The invention relates to a burner arrangement for using in a single combustion chamber or in a can-combustor comprising a center body burner located upstream of a combustion zone, an annular duct with a cross section area, intermediate lobes which are arranged in circumferential direction and in longitudinal direction of the center body. The lobes being actively connected to the cross section area of the annular duct, wherein a cooling air is guided through a number of pipes within the lobes to the center body and cools beforehand at least the front section of the center body based on impingement cooling. Subsequently, the impingement cooling air cools the middle and back face of the center body based on convective and/or effusion cooling. At least the back face of the center body includes on the inside at least one damper.

A gas turbine engine component includes a first surface and a second surface. The component further includes a dilution hole defined by the first surface and the second surface. The component further includes a first effusion hole and a second effusion hole each having an inlet defined by the second surface and an outlet defined by the first surface such that the outlet of the first effusion hole is located nearer to the dilution hole than the outlet of the second effusion hole.

A trapped vortex combustor for use in a gas turbine engine includes an outer vortex chamber wall and a dome attached to, or formed integrally with, the outer vortex chamber wall. The dome, the outer vortex chamber wall, or both define at least in part an outer trapped vortex chamber and a channel. The channel extends along the circumferential direction at a forward end of the outer vortex chamber wall, the channel configured to receive an airflow through or around the outer vortex chamber wall, the dome, or both and provide such airflow as a continuous annular airflow to the inner surface of the outer vortex chamber wall. The dome further defines a fuel nozzle opening, with all openings in the dome outward of the fuel nozzle opening along the radial direction, excepting any effusion cooling holes having a diameter less than about 0.035 inches, being in airflow communication with the channel.

An effusion cooling element for the surface of a hot gas path (HGP) component is disclosed. The effusion cooling element includes a coolant swirling chamber embedded within the body of the HGP component. A coolant delivery passage is in the body and configured to deliver a coolant to the coolant swirling chamber. The coolant swirling chamber imparts a centrifugal force to the coolant. An effusion opening is in the HGP surface and in fluid communication with the coolant swirling chamber, the effusion opening having a smaller width than the coolant swirling chamber. The coolant exits the effusion opening over substantially all of 360° about the effusion opening, creating a coolant film on the HGP surface.

A gas turbine engine combustion chamber includes upstream and downstream ring structures and a plurality of circumferentially arranged combustion chamber segments. Each segment extends the full length of the combustion chamber and each segment is secured to the upstream ring structure and is mounted on the downstream ring structure. A frame structure at the downstream end of each segment has multiple spaced radially extending holes. The downstream end of each segment has an axially upstream extending groove and the downstream ring structure has an annular axially upstream extending hook which locates in the groove of each segment. A portion of the downstream ring structure abuts the frame structure of each segment. The downstream ring structure has multiple holes through the portion abutting the frame structure and segment is removably secured to the downstream ring structure by multiple fasteners locating in the holes in the segments and the downstream ring structure.

A gas turbine combustion chamber with a double-wall embodiment, having an outer cold combustion chamber wall and an inner hot combustion chamber wall which form an intermediate space, with impingement cooling holes in the outer combustion chamber wall, effusion cooling holes in the inner combustion chamber wall, outer mixing holes in the outer combustion chamber wall, and inner mixing holes in the inner combustion chamber wall. Respectively, one tubular mixing element connects the outer mixing hole and the inner mixing hole, wherein the mixing element includes an inflow opening in its area which is arranged inside the intermediate space. The outer mixing hole has a smaller diameter than the inner mixing hole, and the throughflow surface area of the effusion holes that are adjoining the mixing element is reduced by the difference in surface area between the outer mixing hole and the inner mixing hole.

The invention relates to a combustion chamber for a turbomachine, such as an aircraft turbojet engine or turboprop engine, comprising an internal annular shroud and an external annular shroud which at their upstream ends are connected by an annular chamber end wall (48), said chamber comprising deflectors (50) mounted upstream of the annular chamber end wall (48). Injectors (19) are mounted in sleeves (44) at least one of which comprises a radially annular flange (66) which is designed to slide radially between the chamber end wall (48) and the deflector (50) and which is blocked axially between the chamber end wall (48) and the deflector (50).

The invention relates to a mixing arrangement for mixing a fuel with a stream of oxygen containing gas flowing along an axis in an axial channel, especially in the second combustor of a gas turbine with sequential combustion. The mixing is improved and the mixing length reduced by said mixing arrangement comprising an injector with at least one injector ring, which is passed by said stream of gas inside and outside.

A turbomachine component, particularly a gas turbine engine component, has at least one part built in parts from a curved or planar panel, particularly a sheet metal, the part having a plurality of cooling channels via which a cooling fluid, particularly air, is guidable, wherein at least one of the plurality of cooling channels has a continuously tapered section. The at least one of the plurality of cooling channels has a single inlet port from a first surface of the panel and a single outlet port for the cooling fluid to another surface, particularly a surface opposite to the first surface, or to the first surface. Further the panel is built via laser sintering or laser melting or direct laser deposition. A gas turbine engine is equipped with such a component. A method of manufacturing includes incorporating cooling channels having a continuously tapered section.

A combustor of an aircraft engine comprises a liner defining a primary and a dilution zone having a hot surface exposed to a flow of combustion gases traveling from the primary zone downstream to the dilution zone and a cold surface. Dilution holes extending through the liner from the cold to the hot surface delimit the primary from the dilution zone. Effusion holes extending through the liner from the cold to the hot surface direct cooling air into the dilution zone. Two or more rows of effusion holes positioned within three dilution hole diameters downstream of the dilution holes are oriented relative to the liner to direct the cooling air in a cooling direction that is at least one of normal to the direction of the flow of gases passing adjacent the effusion holes, and against the direction of the flow of gases passing adjacent the effusion holes.