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.

Ceramic matrix composite (CMC) components and methods for forming CMC components of gas turbine engines are provided. In one embodiment, a CMC component for a gas turbine engine includes an inner wall defining a first inner surface; an outer wall defining a second inner surface; and a nozzle extending from the inner wall to the outer wall. The inner wall, outer wall, and nozzle are integrally formed from a CMC material such that the inner wall, outer wall, and nozzle are a single unitary component. An exemplary method for forming a CMC component includes laying up a plurality of plies of a CMC material; processing the plurality of plies to form a green state component; firing the green state component; and densifying the fired component to produce a final unitary component. The unitary component comprises a combustor liner portion and a combustor discharge nozzle stage portion.

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.

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. Additionally, 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 and defines a forward section extending from the quench air jets to the dome. The dome and the forward section of the liner are configured to be cooled substantially by one or both of impingement cooling or convective cooling.

A combustion chamber arrangement includes an annular outer and inner walls including at least one row of tiles. Each tile in the row of tiles has a rail extending towards and sealing with the outer wall and lip extending in a downstream direction from the row of tiles. The outer wall has a row of apertures to direct coolant onto the lips of the row of tiles. Each tile has a fastener positioned upstream of the rail and the fastener extends through a corresponding mounting aperture to secure the tile to the outer wall. The rail of each tile defines a plurality of slots with the outer wall and the slots are arranged in a region downstream of the corresponding fastener. None of the apertures in the row of apertures are in a region downstream of the tiles fastener. The arrangement reduces crack generation and propagation in the outer wall.

A method of machining cooling holes includes providing a workpiece in which a cooling hole is to be formed. The cooling hole, once formed, defines distinct first and second sections. The workpiece is secured in a fixture that is mounted in a first machine. In the first machine, a laser is used to drill a through-hole in a wall of the workpiece. The through-hole is spatially common to the first and second sections of the cooling hole. After drilling the through-hole, the fixture with the workpiece secured therein is removed from the first machine and mounted in a second machine. In the second machine, ultrasonic machining is used to expand a portion of the through-hole to form the second section. An abrasive slurry used in the process is drained through the through-hole during the ultrasonic machining.

A composite layer system is presented. The composite layer system includes a metallic substrate, a structured surface, and a thermal protection system. The structured surface may be additively manufactured onto the metallic substrate and includes structured surface features formed to project above the metallic substrate. Each of the structured surface features are separated from adjacent structured surface features by grooves. The thermal protection coating may be thermally sprayed onto the structured surface and is bonded to each of the structured surface features.

A combustion chamber arrangement including an outer wall and an inner wall spaced from the outer wall and the outer wall supports the inner wall. The inner wall including at least one row of circumferentially arranged tiles. At least one tile in the at least one row of tiles having an L, or V, shaped downstream end and the downstream end of the tile in the row of tiles having a first portion extending from the downstream end of the tile towards and sealing with an inner surface of the outer wall and a second portion extending from the first portion in a downstream direction and away from the inner surface of the outer wall. The first portion of the downstream end of the tile having a plurality of apertures to supply coolant over an inner surface of the second portion of the downstream end of the tile.

The gas turbine combustor liner can delimit a combustion chamber, and have at least one monolithic ceramic block having a first face exposed to the combustion chamber and a second face opposite the first face, and a 3D fabric of ceramic fibers partially embedded inside the monolithic ceramic block, and partially extending outside the second face of the monolithic ceramic block, away from the combustion chamber.

A method includes forming a ceramic matrix composite component by infiltrating an array of ceramic-based fibers with a ceramic-based matrix; forming a plurality of cooling holes in the ceramic matrix composite component; applying a slurry of particles in a carrier fluid to the ceramic matrix composite component such that the slurry passes through the cooling holes and wicks into the ceramic matrix composite material; and processing the ceramic matrix composite component to remove the carrier fluid, thereby leaving a filler at a wall surface of the plurality of cooling holes. A component is also disclosed.