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; 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 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 combustor for a turbine engine. The combustor includes a fuel injector and a fluid injection system in fluid communication with the combustor. The fuel injector includes a mixer assembly with a pilot mixer and a main mixer. The pilot mixer and the main mixer operate during high power operation of the turbine engine. The fluid injection system injects a fluid into the combustor during high power operation of the turbine engine. The fluid is shut off during low power operation of the turbine engine and during mid-level power operation of the 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 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 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 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.

Abstract: A mixing assembly includes a pilot mixer and a main mixer. The main mixer includes a main housing, a fuel manifold positioned between a pilot housing and the main housing, a mixer foot extending outward from a forward end of the main housing, and a main swirler body including a plurality of vanes. The main swirler body surrounds the main housing and defines an annular mixing channel between the main housing and the main swirler body, and being coupled to the mixer foot. An annular main fuel ring extends axially aft from the mixer foot into the annular mixing channel. The mixer foot extends farther from the main housing than the main fuel ring. At least one of the mixer foot or the main fuel ring includes a plurality of fuel injection ports positioned to discharge fuel into a central portion of the annular mixing channel.

Abstract: A gas turbine engine including a combustor, a plurality of fuel nozzles, and at least one fuel manifold. The combustor includes a combustion chamber. The plurality of fuel injects fuel into the combustion chamber of the combustor. The gas turbine engine may include a first fuel circuit and a second fuel circuit. The first fuel circuit includes a first fuel manifold fluidly connected to at least one fuel nozzle of the plurality of fuel nozzles to distribute the fuel to the at least one fuel nozzle at a first temperature. The second fuel circuit includes a second fuel manifold fluidly connected to at least one fuel nozzle of the plurality of fuel nozzles to distribute the fuel to the at least one fuel nozzle at a second. The second temperature is less than the first temperature.

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 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 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.

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 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 mixing assembly for a combustor, including: a pilot mixer and a pilot fuel nozzle; a main mixer including: a main housing surrounding the pilot mixer; a fuel manifold positioned between the pilot housing and the main housing; a mixer foot extending outward from a forward end of the main housing; a main swirler body surrounding the main housing such that an annular mixing channel is defined between the main housing and the main swirler body, and coupled to the mixer foot; and a main fuel ring in the mixing channel downstream of the mixer foot connected to the main housing by main fuel vanes, at least one of the main fuel ring and the main fuel vanes including fuel injection ports positioned to discharge fuel into a central portion of the mixing channel, wherein the mixer foot extends farther from the main housing than the main fuel ring.

Abstract: A swirler assembly includes a swirler having primary swirler with a primary swirler venturi, a swirler ferrule plate connected to the primary swirler, and a fuel nozzle disposed in the swirler ferrule plate. The swirler ferrule plate has an aft wall, an annular conical wall, and an annular cavity wall that together form an annular cavity. The annular cavity includes a plurality of inlet orifices, and at least one outlet orifice. A flow of oxidizer through the plurality of inlet orifices into the annular cavity incurs a first pressure drop from a first pressure of a pressure plenum to a second pressure lower than the first pressure, and a flow of the oxidizer from the annular cavity through the at least one outlet orifice into the primary swirler venturi incurs a second pressure drop from the second pressure to a third pressure lower than the second pressure.

Abstract: A swirler assembly includes a swirler having a primary swirler with a primary swirler venturi, a swirler ferrule plate connected upstream to the primary swirler, and a fuel nozzle disposed in the swirler ferrule plate. The swirler ferrule plate has an annular pressure drop cavity with oxidizer inlet orifices in fluid communication with the swirler, and at least one outlet orifice in fluid communication with the primary swirler venturi. A second flow of oxidizer to the swirler incurs a first pressure drop, a third flow of the oxidizer from the swirler to the annular pressure drop cavity incurs a second pressure drop, and a fourth flow of the oxidizer from the annular pressure drop cavity to the primary swirler venturi incurs a third pressure drop.

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 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 gas turbine engine may include a combustion section having a fuel nozzle, a swirler, and a ferrule configured to mount and center the fuel nozzle with the swirler. The combustion section may further include a damper on a cold side of the combustion section. The damper may have an acoustic cavity, a damper neck, and a cavity feed hole. The damper may operate as Helmholtz cavity to absorb a hydrodynamic or acoustic instability present in a region within the swirler.