A system is provided with a fuel sweep system configured to couple to a flow sleeve of a combustor along a first fuel conduit. The flow sleeve is configured to be disposed about a liner of the combustor, and the first fuel conduit is configured to extend along the flow sleeve in a compressor discharge chamber disposed about the flow sleeve. The fuel sweep system includes a first fuel sweep louver adjacent a first fuel sweep opening defined through the flow sleeve.

A liner cooling device for cooling a liner of a gas turbine is provided. The liner cooling device may include a support portion disposed between a liner and a transition piece of a gas turbine and configured to include a cooling flow passage through which cooling air moves to the transition piece. The support portion includes a support member disposed between the liner and the transition piece and an auxiliary support member disposed in the cooling flow passage and having a hole through which the cooling air passes.

A heat shield panel for use in a gas turbine engine combustor is disclosed. In various embodiments, the heat shield panel includes a hot side, a cold side spaced from the hot side, a rail member disposed on the cold side proximate an outer perimeter, the rail member having an outer wall and an inner wall and an orifice extending through the rail member, from the inner wall to the outer wall, the orifice having an entrance portion having an entrance opening positioned on the inner wall and extending at least to an intermediate portion of the rail member and an exit portion having an exit opening positioned on the outer wall and extending at least to the intermediate portion of the rail member, the entrance portion of the orifice being angled relative to the exit portion of the orifice.

An impingement cooling structure is provided. The impingement cooling structure includes a flow channel formed between a first wall and a second wall facing the first wall, a plurality of impingement cooling holes disposed in the first wall such that the plurality of impingement cooling holes are spaced apart from each other along the flow channel, and a flow diverter convexly protruding from a surface of the second wall in each space between injection axes of the plurality of impingement cooling holes.

Methods for manufacturing combustor panels of gas turbine engines and combustor panels are described. The methods include defining a particle deposit near-steady state for at least a portion of a combustor panel, the particle deposit near-steady state representative of a build-up of particles on the at least a portion of the combustor panel during use, generating a template based on the defined particle deposit near-steady state, wherein the template includes one or more augmentation elements based on the representative of build-up of particles, and forming a combustor panel based on the template, wherein the formed combustor panel includes one or more augmentation elements defined in the template.

A liner for a combustor includes a support member, an intermediate member, and a liner member. The intermediate member is positioned intermediate the support member and the liner member and has a plurality of protrusions and a plurality of recesses. The support member is coupled to the intermediate member at a tangent of each protrusion. Additionally, the liner member is comprised of a ceramic matrix composite material. The liner member is coupled to the intermediate member at a tangent of each recess.

Combustors and gas turbines are provided. A combustor includes an axial centerline and an end cover. The combustor further includes at least one fuel nozzle that extends from the end cover and at is least partially surrounded by a combustion liner. The combustion liner extends between the at least one fuel nozzle and an aft frame and that defines a combustion chamber. An outer sleeve is spaced apart from and surrounds the combustion liner such that an annulus is defined therebetween. The outer sleeve defines at least one aperture. A wake energizer is mounted on the outer sleeve. The wake energizer defines at least one passage that is angled with respect to the axial centerline of the combustor. The at least one passage aligns and is in fluid communication with the at least one aperture of the outer sleeve.

The invention is a combustor liner (12) of a dual wall cooling structure including an inner wall section (30) configured to surround a combustion region (13) and in which a plurality of effusion cooling holes (31) are formed, and an outer wall section (20) formed to be spaced apart from the inner wall section (30) and in which a plurality of impingement cooling holes (21) are formed, wherein the inner wall section (30) is constituted by a plurality of plate-shaped members (40), and a support guide member (50) is provided which is configured to guide the plurality of plate-shaped members (40) to enable free insertion and extraction and support the plurality of plate-shaped members (40) at intervals such that deformation by thermal expansion is able to be absorbed.

Disclosed herein are a transition part assembly which is improved in efficiency of cooling a high-temperature region formed on a side surface of a transition part of a gas turbine, and a combustor including the same. The transition part assembly includes a transition part, a collision sleeve, a cooling hole, and a guide which is formed inside the collision sleeve so as to guide air to a side surface of the transition part.

A rich-quench-lean combustor assembly for a gas turbine engine includes a fuel nozzle and a dome, the fuel nozzle attached to the dome. The combustor assembly additionally includes a liner attached to or formed integrally with the dome, the liner and the dome together defining at least in part a combustion chamber. The liner extends between a forward end and an aft end. The liner includes a plurality of quench air jets positioned between the forward end and aft end. The quench air jets include a plurality of primary stage air jets and a plurality of secondary stage air jets. The plurality of primary stage air jets are each spaced from the plurality of secondary stage air jets along the axial direction and together provide the combustion chamber with a quench airflow.