Main fuel nozzle for combustion dynamics attenuation

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.

What is claimed is: 
     
         1 . A combustor assembly for a gas turbine engine, the combustor assembly defining a combustor centerline and a radial direction extended from the combustor centerline, the radial direction and combustor centerline together defining a reference plane, the combustor assembly comprising:
 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 defining a plurality of exit openings in circumferential arrangement relative to the nozzle centerline, the plurality of exit openings defining two or more locations different from one another on the centerbody along the lengthwise direction.   
     
     
         2 . The combustor assembly of  claim 1 , wherein the two or more locations along the lengthwise direction and the circumferential arrangement of the plurality of exit openings defines a waveform. 
     
     
         3 . The combustor assembly of  claim 2 , wherein the waveform is sinusoidal, sawtooth, or box. 
     
     
         4 . The combustor assembly of  claim 1 , wherein the plurality of exit openings defines two or more cross sectional areas, wherein each cross sectional area is different from one another. 
     
     
         5 . The combustor assembly of  claim 4 , wherein the plurality of exit openings provides a fuel to a combustion chamber at two or more flow rates each corresponding to the two or more cross sectional areas of the plurality of exit openings. 
     
     
         6 . The combustor assembly of  claim 1 , wherein the plurality of exit openings defines a first opening and a second opening, the first opening defining a first lengthwise location and the second opening defining a second lengthwise location different from the first lengthwise location. 
     
     
         7 . The combustor assembly of  claim 1 , wherein the plurality of exit openings defines a first opening and a second opening, the first opening defining a first cross sectional area and the second opening defining a second cross sectional area different from the first cross sectional area. 
     
     
         8 . The combustor assembly of  claim 1 , wherein a single exit opening of the plurality of exit openings is defined at each circumferential location along the centerbody from the nozzle centerline. 
     
     
         9 . The combustor assembly of  claim 1 , wherein the plurality of exit openings are defined in asymmetric circumferential arrangement through the centerbody. 
     
     
         10 . A gas turbine engine defining an axial centerline and a radial direction extended from the axial centerline, the radial direction and axial centerline together defining a reference plane, the gas turbine engine comprising:
 a combustor assembly comprising a plurality of fuel nozzles disposed in circumferential arrangement around the axial centerline of the gas turbine engine, wherein each fuel nozzle comprises a centerbody extended along a lengthwise direction, and wherein the fuel nozzle defines a nozzle centerline extended through the centerbody along the lengthwise direction, the centerbody defining a plurality of exit openings in circumferential arrangement relative to the nozzle centerline, the plurality of exit openings defining two or more cross sectional areas different from one another on the centerbody.   
     
     
         11 . The gas turbine engine of  claim 10 , wherein the plurality of exit openings are defined along the same reference plane. 
     
     
         12 . The gas turbine engine of  claim 10 , wherein the plurality of exit openings are defined along two or more locations different from one another along the lengthwise direction. 
     
     
         13 . The gas turbine engine of  claim 12 , wherein the plurality of exit openings provides a fuel to a combustion chamber at two or more flow rates each corresponding to the two or more cross sectional areas of the plurality of exit openings. 
     
     
         14 . The gas turbine engine of  claim 12 , wherein the two or more locations along the lengthwise direction and the circumferential arrangement of the plurality of exit openings defines a waveform along the centerbody. 
     
     
         15 . The gas turbine engine of  claim 14 , wherein the waveform is sinusoidal, sawtooth, or box. 
     
     
         16 . The gas turbine engine of  claim 10 , wherein the plurality of exit openings defines a first opening and a second opening, the first opening defining a first lengthwise location and the second opening defining a second lengthwise location different from the first lengthwise location. 
     
     
         17 . The gas turbine engine of  claim 10 , wherein a single exit opening of the plurality of exit openings is defined at each circumferential location along the centerbody from the nozzle centerline. 
     
     
         18 . The gas turbine engine of  claim 10 , wherein the plurality of fuel nozzles defines a first nozzle and a second nozzle, and wherein the first nozzle defines a plurality of exit openings, the plurality of exit openings comprising a first opening defining a first cross sectional area and a second opening defining a second cross sectional area different from the first cross sectional area, and wherein the second nozzle defines the plurality of exit openings at a plurality of locations along the lengthwise direction. 
     
     
         19 . The gas turbine engine of  claim 18 , wherein the plurality of fuel nozzles are defined in alternating circumferential arrangement. 
     
     
         20 . The gas turbine engine of  claim 18 , wherein at least half of the plurality of fuel nozzles defines the first nozzle and the remainder of the plurality of fuel nozzles defines the second nozzle, and wherein at least two or more of the first fuel nozzle or the second fuel nozzle are in sequential circumferential arrangement.