A fuel nozzle, a fuel nozzle module having the same, and a combustor are provided. The fuel nozzle includes a nozzle cylinder having a space through which fuel flows and a plurality of fuel holes through which the fuel flows in a surface, a shroud spaced apart from the nozzle cylinder and formed to surround the nozzle cylinder in a longitudinal direction of the nozzle cylinder, and a mixing flow path formed between the nozzle cylinder and the shroud to mix the fuel supplied through the plurality of fuel holes and compressed air supplied from a compressor.

The present disclosure is directed to a fuel injector assembly of a gas turbine engine, including a centerbody, an outer sleeve surrounding the centerbody, an inner sleeve disposed between the centerbody and at least a section of the outer sleeve, and at least one axially oriented vane defined between the centerbody and a section of the inner sleeve. A section of the outer sleeve includes at least one radially oriented vane. A portion of the inner sleeve includes a contoured shroud and defines at least one fuel injection port disposed upstream of the contoured shroud. The axially oriented vane is disposed upstream of the fuel injection port.

An air mixer includes an annular body defining a center axis. A plurality of slots are defined in the annular body circumferentially spaced apart from one another. Each slot defines a respective center injection axis extending from an outer surface of the annular body to an inner surface of the annular body. Each respective center injection axis is parallel to a respective plane bisecting the annular body. At least one of the slots is intersected by the respective bisecting plane parallel to its respective center injection axis.

An air mixer includes an annular body defining a center axis. A plurality of slots are defined in the annular body circumferentially spaced apart from one another. Each slot defines a respective center injection axis extending from an outer surface of the annular body to an inner surface of the annular body. Each respective center injection axis is parallel to a respective plane bisecting the annular body. At least one of the slots is intersected by the respective bisecting plane parallel to its respective center injection axis.

A gas turbine facility of an embodiment includes: a combustor casing; a combustor provided in the combustor casing; a cylinder surrounding a periphery of the combustor and dividing a space between the combustor casing and the combustor; a turbine rotated by combustion gas exhausted from the combustor; a heat exchanger cooling the combustion gas exhausted from the turbine; a pipe through which a part of the combustion gas cooled in the heat exchanger passes in the heat exchanger to be heated, the pipe guiding the combustion gas heated in the heat exchanger into the cylinder; and a pipe guiding another part of the combustion gas cooled in the heat exchanger to a space between the combustor casing and the cylinder.

A free-vortex combustor is disclosed that generates vortices which: enhance fuel air mixing, recirculate the air, provide cooling for the combustor walls, and provide low emissions and a substantially uniform exit temperature profile. The combustor is provided fuel or fuel and air through a fuel-injector which atomizes the fuel. A first air swirler couples to the fuel-injector with a prechamber wall abutting the first swirler. A second swirler abuts a downstream end of the prechamber wall. And, a main chamber abuts the second swirler. Each of the first and second swirlers have features that cause the flow to create a vortex in the prechamber and main chamber, respectively. The features creating the swirl are blades or angled orifices. The vortex causes a pressure depression along the centerline and causes backflow along the centerline that improves mixing and improves cooling.

A free-vortex combustor is disclosed that generates vortices which: enhance fuel air mixing, recirculate the air, provide cooling for the combustor walls, and provide low emissions and a substantially uniform exit temperature profile. The combustor is provided fuel or fuel and air through a fuel-injector which atomizes the fuel. A first air swirler couples to the fuel-injector with a prechamber wall abutting the first swirler. A second swirler abuts a downstream end of the prechamber wall. And, a main chamber abuts the second swirler. Each of the first and second swirlers have features that cause the flow to create a vortex in the prechamber and main chamber, respectively. The features creating the swirl are blades or angled orifices. The vortex causes a pressure depression along the centerline and causes backflow along the centerline that improves mixing and improves cooling.

The subject matter of this specification can be embodied in, among other things, a turbine combustor assembly includes a primary combustion chamber in fluid communication with a primary fuel injector and a primary air inlet, and an igniter carried by the primary combustion chamber and including a first igniter stage having an auxiliary combustion chamber housing comprising a mixing chamber and a tubular throat converging downstream of the mixing chamber, an ignition source projecting into the mixing chamber of the auxiliary combustion chamber, and a second igniter stage that includes an auxiliary fuel outlet manifold proximal to an outlet of the tubular throat.

A fuel nozzle for a combustor of an aircraft engine includes a nozzle body defining an a fuel passage, extending therethrough between a fuel inlet and a fuel outlet located at the outlet end that at least partially defines a nozzle tip, for directing a fuel flow into the combustor via the nozzle tip. A core air passage extends through the nozzle body for directing a core air flow into the combustor via the nozzle tip. At least two flow restrictors are disposed in series within the core air passage, the flow restrictors including an upstream flow restrictor and a downstream flow restrictor each having an orifice therein. The restricted air flow passage having a cross-sectional area smaller than that of the core air passage. The orifice in the upstream flow restrictor being at least partially offset from the orifice in the downstream flow restrictor.

Combustor dome assemblies having combustor deflectors are provided. For example, a combustor dome assembly comprises a combustor dome defining an opening; a ceramic matrix composite (CMC) deflector positioned adjacent the combustor dome on an aft side of the assembly; a fuel-air mixer defining a groove about an outer perimeter thereof; and a seal plate including a key. The CMC deflector includes a cup extending forward through the opening in the combustor dome that defines one or more bayonets and a slot. The bayonets are received in the fuel-air mixer groove, and the seal plate key is received in the CMC deflector slot. In another embodiment, where the seal plate may be omitted, a spring is positioned between the fuel-air mixer and the CMC deflector to hold the CMC deflector in place with respect to the combustor dome. Methods of assembling combustor dome assemblies having CMC deflectors also are provided.