Abstract: A gas turbine combustion system for burning air and fuel into exhaust gases. The gas turbine combustion system may include a combustor, a turbine nozzle integral with the combustor for providing the air to the combustor, and a fuel injector for providing the fuel to the combustor. The turbine nozzle and the fuel injector are positioned within the combustor such that a mixture of the air and the fuel flows in a first direction and the exhaust gases flow in a second direction.

Abstract: A crossfire tube assembly is configured for connecting adjacent combustion cans in a gas turbine, and includes a first tube segment having a first end and an opposite female end. A second tube segment has a first end and an opposite male end fitted concentrically within the female end with an overlap region between the female and male ends. Each of the first ends of the tube segments is configured for securing to a liner of a respective combustion can. Oppositely oriented first and second impingement sleeves extend from the female end of the first tube segment to the respective first ends of the tube segments, Combustion cooling air is directed through metering holes in the impingement sleeves and flows axially along concentric cavities defined between the impingement sleeves and the first and second tube segments. The combustion cooling air vents from the cavities into an axial combustion air flow stream between the respective combustion can liners and sleeves.

Abstract: A direct-fired coal combustion system includes a swirl chamber having an input configured to receive a coal-water slurry and causing the coal-water slurry to mix with discharge air from a compressor to gasify the coal-water slurry and create a synthesis gas. The system also includes a cyclone separator directly coupled to the second end of the swirl chamber and a second stage combustion input coupled to an output of the cyclone separator.

Abstract: The structure and operation of a fuel nozzle for a gas turbine combustor is disclosed where the fuel nozzle provides for flameholding protection and, more specifically, to such a nozzle that provides for nondestructive protection from flamebolding. The nozzle provides for differential thermal expansion between tubes forming fuel passages to allowing for the nondestructive venting of fuel during a flameholding condition. Upon extinguishing the flameholding condition, the nozzle returns to normal operating condition.

Abstract: A diffuser for a turbine engine, including an annular casing to seal compressor discharge air and struts attached to the casing to delimit an annular array of intra-strut regions, is provided and includes a ring portion, including a ring shaped annular body with a forward face disposed on a forward side thereof, the ring portion being supportable by one or more of the struts, and a splitter portion, including a ring shaped annular body with an aft face disposed on an aft side thereof to mate with the forward face of the ring portion body, the splitter portion further including a lead edge, on a forward side of the splitter portion body, which is extendable into a flow path of the compressor discharge air to split the flow path into secondary flow paths defined radially outside of and inside of the splitter portion body.

Abstract: A system and method for reintroducing gas turbine combustion bypass flow. The system may include a combustor body, wherein the combustor body includes a reaction zone for primary combustion of fuel and air, and a casing enclosing the combustor body and defining an annular passageway for carrying compressor discharge air into the combustor body at one end. The system further may include a reintroduction manifold for receiving combustor bypass air extracted from the compressor discharge air in the annular passageway, and one or more reintroduction slots in communication with the reintroduction manifold for injecting the combustor bypass air into the combustor body downstream of the reaction zone. The method may include extracting combustor bypass air from the annular passageway, transporting the combustor bypass air to a reintroduction manifold, and reintroducing the combustor bypass air into the combustor body through one or more reintroduction slots in communication with the reintroduction manifold.

Abstract: A fuel nozzle manifold comprising a flange, a stem and a swirler is provided. The flange has a first fluid inlet fluidly connected to a radially extending first flow passage, the stem includes at least a first axially extending and only partially circumferentially extending flow channel, and the swirler has at least a first radially extending premix passage. The flange and the stem can comprise a homogeneous component, or two separate components fluidly connecting the first axially extending flow channel to the first flow passage, to form a fluid connection between the flange and the stem, the swirler comprises another component fitted together with the flange and stem component and fluidly connecting the first premix passage and the first flow channel.

Abstract: A direct-fired coal combustion system includes a swirl chamber having an input configured to receive a coal-water slurry and causing the coal-water slurry to mix with discharge air from a compressor to gasify the coal-water slurry and create a synthesis gas. The system also includes a cyclone separator directly coupled to the second end of the swirl chamber and a second stage combustion input coupled to an output of the cyclone separator.

Abstract: A fuel nozzle assembly for a combustor of a gas turbine including: a nozzle body having a front and an inner tube defining a fuel passage extending through the nozzle body, wherein the front proximate to a combustion section of the combustor; an outer casing around the inner tube, wherein an air passage is defined between the outer casing and the inner tube; a gas conduit arranged in the air passage and having an outlet proximate to the front of the nozzle body, wherein fuel starts flowing through the expandable conduit only after a flashback condition occurs in the combustor, and a premix fuel passage and port discharging fuel to a premix section of the combustor, wherein the gas conduit has an inlet open to the premix fuel passage.

Abstract: A secondary combustion system for a stage one turbine nozzle. The secondary combustion system may include a supply tube extending into the stage one nozzle, a number of injectors extending from the supply tube to an outer surface of the stage one nozzle, and an air gap surrounding each of the number of injectors.

Abstract: A nozzle is provided and includes an outer annulus defined by an exterior wall and an interior wall and including air inlets through which air flows to a fuel mixing zone and a combustion zone, an inner annulus disposed within the interior wall and including a fuel volume into which fuel is fed to a distal end thereof, which is adjacent to and isolated from the combustion zone and an airflow line, disposed between the fuel volume and the interior wall, through which air flows to the combustion zone with the airflow line and the combustion zone isolated from the fuel volume, and a fuse configured to melt during a flameholding incident and to form a breach through which fuel flows from the fuel volume, bypassing the fuel mixing zone, to a fuel burning zone downstream from the fuel mixing zone.

Abstract: A combustor for a gas turbine includes a combustion chamber and a plurality of radially outer nozzles surrounding a single, center nozzle, the radially outer nozzles configured to supply only gas fuel to the combustion chamber and the center nozzle configured to supply both gas and liquid fuel to the combustion chamber.

Abstract: A method and apparatus for cooling a combustor liner and transitions piece of a gas turbine include a combustor liner with a plurality of 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, 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 eff

Abstract: A gas-air premixing burner for gas turbines includes an air swirler and an annular burner tube surrounding a bluff centerbody. The bluff body serves to stabilize the flame by defining a recirculating vortex. Cooling air is directed to impinge against the bluff face of the centerbody and the spent impingement cooling air flows in a reverse direction towards the air swirler within the centerbody and is discharged through holes at the outer diameter of the centerbody, where it mixes with the fuel/air mixture prior to reaching the flame zone.

Abstract: A combustor liner for a gas turbine includes a body having a plurality of angled strips on an outside surface of the combustor liner. The plurality of angled strips are arranged in an array about the outside surface. A space is disposed between each of said plurality of angled strips so as to create vortices in a cooling air flowing in a longitudinal direction across said outside surface of said combustor liner. A method of fabricating a combustor liner includes forming a plurality of angled strips on an outside surface of the combustor liner and arranged in an array about the outside surface, each of the angled strips is disposed so as to be spaced apart to create vortices in a cooling air flowing across the outside surface of the combustor liner.

Abstract: A secondary combustion system for a stage one turbine nozzle. The secondary combustion system may include a supply tube extending into the stage one nozzle, a number of injectors extending from the supply tube to an outer surface of the stage one nozzle, and an air gap surrounding each of the number of injectors.

Abstract: A method of operating a gas turbine engine includes providing an inlet flow conditioner (IFC). The IFC has an annular chamber defined therein by at least one wall wherein the wall includes a plurality of perforations extending therethrough. The perforations are spaced in at least two axially-spaced rows that extend circumferentially about the wall. The method also includes channeling a fluid into the IFC and discharging the fluid from the IFC with a substantially uniform flow profile.

Abstract: A combustor includes a center nozzle surrounded by a plurality of outer nozzles, the center nozzle and each of the outer nozzles having a fuel passage and an air passage, with a swirler surrounding the fuel passage and having a plurality of vanes projecting radially within the air passage, each vane having a trailing edge arranged at a swirl angle relative to a longitudinal axis of the nozzle, wherein the swirl angle for the swirler in the center nozzle is different than the swirl angle for the swirlers in the plurality of outer nozzles.

Abstract: A can combustor for an industrial turbine includes a transition piece transitioning directly from a combustor head-end to a turbine inlet using a single piece transition piece. In an exemplary embodiment, the transition piece is jointless.

Abstract: A gas-air premixing burner for gas turbines includes an air swirler and an annular burner tube surrounding a bluff centerbody. The bluff body serves to stabilize the flame by defining a recirculating vortex. Cooling air is directed to impinge against the bluff face of the centerbody and the spent impingement cooling air flows in a reverse direction towards the air swirler within the centerbody and is discharged through holes at the outer diameter of the centerbody, where it mixes with the fuel/air mixture prior to reaching the flame zone.