Abstract: A method of assembling a combustor assembly for use in a turbine engine is provided. A combustor liner that defines a combustion chamber therein is provided. The liner includes a forward end and an aft end, wherein the aft end includes at least one channel extending therethrough. The channel is aligned obliquely with respect to a centerline extending through the aft end. A plurality of fuel nozzles are coupled to the forward end such that the fuel nozzles extend through the forward end. An annular sleeve that includes at least one opening extending radially therethrough is coupled to the aft end, wherein the sleeve substantially circumscribes the aft end such that a cavity is defined therebetween and such that a fluid channeled through the at least one opening impinges against a surface of the aft end prior to the fluid being channeled into the combustion chamber.

Abstract: A fastener assembly configured for use within an opening defined in a transition piece bracket is disclosed. The fastener assembly may generally include a fastener having a head configured to be positioned proximal to a first end of the opening and a body. The body may include a shank portion, a threaded portion opposite the shank portion and a recessed portion defined between the shank portion and the threaded portion. The fastener assembly may also include a retention member configured to be positioned within the opening adjacent the recessed portion such that the retention member engages at least one of the shank portion or the threaded portion as the fastener is moved within the opening. In addition, the fastener assembly may include a biasing member engaged against the retention member. The biasing member may be configured to bias the head of the fastener away from the first end of the opening.

Abstract: A cross fire retention system for a gas turbine engine includes a retention system housing operably coupled to a radially outer surface of a flow sleeve surrounding a combustion chamber, wherein the retention system housing includes a central aperture relatively aligned with a flow sleeve channel and is configured to receive a cross fire tube. Also included is a locking element extending through a side aperture of the retention system housing and having a first end configured to fittingly engage a first portion of an outer surface of the cross fire tube. Further included is a resilient member disposed within the retention system housing and configured to engage a second portion of the outer surface of the cross fire tube.

Abstract: 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.

Abstract: 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.

Abstract: A bracket for a combustor, and a method for mounting a transition piece in a combustor, are disclosed. The combustor has an impingement sleeve at least partially surrounding a transition piece and an outer casing at least partially surrounding the impingement sleeve. The bracket is mounted to the transition piece and connected to the outer casing. The method includes mounting a bracket to the transition piece, extending the bracket through an impingement sleeve, the impingement sleeve at least partially surrounding the transition piece, and connecting the bracket to an outer casing, the outer casing at least partially surrounding the impingement sleeve.

Abstract: A combustion system is provided having a liner, a flow sleeve, a flow-obstructing element, and a venturi. The liner is disposed around a combustion region. The flow sleeve is disposed around the liner. The liner and the flow sleeve cooperate to create an air passage having an airflow located between the liner and the flow sleeve. The flow-obstructing element is disposed within the air passage, and generally obstructs the airflow in the air passage to create wakes in the airflow. The venturi is disposed downstream from the flow-obstructing element, and generally restricts and diffuses the airflow in the air passage to reduce wakes in the airflow.

Abstract: A combustion liner assembly is provided and includes a combustion liner and a transition piece. A portion of the transition piece is circumferentially disposed around a portion of the combustion liner. A seal is attached to the transition piece, and the seal is configured to apply a compressive force to an aft end of the combustion liner.

Abstract: According to one aspect of the invention, a turbine combustor includes an outer member coupled to a wall of the combustor, wherein there is at least one damping hole formed in outer member. The turbine combustor further includes at least one temperature control hole formed in the wall wherein the at least one temperature control hole is formed at an angle with respect to a line perpendicular to a hot gas path in the combustor.

Abstract: A system includes a gas turbine combustor, which includes a combustion liner disposed about a combustion region, a flow sleeve disposed about the combustion liner, an air passage between the combustion liner and the flow sleeve, a fuel injector disposed downstream of the combustion liner and the flow sleeve, a liner mount extending between the combustion liner and the flow sleeve, and a crossfire tube extending between the combustion liner and the flow sleeve. The fuel injector, the liner mount, and the crossfire tube are aligned with one another in a flow enhancing arrangement along a common axis in an axial direction relative to an axis of the gas turbine combustor. The flow enhancing arrangement reduces an air flow disturbance caused by the fuel injector, the liner mount, and the crossfire tube.

Abstract: A system includes a combustor transition piece configured to mount between a combustor and a turbine of a gas turbine engine. The combustor transition piece includes a hollow body with an internal flow passage extending from an upstream end portion to a downstream end portion. The combustor transition piece also includes an aft frame coupled to the downstream end portion of the hollow body, an aft flange coupled to the aft frame, and a mounting lug removably coupled to the aft frame at a first joint. The mounting lug is configured to removably couple to a mounting bracket at a second joint.

Abstract: A turbomachine combustor assembly includes a combustor housing and a combustor body. The combustor body defines a combustor liner having a first end portion that extends to a second end portion through a combustion chamber. A cap assembly is mounted at the combustor housing. The cap assembly includes an endcover, a plurality of fuel nozzles supported by the end cover, and an outer barrel member. The outer barrel member extends about the plurality of fuel nozzles. A liner stop is arranged on one of the outer barrel member and the first end portion of the combustor liner. The liner stop includes a stepped lip portion that receives the other of the outer barrel member and the first end portion of the combustor liner to form substantially smooth liner to outer barrel interface. The liner stop restricts axial movement of the combustor liner relative to the outer barrel member.

Abstract: 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 combustion 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.

Abstract: A combustor for a turbine is provided. The combustor includes a plurality of fuel nozzles and a combustion zone is aligned with a combustion process associated with each of the fuel nozzles. A combustion liner includes a plurality of turbulator groups, and each of the turbulator groups has or more individual turbulators. Each of the turbulator groups is aligned with a hot streak caused by the combustion zone associated with the fuel nozzle. Each of the turbulator groups are circumferentially spaced from a neighboring turbulator group.

Abstract: A seal assembly is provided having a retention assembly attached to a first component, and a seal element secured by the retention assembly. The seal element comprises at least one leaf seal, and the leaf seal is comprised of a plurality of tabs in series with a plurality of slots. The seal assembly forms a flexible seal between the first component and a second component.

Abstract: An apparatus is disclosed. The apparatus may include a body configured to flow hot gases of combustion between a forward end and an aft end. Additionally, the body may include a plurality of raised sections spaced apart circumferentially around an outer perimeter of the body. The raised sections may generally extend lengthwise between the forward and aft ends.

Abstract: An injector apparatus is provided and includes an annular inlet in which a first fluid traveling in a first direction is mixable with a second fluid to form a mixture, an annular outlet disposed downstream from the inlet from which the mixture is injectable into a main flow in a third direction and an annular intermediate section fluidly interposed between the inlet and the outlet and along which the mixture is re-directable from the inlet to the outlet.

Abstract: An impingement sleeve and methods for designing and forming an impingement sleeve are disclosed. In one embodiment, the impingement sleeve includes a body configured to at least partially surround a transition piece of the combustor. The impingement sleeve further includes a plurality of cooling holes defined in the body, the plurality of cooling holes having a cooling hole pattern configured to provide a desired operational value for the transition piece. At least one of the plurality of cooling holes has a chamfer extending at least partially between an inlet and an outlet of the at least one of the plurality of cooling holes. At least a portion of the plurality of cooling holes are generally longitudinally asymmetric.

Abstract: A system includes a gas turbine combustor, which includes a combustion liner disposed about a combustion region, a flow sleeve disposed about the combustion liner, an air passage between the combustion liner and the flow sleeve, and an airflow guide vane disposed in the air passage. The airflow guide vane includes an upstream vane portion and a downstream vane portion. The upstream vane portion is oriented at an angle from an axial axis of the gas turbine combustor. The downstream vane portion is aligned with the axial axis. The airflow guide vane is configured to remove a circumferential swirl of an airflow upstream of the airflow guide vane to straighten the airflow downstream of the airflow guide vane along the axial axis.

Abstract: A system includes a gas turbine combustor, which includes a combustion liner disposed about a combustion region, a flow sleeve disposed about the combustion liner, an air passage between the combustion liner and the flow sleeve, a fuel injector disposed downstream of the combustion liner and the flow sleeve, a liner mount extending between the combustion liner and the flow sleeve, and a crossfire tube extending between the combustion liner and the flow sleeve. The fuel injector, the liner mount, and the crossfire tube are aligned with one another in a flow enhancing arrangement along a common axis in an axial direction relative to an axis of the gas turbine combustor. The flow enhancing arrangement reduces an air flow disturbance caused by the fuel injector, the liner mount, and the crossfire tube.