A wake reducing structure for a turbine system includes a combustor liner having an inner surface and an outer surface, the inner surface defining a combustor chamber. Also included is an airflow path located along the outer surface of the combustor liner. Further included is a wake generating component disposed in the airflow path and proximate the combustor liner, wherein the wake generating component generates a wake region located downstream of the wake generating component. Yet further included is an airfoil at least partially disposed in the wake region, the airfoil comprising at least one airfoil hole.

An assembly is provided for an engine. This engine assembly includes an engine structure and a fuel injector. The engine structure includes an injector aperture and a fuel supply passage. The injector aperture extends longitudinally through the engine structure along a centerline. The fuel supply passage extends laterally within the engine structure to the injector aperture. The fuel injector is mated with the injector aperture. The fuel injector includes a nozzle passage and a nozzle orifice. The nozzle passage extends longitudinally along the centerline within the fuel injector to the nozzle orifice. The nozzle passage is fluidly coupled with the fuel supply passage.

A combustor for a gas turbine engine includes a combustor shell, a heat shield and a burner seal. The combustor shell includes metallic materials and is formed to define an interior combustion space. The heat shield includes ceramic matrix composite materials and is configured to shield a portion of the combustor shell from the interior combustion space. The burner seal includes ceramic matrix composite materials and is configured to extend through apertures formed in the combustor shell and the heat shield.

The disclosure is directed to a swirler body for a combustor of a gas turbine engine, where the swirler body includes an annular mount face which defines at least one pocket. The disclosure is directed to a swirler assembly for a combustor of a gas turbine engine, where the swirler assembly includes a swirler first body with an annular first mount face which defines at least one first pocket and a swirler second body with an annular second mount face which abuts said annular first mount face, where said second annular mount face defines at least one second pocket. The disclosure is directed to a method of lightening a swirler assembly for a combustor of a gas turbine engine, where the method includes defining at least one pocket within an annular mount face of a swirler body.

An insert for an impingement plate and an impingement plate including the insert are disclosed. The impingement plate includes a plurality of cooling holes therein. The insert includes a body including an opening extending longitudinally therethrough. The body also includes a discharge end configured for positioning in an impingement air plenum between the impingement plate and an upstream surface of a combustor cap plate; a flexible insertion end configured for insertion into a respective cooling hole of the plurality of cooling holes; and a fixation element between the discharge end and the flexible insertion end. The fixation element has an outer dimension configured to fixedly couple the body in the respective cooling hole.

A combustor cap assembly, combustor and related method are disclosed. The combustor cap assembly includes: an impingement plate defining a plurality of impingement cooling holes with a first side of the impingement plate in fluid communication with a cooling air plenum. The assembly also includes a combustor cap plate coupled to the impingement plate, such that an impingement air plenum is defined between a second side of the impingement plate and the combustor cap plate. Tubes extend from at least a portion of the plurality of impingement cooling holes at the second side of the impingement plate and extend partially towards the combustor cap plate through the impingement air plenum. The plurality of impingement cooling holes provides for fluid communication between the cooling air plenum and the impingement air plenum through the tubes.

A method for replacing a swirler (2) of a premixing or main burner of a burner arrangement (1), wherein the swirler (2) is connected to a distributor flange plate (5) of the burner arrangement (1) via a fuel gas feed pipe (3) and a fuel oil feed line (4) accommodated therein. The fuel gas feed line and the fuel oil feed line are severed to enable removal of the separated end portion of the fuel gas feed line. A sleeve is installed to replace the removed end portion.

The invention relates to a device for installing and removing a component on a gas turbine, having a rail system on which in particular a rotor can be displaced along a predetermined displacement axis, and having a terminal positioning part which is fixedly connected to the rail system. The positioning part has at least one guide portion which is designed to co-operate mechanically with a respective counterpart guide portion on the housing of the gas turbine, in order to achieve a suitable orientation of the device with respect to the housing of the gas turbine. The positioning part also has at least one connecting section which is designed in each case to form a fixed but releasable connection to a respective counterpart connecting section on the housing of the gas turbine construct.

A system includes an internal manifold ring defining at least one circumferentially extending flow channel. A plurality of feed arms extends outward from the manifold ring. A circumferentially segmented outer ring can be supported from the feed arms, outboard of the feed arms. Each feed arm can include a plurality of branches extending therefrom, wherein each branch is in fluid communication with the manifold ring through a respective one of the feed arms. A plurality of nozzles can be included, each nozzle connected to a respective one of the branches, wherein the system is devoid of nozzles radially inward from the manifold ring.

An apparatus and method for mounting a combustion liner within a flow sleeve of a gas turbine combustion system is disclosed. A mounting system comprises a plurality of low-profile mounting tabs secured to a combustion liner where each of the mounting tabs are placed within slots of flow sleeve pegs when the combustion liner is installed in a flow sleeve. A plurality of liner stop brackets are removably secured to a flange of the flow sleeve and have an arm extending to be adjacent to a top contact surface of the mounting tabs. The mounting system reduces blockage to the surrounding airflow.