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 multi-stage vortex mixer for a combustor of a gas turbine engine includes a vortex amplifier stage in communication with a first stage amplifier, the vortex amplifier stage in communication with a dilution hole.

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 shell for a combustor liner includes a cold side, a hot side, a row of cooling holes and a jet wall. The jet wall projects from the hot side for creating a wall shear jet of increased velocity cooling flow in a tangential direction away from the row of cooling holes and along an adjacent heat shield cold side wall.

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 support shell with a convex profile which faces the heat shield. A further embodiment of the foregoing embodiment of the present disclosure is where the convex profile is defined by a hyperbolic cosine function. A further embodiment of any of the foregoing embodiments of the present disclosure is where the convex profile provides an approximate 4.5 inlet-to-exit area ratio. A further embodiment of any of the foregoing embodiments, of the present disclosure wherein the convex profile provides a flow acceleration toward approximately 0.5 Mach towards an end of a convergent section.

Abstract: A combustor having U-shaped cooling channels is disclosed. The combustor may include a shell having an impingement hole, a liner spaced from the shell and having an effusion hole; a first partition spanning between the shell and the liner, a second partition spaced from the first partition and spanning between the shell and the liner; and a U-shaped channel defined between the shell and the liner and defined in part by the wall, the channel having upstream and downstream ends both adjacent the first partition and separated by the wall, wherein the impingement hole communicates with the upstream end and the effusion hole communicates with the downstream end.

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 combustor liner which reduces cooling flow to a combustion chamber and augments pressure drop split between impingement holes and effusion holes is disclosed. The combustor liner may further include accelerating channels, trip strips, pedestals, and cone-shaped effusion holes to provide further cooling of the liner. The combustor liner may reduce NOx production and the temperature of the combustion chamber of a gas turbine engine or the like.

Abstract: A wall assembly of a combustor for a turbine engine includes a shell and a liner and a sealing interface structure. A cooling cavity is defined between the shell and liner and the sealing interface structure functions to minimize leakage out of the cooling cavity. The sealing interface structure may have a sealing width that is equal to or greater than twice the thickness of the liner and/or may include a chamber located between longitudinal segments of the sealing interface structure that inhibits leakage.

Abstract: A grommet assembly and method of design to enhance the flow coefficient, thereof, includes a shell having a first side and an opposite second side, and a chamfered grommet projecting through the shell along a centerline and including an annular first end surface spaced outward from the first side and a conical face spanning axially and radially inward from the annular first end surface and axially beyond the second side. The assembly may further include a panel spaced from the shell and defining a cooling cavity therebetween with the conical surface defining at least in-part a hole in fluid communication through the shell and panel and isolated from the cooling cavity. A plurality of cooling channels in the grommet are in fluid communication with the cooling cavity and communicate through the panel. The combination of the conical face and the cooling channels improve the discharge coefficient of the grommet while enhancing grommet cooling.

Abstract: A combustor component of a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a heat shield panel. The heat shield panel defines a bend and a microcircuit flow path within a thickness of the heat shield panel. The microcircuit flow path includes an inlet and an outlet radially outward of the inlet. The microcircuit flow path at the bend is positioned radially between the inlet and the outlet, and the microcircuit flow path follows the bend. A method of cooling a combustor of a gas turbine engine is also disclosed.

Abstract: A combustor component of a gas turbine engine including a heat shield panel. The heat shield panel a defines microcircuit flow path within a thickness of said heat shield panel.

Abstract: A combustor component of a gas turbine engine includes a refractory metal core (RMC) microcircuit for self-regulating a cooling flow.

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

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

Abstract: A combustor liner which reduces cooling flow to a combustion chamber and augments pressure drop split between impingement holes and effusion holes is disclosed. The combustor liner may further include accelerating channels, trip strips, pedestals, and cone-shaped effusion holes to provide further cooling of the liner. The combustor liner may reduce NOx production and the temperature of the combustion chamber of a gas turbine engine or the like.

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.

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.