General Electric Company

Exxonmobil Upstream Research Company

William Lawrence Byrne

Patrick Benedict Melton

David William Cihlar

Carolyn Ashley Antoniono

Elizabeth Angelyn Fadde

Lucas John Stoia

Wei A Chen

Donald Mark Bailey

James Anthony West

John Charles Intile

James Harper

Christopher Paul Willis

Donald Timothy Lemon

Mert Enis Berkman

Xiaoguang Yu

Devin Patrick Perkins

Patrick Bendict Melton

Willis Christopher Paul Paul

A system includes a turbine combustor that includes a head end portion having a head end chamber, a combustion portion having a combustion chamber disposed downstream from the head end chamber, a cap disposed between the head end chamber and the combustion chamber, and a flow distributor configured to distribute an exhaust flow circumferentially around the head end chamber. The flow distributor includes at least one exhaust gas flow path.

System and method for a turbine combustor

A system includes a turbine combustor that includes a head end portion having a head end chamber, a combustion portion having a combustion chamber disposed downstream from the head end chamber, a cap disposed between the head end chamber and the combustion chamber, and a flow distributor configured to distribute at least one of an exhaust flow, an oxidant flow, an oxidant-exhaust mixture, or any combination thereof circumferentially around the head end chamber.

A system includes a turbine combustor that includes a head end portion having a head end chamber, a combustion portion having a combustion chamber disposed downstream from the head end chamber, a cap disposed between the head end chamber and the combustion chamber, and a flow separator configured to separate a first exhaust flow from an oxidant flow. The flow separator is configured to direct the first exhaust flow into the head end chamber. The turbine combustor also includes a mixing region configured to mix the first exhaust flow with the oxidant flow to provide an oxidant-exhaust mixture.

The present disclosure relates to an apparatus having an elongated body configured to be inserted into a tubular structure extending between a first combustor and a second combustor of a gas turbine engine. A movable arm may be positioned proximate to a first end of the elongated body, and the movable arm may be configured to engage a surface of the first combustor when the elongated body is placed within the tubular structure. A drive member may be accessible near a second end of the elongated body, and the drive member may be configured to rotate within the elongated body and to drive the movable arm along a longitudinal axis of the elongated body toward the second end to separate the tubular structure from the second combustor.

Apparatus and method for servicing gas turbine engines

A transition piece aft frame assembly is provided, and includes a transition piece aft frame and a heat shield. The transition piece aft frame has an aft face. At least a portion of the aft face is exposed to an exhaust gas stream. The heat shield is connected to the transition piece aft frame. The heat shield is oriented to generally deflect the exhaust gas stream away from the aft face of the transition piece aft frame.

Transition piece aft frame assembly having a heat shield

A late lean injection sleeve assembly allows the injection of fuel at the aft end of a gas turbine liner, before the transition piece, into the combustion gases downstream of a turbine combustor's fuel nozzles. The late lean injection enables fuel injection downstream of the fuel nozzles to create a secondary/tertiary (with quaternary injection upstream of the fuel nozzles) combustion zone while reducing/eliminating the risk of fuel leaking into the combustor discharge case. The fuel is delivered by the flow sleeve into one or more nozzles that mix the fuel with CDC air before injecting it into the combustor's liner.

Integrated late lean injection on a combustion liner and late lean injection sleeve assembly

A combustor nozzle includes a single-piece swirler. The single-piece swirler includes a center body extending axially along the single-piece swirler, a first fuel passage inside the center body, and a plurality of vanes extending radially from the center body. A method for fabricating a combustor nozzle includes casting a single-piece swirler having a center body and a plurality of vanes extending radially from the center body.

Combustor nozzle and method for fabricating the combustor nozzle

A method and apparatus for a diffusion tip for use with a fuel nozzle is described. The diffusion tip has a substantially circular body including an outer surface and an opposite inner surface. The diffusion tip body extends from a discharge end to an inlet end. The diffusion tip includes an inlet surface adjacent to the discharge end and defined within the body. A discharge surface is defined opposite the inlet surface. A plurality of diffusion apertures each extend between the discharge surface and the inlet surface, each aperture is oriented relative to the body to discharge a diffusion flow outward therefrom at an angle &#x3b3; (gamma) measured in an X-Z plane between a centerline of the aperture and an X-axis extending tangentially to the outer surface, and at an angle &#x3b8; (theta) measured in a Y-Z plane between the centerline of the aperture and a Y-axis extending radially outward from the centerline.

Method and apparatus to facilitate cooling of a diffusion tip within a gas turbine engine

Cooling is provided for a gas turbine combustion system that includes a first member, such as a liner, having exterior and interior surfaces that define a wall therebetween and a central area. A second member, such as a cap, is located adjacent that interior surface. A source of cooling gas is in fluid communication with the exterior surface of the first member. An open space, in which is located a hula seal, is located between the interior surface of the first member and the second member. At least one opening in the wall of the first member provides a passageway for the cooling gas from the exterior surface of the first member to the open space. The passageway has a directional axis along which the cooling gas flows and is discharged into the open space. The directional axis is substantially oriented in a direction other than a direction towards the central area of the first member.

Gas turbine combustion system cooling arrangement

A method and apparatus for cooling a combustor liner and transitions piece of a gas turbine include a combustor liner with a plurality of circular ring turbulators arranged in an array axially along a length defining a length of the combustor liner and located on an outer surface thereof; a first flow sleeve surrounding the combustor liner with a first flow annulus therebetween including a plurality of axial channels (C) extending over a portion of an aft end portion of the liner parallel to each other, the cross-sectional area of each channel either constant or varying along the length of the channel, the first flow sleeve having a plurality of rows of cooling holes formed about a circumference of the first flow sleeve for directing cooling air from the compressor discharge into the first flow annulus; a transition piece connected to the combustor liner and adapted to carry hot combustion gases to a stage of the turbine; a second flow sleeve surrounding the transition piece a second plurality of rows of cooling apertures for directing cooling air into a second flow annulus between the second flow sleeve and the transition piece; wherein the first plurality of cooling holes and second plurality of cooling apertures are each configured with an effective area to distribute less than 50% of compressor discharge air to the first flow sleeve and mix with cooling air from the second flow annulus.

Growing community of inventors

Simpsonville, SC, United States of America

William Lawrence Byrne