General Electric Company

Ely E Halila

Michael J Anderson

John David Bibler

Paul V Heberling

James A Martus

Christopher Charles Glynn

Robert Joseph Orlando

Jack Rogers Taylor

Douglas Melton Carper

Monty Lee Shelton

Thomas Tracy Wallace

Willard J Dodds

John A Kastl

Jeffrey Lynn Myers

Howard L Foltz

Valentine R Boehm, Jr

Thomas J Dunyak

Joseph W Savage

David Byron Morris

Earl D Deaton

Stuart C Greenfield

Robert W Aho

A combustor for a gas turbine engine including a pilot fuel injection assembly that includes a replaceable heat shield is described. The pilot fuel injection assembly also includes a fuel stem and a mixer divider. The fuel stem includes a first end, a second end, and a body extending therebetween. An alignment slot is defined within a portion of the interface between the fuel stem body and the second end. The mixer divider is attached to the fuel stem second end and the heat shield is removably coupled to the mixer divider with a plurality of transfer tubes.

Method of fabricating gas turbine fuel injection

Concentric installation of a pilot fuel assembly in an opening in a gas turbine combustor casing is achieved by providing a boss having at least two flat surfaces which are perpendicular to each other on the combustor casing surrounding the opening and a mounting flange having at least two flat surfaces which are perpendicular to each other on the pilot fuel assembly. The pilot fuel assembly is concentrically installed to the combustor casing by inserting the assembly into the combustor casing opening, and moving the pilot fuel assembly as far as it will go in a first direction substantially parallel to one of the flat boss surfaces. The distance between the other flat boss surface and one of the flat flange surfaces is then taken. Next, the pilot fuel assembly is moved in the direction opposite the first direction, at which point, the distance between the same two flat surfaces is again measured. Lastly, the pilot fuel assembly is located at a position where the distance between the two measuring surfaces is equal to the average of the first and second measurements. If desired, these steps can be repeated back and forth along an axis perpendicular to the first and second directions.

Method for providing concentricity of pilot fuel assembly in a combustor

Apparatus for injecting fuel into gas turbine engines

A combustor for a gas turbine engine having a longitudinal axis therethrough, including an outer liner having a forward end and an aft end, an inner liner having a forward end and an aft end, a first dome formed upstream of the outer liner forward end so as to define a first combustion zone radially oriented to the longitudinal axis, and a dome plate having an outer portion connected to an upstream end of the first dome and an inner portion connected to the inner liner forward end, wherein a second combustion zone is defined by the dome plate, the outer liner, and the inner liner substantially perpendicular to the first combustion zone.

Multi-stage radial axial gas turbine engine combustor

A fuel air mixer for a gas turbine engine combustor having a longitudinal axis therethrough, wherein the fuel air mixer is configured for use in a dome oriented substantially radial to the longitudinal axis. The fuel air mixer includes a fuel injection assembly having a first end, a second end, a fuel passage extending therethrough, and a flange portion having a plurality of spaced openings formed therein which extends from the first end. The fuel air mixer also includes a mixer assembly having a first end, a second end, a cavity formed in a central portion thereof, and a flange portion having a plurality of spaced openings formed therein which extends from the first end. The mixer assembly is configured to receive the fuel injection assembly in the cavity so that the fuel injection assembly and the mixer assembly are able to be connected to an outer casing of the combustor by means of the respective flange portions.

Fuel air mixer for a radial dome in a gas turbine engine combustor

A combustor dome module includes a mixer tube having a hollow heat shield sealingly joined around the outlet end thereof. The modules may then be assembled in an array for defining the combustor dome, with each module being individually removable therefrom.

Modular combustor dome

A variable area turbine nozzle includes a plurality of circumferentially adjoining nozzle segments. Each nozzle segment includes outer and inner bands, with a plurality of first vane segments fixedly joined therebetween. A plurality of second vane segments adjoin respective ones of the first vane segments to define therewith corresponding vanes which are spaced apart to define respective throats of minimum flow area for channeling therethrough combustion gas. The second vane segments are pivotable to selectively vary the throat area.

Variable area turbine nozzle

A gripping system for use in uniaxial testing of test specimens at elevated temperatures, typically above 900.degree.-1000.degree. F. The uniaxial mechanical testing is typically tensile testing conducted on high temperature engineering materials which operate under severe conditions at elevated temperatures. The gripping system includes a gripping element having a first section adapted to contact the test specimen wherein a heater is received in the second section for heating the gripping surface. The heater is designed to heat the gripping surface and the test specimen to the desired temperature.

Hot grip assembly

A film starter structure for a combustor of a gas turbine engine which includes a plurality of circumferentially spaced, axially extending ribs formed on a radially inner surface of a forward section of an outer combustor liner adjacent a combustor dome. An annular ring overlays the ribs for defining a plurality of air passages. A support extends from the combustor dome and supports the outer liner about the dome. Compressor discharge air is introduced into the air passages and exits the air passages along the inner surface of the outer liner for establishing a cooling film barrier on the outer combustor liner surface. A spring seal between the combustor dome and the inner ring seats the dome within the ring and establishes a seal for preventing leakage air therebetween and allowing independent radial expansion of the liner and dome by compressing the spring seal. The liner and dome structure are further arranged to allow assembly using flanges and split rings so as to eliminate placement of bolted connections in critical air flow paths.

Film cooling starter geometry for combustor liners

A gas turbine engine combustor cooling system for imperforate non-metallic combustor liners has a wall positioned adjacent to the liners forming therewith a cavity. The wall has a plurality of inlets for admitting cooling air into the cavity and a plurality of outlets for exhausting the cooling air into a separate passageway after it impinges the liners. The exhausted cooling air is transferred upstream of the liner where it is combined with fuel for burning rather than discharged downstream as cooling film.

Combustor liner cooling system

A combustor liner segment includes a panel having four sidewalls forming a rectangular outer perimeter. A plurality of integral supporting lugs are disposed substantially perpendicularly to the panel and extend from respective ones of the four sidewalls. A plurality of integral bosses are disposed substantially perpendicularly to the panel and extend from respective ones of the four sidewalls, with the bosses being shorter than the lugs. In one embodiment, the lugs extend through supporting holes in an annular frame for mounting the liner segments thereto, with the bosses abutting the frame for maintaining a predetermined spacing therefrom.

Segmented combustor

Film cooling starter geometry for combustor lines

A seal assembly includes first and second segments spaced from a frame and defining between each other an elongate lateral gap. A plurality of radially adjoining leaf seals bridge the lateral gap between the segments and are supported by a plurality of spaced apart pins, with each pin having an enlarged head on one side of the frame and a root on an opposite side of the frame fixedly joined to a bottom one of the leaf seals. A compression spring is disposed around each of the pins between the frame and a top one of the leaf seals for urging the top leaf seal toward the bottom leaf seal to seal flow radially therebetween and between adjacent support pins, and for urging the bottom leaf seal in sealing contact with the first and second segments to seal flow radially therebetween through the lateral gap.

Seal assembly

A support assembly for a gas turbine engine combustor includes an annular frame having a plurality of circumferentially spaced apart tenons, and an annular combustor liner disposed coaxially with the frame and including a plurality of circumferentially spaced apart tenons circumferentially adjoining respective ones of the frame tenons for radially and tangentially supporting the liner to the frame while allowing unrestrained differential thermal radial movement therebetween.

Combustor liner support assembly

A mounting assembly includes an annular supporting flange disposed coaxially about a centerline axis which has a plurality of circumferentially spaced apart supporting holes therethrough. An annular liner is disposed coaxially with the supporting flange and includes a plurality of circumferentially spaced apart mounting holes aligned with respective ones of the supporting holes. Each of a plurality of mounting pins includes a proximal end fixedly joined to the supporting flange through a respective one of the supporting holes, and a distal end disposed through a respective one of the liner mounting holes for supporting the liner to the supporting flange while unrestrained differential thermal movement of the liner relative to the supporting flange.

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Cincinnati, OH, United States of America

Ely E Halila