Combustor liner cooling system

Abstract

A turbine engine with a combustor that includes a hollow wall about a combustor liner. The combustor liner includes an inner surface facing inwardly toward a combustion chamber. The turbine engine includes a first air flow path in an upstream direction through the hollow wall toward a head end of the combustor. The first air flow path includes a plurality of bypass openings extending through the combustor liner to the inner surface to supply a first cooling film to a downstream end portion of the combustor liner. The turbine engine further includes a second flow path in a second direction opposite the upstream direction through the hollow wall. The second flow path includes a plurality of film holes extending through the combustor liner to the inner surface to supply a second cooling film to the downstream end portion of the combustor liner downstream of the first cooling film.

1 . A system comprising:
 a turbine engine comprising:
 a combustor comprising a hollow wall having a sleeve disposed about a combustor liner, wherein the combustor liner comprises an inner surface facing inwardly toward a combustion chamber; 
 a first air flow path in an upstream direction through the hollow wall toward a head end of the combustor, wherein the first air flow path comprises a plurality of bypass openings extending through the combustor liner to the inner surface, wherein the plurality of bypass openings is configured to supply a first cooling film to a downstream end portion of the combustor liner; and 
 a second air flow path in a downstream direction opposite the upstream direction through the hollow wall, wherein the second flow path comprises a plurality of film holes extending through the combustor liner to the inner surface, and the plurality of film holes is configured to supply a second cooling film to the downstream end portion of the combustor liner downstream of the first cooling film. 
   
     
     
         2 . The system of  claim 1 , wherein the plurality of bypass openings extending through the combustor liner is disposed upstream of the plurality of film holes extending through the combustor liner. 
     
     
         3 . The system of  claim 2 , wherein the plurality of bypass openings extend through the downstream end portion of the combustor liner, and an axial length of the downstream end portion is less than or equal to approximately 35 percent of a total axial length of the combustor liner. 
     
     
         4 . The system of  claim 1 , wherein the second air flow path is defined by a passage formed by an axial cooling channel on the downstream end portion of the combustor liner and an inner surface of a wrapper coaxially disposed generally about the downstream end portion. 
     
     
         5 . The system of  claim 4 , wherein the wrapper comprises one or more radial openings configured to supply a portion of an air flow along the first air flow path into the axial cooling channel. 
     
     
         6 . The system of  claim 5 , wherein the plurality of film holes extends radially through the axial cooling channel downstream of the one or more radial openings, and the plurality of film holes is configured to direct the portion of the air flow to supply the second cooling film to the downstream end portion of the combustor liner. 
     
     
         7 . The system of  claim 1 , wherein the first air flow path is at least partially defined by a first passage between a transition piece and a transition sleeve that surrounds the transition piece. 
     
     
         8 . The system of  claim 7 , wherein the first passage is fluidly coupled to a second passage between the combustor liner and the transition sleeve, wherein the second passage extends in the upstream direction from the first passage, the first passage comprises a first plurality of inlets to receive a first portion of air that flows through the first passage and the second passage in the upstream direction, and the second passage comprises a second plurality of inlets to receive a second portion of air that flows through the second passage in the upstream direction. 
     
     
         9 . The system of  claim 8 , wherein the turbine engine comprises one or more fuel nozzles, wherein the fuel nozzles are configured to receive the first portion of air and the second portion of air flowing in the upstream direction through the second passage and to mix the first portion of air and the second portion of air with a fuel, and wherein the fuel nozzles are configured to output a resulting air-fuel mixture into the combustion chamber surrounded by the combustor liner for combustion. 
     
     
         10 . A system comprising:
 a turbine combustor liner comprising:
 a plurality of axial cooling channels arranged circumferentially about a downstream end portion relative to a downstream direction of combustion along a longitudinal axis of the turbine combustor liner; 
 an inner surface facing inwardly toward a combustion chamber; and 
 a plurality of bypass openings arranged circumferentially about the downstream end portion upstream of the plurality of axial cooling channels, wherein the plurality of bypass openings is configured to supply a first cooling film to the inner surface of the combustor liner, and each of the plurality of axial cooling channels comprises a plurality of film holes configured to supply a second cooling film to the inner surface of the combustor liner at the downstream end portion. 
   
     
     
         11 . The system of  claim 10 , wherein an interior of the turbine combustor liner has a combustion path with a downstream direction of flow of combustion gases, an exterior of the turbine combustor liner has a first air path with an upstream direction of flow opposite to the downstream direction, and the exterior of the turbine combustor liner has the plurality of cooling channels with a second air path in the downstream direction. 
     
     
         12 . The system of  claim 11 , comprising a first flow sleeve disposed about the turbine combustor liner to define a first hollow wall, and a second flow sleeve disposed about a transition piece to define a second hollow wall, wherein the first and second hollow walls are coupled to one another at the downstream end portion, the first and second hollow walls define the first air path with the upstream direction, and the second air path in the downstream direction is disposed radially between the plurality of cooling channels and the transition piece. 
     
     
         13 . The system of  claim 10 , wherein the plurality of cooling channels is defined by alternating axial grooves and axial protrusions about a circumference of the turbine combustor liner, and the plurality of film holes extend radially through the axial grooves into an interior of the turbine combustor liner. 
     
     
         14 . The system of  claim 10 , wherein the plurality of bypass openings extend radially through the combustor liner at an angle of approximately 90 degrees relative to the inner surface. 
     
     
         15 . The system of  claim 10 , wherein the plurality of bypass openings extend radially through the combustor liner at an angle between approximately 30 to 60 degrees. 
     
     
         16 . The system of  claim 10 , wherein at least one bypass opening of the plurality of bypass openings has a geometry that converges or diverges through the combustor liner into the interior of the turbine combustor liner. 
     
     
         17 . The system of  claim 10 , wherein each bypass opening of the plurality of bypass openings is disposed in an upstream direction of combustion along the longitudinal axis from an axial cooling channel of the plurality of axial cooling channels. 
     
     
         18 . The system of  claim 10 , wherein the bypass openings are arranged in two or more axially spaced sets, wherein each set is disposed circumferentially about the downstream end portion. 
     
     
         19 . The system of  claim 10 , wherein an axial length of the downstream end portion is less than or equal to approximately 35 percent of a total axial length of the turbine combustor liner, an axial channel length of each of the plurality of cooling channels is less than or equal to the axial length of the downstream end portion, and the cooling channels have a depth of approximately 0.05 to 0.30 inches and a width of approximately 0.25 to 1.0 inches. 
     
     
         20 . A system comprising:
 a turbine engine comprising:
 a combustor comprising:
 a flow sleeve; and 
 a combustor liner surrounded by the flow sleeve and defining a flow path therebetween configured to receive an air flow in a first direction toward a head end chamber, wherein the combustor liner comprises:
 an inner surface facing inwardly toward a combustion chamber; 
 a plurality of axial cooling channels arranged circumferentially about a downstream end portion of the combustor liner; 
 a plurality of bypass holes arranged in two or more axially spaced sets at the downstream end portion, wherein each bypass opening is disposed upstream in the first direction relative to an axial cooling channel of the plurality of axial cooling channels, and each bypass opening is configured to supply a first cooling film to the inner surface of the combustor liner at the downstream end portion; 
 one or more fuel nozzles disposed in the head end chamber of the combustor; and 
 wherein each of the plurality of axial cooling channels comprises a plurality of film holes extending through the combustor liner to the inner surface, each of the plurality of axial cooling channels is configured to receive a portion of the air flow from the flow path, to direct a first portion of the received air along an axial length of the axial cooling channel in a second direction away from the head end chamber, and to direct a second portion of the received air through the plurality of film holes to supply a second cooling film to the inner surface of the combustor liner at the downstream end portion.