Abstract: A turbine engine and method for controlling nitrogen oxides present within a combustor of the turbine engine. The turbine engine having a compressor section, a combustion section, and a turbine section in serial flow arrangement along an engine centerline. The combustion section having a combustor liner having a first end, a second end, opposing the first end, and at least partially defining a combustion chamber extending between the first and second ends. A dome assembly is mounted to the combustor liner at the first end and defines a dome inlet of the combustion chamber. There are multiple sets of dilution holes including a first set of dilution holes provided in the combustor liner downstream from the dome inlet and a second set of dilution holes provided in the combustor liner between the first set of dilution holes and the dome inlet.

Abstract: An engine can utilize a combustor to combust fuel to drive the engine. A fuel nozzle assembly can supply fuel to the combustor for combustion or ignition of the fuel. The fuel nozzle assembly can include a swirler and a fuel nozzle to supply a mixture of fuel and air for combustion, which can supply a primary fuel supply and a secondary fuel supply. Increasing efficiency and reducing emission require the use of alternative fuels, which combust at higher temperatures or burn at faster burn speeds than traditional fuels, requiring improved fuel introduction without the occurrence of flame holding or flashback.

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 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 turbine engine is provided, the combustor includes an outer liner, an inner liner and a dome that together define a combustion chamber; a diffuser positioned upstream of the combustion chamber, the diffuser being configured to receive air flow from a compressor section and to provide a flow of compressed air to the combustion chamber; and an outer cowl and an inner cowl located upstream of the combustion chamber, the outer cowl and the inner cowl being configured to direct a portion of air flow from the diffuser to the combustion chamber. The diffuser is configured to output air flow having an amount of air pressure maximized at a center of the air flow so as to optimize total air pressure fed to the combustion chamber through the dome.

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 pilot fuel nozzle assembly includes a fuel nozzle, a swirler, and a vented pilot venturi. The vented pilot venturi has an annular wall with an oxidizer flow passage therein. An expansion flow surface portion of the venturi has a larger diameter at an outlet than at a throat of the venturi. A plurality of venturi oxidizer outlet ports extend through the expansion flow surface to the oxidizer flow passage within the annular wall to provide a flow of oxidizer through the venturi wall into a mixing cavity of the venturi and at an outlet end of the venturi. The oxidizer outlet ports are circumferentially spaced about a circumference of the expansion flow surface, and may be arranged in a plurality of rows. The oxidizer outlet ports may be angled with respect to the expansion flow surface and may angled circumferentially in a co-swirl direction with the swirler.

Abstract: A combustor for a turbine engine is provided, the combustor includes an outer liner, an inner liner and a dome that together define a combustion chamber; a diffuser positioned upstream of the combustion chamber, the diffuser being configured to receive air flow from a compressor section and to provide a flow of compressed air to the combustion chamber; and an outer cowl and an inner cowl located upstream of the combustion chamber, the outer cowl and the inner cowl being configured to direct a portion of air flow from the diffuser to the combustion chamber. The diffuser is configured to output air flow having an amount of air pressure maximized at a center of the air flow so as to optimize total air pressure fed to the combustion chamber through the dome.

Abstract: An engine may include an integrated diffuser-bleed baffle assembly, a diffuser, and a bleed port. The integrated diffuser-bleed baffle assembly fluidly coupled between the diffuser and the bleed port. The integrated diffuser-bleed baffle assembly is configured to flow a boundary layer flow from the diffuser to the bleed port. A baffle hole may be included in the integrated diffuser-bleed baffle assembly such that the assembly may function to dampen acoustic instabilities in the engine.

Abstract: An engine may include an integrated diffuser-bleed baffle assembly, a diffuser, and a bleed port. The integrated diffuser-bleed baffle assembly fluidly coupled between the diffuser and the bleed port. The integrated diffuser-bleed baffle assembly is configured to flow a boundary layer flow from the diffuser to the bleed port. A baffle hole may be included in the integrated diffuser-bleed baffle assembly such that the assembly may function to dampen acoustic instabilities in the engine.

Abstract: A venturi for a swirler assembly of a combustor for a gas turbine engine. The venturi has an annular wall having internal diameter of a forward end of the annular wall that is larger than an internal diameter of a middle portion of the annular wall, which is smaller than an internal diameter of an aft end of the annular wall. The internal surface of the annular wall has a plurality of grooves extending in a longitudinal direction from the forward end of the annular wall to the aft end of the annular wall. The plurality of grooves are angled with respect to a centerline axis of the annular wall and may be spiraled about the internal surface of the annular wall.

Abstract: A swirler with integrated damper may include a primary swirler vane having a primary air passage, a secondary swirler vane having a secondary air passage, and a damper within the primary swirler vane, the secondary swirler vane, or both the primary swirler vane and the secondary swirler vane. The damper may include a series of cavities. The damper is configured to absorb one or more frequencies present in an air flow through the primary air passage, the secondary air passage, or both the primary air passage and the secondary air passage.

Abstract: An annular gas turbine engine combustor liner has a single piece annular shell circumscribed about a first axis of revolution. The shell has a hot side and a cold side, an annular dilution hole section, and a plurality of asymmetrical cylindrical dilution holes extending through the section. Pockets are circumferentially interdigitated between the dilution holes and extend into the shell from the cold side such that the pockets form dilution hole bosses between the pockets through which the dilution holes extend. Each of the dilution holes is circumscribed about a second axis of revolution which is not parallel to the first axis of revolution. The liner may further include forward and aft annular cooling nuggets located forward and aft of the dilution hole section and having annular film cooling slots which are open in the axially aftwardly direction.

Abstract: An annular gas turbine engine combustor liner has a single piece annular shell circumscribed about a first axis of revolution. The shell has a hot side and a cold side, an annular dilution hole section, and a plurality of asymmetrical cylindrical dilution holes extending through the section. Pockets are circumferentially interdigitated between the dilution holes and extend into the shell from the cold side such that the pockets form dilution hole bosses between the pockets through which the dilution holes extend. Each of the dilution holes is circumscribed about a second axis of revolution which is not parallel to the first axis of revolution. The liner may further include forward and aft annular cooling nuggets located forward and aft of the dilution hole section and having annular film cooling slots which are open in the axially aftwardly direction.