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 including a dilution slot in fluid communication with the combustion chamber; and one or more guides configured to redirect cooling medium into the dilution slot.

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.

Abstract: A system and method for operating a combustion system comprising a fuel nozzle defining at least one main fuel circuit and at least one pilot fuel circuit is generally provided. The method includes determining an overall flow of fuel, the overall flow of fuel defining a sum total fuel through the main fuel circuit and the pilot fuel circuit; determining a plurality of ranges of ratios of main fuel flow through the main fuel circuit versus pilot fuel flow through the pilot circuit from the overall flow of fuel, wherein each range of ratios is based on a combustion criterion different from one another; determining a resultant range of ratios of main fuel flow versus pilot fuel flow based on a hierarchy of combustion criteria, wherein the hierarchy of combustion criteria provides a priority ranking of the combustion criterion; and flowing the overall flow of fuel to the main fuel circuit and the pilot fuel circuit based on the resultant range of ratios of main fuel flow versus pilot fuel flow.

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 system and method for operating a combustion system comprising a fuel nozzle defining at least one main fuel circuit and at least one pilot fuel circuit is generally provided. The method includes determining an overall flow of fuel, the overall flow of fuel defining a sum total fuel through the main fuel circuit and the pilot fuel circuit; determining a plurality of ranges of ratios of main fuel flow through the main fuel circuit versus pilot fuel flow through the pilot circuit from the overall flow of fuel, wherein each range of ratios is based on a combustion criterion different from one another; determining a resultant range of ratios of main fuel flow versus pilot fuel flow based on a hierarchy of combustion criteria, wherein the hierarchy of combustion criteria provides a priority ranking of the combustion criterion; and flowing the overall flow of fuel to the main fuel circuit and the pilot fuel circuit based on the resultant range of ratios of main fuel flow versus pilot fuel flow.

Abstract: A laser ignition system for an internal combustion engine, and more specifically a gas turbine engine, is provided. The system including a laser light source configured to generate a laser beam, an ignition port configured to provide optimized optical access of the laser beam to a combustion chamber and an optical beam guidance component disposed between the laser light source and the ignition port. The optical beam guidance component is configured to include optimized optic components to transmit the laser beam to irradiate on a fuel mixture supplied into the combustion chamber to generate a combustor flame in a flame region. A method for igniting a fuel mixture in an internal combustion engine is also presented.

Abstract: A laser ignition system for an internal combustion engine, and more specifically a gas turbine engine, is provided. The system including a laser light source configured to generate a laser beam, an ignition port configured to provide optimized optical access of the laser beam to a combustion chamber and an optical beam guidance component disposed between the laser light source and the ignition port. The optical beam guidance component is configured to include optimized optic components to transmit the laser beam to irradiate on a fuel mixture supplied into the combustion chamber to generate a combustor flame in a flame region. A method for igniting a fuel mixture in an internal combustion engine is also presented.