A pilot burner for a combustor includes an inner conduit configured to deliver a fuel, and an outer conduit concentric with the inner conduit and configured to deliver air. An inner wall defines an inner plenum, and a partition wall is radially outward of the inner wall and defines an intermediate plenum with at least a portion of the inner wall. Exit passages fluidly couple the inner plenum to the intermediate plenum. An outer wall defines an outer plenum with at least a portion of the partition wall. A crossover section includes passages fluidly coupling the inner conduit to the outer plenum, and passages fluidly coupling the outer conduit to the inner plenum. An end plate includes openings to direct fuel, air for combustion, and air for cooling from the respective plenums.

A combustor liner for a gas turbine includes a liner that at least partially defines a combustion chamber, and that a plurality of dilution openings therethrough. Each dilution opening includes an outer wall defining an outer perimeter of the dilution opening and defining a dilution opening centerline axis through the dilution opening. A plurality of swirl vanes extend from the outer wall into a dilution airflow passage that extends through the dilution opening. Each of the plurality of swirl vanes extends from the outer wall into the dilution airflow passage at a respective swirl vane angle with respect to the outer wall. The plurality of swirl vanes are arranged in a successive arrangement about the outer wall, and successive respective ones of the plurality of swirl vanes extend from the outer wall at a different swirl vane angle.

A combustor liner for a gas turbine includes a liner that at least partially defines a combustion chamber, and that a plurality of dilution openings therethrough. Each dilution opening includes an outer wall defining an outer perimeter of the dilution opening and defining a dilution opening centerline axis through the dilution opening. A plurality of swirl vanes extend from the outer wall into a dilution airflow passage that extends through the dilution opening. Each of the plurality of swirl vanes extends from the outer wall into the dilution airflow passage at a respective swirl vane angle with respect to the outer wall. The plurality of swirl vanes are arranged in a successive arrangement about the outer wall, and successive respective ones of the plurality of swirl vanes extend from the outer wall at a different swirl vane angle.

A fuel nozzle device for a gas turbine engine includes a nozzle body extending in an axial direction and projecting into the combustion chamber, and swirler including a tubular portion concentrically surrounding the nozzle body, a deflector formed as a hollow shell concentrically surrounding the tubular portion, and a plurality of swirl vanes extending between the tubular portion and the deflector. Each swirl vane extends along a logarithmic spiral around an axial center of the nozzle body between the tubular portion and the deflector with a certain meridian crossing angle α and a certain twist angle Φ such that the outer peripheral inclination angle θ is smaller than the stacking angle β of the swirl vane when the swirl vane is manufactured by an additive manufacturing process with a front end up orientation.

A fuel nozzle device for a gas turbine engine includes a nozzle body extending in an axial direction and projecting into the combustion chamber, and swirler including a tubular portion concentrically surrounding the nozzle body, a deflector formed as a hollow shell concentrically surrounding the tubular portion, and a plurality of swirl vanes extending between the tubular portion and the deflector. Each swirl vane extends along a logarithmic spiral around an axial center of the nozzle body between the tubular portion and the deflector with a certain meridian crossing angle α and a certain twist angle Φ such that the outer peripheral inclination angle θ is smaller than the stacking angle β of the swirl vane when the swirl vane is manufactured by an additive manufacturing process with a front end up orientation.

A fuel nozzle includes a center body that extends from an upstream end to a downstream end. A nozzle tip, which is fixedly coupled to the downstream end of the center body, includes an outer annular wall that has a forward end and an aft end. A solid aft wall is disposed at the aft end of the outer annular wall. The solid aft wall and the outer annular wall at least partially define a plenum. A nozzle portion, which is disposed within the plenum, includes an inner annular wall and an impingement wall. The impingement wall is spaced apart from the solid aft wall and defines a plurality of apertures configured to direct fluid to impinge upon the solid aft wall. A plurality of outlets is defined in the outer annular wall forward of the solid aft wall to direct post-impingement fluid from the nozzle tip.

A fuel injection system for a gas turbine engine includes a first plurality of fuel nozzles arrayed in a circular pattern. Each of the nozzles in the first plurality of fuel nozzles includes a first airflow area defined therethrough. A second plurality of fuel nozzles radially inward from the first plurality of fuel nozzles. Each of the nozzles in the second plurality of fuel nozzles includes a second airflow area defined therethrough. The first airflow area is larger than the second airflow area. A third plurality of fuel nozzles can be radially inward from the second plurality of fuel nozzles. Each of the nozzles in the third plurality of fuel nozzles can include a third airflow area defined therethrough. The second airflow area can be larger than the third airflow area.

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.

A fuel and air injection handling system for a rotating detonation engine is provided. The system includes a compressor configured to compress air received via a compressor inlet and configured to output the air that is compressed as swirling, compressed air through a compressor outlet. The system also includes an annular rotating detonation combustor fluidly coupled with the compressor outlet. The annular rotating detonation combustor has a detonation cavity that extends around an annular axis, the annular rotating detonation combustor configured to combust the compressed air from the compressor in detonations that rotate within the detonation cavity around the annular axis of the annular rotating detonation combustor. The annular rotating detonation combustor is fluidly and directly coupled with the compressor outlet.

A combustor may comprise an outer combustor panel and an inner combustor panel radially inward of the outer combustor panel. A bulkhead panel may extend radially between the outer combustor panel and the inner combustor panel. An outer spacer may be located between an outer flange of the bulkhead panel and the outer combustor panel. An inner spacer may be located between an inner flange of the bulkhead panel and the inner combustor panel.