Turbine engine fuel injector with non-circular nozzle passage

An apparatus for a turbine engine, comprising:

• a combustor wall at least partially forming a combustion chamber;
• a fuel conduit including a fuel conduit centerline extending circumferentially about a centerline of the turbine engine when viewed in a cross-sectional reference plane of the turbine engine orthogonal to the centerline of the turbine engine;
• a plurality of fuel nozzles arranged circumferentially about the centerline of the turbine engine, each of the plurality of fuel nozzles fluidly coupled with the fuel conduit, the plurality of fuel nozzles comprising a first fuel nozzle, the first fuel nozzle comprising a nozzle tube, a nozzle support gusset, a nozzle passage and a nozzle orifice, the first fuel nozzle configured to receive fuel from the fuel conduit within the nozzle passage, the nozzle tube extending from the fuel conduit to the nozzle orifice at a distal end of the nozzle tube, and the first fuel nozzle configured to direct the fuel out of the nozzle passage through the nozzle orifice;
• an apparatus base radially spaced apart from the combustor wall relative to the centerline of the turbine engine and supporting the fuel conduit and the first fuel nozzle;
• and
• an air tube comprising an air passage, the air tube extending radially along a central axis from the combustor wall into the combustion chamber when viewed in a second reference plane parallel with and including the centerline of the turbine engine, and the nozzle tube extending radially relative to the fuel conduit centerline into the air passage to the nozzle orifice;
• the nozzle support gusset extending between the apparatus base and the nozzle tube:
• the nozzle passage extending longitudinally along a centerline of the nozzle passage from the fuel conduit, through the first fuel nozzle, to the nozzle orifice when viewed in the second reference plane parallel with and including the centerline of the turbine engine, and the centerline of the nozzle passage angularly offset from the centerline of the turbine engine when viewed in the second reference plane parallel with and including the centerline of the turbine engine such that the nozzle tube extends upstream towards a compressor section of the turbine engine;
• and
• the nozzle passage having a solid polygonal cross-sectional geometry as the nozzle passage extends longitudinally along the centerline of the nozzle passage from the fuel conduit to the nozzle orifice.

An apparatus for a turbine engine, comprising:

• a fuel conduit comprising a fuel supply passage;
• a fuel nozzle fluidly coupled with the fuel conduit, the fuel nozzle comprising a nozzle support gusset, a nozzle passage and a nozzle orifice, the fuel nozzle extending along a nozzle tube from the fuel conduit to a distal end of the fuel nozzle, and the nozzle passage extending longitudinally along a centerline within the fuel nozzle to the nozzle orifice;
• an apparatus base supporting the fuel conduit and the fuel nozzle;
• an air tube including an air passage, the nozzle tube projecting into the air passage to the distal end;
• and
• a combustor wall at least partially forming a combustion chamber, the air tube connected to the combustor wall and projecting radially from the combustor wall into the combustion chamber relative to a centerline of the turbine engine, and the combustor wall being radially spaced apart from the apparatus base relative to the centerline of the turbine engine;
• wherein the nozzle passage has a solid polygonal cross-sectional geometry at the nozzle orifice and the fuel nozzle is configured to direct fuel out of the nozzle passage through the nozzle orifice into the air passage;
• wherein the nozzle support gusset extends between the apparatus base and an upstream side of the nozzle tube:
• wherein a first portion of the nozzle passage includes convergent sidewall surfaces tapering inward towards the centerline of the nozzle passage as the nozzle passage extends longitudinally along the centerline of the nozzle passage, and a slope of the convergent sidewall surfaces has a rise to run ratio equal to about 0.6;
• wherein the fuel nozzle is configured to direct at least a portion of the fuel against a surface of the air tube;
• and
• wherein the centerline of the fuel nozzle is angularly offset from a centerline of the air tube when viewed in a reference plane perpendicular to the centerline of the turbine engine.