General Electric Company

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

Harold Ray Hansel

Christopher Charles Glynn

Mark Eugene Noe

David Edward Bulman

Vincent Marion Drerup

Krista Anne Mitchell

John David Bibler

Thomas Allen Wells

Ching-Pang Lee

Toby George Darkins, Jr

George E Cook

Craig Patrick Burns

Joseph John Charneski

James Dale Steibel

Steven Douglas Ward

Wayne Garcia Edmondson

James Anthony Ketzer

Charles A Snyder

Christopher Charles Glyn

Jr Toby George Darkins

A method for assembling a gas turbine engine is provided. The method comprises coupling a first turbine nozzle within the engine, coupling a second turbine nozzle circumferentially adjacent the first turbine nozzle such that a gap is defined between the first and second turbine nozzles and providing at least one spline seal including a substantially planar body. The method also comprises forming at least one catch to extend outward from the body portion of the at least one spline seal, and inserting the at least one spline seal into a slot defined in at least one of the first and second turbine nozzles to facilitate reducing leakage through said gap, such that a portion of the at least one spline seal is received within a recess defined within the turbine nozzle slot to facilitate retaining the at least one spline seal within the turbine nozzle slot.

Methods and apparatus for assembling gas turbine engine stator assemblies

A method for assembling a gas turbine engine is provided. The method comprises coupling a first turbine nozzle within the engine, coupling a second turbine nozzle circumferentially adjacent the first turbine nozzle such that a gap is defined between the first and second turbine nozzles and providing at least one spline seal including a substantially planar body. The method also comprises forming at least one retainer tab to extend outward from the body portion of the at least one spline seal, and inserting the at least one spline seal into a slot defined in at least one of the first and second turbine nozzles to facilitate reducing leakage through the gap, such that the at least one retainer tab facilitates retaining the retainer tab within the turbine nozzle slot.

Methods and apparatus for assembling turbine engines

An assembly for mounting an igniter in a gas turbine engine combustor between an outer casing and an outer liner, wherein a longitudinal centerline axis extends through the gas turbine engine. The igniter mounting assembly includes a first spring member encircling a portion of said igniter and being positioned between a surface adjacent the outer casing and an outer surface of the outer liner, a first ring member connected to a first end of the first spring member adjacent the outer surface of the outer liner, and a second ring member connected to a second end of the first spring member adjacent the surface adjacent the outer casing. Accordingly, the igniter is able to maintain substantial alignment with respect to an opening in the outer liner while moving radially and/or axially with respect to the outer liner as the outer casing experiences thermal growth greater than the outer liner.

Mounting assembly for igniter in a gas turbine engine combustor having a ceramic matrix composite liner

A mounting assembly for a forward end of a liner in a combustor of a gas turbine engine including a dome and a cowl, wherein a longitudinal centerline axis extends through the gas turbine engine. The mounting assembly includes a pin member extending through each one of a plurality of circumferentially spaced openings formed in the forward end of the liner, an aft portion of the cowl, and a portion of the dome, with each pin member including a head portion at one end thereof. A nut is adjustably connected to an end of each pin member opposite the head portion. A bushing is located on each pin member at a position intermediate the head portion and the nut, wherein the openings in the liner forward end are sized to fit around the bushings. In this way, the cowl aft portion and the dome portion are fixedly connected together between the bushing and the nut so that the bushings are able to slide radially through the openings in the liner forward end as the cowl and the dome experience thermal growth greater than the liner.

Mounting assembly for the forward end of a ceramic matrix composite liner in a gas turbine engine combustor

A mounting assembly for an aft end of a liner of a gas turbine engine combustor including a support member, wherein a longitudinal centerline axis extends through the gas turbine engine. The mounting assembly includes a pin member extending through each one of a plurality of circumferentially spaced openings in a portion of the support member for the combustor and into a plurality of partial openings formed in the aft end of the liner, with each pin member including a head portion at one end thereof, and a device positioned within each opening in the support member so as to retain the pin members therein. The pin members and the support member are able to slide radially and/or axially with respect to the liner aft end as the support member experiences thermal growth greater than the liner.

Mounting assembly for the aft end of a ceramic matrix composite liner in a gas turbine engine combustor

An assembly for providing a seal at an aft end of a combustor liner for a gas turbine engine including a longitudinal centerline axis extending therethrough. The sealing assembly includes a substantially annular first sealing member positioned between an aft portion of a support member and the liner aft end so as to seat on a designated surface portion of the liner aft end and a substantially annular second sealing member positioned between the support member aft portion and a turbine nozzle located downstream of the liner aft end so as to seat on a designated surface portion of the support member aft portion. Accordingly, the first sealing member is maintained in its seated position as the support member aft portion moves radially with respect to the liner aft end and the second sealing member is maintained in its seated position as the support member aft portion moves axially with respect to the turbine nozzle. The first and second sealing members are also maintained in their respective seating positions as the support member aft portion moves axially with respect to the liner aft end and radially with respect to the turbine nozzle.

Sealing assembly for the aft end of a ceramic matrix composite liner in a gas turbine engine combustor

A support assembly for a gas turbine engine combustor including an inner liner and an inner casing spaced therefrom, wherein a longitudinal centerline axis extends through the gas turbine engine. The support assembly includes an annular inner support cone located adjacent an aft end of said inner liner, an annular nozzle support connected to the inner support cone, and a plurality of support members connected at a first end to a forward end of the inner liner and connected at a second end to the inner support cone.

Support assembly for a gas turbine engine combustor

A combustor having liners made from ceramic matrix composite materials (CMC's) that are capable of withstanding higher temperatures than metallic liners. The ceramic matrix composite liners are used in conjunction with mating components that are manufactured from superalloy materials. To permit the use of a combustor having liners made from CMC materials in conjunction with metallic materials used for the mating forward cowls, and aft seals with attached seal retainer over the broad range of temperatures of a combustor, the combustor is designed to allow for the differential thermal expansion of the differing materials at their interfaces in a manner that does not introduce stresses into the liner as a result of thermal expansion and also balances the flow of cooling air as a result of the thermal expansion.

Conbustor having a ceramic matrix composite liner

A swirler is provided for mixing air from a compressor and fuel from a fuel injector for discharge into a dome of a gas turbine engine combustor. The swirler includes a tubular ferrule for coaxially receiving the fuel injector. A plurality of circumferentially spaced apart swirl vanes are fixedly joined coaxially with the ferrule. An outlet tube is fixedly joined coaxially with the swirl vanes in flow communication therewith for receiving air from the swirlers and fuel from the fuel injector. An annular collar is fixedly joined around the outlet tube and has a convex spherical outer surface. An annular mounting flange for mounting the swirler to the combustor dome has a concave spherical inner surface disposed coaxially around the collar outer surface in a sliding fit therewith to define a ball joint for allowing relative rotation therebetween for self-aligning the fuel injector with the swirler.

Self-aligning swirler with ball joint

A gas turbine engine fuel injection assembly includes a combustor case surrounding a combustor having a dome. An arcuate manifold is disposed inside the combustor case adjacent to the dome, and includes a plurality of fuel injectors disposed in flow communication therewith for receiving fuel for injection into the combustor. An inlet stem is joined in flow communication with the manifold for channeling fuel thereto, and includes an inner fitting disposed in a mounting port in the combustor case. A mounting adaptor is joined to the combustor case at the mounting port, and circumferentially engages the inner fitting to restrain torsional movement thereof. An outer fitting extends in the adaptor and engages the inner fitting in flow communication therewith for supplying fuel thereto.

Internal manifold fuel injection assembly for gas turbine

An apparatus and method for effecting the disassembly of an annular combustor for the maintenance of very large power generation gas generators such as those used in military aircraft engines, commercial aircraft engines, and smaller power generation turbines. The apparatus utilizes rollers which are attachable to a lower half outer casing. An inversion ring located radially inward from the lower half outer casing is supported by the rollers. By disconnecting a removable upper half outer casing, the combustor segments of a first section of the combustor can be removed. Supports connected to the first section of the combustor are secured by pins to a tool ring which is in turn secured to the inversion ring so as to create a closed path. By removing pins which secure the inversion ring and lower half outer casing to supports connected to a second section of combustor segments, the combustor can be rotated so that the second section of formerly inaccessible segments can be easily removed.

Method and system for the disassembly of an annular combustor

An air transfer assembly for a gas turbine engine is provided. The assembly transfers temperature control air from an annular plenum to an annular manifold through an air transfer tube interposed therebetween. The air transfer tube prevents temperature control air leakage even though it is permitted to slide between limits set by an air transfer bushing assembly and a manifold cup formed in the annular manifold. The air transfer bushing is removable and is installed in an aperature in an outer support. The manifold cup is formed in the annular manifold.

Air transfer bushing

The invention comprises a borescope plug for sealing a borescope inspection port on a turbine engine having a fluid flowpath and multiple casings. The borescope plug comprises a plug body having an interlocking drive interface, a plunger portion positioned for sliding coaxial movement relative to the plug body and spring means engaging both the plug body and the plunger. The spring means presses the plunger towards a fully extended position relative to the plug body. The plunger comprises a fixedly attached tip portion shaped to conform with the fluid flowpath. Key means is provided for aligning the tip to conform with the fluid flowpath and the tip is cooled with a cooling airflow.

Borescope plug

The invention comprises a boltless turbine nozzle and nozzle support assembly that includes a turbine nozzle mounting flange seated in a nozzle seat forming a part of the nozzle support. Pins are used for affixing the nozzle mounting flange to the nozzle support seat and baffles are provided for maintaining the pins in their desired locations. A retaining ring is preferrably used to retain the baffles adjacent to the pins. A pressure actuated pliable seal is affixed to the turbine nozzle adjacent to the nozzle seat in order to provide an air seal across the completed assembly.

Growing community of inventors

Mason, OH, United States of America

Harold Ray Hansel