A cooled gas turbine engine component comprises a wall which has a plurality of effusion cooling apertures extending there-through from a first surface to a second surface. The apertures are arranged at an angle to the second surface and each aperture has an inlet in the first surface and an outlet in the second surface. Each aperture has a metering portion and a diffusing portion arranged in flow series and each metering portion is elongate and the width is greater than the length of the metering portion. Each diffusing portion increases in dimension in the length from the metering portion to the outlet. Each outlet has a rectangular shape in the second surface of the wall. Each inlet has an elongate shape in the first surface of the wall and the inlet in the wall is arranged substantially diagonally with respect to the outlet in the wall.

A combustor wall is provided for a turbine engine. The combustor wall includes a combustor shell and a combustor heat shield that is attached to the shell. The heat shield includes a first panel and a second panel that sealingly engages the first panel in an overlap joint. A cooling cavity extends between the shell and the heat shield and fluidly couples a plurality of apertures in the shell with a plurality of apertures in the heat shield.

An article may include a substrate, a plurality of cooling holes in the substrate, wherein each of the plurality of cooling holes defines substantially the same diameter measured parallel to an outer surface of the substrate, and a coating on the surface of the substrate. In accordance with these examples, the coating covers and substantially blocks a first set of cooling holes from the plurality of cooling holes and leaves a second set of cooling holes from the plurality of cooling holes substantially uncovered. In some examples, an article including a substrate, a plurality of cooling holes in the substrate, and a coating on the substrate. In accordance with these examples, the coating covers and partially occludes each cooling hole of the plurality of cooling holes, and the coating does not extend into any of the plurality of cooling holes.

In a combustion cylinder, an outer-side region of a cooling part on an outer side with regard to a reference line orthogonal to a radial direction and an axial direction of a gas turbine and passing through a center of a combustor basket, and an inner-side region of the cooling part on an inner side with regard to the reference line are set. A connection angle at an intersection between an extension line of an outer surface of the combustor basket along the axial direction and an inner surface of a combustor transition piece is set. First regions are set at positions near the reference line, and second regions, with a larger connection angle than the first regions, are set at positions farther from the reference line than the first regions. The second regions are set to have a higher flow rate of a cooling medium than the first regions.

A liner for a combustor includes a first wall and a second wall extending around at least a portion of the first wall to form a liner cavity with the first wall. The first wall defines a first wall orifice and the second wall defines a second wall orifice. The liner further includes a first insert mounted on the second wall within the second wall orifice and extending through the first wall orifice. The first insert is configured to direct a first air jet through the second wall and the first wall. The first insert is formed by a tubular body portion extending through the first wall orifice and the second wall orifice, and a shoulder extending around the body portion that abuts the first side of the second wall. The shoulder has a first diameter and the first wall orifice has a second diameter, greater than the first diameter.

A liner panel for use in a combustor of a gas turbine engine, the liner panel including a liner panel with a perimeter rail and a multiple of intermediate rails. A combustor for a gas turbine engine including a support shell and a multiple of liner panels mounted to the support shell via a multiple of studs, each of the multiple of liner panels having a multiple of intermediate rails to form a multiple of circumferential cavities.

A gasket for installation within a cavity between a support shell and a liner panel of a combustor wall assembly, the support shell having a multiple of impingement holes, and the liner panel having a multiple of film holes, the gasket including a multiple of cells arranged with respect to the multiple of impingement holes and the multiple of film holes, wherein at least one of the multiple of cells segregate a respective subset of the multiple of impingement holes and a subset of the multiple of film holes.

A combustor for a gas turbine engine including a liner panel mounted to a support shell via a multiple of studs, the liner panel including a forward section and an aft section that defines an arcuate surface section therebetween in the axial profile between the forward section and the aft section.

A combustor for a gas turbine engine including a liner panel mounted to a support shell via a multiple of studs, the liner panel including a forward section and an aft section that defines an angle therebetween in the axial profile between the forward section and the aft section.

A combustor for a gas turbine engine includes an annular shell, an annular bulkhead connected to the shell, and a heat shield panel. The heat shield panel has a first surface facing a combustion chamber and a second surface opposite the first surface. The heat shield panel is mounted to the bulkhead and defines a cooling chamber between the bulkhead and the heat shield panel. The heat shield panel has a wall extending from the heat shield panel toward the bulkhead around at least a portion of a periphery of the heat shield panel. The wall includes a circumferential wall portion including at least one cooling air passage extending between the cooling chamber and a cavity defined between the circumferential wall portion and the shell. The at least one cooling air passage is configured to purge the cavity by directing a first cooling air stream from the cooling chamber into the cavity.