A combustor may comprise an outer wall defining, at least, a portion of a combustion chamber. A dilution chute may extend from an interior surface of the outer wall. A support structure may extend between the dilution chute and the interior surface of the outer wall. A spall plate may extend from the interior surface of the outer wall. The spall plate may be located between the support structure and an outlet of the combustion chamber.

A method for producing a diffusion cooling hole extending between a wall having a first wall surface and a second wall surface includes forming a cooling hole inlet at the first wall surface, forming a cooling hole outlet at the second wall surface, forming a metering section downstream from the inlet and forming a multi-lobed diffusing section between the metering section and the outlet. The inlet, outlet, metering section and multi-lobed diffusing section are formed by laser drilling, particle beam machining, fluid jet guided laser machining, mechanical machining, masking and combinations thereof.

An article has a metallic substrate having a plurality of recesses. A first coating is at least at the recesses and has: a splatted layer; and a columnar layer atop the splatted layer. A second coating is away from the recesses and has: a columnar layer atop the substrate without an intervening splatted layer.

A self-cooled orifice structure, that may be for a combustor of a gas turbine engine includes a hot side panel, a cold side panel spaced from the hot side panel, and a continuous first wall extending axially between the hot and cold side panels and spaced radially outward from a centerline. The structure may further include a first plurality of helical vanes projecting laterally, radially, inward from the first wall for flowing cooling air in a spiraling fashion through the cold side panel, then through the hot side panel.

The present disclosure provides a gas turbine combustor including a combustion structure (10) having a combustor liner (14) and a flow sleeve (12). The combustor liner (14) includes inner and outer surfaces (31, 30) and defines a combustion zone (15). The gas turbine combustor further includes a plurality of hollow airfoil-shaped structures (22) affixed to the combustor liner (14) and extending radially outwardly into an airflow space (18) defined radially between the flow sleeve (12) and the combustor liner (14). Each hollow structure (22) includes at least one metering tube (26) providing acoustic communication between the combustion zone (15) and the hollow structure (22). The metering tubes (26) are detachably coupled to the combustor liner (14) for permitting interchanging of the metering tube (26) with at least one additional metering tube having at least one different dimension to effect a change in an acoustic characteristic of the hollow structure (22).

The present disclosure provides a gas turbine combustor including a combustion structure (10) having a combustor liner (14) and a flow sleeve (12). The combustor liner (14) includes inner and outer surfaces (31, 30) and defines a combustion zone (15). The gas turbine combustor further includes a plurality of hollow airfoil-shaped structures (22) affixed to the combustor liner (14) and extending radially outwardly into an airflow space (18) defined radially between the flow sleeve (12) and the combustor liner (14). Each hollow structure (22) includes at least one metering tube (26) providing acoustic communication between the combustion zone (15) and the hollow structure (22). The metering tubes (26) are detachably coupled to the combustor liner (14) for permitting interchanging of the metering tube (26) with at least one additional metering tube having at least one different dimension to effect a change in an acoustic characteristic of the hollow structure (22).

Attaching a fastener to a ceramic matrix composite or ceramic body by inserting a portion of the fastener into a compartment of the body, the compartment having a first portion and a second portion, the second portion outboard of and narrower than the first portion; and delivering a fluid into the fastener to expand a portion of the fastener within the blind compartment first portion.

A combustor for a gas turbine engine includes a combustion chamber and a fuel igniter assembly. The combustion chamber is defined by an annular inner combustor liner and an annular outer combustor liner. The fuel igniter assembly is coupled to the combustor and extends radially outward from the outer combustor liner. The fuel igniter assembly includes an igniter housing configured to house a fuel igniter therein, and a heat-dissipating element coupled to the igniter housing. The heat-dissipating element includes a plurality of fins configured to dissipate heat from the fuel igniter assembly.

A premix pilot nozzle includes a tip portion having a downstream surface that extends between a downstream end of an inner wall of the tip portion and a downstream end of an outer wall of the tip portion. The downstream end of the inner wall terminates axially upstream from the downstream end of the outer wall. At least a portion of the downstream surface is curvilinear. The tip portion further comprises a plurality of axially extending premix tubes annularly arranged about the tip portion. Each premix tube defines a premix flow passage through the tip portion. Each premix tube also includes an outlet that is axially offset from the downstream surface.

A grommet for a combustor of a gas turbine engine is disclosed. The grommet includes a lower platform and a raised platform. The lower platform includes a plurality of top slots located in the top surface and a plurality of bottom slots located in the bottom surface. Each top slot and bottom slot may extend radially along the lower platform. The plurality of top slots may be spaced equidistantly from one another along the circumference of the lower platform. The plurality of bottom slots may be spaced equidistantly from one another along the circumference of the lower platform.