which can reduce exhaust emissions increase engine performance and extend a working life of the combustor assembly.

Combustor rear facing step hot side contour method

A combustor liner has a stepped combustor liner surface and an overhang portion forming an air cooling slot. A contoured rear facing edge of the overhang portion reduces turbulence of combustion gas flow and reduces a combustor liner surface area exposed to combustion gases. A thermal barrier coating is also applied to the contoured rear facing edge, reducing heat flow into the overhang portion and hence reducing the operating temperature of the combustor liner. Thus, the amount of cooling air is reduced, which can reduce exhaust emissions, increase engine performance and extend a working life of the combustor assembly.

Combustor rear facing step hot side contour method and apparatus

A method and apparatus for aiding in positioning during assembly of a locomotive wheel with respect to a drive gear mounted adjacent the wheel during assembly and for maintaining the position of the drive gear during locomotive operation. The apparatus includes a lateral stop including a spacer formed of a semi-elastomeric material having a low coefficient of friction and specially adapted to fit into a stress relief groove of an axle between a hub of the drive gear and a hub of the wheel. The spacer is so designed as to be installable on the locomotive axle after installation of the drive gear, to be self-centering on the axle and to inhibit rotation of the spacer with respect to the axle such that the axle will not be marked circumferentially by spacer rotation.

Lateral support for a drive gear and wheel on a locomotive axle

A stator frame comprises a connection end (CE) casting having at least one CE integral partial tie bar extending from an inboard side of the CE casting and a pinion end (PE) casting having at least one PE integral partial tie bar extending from an inboard side of the PE casting. Each PE partial tie bar is welded to a respective CE partial tie bar.

Stator including partial tie bars extending from end castings and welded

An apparatus for modulating a flow of cooling air from a compressor to a turbine in a gas turbine engine is disclosed. The apparatus includes a support member which forms a channel for receiving the cooling flow and has a throat at a downstream end with a radial orientation. A radial flow inducer is connected to the support member adjacent to the channel downstream end for accelerating the cooling flow through the throat of the support member. Apparatus for modulating the quantity of the cooling air flowing through the throat and into the turbine is provided in the form of an axially translating endwall, wherein variation of the throat area is accomplished through modulation of the endwall with respect to the support member. The support member may also house either a stationary, unmodulated axial flow inducer or a second modulated radial flow inducer in addition to the modulated radial flow inducer.

Turbine cooling flow modulation apparatus

Stator frame fabrication

An insulative panel is attachable to a structure to thermally protect the structure. The insulative panel includes a piece of a compliant, porous material. A support base is bonded to a first side of the piece of porous material, the support base including means for attaching the support base to the structure that is to be protected. A skin made of a thermally resistant material is bonded to a second side of the piece of porous material. Preferably, the porous material is a piece of porous ceramic foam having an apparent density of less than about 25 percent, the support base is a piece of nonporous ceramic material, and the skin is a ceramic thermal barrier coating that is plasma sprayed onto the porous material.

Thermal barrier structure

Thermally actuated flow diverters for gas turbine engines are provided which include an arrangement of vanes in spaced overlapping array, which direct, e.g. coolant air flow to rotating engine components, each vane having 3 foil components joined at 3 hinge points. That is, a leading foil positioned on the suction side of the vane, a trailing foil pivotably mounted to a rearward portion of such leading foil at a first pivot axis and a third foil positioned on the pressure side of the vane connecting a forward portion of such leading foil and such trailing foil at second and third pivot axes respectively. Either the third foil or the leading foil is fixedly mounted to engine walls of the same material, the other of these two foils defining an actuation link that is expandable and contractable relative to such walls; the third foil having a coefficient of thermal expansion greater than that of the leading foil. The vanes are positioned so that upon temperature change, the third foil contracts or expands relative to the leading foil and pivots the trailing foil toward or away from the leading foil of a downstream vane so as to control the flow of engine coolant therebetween.

Thermally actuated vane flow control

A relatively stationary segmented turbine engine shroud, nozzle or frame assembly is sealed against the passage of cooling air into the exhaust gas flowpath by a plurality of spline seals which cooperate with an annular stationary brush seal. The spline seals provide sealing between axially extending confronting edges of adjacent segments while the brush seal provides continuous circumferential sealing around the segments and spline seals.

Growing community of inventors

Mason, OH, United States of America

Michael Philip Hagle