A combustion chamber shingle arrangement of a gas turbine, including a combustion chamber shingle, at least one threaded bolt configured for attaching the combustion chamber shingle, and a base body connecting the threaded bolt to the combustion chamber shingle, wherein the base body comprises a platform and a connection area that connects the platform to the combustion chamber shingle, wherein the threaded bolt is arranged at the platform, and wherein a hollow space is embodied between the platform and the combustion chamber shingle.

A combustor for a gas turbine engine includes a support shell; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel such that the rail is non-perpendicular to the cold side and includes a concave surface to at least partially form a curved interface passage; and a second liner panel mounted to the support shell via a multiple of studs, the first liner panel including a second rail that extends from a cold side of the second liner panel and includes a convex surface to at least partially form the curved interface passage.

A combustor includes an inner liner and an outer liner defining a combustion chamber. The inner liner includes an inner mesh structure, and a plurality of inner planks mounted to the inner mesh structure. The outer liner includes an outer mesh structure, and a plurality of outer planks mounted to the outer mesh structure. Each of the plurality of inner planks and outer planks includes an inner wall, an outer wall, and lateral walls defining a cavity to allow circulation of airflow within the cavity to cool down the inner wall.

A combustor heat shield comprises a panel body having a front surface and a back surface. The back surface has rail-less edges and a turbulator band extending inwardly from the rail-less edges. The band includes a plurality of turbulators defining tortuous cooling paths up to the rail-less edges. Effusion holes are distributed over the central area.

An annular combustor for a gas turbine engine including inner and outer combustor walls, wherein each wall defines an annulus and the inner wall is radially inward of the outer wall. The combustor includes a primary zone where the inner and outer combustor walls converge in a downstream direction, and a secondary zone downstream of the primary zone. In the secondary zone, the inner and outer walls are arranged to converge at a different rate to the primary zone, are non-convergent or are divergent in a downstream direction, a rate of change of radial width of the combustor is different in the zones. A transition is provided from the primary zone to the secondary zone. A plurality of combustor cooling tiles lines the inner and outer walls. One or more of the tiles are arranged to extend from the primary to secondary zone and across the transition from the zones.

A liner panel for use in a combustor of a gas turbine engine, the liner panel includes a radiused edge blended into a hot side of the linear panel. 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 radiused edge blended into a hot side of the linear panel, the hot side including a thermal barrier coating. A method of manufacturing including casting a radiused edge blended into a hot side of a linear panel; and applying a thermal barrier coating to the hot side of the liner panel.

A combustor may include a combustor shell, a particle collection panel, and a combustor panel. The combustor shell may define a plurality of first impingement holes, the particle collection panel may include a plurality of particle collection chambers and may define a plurality of second impingement holes, and the combustor panel may define a plurality of effusion holes. The particle collection panel may be disposed inward of the combustor shell and the combustor panel may be disposed inward of the particle collection panel. Each particle collection chamber may have a closed inward end and an opening defined in an outward end. The particle collection chambers may be configured to entrap particulates.

A gas turbine engine component assembly including: a first component having a first surface and second surface opposite the first surface, the first component includes a cooling hole extending from second surface to first surface; a second component having a first surface and second surface, the first surface of first component and second surface of second component defining a cooling channel therebetween in fluid communication with the cooling hole for cooling second surface of second component; and an aperture extension feature in the second component and extending into cooling channel, the aperture extension feature includes a passageway fluidly connecting cooling channel to an area located proximate to the first surface of second component, the aperture extension feature includes a face and inlet located in the first face to fluidly connect passageway to cooling channel, the first face being oriented between 0 and 120 relative to second surface of second component.

A combustor for a gas turbine engine includes a support shell; a first liner panel mounted to the support shell via a multiple of studs, the first liner panel including a first rail that extends from a cold side of the first liner panel; a second liner panel mounted to the support shell via a multiple of studs, the second liner panel including a second rail that extends from a cold side of the second liner panel adjacent to the first rail to form an interface passage; and at least one heat transfer feature within the interface passage.

A liner panel for a combustor of a gas turbine engine includes a rail which at least partially defines an impingement cavity. The rail includes a notch which faces toward the impingement cavity. A method of cooling a wall assembly within a combustor for of a gas turbine engine includes directing air through a support shell and a liner panel to form a pressure drop across the support shell that is less than about 80% of a pressure drop across the combustor and to also form a pressure drop across the liner panel greater than about 20% of the pressure drop across the combustor.