Abstract: A combustor panel arrangement for a gas turbine engine. The combustor panel arrangement includes a first combustor panel that has a first edge. A second combustor panel has a second edge facing the first edge. A first plurality of effusion holes extend through the first edge towards the second edge along a corresponding one of a first plurality of flow paths. A second plurality of effusion holes extend through the second edge along a corresponding one of a second plurality flow paths towards the first edge. The first plurality of flow paths and the second plurality of flow paths are non-intersecting.

Abstract: A combustor panel arrangement for a gas turbine engine. The combustor panel arrangement includes a first combustor panel that has a first edge. A second combustor panel has a second edge facing the first edge. A first plurality of effusion holes extend through the first edge towards the second edge along a corresponding one of a first plurality of flow paths. A second plurality of effusion holes extend through the second edge along a corresponding one of a second plurality flow paths towards the first edge. The first plurality of flow paths and the second plurality of flow paths are non-intersecting.

Abstract: A wall assembly for a combustor of a gas turbine engine is provided. The wall assembly includes a support shell with a contoured region to define at least one convergent passage between the support shell and a multiple of liner panels. A method of cooling a wall assembly for a combustor of a gas turbine engine is also provided. This method includes a step of directing cooling air into at least one convergent passage between the support shell and a multiple of liner panels.

Abstract: A bulkhead assembly for a combustor of a gas turbine engine includes a bulkhead support shell with a multiple of swirler openings. Each of the multiple of swirler openings is surrounded by a first multiple of cooling impingement passages that define an angle ? with respect to a hot side of the bulkhead support shell and a second multiple of cooling impingement passages that define an angle ? with respect to the hot side of the bulkhead support shell. The angle ? is different than the angle ?.

Abstract: A heat shield for a combustor of a gas turbine engine includes a first edge with a first set of cantilevered members and a second edge with a second set of cantilevered members.

Abstract: A liner assembly for a combustor of a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a grommet with a multiple of grommet cooling passages.

Abstract: A combustor wall is provided for a turbine engine and includes a combustor shell and a heat shield. The heat shield is attached to the shell with first and second cavities extending between the shell and the heat shield. The first cavity fluidly couples apertures defined in the shell with the second cavity. The second cavity fluidly couples the first cavity with apertures defined in the heat shield. The shell and the heat shield converge toward one another about the second cavity.

Abstract: A combustor section is provided for a gas turbine engine. The combustor section may include an outer liner panel, an inner liner panel and a bulkhead, which is arranged with the outer and the inner liner panels to form an annular combustion chamber. The combustor section may also include a swirler assembly and a combustor vane. The swirler assembly may be configured with the bulkhead. The combustor vane may extend at least partially into said combustion chamber, wherein the combustor vane is circumferentially aligned with the swirler assembly.

Abstract: A combustor wall is provided for a turbine engine and includes a combustor shell and a heat shield. The heat shield is attached to the shell with first and second cavities extending between the shell and the heat shield. The first cavity fluidly couples apertures defined in the shell with the second cavity. The second cavity fluidly couples the first cavity with apertures defined in the heat shield. The shell and the heat shield converge toward one another about the second cavity.

Abstract: A combustor section for a gas turbine engine includes a combustor vane which extends at least partially into a combustion chamber.

Abstract: A wall assembly for a combustor of a gas turbine engine is provided. The wall assembly includes a support shell with a contoured region to define at least one convergent passage between the support shell and a multiple of liner panels. A method of cooling a wall assembly for a combustor of a gas turbine engine is also provided. This method includes a step of directing cooling air into at least one convergent passage between the support shell and a multiple of liner panels.

Abstract: A bulkhead assembly for a combustor of a gas turbine engine includes a bulkhead support shell with a multiple of swirler openings. Each of the multiple of swirler openings is surrounded by a first multiple of cooling impingement passages that define an angle ? with respect to a hot side of the bulkhead support shell and a second multiple of cooling impingement passages that define an angle ? with respect to the hot side of the bulkhead support shell. The angle ? is different than the angle ?.

Abstract: A swirler assembly for a gas turbine engine includes an outer annular injector which at least partially surrounds an inner injector. In sonic embodiments, a combustor section for a gas turbine engine comprises an inner injector which defines an axis, an outer annular injector which surrounds said inner injector, and a combustor vane along said axis.

Abstract: A combustor liner is arcuate in shape and defines an axis and a circumferential direction. The combustor liner includes a first row of dilution openings and a second row of dilution openings. The first row runs in the circumferential direction. The second row runs parallel to the first row and is axially spaced from the first row. Each dilution opening of the second row overlaps in an axial direction a portion of each of two adjacent dilution openings of the first row.

Abstract: A heat shield is disclosed. The heat shield may comprise a body having a back surface and an opposite front surface, wherein an opening in the body communicates through the front and back surfaces. The heat shield may further comprise at least one radial rail disposed on the back surface and extending radially outward from the opening for directing cooling air flow.

Abstract: A swirler assembly for a gas turbine engine includes an outer annular injector which at least partially surrounds an inner injector and an air-assist atomizer upstream of said inner injector.

Abstract: A liner assembly for a combustor of a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a grommet with a multiple of grommet cooling passages.

Abstract: A heat shield for a combustor liner includes first linear film cooling slots through the heat shield and second linear film cooling slots through the heat shield. The first linear film cooling slots are run in a row and each of the first linear film cooling slots is angled from the row in a first direction. The second linear film cooling slots also run in the row and each of the second linear film cooling slots is angled from the row in a second direction opposite the first direction. The second linear film cooling slots alternate with the first linear film cooling slots in the row. The first and second linear film cooling slots are connected to form a single, multi-cornered film cooling slot.

Abstract: A combustor liner includes a heat shield, a shell, a series of trip strips, and a series of projecting walls. The heat shield has a shield cold side. The shell is attached to the heat shield and includes a shell hot side facing the shield cold side, a shell cold side facing away from the shield cold side, and a row of cooling holes. The trip strips run parallel to each other and all project from the shield cold side the same distance. Each projecting wall runs parallel to, and opposite of, a corresponding trip strip and projects from the shell hot side such that the distance to which each projecting wall projects is greater for projecting walls farther from the row of cooling holes, creating successive gaps between the projecting walls and corresponding trip strips that decrease from the row of cooling holes to create a convergent channel.

Abstract: A heat shield for a combustor of a gas turbine engine includes a first edge with a first set of cantilevered members and a second edge with a second set of cantilevered members.