Abstract: A fuel injector for a turbine engine may include a body member disposed about a longitudinal axis, and a barrel member located radially outwardly from the body member. The fuel injector may also include an annular passageway extending between the body member and the barrel member from a first end to a second end. The first end may be configured to be fluidly coupled to a compressor of the turbine engine and the second end may be configured to be fluidly coupled to a combustor of the turbine engine. The fuel injector may also include a perforated plate positioned proximate the first end of the passageway. The perforated plate may be configured to direct compressed air into the annular passageway with a first pressure drop. The fuel injector may also include at least one fuel discharge orifice positioned downstream of the perforated plate. The at least one fuel discharge orifice may be configured to discharge a fuel into the annular passageway with a second pressure drop.

Abstract: A gas turbine engine may include a plurality of pilot fuel supply lines configured to supply liquid fuel. Each pilot fuel supply line may be coupled to a respective fuel injector. The turbine engine may include a plurality of main fuel supply lines configured to supply liquid fuel. Each main fuel supply line may be coupled to a respective fuel injector. The turbine engine may also include a flow restriction provided in a first plurality of the plurality of main fuel supply lines. A second plurality of the plurality of main fuel supply lines may be free from the flow restriction.

Abstract: A gas fuel turbine engine may include a gaseous pilot fuel supply, a first main fuel supply, and a second main fuel supply. The first main fuel supply may provide gaseous fuel to a first plurality of fuel injectors, and the second main fuel supply may provide gaseous fuel to a second plurality of fuel injectors. The turbine engine may also include a flow restriction provided in the first main fuel supply. The second main fuel supply may be free of the flow restriction.

Abstract: Systems and methods for operating a turbine engine having a plurality of fuel injectors arranged circumferentially in a combustor, with each fuel injector having a main fuel supply and a pilot fuel supply, includes supplying fuel to the plurality of fuel injectors through the main fuel supply to create a circumferential thermal gradient in the combustor.

Abstract: A fuel nozzle for a turbine engine has a central body member with a pilot, a surrounding barrel housing, a mixing duct and an air inlet duct. The fuel nozzle additionally has a main fuel injection device located between the air inlet duct and the mixing duct. The main fuel injection device is configured to introduce a flow of fuel into the barrel member to create a fuel/air mixture which is then premixed with a swirler. The fuel/air mixture then further mixes in the mixing duct and exits the nozzle into a combustor for combustion. The geometry of the fuel nozzle ensures that pressure waves from the combustor do not create a time varying fuel to air equivalence ratio in the flow through the nozzle that achieves a resonance with the pressure waves.

Abstract: A fuel injector for a turbine engine may include a body member disposed about a longitudinal axis, and a barrel member located radially outwardly from the body member. The fuel injector may also include an annular passageway extending between the body member and the barrel member from a first end to a second end. The first end may be configured to be fluidly coupled to a compressor of the turbine engine and the second end may be configured to be fluidly coupled to a combustor of the turbine engine. The fuel injector may also include a perforated plate positioned proximate the first end of the passageway. The perforated plate may be configured to direct compressed air into the annular passageway with a first pressure drop. The fuel injector may also include at least one fuel discharge orifice positioned downstream of the perforated plate. The at least one fuel discharge orifice may be configured to discharge a fuel into the annular passageway with a second pressure drop.

Abstract: A fuel nozzle for a turbine engine has a central body member with a pilot, a surrounding barrel housing, a mixing duct and an air inlet duct. The fuel nozzle additionally has a main fuel injection device located between the air inlet duct and the mixing duct. The main fuel injection device is configured to introduce a flow of fuel into the barrel member to create a fuel/air mixture which is then premixed with a swirler. The fuel/air mixture then further mixes in the mixing duct and exits the nozzle into a combustor for combustion. The geometry of the fuel nozzle ensures that pressure waves from the combustor do not create a time varying fuel to air equivalence ratio in the flow through the nozzle that achieves a resonance with the pressure waves.

Abstract: A system and method for modifying the supply of fuel to injectors to attenuate combustion oscillations in a gas turbine engine. The gas turbine engine may comprise a combustor, a plurality of injectors and a manifold. The plurality of injectors may be operable to provide fuel to the combustor. The manifold may be configured to supply fuel to all of the plurality of injectors or to only a portion of the plurality of injectors in reaction to a determination of an existence of combustion oscillations.

Abstract: A fuel nozzle for a turbine engine has a central body member with a pilot, a surrounding barrel housing, a mixing duct and an air inlet duct. The fuel nozzle additionally has a main fuel injection device located between the air inlet duct and the mixing duct. The main fuel injection device is configured to introduce a flow of fuel into the barrel member to create a fuel/air mixture which is then premixed with a swirler. The fuel/air mixture then further mixes in the mixing duct and exits the nozzle into a combustor for combustion. The geometry of the fuel nozzle ensures that pressure waves from the combustor do not create a time varying fuel to air equivalence ratio in the flow through the nozzle that achieves a resonance with the pressure waves.

Abstract: A fuel nozzle for a turbine engine has a central body member with a pilot, a surrounding barrel housing, a mixing duct and an air inlet duct. The fuel nozzle additionally has a main fuel injection device located between the air inlet duct and the mixing duct. The main fuel injection device is configured to introduce a flow of fuel into the barrel member to create a fuel/air mixture which is then premixed with a swirler. The fuel/air mixture then further mixes in the mixing duct and exits the nozzle into a combustor for combustion. The geometry of the fuel nozzle ensures that pressure waves from the combustor do not create a time varying fuel to air equivalence ratio in the flow through the nozzle that achieves a resonance with the pressure waves.

Abstract: A splash plate dome assembly for a combustion liner of a turbine engine is disclosed. The splash plate may include an outer periphery having a plurality of corners and an inner periphery defining an aperture and an annular flange. The splash plate may include a plurality of first flow guides extending from the outer periphery to the inner periphery. The splash plate may further include a plurality of second flow guides extending from the outer periphery to a location intermediate the inner periphery and the outer periphery. The splash plate, annular flange, first flow guides, and second flow guides may be integrally formed by casting. The splash plate may be mounted to a combustion dome having a plurality of distributed through holes for forming impingement jets on an upstream face of the splash plate.

Abstract: A combustor for a gas turbine engine uses compressed air to cool a combustor liner and uses at least a portion of the same compressed air for combustion air. A flow diverting mechanism regulates compressed air flow entering a combustion air plenum feeding combustion air to a plurality of fuel nozzles. The flow diverting mechanism adjusts combustion air according to engine loading.

Abstract: Existing combustors have the tendency to emit emissions and require a large quantity of cooling air to retain or extend the life of the components to a reasonable life expectancy. The present combustor reduces the emissions emitted therefrom, requires a reduced quantity of cooling air while resulting in a high heat transfer cooling rate extending the life expectancy of the components. The combustor construction includes an interior liner having a plurality of angled holes extending therethrough arranged in a preestablished pattern defining a centroid and an exterior liner having a plurality of holes extending therethrough at about a 90 degree. At least a portion of the plurality of holes in the exterior liner being radially aligned with the centroid of the plurality of holes in the interior line.

Abstract: The present method of making a workpiece having a plurality of holes defined therein includes the plurality of holes defining a specific flow at a specific flow condition.

Abstract: A pulse-cooled gas turbine engine assembly includes a hollow airfoil and a fluidic oscillator joined thereto. The fluidic oscillator discharges pressurized cooling air inside the airfoil at a predetermined pulsation frequency for cooling the airfoil.

Abstract: The uprating of existing jet engines increases the operating temperatures within the engines and also increases the tendency for flashback in some afterburner type engines. Afterburner type engines include a pilot fuel delivery pipe which is included between inner and outer annular members of a flameholder. Downstream from the pilot, partitions direct an ignited fuel/air mixture into the afterburner chamber. The afterburner flame is downstream from the flameholder, but under conditions of increased temperature the afterburner flame will tend to migrate upstream toward the flameholder. To avoid this tendency, the flameholder includes means for discharging gas which cools the flameholder metal itself while also decreasing the downstream air temperature, thereby decreasing the potential for flashback. In one embodiment, fuel and oxygen depleted turbine discharge gas are used to further decrease the tendency toward flashback.

Abstract: A diffuser is provided for achieving efficient conversion of the dynamic head associated with pressurized fluid into static pressure. The diffuser includes structure for reducing the boundary layer accumulated by the fluid during flow in the diffuser and for turning the fluid within the diffuser flowpath.