Abstract: A method of optimizing a premix fuel nozzle for a gas turbine is provided including the step of providing a nozzle that provides, when the gas turbine is in operation, an axial flow field of an air and fuel mixture flows through the burner tube and around the nozzle tip, and at least two recirculation zones with at least two differing radial extents as part of a toroidal vortex generated on the nozzle tip to provide strong flame holding and flame propagation.

Abstract: A nozzle tip for a premix fuel nozzle includes an outer body having an outer body external face facing the downstream end of the burner tube, the outer body external face having a smaller cross-sectional area than the cross-sectional area of the burner tube; and at least one segment radiating radially outwardly toward the internal wall of the burner tube from the outer body, wherein each segment has a proximal end disposed adjacent to the outer body external face and a distal end disposed in a direction toward the burner tube, wherein each segment has a segment downstream face angled relative to the longitudinal axis of the burner tube towards the downstream end of the burner tube. When the gas turbine is in operation, an axial flow field of an air and fuel mixture flows through the burner tube and around the nozzle tip, and two or more recirculation zones of differing radial extent are generated on the nozzle tip by the segments to provide strong flame holding and flame propagation.

Abstract: A method of optimizing a premix fuel nozzle for a gas turbine is provided including the step of providing a nozzle that provides, when the gas turbine is in operation, an axial flow field of an air and fuel mixture flows through the burner tube and around the nozzle tip, and at least two recirculation zones with at least two differing radial extents as part of a toroidal vortex generated on the nozzle tip to provide strong flame holding and flame propagation.

Abstract: A nozzle tip for a premix fuel nozzle includes an outer body having an outer body external face facing the downstream end of the burner tube, the outer body external face having a smaller cross-sectional area than the cross-sectional area of the burner tube; and at least one segment radiating radially outwardly toward the internal wall of the burner tube from the outer body, wherein each segment has a proximal end disposed adjacent to the outer body external face and a distal end disposed in a direction toward the burner tube, wherein each segment has a segment downstream face angled relative to the longitudinal axis of the burner tube towards the downstream end of the burner tube. When the gas turbine is in operation, an axial flow field of an air and fuel mixture flows through the burner tube and around the nozzle tip, and two or more recirculation zones of differing radial extent are generated on the nozzle tip by the segments to provide strong flame holding and flame propagation.

Abstract: A turbine engine having a plenum for passing fluids from an outlet of a compressor to an inlet of a combustor that may increase the efficiency of the turbine engine. The turbine engine may include a combustor, a compressor positioned upstream of the combustor, a transition channel extending from the compressor to the combustor, and a shell extending between the compressor and a combustor portal and positioned around the at least one transition channel. The turbine engine may also include an axial diffusor in the shell near the at least one transition channel, wherein the axial diffusor may include a fluid flow recess in a trailing edge of the axial diffusor. The turbine engine may also include a wave protrusion extending from a surface positioned radially inward of the axial diffusor. The fluid flow recess and the wave protrusion may reduce fluid flow loss within the shell.

Abstract: A method of assembling a radiant cooler includes providing a vessel shell that defines a gas flow passage therein that extends generally axially through the vessel shell, forming a tube cage from coupling a plurality of cooling tubes together to form a tube cage defined by a plurality of chevron-shaped projections that extend circumferentially about a center axis of the tube cage, each chevron-shaped projection includes a first side and a second side coupled together a tip, circumferentially-adjacent pairs of projections coupled together such that a valley is defined between each pair of circumferentially-spaced projections, each of the projection tips is positioned radially outward from each of the valleys, and orienting the tube cage within the vessel shell such that the tube cage is in flow communication with the flow passage.

Abstract: An aft liner seal for a combustor for a gas turbine includes an inner shell having an inner and outer surface and central axis. The aft liner seal has an outer shell positioned over the inner shell that has an inner and outer surface and central axis coaxial with the inner shell. One of the outer surface of the inner shell or the inner surface of the outer shell has grooves angled relative to the central axis. The angled grooves and adjacent portions of the inner surface of the outer shell or the outer surface of the inner shell form cooling passages. The cooling air exits the cooling passages at an exit angle that is matched to a swirl angle of the combustor flow. Matching of the exit angle and the swirl angle minimizes shear of cooling air with respect to flow exiting the combustor liner.

Abstract: A secondary nozzle is provided for a gas turbine. The secondary nozzle includes a flange and an elongated nozzle body extending from the flange. At least one premix fuel injector is spaced radially from the nozzle body and extends from the flange generally parallel to the nozzle body. At least one second nozzle tube is fluidly connected to the fuel source and spaced radially outward from the first nozzle tube with a proximal end fixed to the flange. The second nozzle tube has a distal end, spaced from the proximal end, with at least one aperture therein. A passageway extends between the proximal end and the distal end of the second nozzle tube, with the passageway fluidly connecting to the fuel source and the aperture.

Abstract: A burner (27) of a gas turbine engine (10) includes a cylindrical basket (60) comprising an air flow reversal region (86). The flow reversal region ends at an air inlet plane (84) of the basket. The burner also includes a flow conditioner (90) disposed in the flow reversal region transecting an air flow (80) flowing non-uniformly through the flow reversal region, the flow conditioner being effective to mitigate variation of the air flow entering the basket across the inlet plane.

Abstract: A combustor for a gas turbine is provided having a nozzle assembly located at one end and a combustion chamber defining a second end of the combustor. A venturi is positioned within the combustor, between the nozzle and the combustion chamber. The venturi defines a passageway therein having a first side facing the nozzle and a second side facing the combustion chamber. Compressed air is directed into an inlet in fluid communication with the first and second sides of the venturi passageway. The venturi passageway directs the compressed air from the inlet in opposite directions within the first and second sides of the passageway for cooling the venturi.

Abstract: A combustor for a gas turbine is provided having a nozzle assembly located at one end and a combustion chamber defining a second end of the combustor. A venturi is positioned within the combustor, between the nozzle and the combustion chamber. The venturi defines an internal passageway therein having a first side facing the nozzle and a second side facing the combustion chamber. Compressed air is directed into an inlet in fluid communication with the first and second sides of the internal passageway. The internal passageway directs the compressed air from the inlet in opposite directions within the first and second sides of the internal passageway for cooling the venturi.

Abstract: A secondary nozzle is provided for a gas turbine. The secondary nozzle includes a flange and an elongated nozzle body extending from the flange. At least one premix fuel injector is spaced radially from the nozzle body and extends from the flange generally parallel to the nozzle body. At least one second nozzle tube is fluidly connected to the fuel source and spaced radially outward from the first nozzle tube with a proximal end fixed to the flange. The second nozzle tube has a distal end, spaced from the proximal end, with at least one aperture therein. A passageway extends between the proximal end and the distal end of the second nozzle tube, with the passageway fluidly connecting to the fuel source and the aperture.

Abstract: A turbine engine having a plenum for passing fluids from an outlet of a compressor to an inlet of a combustor that may increase the efficiency of the turbine engine. The turbine engine may include a combustor, a compressor positioned upstream of the combustor, a transition channel extending from the compressor to the combustor, and a shell extending between the compressor and a combustor portal and positioned around the at least one transition channel. The turbine engine may also include an axial diffusor in the shell near the at least one transition channel, wherein the axial diffusor may include a fluid flow recess in a trailing edge of the axial diffusor. The turbine engine may also include a wave protrusion extending from a surface positioned radially inward of the axial diffusor. The fluid flow recess and the wave protrusion may reduce fluid flow loss within the shell.

Abstract: A method of assembling a radiant cooler includes providing a vessel shell that defines a gas flow passage therein that extends generally axially through the vessel shell, forming a tube cage from coupling a plurality of cooling tubes together to form a tube cage defined by a plurality of chevron-shaped projections that extend circumferentially about a center axis of the tube cage, each chevron-shaped projection includes a first side and a second side coupled together a tip, circumferentially-adjacent pairs of projections coupled together such that a valley is defined between each pair of circumferentially-spaced projections, each of the projection tips is positioned radially outward from each of the valleys, and orienting the tube cage within the vessel shell such that the tube cage is in flow communication with the flow passage.

Abstract: A burner (27) of a gas turbine engine (10) includes a cylindrical basket (60) comprising an air flow reversal region (86). The flow reversal region ends at an air inlet plane (84) of the basket. The burner also includes a flow conditioner (90) disposed in the flow reversal region transecting an air flow (80) flowing non-uniformly through the flow reversal region, the flow conditioner being effective to mitigate variation of the air flow entering the basket across the inlet plane.

Abstract: A turbine engine having a plenum for passing fluids from an outlet of a compressor to an inlet of a combustor that may increase the efficiency of the turbine engine. The turbine engine may include a combustor, a compressor positioned upstream of the combustor, a transition channel extending from the compressor to the combustor, and a shell extending between the compressor and a combustor portal and positioned around the at least one transition channel. The turbine engine may also include an axial diffusor in the shell near the at least one transition channel, wherein the axial diffusor may include a fluid flow recess in a leading edge of the axial diffusor. The turbine engine may also include a wave protrusion extending from a surface positioned radially inward of the axial diffusor. The fluid flow recess and the wave protrusion may reduce fluid flow loss within the shell.

Abstract: Aspects according to the invention relate to a catalytic combustor system for a turbine engine and an associated method. Catalytic combustors are used in connection with turbine engines because they can minimize the formation of oxides of nitrogen during combustion. Despite this emissions advantage, catalytic combustion systems can increase the level of CO in the turbine exhaust. According to aspects of the invention, vortex formation devices includes vortex generators, swirlers and mixers can be placed downstream of each catalytic module surrounding the pilot nozzle so as to form one or more vortices in the otherwise substantially laminar flow exiting the modules. The vortices can create a suction so that a portion of the flow exiting the pilot nozzles is mixed with the flow exiting the catalyst modules. The introduction of the higher temperature pilot flow can accelerate the catalytic reaction time, promoting burnout of the CO formed during combustion.

Abstract: A system and method for providing continuous measurement and control of a combustion device by altering the fuel composition delivered thereto. The system includes devices for sensing related information, such as fuel characteristics, combustion characteristics, or other device characteristics, and controlling the performance of the combustion device based on the sensed information. Performance control occurs via addition of one or more additives to the fuel to adjust combustion characteristics. Via such sensing and performance control, consistent combustion device performance may be maintained, despite varying fuel characteristics. In one variation, sensing occurs for the fuel delivered to the combustion device, and one or more additives are added to the fuel, based on the composition and flow rate for the fuel.

Abstract: Aspects according to the invention relate to a catalytic combustor system for a turbine engine and an associated method. Catalytic combustors are used in connection with turbine engines because they can minimize the formation of oxides of nitrogen during combustion. Despite this emissions advantage, catalytic combustion systems can increase the level of CO in the turbine exhaust. According to aspects of the invention, vortex formation devices includes vortex generators, swirlers and mixers can be placed downstream of each catalytic module surrounding the pilot nozzle so as to form one or more vortices in the otherwise substantially laminar flow exiting the modules. The vortices can create a suction so that a portion of the flow exiting the pilot nozzle is mixed with the flow exiting thee catalyst modules. The introduction of the higher temperature pilot flow can accelerate the catalytic reaction time, promoting burnout of the CO formed during combustion.

Abstract: A method and apparatus for reducing total pressure loss in the exhaust of a combustion device via an axial diffuser having no turning vanes, support struts, rods, or other obstructions to exhaust flow. The axial diffuser includes an inner wall and an outer wall each having a turn radius and forming an annular cross section having increasing cross-sectional area in the downstream direction of flow, the flow being transitioned from axial to radial direction thereby. The diffuser is supported in the plenum by rods, struts or a cradle; from a stand surrounding the axial diffuser via struts or rods; or from existing diffuser support structure by web stiffeners. The diffuser flow path turn radius of the outer wall ranges from about 50-95% of the radial distance from the inner wall to a wall of the exhaust plenum, and of the inner wall ranges from about 10-70% of this distance.