Abstract: A System for Aerodynamic Premixer for Reduced Emissions comprising a premixer is generally cylindrical in form and defined by the relationship in physical space between a first ring, a second ring, and a plurality of radial vanes. The first and second rings are found to be generally equidistant, one from the other, at all points along their facing surfaces. Radial vanes connect the first ring to the second ring and thereby form the premixer.

Abstract: A System for Aerodynamic Premixer for Reduced Emissions comprising a premixer is generally cylindrical in form and defined by the relationship in physical space between a first ring, a second ring, and a plurality of radial vanes. The first and second rings are found to be generally equidistant, one from the other, at all points along their facing surfaces. Radial vanes connect the first ring to the second ring and thereby form the premixer.

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 System for Aerodynamic Premixer for Reduced Emissions comprising a premixer is generally cylindrical in form and defined by the relationship in physical space between a first ring, a second ring, and a plurality of radial vanes. The first and second rings are found to be generally equidistant, one from the other, at all points along their facing surfaces. Radial vanes connect the first ring to the second ring and thereby form the premixer.

Abstract: A fuel nozzle for a combustor assembly of a heat engine is generally provided. The fuel nozzle defines a fuel nozzle centerline therethrough and a radial direction extended from the nozzle centerline. The fuel nozzle includes a first wall extended along the radial direction. The first wall defines a first cooling passage and a second cooling passage each extended therethrough. The first cooling passage is defined at a first acute angle relative to the nozzle centerline. The second cooling passage is defined at a second acute angle 90 degrees or less relative to the first acute angle.

Abstract: A System for Aerodynamic Premixer for Reduced Emissions comprising a premixer is generally cylindrical in form and defined by the relationship in physical space between a first ring, a second ring, and a plurality of radial vanes. The first and second rings are found to be generally equidistant, one from the other, at all points along their facing surfaces. Radial vanes connect the first ring to the second ring and thereby form the premixer.

Abstract: A fuel nozzle apparatus for a gas turbine engine includes: a first pilot fuel injector disposed on a centerline axis of the fuel nozzle which defines a direction of air flow through the fuel nozzle, the first pilot fuel injector being of a pressure atomizing type; an annular second pilot fuel injector at least partially surrounding the first pilot fuel injector, the second pilot fuel injector being of an air blast type and having a fuel outlet disposed axially downstream and radially outboard of the first pilot fuel injector; an annular venturi surrounding the first and second pilot fuel injectors, the venturi including a throat of minimum diameter; an array of inner swirl vanes extending between the first pilot fuel injector and the second pilot fuel injector; and an array of outer swirl vanes extending between the second pilot fuel injector and the venturi.

Abstract: A System for Aerodynamic Premixer for Reduced Emissions comprising a premixer is generally cylindrical in form and defined by the relationship in physical space between a first ring, a second ring, and a plurality of radial vanes. The first and second rings are found to be generally equidistant, one from the other, at all points along their facing surfaces. Radial vanes connect the first ring to the second ring and thereby form the premixer.

Abstract: A combustion section for gas turbine engine includes an inner liner, an outer liner, and a dome attached to the inner liner and to the outer liner. A heat shield is attached to the dome, with the dome, the heat shield, or both define an opening. The combustion section also includes a fuel nozzle extending at least partially into the opening, the fuel nozzle defining a fuel nozzle axis and a radial direction relative to the fuel nozzle axis. The fuel nozzle defines an aft end, the aft end of the fuel nozzle positioned forward of the aft surface of the heat shield along the fuel nozzle axis such that the aft end of the fuel nozzle defines a minimum separation from the aft surface of the heat shield along the fuel nozzle axis of at least about 0.15 inches.

Abstract: The present disclosure is directed to a combustor assembly for a gas turbine engine comprising a bulkhead wall extended generally along a radial direction relative to a combustor centerline and a mixer disposed concentric to the nozzle centerline. The bulkhead wall defines one or more openings therethrough adjacent to a combustion chamber. Each of the openings defines a nozzle centerline extended along a lengthwise direction. The mixer comprises a retaining wall extended along a radial direction relative to the nozzle centerline and coupled to the upstream side of the bulkhead wall. The mixer further comprises an annular shroud extended at least partially through the one or more openings of the bulkhead wall. The mixer defines a cooling passage through the retaining wall and the annular shroud. The cooling passage defines an inlet opening through the retaining wall and an outlet opening through the annular shroud.

Abstract: A fuel nozzle for a gas turbine engine includes an outer body extending generally along a centerline axis and defining a plurality of openings in an exterior surface. The fuel nozzle additionally includes a main injection ring disposed at least partially inside the outer body, the main injection ring including a fuel post extending into or through one of the plurality of openings of the outer body. The fuel post defines a spray well and a main fuel orifice, the spray well defining a bottom surface, a side wall, and a taper in the bottom surface extending from the main fuel orifice towards the side wall.

Abstract: A combustor assembly for a gas turbine engine includes a dome and a deflector positioned adjacent to the dome. One or both of the dome and the deflector define an opening, a component axis extending through the opening, and a radial direction relative to the component axis. The combustor assembly also includes a retainer having an outer member contacting the dome, the deflector or both. The outer member of the retainer defines at least in part a retainer cavity inward of the outer member along the radial direction. The dome, the deflector, or both define a plurality of cooling holes for providing a cooling airflow from the retainer cavity to the opening.

Abstract: The present disclosure is directed to a fuel nozzle assembly for a gas turbine engine. The fuel nozzle assembly includes a centerbody extended along a nozzle centerline axis and generally concentric thereto and an outer sleeve surrounding the centerbody and extended along the nozzle centerline axis and generally concentric thereto. The centerbody defines an outer wall extended at least partially along the nozzle centerline axis in which the centerbody defines a first fuel passage therewithin and one or more first fuel exit openings through the outer wall. Each first fuel exit opening is discrete from another along the outer wall. The outer sleeve and centerbody together define a first air passage therebetween. The first fuel passage and the first fuel exit opening are in fluid communication with the first air passage.

Abstract: A fuel nozzle for a gas turbine engine includes an outer body defining a plurality of openings in an exterior surface. The fuel nozzle also includes a main injection ring disposed at least partially inside the outer body. The main injection ring includes a plurality of fuel posts extending into or through the plurality of openings of the outer body. The plurality of fuel posts include an LP fuel post defining a main fuel orifice, a top surface, and a scarf, the scarf of the LP fuel post extending in the top surface in a first direction relative to the centerline axis away from the main fuel orifice. The plurality of fuel posts also include an HP fuel post defining a main fuel orifice, a top surface, and a scarf, the scarf of the HP fuel post extending in the top surface f in a second direction relative to the centerline axis away from the main fuel orifice, the second direction being at least ninety degrees different than the first direction.

Abstract: A fuel nozzle apparatus for a gas turbine engine includes: an annular outer body extending parallel to a centerline axis, with an exterior surface having a plurality of openings; a main injection ring disposed inside the outer body and including: a circumferential main fuel gallery; and a plurality of main fuel orifices, each main fuel orifice communicating with the main fuel gallery and aligned with one of the openings; a venturi disposed inside the main injection ring; an annular splitter disposed inside the venturi, and including inner and outer walls spaced-apart from each other defining a splitter cavity, and a plurality of discharge holes communicating with the splitter cavity; an array of outer swirl vanes extending between the venturi and the splitter; a pilot fuel injector within the splitter; and an array of inner swirl vanes extending between the splitter and the pilot fuel injector.

Abstract: A fuel nozzle for a combustor assembly of a heat engine is generally provided. The fuel nozzle defines a fuel nozzle centerline therethrough and a radial direction extended from the nozzle centerline. The fuel nozzle includes a first wall extended along the radial direction. The first wall defines a first cooling passage and a second cooling passage each extended therethrough. The first cooling passage is defined at a first acute angle relative to the nozzle centerline. The second cooling passage is defined at a second acute angle 90 degrees or less relative to the first acute angle.

Abstract: A fuel injection system for use in a combustor of a turbine engine includes a mixer assembly including a mixer housing and a fuel nozzle assembly positioned radially inward of the mixer housing. The fuel nozzle assembly includes a substantially annular fuel injection housing and a substantially annular main fuel injector coupled to the fuel injection housing. The main fuel injector includes a body, a fuel delivery passage defined in the body, a swirl chamber defined in the body downstream of the fuel delivery passage, and a plurality of circumferentially-spaced fuel metering slots defined in the body and coupled in flow communication with and between the fuel delivery passage and the swirl chamber.

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: The present disclosure is directed to a combustor assembly for a gas turbine engine. The combustor assembly includes a fuel nozzle comprising a centerbody extended along a lengthwise direction, wherein the fuel nozzle defines a nozzle centerline extended through the centerbody along the lengthwise direction. The centerbody defines a plurality of exit openings in circumferential arrangement relative to the nozzle centerline. The plurality of exit openings defines two or more locations different from one another on the centerbody along the lengthwise direction.

Abstract: The present disclosure is directed to a combustor assembly for a gas turbine engine comprising a bulkhead wall extended generally along a radial direction relative to a combustor centerline and a mixer disposed concentric to the nozzle centerline. The bulkhead wall defines one or more openings therethrough adjacent to a combustion chamber. Each of the openings defines a nozzle centerline extended along a lengthwise direction. The mixer comprises a retaining wall extended along a radial direction relative to the nozzle centerline and coupled to the upstream side of the bulkhead wall. The mixer further comprises an annular shroud extended at least partially through the one or more openings of the bulkhead wall. The mixer defines a cooling passage through the retaining wall and the annular shroud. The cooling passage defines an inlet opening through the retaining wall and an outlet opening through the annular shroud.