Abstract: A combustor liner for a combustor of a gas turbine includes an outer liner extending circumferentially about a combustor centerline, and an inner liner extending circumferentially about the combustor centerline, where the outer liner and the inner liner define a combustion chamber therebetween. At least one of the outer liner and the inner liner includes a dilution flow assembly comprising, (a) an annular slot dilution opening, and (b) a dilution fence extending between an upstream side of the annular slot dilution opening to a downstream side of the annular slot dilution opening, and extending into the combustion chamber, the dilution fence including a plurality of dilution openings therethrough for providing a flow of an oxidizer through the dilution fence into the combustion chamber.

Abstract: A liner for a combustor in a gas turbine engine and a related method. The liner includes a liner body having a cold side and a hot side. The liner includes a dilution passage having a concatenated geometry extending through the liner body. The dilution passage is configured (i) to integrate a first dilution air flow flowing through the dilution passage from the cold side to the hot side and a second dilution air flow flowing through the dilution passage from the cold side to the hot side into an integrated dilution air flow, and (ii) to inject the integrated dilution air flow into a core primary combustion zone of the combustor to attain a predetermined combustion state of the combustor.

Abstract: A liner for a combustor in a gas turbine engine and a related method. The liner includes a liner body having a cold side and a hot side. The liner includes a dilution passage having a concatenated geometry extending through the liner body. The concatenated geometry has a plurality of discrete dilution holes, an annular slot, and a plurality of dilution inserts. The dilution passage is configured (i) to integrate a first dilution air flow flowing through the dilution passage from the cold side to the hot side and a second dilution air flow flowing through the dilution passage from the cold side to the hot side into an integrated dilution air flow, and (ii) to inject the integrated dilution air flow into a core primary combustion zone of the combustor to attain a predetermined combustion state of the combustor.

Abstract: A liner for a combustor in a gas turbine engine. The liner includes a liner body having a cold side and a hot side. The liner includes a dilution passage having a concatenated geometry extending through the liner body. The concatenated geometry has a plurality of discrete dilution holes, an annular slot, and a fence concatenated with the plurality of discrete dilution holes. The dilution passage is configured (i) to integrate a first dilution air flow flowing through the plurality of discrete dilution holes from the cold side to the hot side and a second dilution air flow flowing through the annular slot from the cold side to the hot side into an integrated dilution air flow, and (ii) to inject the integrated dilution air flow into a core primary combustion zone of the combustor to attain a predetermined combustion state of the combustor.

Abstract: A liner for a combustor in a gas turbine engine and a related method. The liner includes a liner body having a cold side and a hot side. The liner includes a dilution array having a plurality of dilution passages, each dilution passage of the plurality of dilution passages having a concatenated geometry repeating in a predetermined pattern and extending circumferentially around the liner body. The dilution passage integrates a first dilution air flow flowing through the dilution passage from the cold side to the hot side and a second dilution air flow flowing through the dilution passage from the cold side to the hot side into an integrated dilution air flow and injects the integrated dilution air flow into a core primary combustion zone of the combustor to attain a predetermined combustion state of the combustor. The dilution array is repeated along an axial length of the liner body.

Abstract: A combustor liner for a combustor of a gas turbine includes an outer liner having a plurality of outer liner segments, and an inner liner having a plurality of inner liner segments. Each segment of the inner liner and the outer liner includes at least one slotted dilution opening therethrough extending in a circumferential direction, and each slotted dilution opening includes a deflector wall extending radially from the respective liner into a dilution zone of a combustion chamber between the outer liner and the inner liner. The at least one slotted dilution opening may be a curved slot (either concave or convex) dilution opening, and the curved slot dilution openings for each segment of the outer liner and the inner liner may be connected so as to provide a wavy slotted dilution opening extending annularly through the outer liner and the inner liner.

Abstract: A swirler assembly includes a swirler having a primary swirler and a secondary swirler, and a flare connected the secondary swirler, and a dome disposed radially outward of the flare. A flare oxidizer flow passage is defined by a flare outer surface and a dome inner surface, and the flare oxidizer flow passage includes a swirl inducing member therewithin that induces a swirled flow of an oxidizer passing through the flare oxidizer flow passage into a combustion chamber.

Abstract: A combustor for a turbomachine engine includes a dome made of a ceramic matrix composite (CMC) material, the dome being secured within a support structure. The combustor includes an outer liner made of the CMC material, the outer liner being secured to the dome within the support structure. The combustor also includes an inner liner made of the CMC material, the inner liner being secured to the dome within the support structure.

Abstract: A combustor for a gas turbine includes a ceramic matrix composite (CMC) dome with a swirler mounting wall formed integral with the CMC dome, and a swirler assembly including a plurality of clevis dome attachment members for connecting the swirler assembly to the CMC dome. Bushings are arranged within a plurality of dome-side swirler assembly mounting openings of the swirler mounting wall, and the CMC dome is arranged within respective ones of the plurality of clevis dome attachment members of the swirler assembly. A swirler-dome connecting member is disposed through each of the clevis dome attachment members so as to mount the swirler assembly to the CMC dome.

Abstract: A combustor for a gas turbine includes a ceramic matrix composite (CMC) dome including a swirler opening therethrough with a flare interface surface surrounding the swirler opening, a swirler assembly including (a) a secondary swirler having a threaded flare attachment portion, and (b) a flare having (i) a threaded secondary swirler attachment portion, and (ii) a dome interface wall that interfaces with the flare interface surface of the CMC dome, and a swirler-dome attachment member. The flare is connected to the secondary swirler via the threaded flare attachment portion and the threaded secondary swirler attachment portion, and the swirler-dome attachment member applies a force to the CMC dome to engage the dome interface wall and the flare interface surface so as to connect the CMC dome and the swirler assembly.

Abstract: A swirler assembly includes a primary swirler having a primary swirler connecting portion, a fuel nozzle connecting portion for connecting a fuel nozzle to the primary swirler, and a primary oxidizer flow passage that extends at least partially in the longitudinal direction, and a secondary swirler having a secondary swirler connecting member, where the primary swirler connecting portion and the secondary swirler connecting member engage to connect the primary swirler and the secondary swirler in the longitudinal direction, and to permit radial movement of the primary swirler with respect to the secondary swirler.

Abstract: A combustor liner for a combustor of a gas turbine includes an outer liner and an inner liner. The outer liner has an annular outer liner slot dilution opening, with a plurality of outer liner swirl vanes disposed within the annular outer liner slot dilution opening. The inner liner of the combustor liner includes an annular inner liner slot dilution opening, and also includes a plurality of inner liner swirl vanes disposed within the annular inner liner slot dilution opening.

Abstract: A swirler assembly includes a swirler having a primary swirler and a secondary swirler, and a flare connected the secondary swirler, and a dome disposed radially outward of the flare. A flare oxidizer flow passage is defined by a flare outer surface and a dome inner surface, and the flare oxidizer flow passage includes a swirl inducing member therewithin that induces a swirled flow of an oxidizer passing through the flare oxidizer flow passage into a combustion chamber.

Abstract: A combustor liner for a combustor of a gas turbine includes an outer liner and an inner liner. At least one of the outer liner and the inner liner has an upstream liner portion and a downstream liner portion with an annular gap therebetween. At least one dilution flow assembly is arranged to bridge across the annular gap. The at least one dilution flow assembly includes an upstream liner panel and a downstream liner panel with a dilution opening provided between the upstream liner panel and the downstream liner panel, and a plurality of dilution fence members arranged within the dilution opening and extending into a combustion chamber. The plurality of dilution fence members are arranged in successive arrangement in a circumferential direction and are arranged such that successive dilution fence members are alternatingly offset in an axial direction.

Abstract: A swirler-ferrule assembly includes a radial swirler, a ferrule, a fuel nozzle, and a surface feature. The radial swirler includes a primary swirler vane having a primary air passage and a secondary swirler vane having a secondary air passage. The ferrule may be connected to the radial swirler. The surface feature may be located on the primary swirler vane and/or the ferrule. The surface feature may be configured to direct an air flow through the primary air passage away from a recirculation zone located upstream of the primary swirler vane. The surface feature has a trailing end and a distal end, and the fuel nozzle is axially aligned with the trailing end of the surface feature or is located axially downstream of the trailing end of the surface feature. The surface feature may have a plurality of grooves.

Abstract: A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a liner at least partially defining a combustion chamber of the gas turbine engine, wherein the liner comprises a looped feature.

Abstract: A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a liner at least partially defining a combustion chamber, the liner further defining a gap; and an intermediate member disposed at least partially within the gap of the liner and in fluid communication with the combustion chamber, wherein the intermediate member is retained at a relatively fixed position with respect to the liner by a support extending between the intermediate member and a case of the gas turbine engine.

Abstract: A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a forward liner segment; an aft liner segment disposed downstream from the forward liner segment relative to a direction of flow through the combustor, the forward and aft liner segments at least partially defining a combustion chamber; and an intermediate member disposed at least partially between the forward and aft liner segments and extending in the circumferential direction.

Abstract: A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a forward liner segment; and an aft liner segment disposed downstream from the forward liner segment relative to a direction of flow through the combustor, the forward and aft liner segments at least partially defining a combustion chamber, wherein the forward and aft liner segments are coupled together at a moveable interface.

Abstract: A combustor for a gas turbine engine, the gas turbine engine defining a longitudinal centerline extending in a longitudinal direction, a radial direction extending orthogonally outward from the longitudinal centerline, and a circumferential direction extending concentrically around the longitudinal centerline, the combustor including: a forward liner segment; an aft liner segment disposed downstream from the forward liner segment relative to a direction of flow through the combustor, the forward and aft liner segments at least partially defining a combustion chamber; and a fence disposed between the forward and aft liner segments, wherein the fence extends in the circumferential direction, and wherein the fence extends into the combustion chamber along the radial direction.