Abstract: Methods and systems are provided for inducing combustion dynamics within turbine engines to remove combustion deposits within the turbine engine during operation of the turbine engine.

Abstract: The present application provides a micro-mixer combustion nozzle for mixing a flow of fuel and a flow of air in a gas turbine engine. The micro-mixer combustion nozzle may include a fuel plate with a number of fuel plate apertures and a fuel plate passage in communication with the flow of fuel and an air plate with a number of air plate apertures and an air plate passage in communication with the flow of air. The fuel plate passage and the air plate passage may align to mix in part the flow of fuel and the flow of air.

Abstract: A gas turbine includes a compressor and at least one combustor downstream from the compressor. The combustor includes a burner having an inner shroud extending axially along at least a portion of the burner, an outer shroud radially separated from the inner shroud and extending axially along at least a portion of the burner, and a plurality of stator vanes extending radially between the inner shroud and the outer shroud. The stator vanes have an inner end proximate the inner shroud and an outer end proximate the outer shroud. The burner further includes a vortex tip at one of either the inner end or the outer end of the stator vanes. The vortex tip provides a gap between the inner end and the inner shroud or the outer end and the outer shroud, and the vortex tip includes a plurality of fuel ports. The gas turbine further includes a turbine downstream from the combustor.

Abstract: Turbine systems are provided. A turbine system includes a transition duct comprising an inlet, an outlet, and a duct passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The duct passage includes an upstream portion and a downstream portion. The upstream portion extends from the inlet between an inlet end and an aft end. The downstream portion extends from the outlet between an outlet end and a head end. The turbine system further includes a joint coupling the aft end of the upstream portion and the head end of the downstream portion together. The joint is configured to allow movement of the upstream portion and the downstream portion relative to each other about or along at least one axis.

Abstract: An aft frame of a turbine engine transition piece body is provided and includes an annular body disposed within a first annular space defined between an impingement sleeve and a compressor discharge casing and aft of a second annular space defined between the transition piece body and the impingement sleeve and including a main portion with a first surface facing the first annular space and a second surface facing the forward annular space. The main portion has an impingement hole extending therethrough from an inlet at the first surface of the annular body to an outlet at the second surface of the annular body to define a fluid path along which the first and second annular spaces communicate with one another.

Abstract: A turbine system is disclosed. In one embodiment, the turbine system includes a transition duct. The transition duct includes an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The transition duct further includes an interface member for interfacing with a turbine section. The turbine system further includes a flexible metallic seal contacting the interface member to provide a seal between the interface member and the turbine section.

Abstract: An injector seal for a gas turbomachine includes a seal body having a first surface, an opposing second surface, a central opening configured and disposed to receive a gas turbomachine injector, an outer edge, and a plurality of passages extending through the seal body. The plurality of passages are configured and disposed to guide a fluid at the first surface through the second surface.

Abstract: A support assembly for a turbine system is disclosed. The support assembly includes a transition duct extending between a fuel nozzle and a turbine section. The transition duct has an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis. The support assembly further includes an annular contact member extending from the transition duct. The contact member is configured to allow movement of the transition duct about at least one axis.

Abstract: An injector seal for a gas turbomachine includes a seal body having a first surface, an opposing second surface, a central opening configured and disposed to receive a gas turbomachine injector, an outer edge, and a plurality of passages extending through the seal body. The plurality of passages are configured and disposed to guide a fluid at the first surface through the second surface.

Abstract: A system comprising a fuel nozzle. The fuel nozzle includes a mounting base and an inlet flow conditioner extending directly from the mounting base in a downstream direction. Moreover, the inlet flow conditioner structurally supports the fuel nozzle without a central support member extending directly from the mounting base inside the inlet flow conditioner.

Abstract: A combustor liner is provided. The combustor liner may include an upstream portion and a downstream end portion. The upstream portion may have a radius and a length along a generally longitudinal axis. The downstream end portion may have a radius and a length along the generally longitudinal axis. The downstream end portion may define a plurality of channels. Each of the plurality of channels may extend helically through the length of the downstream end portion. Each of the plurality of channels may be configured to flow an air flow therethrough, cooling the downstream end portion.

Abstract: A turbine system is disclosed. In one embodiment, the turbine system includes a transition duct. The transition duct includes an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The transition duct further includes an interface feature for interfacing with an adjacent transition duct. The turbine system further includes a convolution seal contacting the interface feature to provide a seal between the interface feature and the adjacent transition duct.

Abstract: A combustor includes an end cap that extends radially across at least a portion of the combustor, wherein the end cap comprises an upstream surface axially separated from a downstream surface. A shroud circumferentially surrounds at least a portion of the end cap, wherein the shroud at least partially defines a fuel plenum between the upstream surface and the downstream surface. A combustion chamber downstream from the end cap defines a longitudinal axis. A plurality of tubes extend from the upstream surface through the downstream surface of the end cap to provide fluid communication through the end cap to the combustion chamber. A transition duct circumferentially surrounds at least a portion of the combustion chamber downstream from the end cap and curves tangentially from the longitudinal axis.

Abstract: A turbine system is disclosed. In one embodiment, the turbine system includes a transition duct. The transition duct includes an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The transition duct further includes an interface feature for interfacing with an adjacent transition duct. The turbine system further includes a flexible seal contacting the interface feature to provide a seal between the interface feature and the adjacent transition duct. The flexible seal includes a sheet having a first surface, an opposing second surface, and a peripheral edge therebetween.

Abstract: A system for supplying an injection fluid to a combustor is disclosed. The system includes a transition duct comprising an inlet, an outlet, and a passage extending between the inlet and the outlet and defining a longitudinal axis, a radial axis, and a tangential axis. The outlet of the transition duct is offset from the inlet along the longitudinal axis and the tangential axis. The passage defines a combustion chamber. The system further includes a tube providing fluid communication for the injection fluid to flow through the transition duct and into the combustion chamber.

Abstract: A combustion system including a transition piece and method of forming the transition piece using a cast nickel-based superalloy is provided. The transition piece includes a body defining a flowpath and enclosure, the body having a circular inlet section for receiving combustion product from the combustor and an outlet end for flowing the combustor products into a first stage nozzle of a gas turbine. The transition piece comprises a one-piece single casting construction of the circular inlet section, the body, and the outlet end. The transition piece is formed from a nickel-based superalloy. The transition piece has a thick-to-thin ratio of wall thickness of thick transition to wall thickness of thin transition of about 1.1 to about 2.5.

Abstract: The present application provides a combustion system for use with a cooling flow. The combustion system may include a head end, an aft end, a transition nozzle extending from the head end to the aft end, and an impingement sleeve surrounding the transition nozzle. The impingement sleeve may define a first cavity in communication with the head end for a first portion of the cooling flow and a second cavity in communication with the aft end for a second portion of the cooling flow. The transition nozzle may include a number of cooling holes thereon in communication with the second portion of the cooling flow.

Abstract: The present application provides a combustor for use with a gas turbine engine. The combustor may include a head end with a non-circular configuration, a number of fuel nozzles positioned about the head end, and a transition piece extending downstream of the head end.

Abstract: A system comprising a fuel nozzle. The fuel nozzle includes a mounting base and an inlet flow conditioner extending directly from the mounting base in a downstream direction. Moreover, the inlet flow conditioner structurally supports the fuel nozzle without a central support member extending directly from the mounting base inside the inlet flow conditioner.

Abstract: A system comprises a turbine combustor cap. The turbine combustor cap includes a plurality of segments. Each segment of the plurality of segments has edges abutting at least two fuel nozzle receptacles. Moreover, each segment of the plurality of segments does not completely surrounding any one fuel nozzle receptacle.