Method and apparatus for assembling gas turbine engine combustors

Abstract

A method enables the operation of a gas turbine engine. The method comprises channeling airflow into a cooling passageway defined between the combustor casing and an inner liner of the combustor, wherein the inner liner is fabricated from a plurality of panels coupled together, channeling airflow into a cooling passageway defined between the combustor casing and an outer liner of the combustor; wherein the outer liner is fabricated from a plurality of panels coupled together, and channeling dilution airflow into a combustion chamber defined between the inner and outer liners, through a plurality of openings formed within at least one panel within at least one of the inner liner panels and the outer liner panels, wherein the plurality of openings are non-circular.

1 . A method for operating a gas turbine engine, said method comprising: 
 channeling airflow into a cooling passageway defined between the combustor casing and an inner liner of the combustor, wherein the inner liner is fabricated from a plurality of panels coupled together;    channeling airflow into a cooling passageway defined between the combustor casing and an outer liner of the combustor; wherein the outer liner is fabricated from a plurality of panels coupled together; and    channeling dilution airflow into a combustion chamber defined between the inner and outer liners, through a plurality of openings formed within at least one panel within at least one of the inner liner panels and the outer liner panels, wherein the plurality of openings are non-circular.    
   
   
       2 . A method in accordance with  claim 1  wherein channeling dilution airflow into a combustion chamber further comprises channeling dilution airflow into the combustion chamber to facilitate controlling an exit temperature profile of the combustor.  
   
   
       3 . A method in accordance with  claim 1  wherein channeling dilution airflow into a combustion chamber further comprises channeling dilution airflow into the combustion chamber through the plurality of openings, wherein the openings are shaped to enable cooling air to penetrate into the combustion chamber to facilitate achieving a desired radial temperature profile within the combustion chamber.  
   
   
       4 . A method in accordance with  claim 1  wherein channeling dilution airflow into a combustion chamber further comprises channeling dilution airflow into the combustion chamber through the plurality of openings, wherein the openings are generally elliptically shaped.  
   
   
       5 . A method in accordance with  claim 1  wherein channeling dilution airflow into a combustion chamber further comprises channeling dilution airflow into the combustion chamber through the plurality of openings, wherein the openings are defined by a pair of substantially parallel walls that are connected together by a pair of opposed arcuate sidewalls formed with a predetermined radius of curvature.  
   
   
       6 . A combustor for a gas turbine engine, said combustor comprising: 
 an inner liner comprising a plurality of panels coupled together;    an outer liner comprising a plurality of panels coupled together; and    a combustion chamber defined between said inner and outer liners, at least one of said plurality of inner liner panels and said plurality of outer liner panels comprises a plurality of openings extending therethrough for channeling dilution airflow into said combustion chamber, said plurality of openings are non-circular.    
   
   
       7 . A combustor in accordance with  claim 6  wherein said plurality of openings facilitate controlling an exit temperature profile of said combustor.  
   
   
       8 . A combustor in accordance with  claim 6  wherein said plurality of openings are each substantially elliptically-shaped.  
   
   
       9 . A combustor in accordance with  claim 6  wherein said plurality of openings are shaped to enable cooling air to penetrate into said combustion chamber to facilitate achieving a desired radial temperature profile within said combustion chamber.  
   
   
       10 . A combustor in accordance with  claim 6  wherein said plurality of openings are defined by a pair of opposed substantially parallel sidewalls connected together by a pair of opposed arcuate walls formed with a pre-determined radius.  
   
   
       11 . A combustor in accordance with  claim 10  wherein adjacent of said plurality of openings are separated by a distance that is approximately equal to twice the diameter of said arcuate walls.  
   
   
       12 . A combustor in accordance with  claim 6  wherein said at least one panel comprises a pair of opposed circumferential edges coupled together by a leading edge and a side edge, said plurality of openings comprises at least three openings spaced approximately equi-distantly between said pair of opposed circumferential edges.  
   
   
       13 . A gas turbine engine comprising a combustor comprising an inner liner, an outer liner, and a combustion chamber defined between said inner and outer liners, each of said inner and outer liners comprises a plurality of panels coupled together, at least one of said panels within at least one of said inner liner and said outer liner comprises a plurality of openings extending therethrough for channeling dilution air into said combustion chamber, said plurality of openings are non-circular.  
   
   
       14 . A gas turbine engine in accordance with  claim 13  wherein said combustor plurality of openings extending through said at least one panel facilitate controlling an exit temperature profile of said combustor.  
   
   
       15 . A gas turbine engine in accordance with  claim 14  wherein said combustor plurality of openings extending through said at least one panel are each generally elliptically-shaped.  
   
   
       16 . A gas turbine engine in accordance with  claim 14  wherein said combustor plurality of openings extending through said at least one panel are shaped to enable cooling air to penetrate into said combustion chamber to facilitate achieving a desired radial temperature profile within said combustion chamber.  
   
   
       17 . A gas turbine engine in accordance with  claim 14  wherein said combustor plurality of openings extending through said at least one panel are defined by a pair of opposed substantially parallel sidewalls that are connected together by a pair of opposed arcuate walls formed with a pre-determined radius.  
   
   
       18 . A gas turbine engine in accordance with  claim 17  wherein adjacent of said plurality of openings extending through said at least one panel are separated within said panel by a distance that is approximately equal to twice the diameter of said arcuate walls.  
   
   
       19 . A gas turbine engine in accordance with  claim 14  wherein each of said plurality of panels comprises a pair of opposed circumferential edges coupled together by a leading edge and a side edge, said plurality of openings extending through said at least one panel comprises at least three openings spaced approximately equi-distantly between said pair of opposed circumferential edges.