Combustor liner

Abstract

The present invention relates to a combustor liner for a gas turbine, the combustor liner having substantially cylindrical shape and comprising a first section and a second section wherein the first section is upstream of the second section with respect to the hot gas flow during operation, wherein the first section is ring shaped and comprises a rounded lip section and a trailing section, wherein an inner radius of the trailing section is increasing along a centerline of the liner in the direction of the hot gas flow during operation. The invention also relates to a combustor and a gas turbine comprising such combustion liner.

1 . A combustor liner for a gas turbine, the combustor liner having substantially cylindrical shape and comprising a first section and a second section wherein the first section is upstream of the second section with respect to the hot gas flow during operation, wherein the first section is ring shaped and comprises a rounded lip section and a trailing section, wherein an inner radius of the trailing section is increasing along a centerline of the liner in the direction of the hot gas flow during operation. 
     
     
         2 . The combustor liner of  claim 1 , wherein the inner radius (R 1 ) of the trailing section is increasing monotonically along the centerline of the liner. 
     
     
         3 . The combustor liner of  claim 1 , wherein the length in axial direction of the first section is in the range from 20 percent to 80 percent of the total length in axial direction of the liner. 
     
     
         4 . The combustor liner  claim 1 , wherein an angle (α) between the trailing section and an outer surface of the combustion liner is in the range of 5 to 15 degrees. 
     
     
         5 . The combustor liner of  claim 1 , wherein a radius of an outer surface of the combustion liner is substantially constant along the centerline of the liner. 
     
     
         6 . The combustor liner of  claim 1 , wherein a radius (R 2 ) of the second section is substantially constant along the centerline of the liner. 
     
     
         7 . The combustor liner of  claim 1 , wherein the first section is substantially hollow. 
     
     
         8 . The combustor liner of  claim 1 , wherein the first section comprises a damper device, preferably the damper device is Helmholtz damper. 
     
     
         9 . The combustor liner of  claim 1 , wherein the first section comprises means for a liquid fuel injection. 
     
     
         10 . A combustor for a gas turbine wherein it comprises the combustor liner according to  claim 1  and a combustion zone delimited by the combustion liner. 
     
     
         11 . The combustor  claim 10  further comprising:
 a substantially cylindrical flow sleeve, wherein the combustion liner is located at least partially within the flow sleeve thereby forming a first passage between the flow sleeve and the combustion liner; 
 a dome located forward of the flow sleeve and encompassing at least partly a first section of the combustion liner, the dome having a substantially rounded head end thereby forming a turning passage between the rounded lip section of the first section of combustion liner and the dome; 
 at least one pilot channel comprising a means for supplying a pilot fuel and a first swirling device. 
 
     
     
         12 . The combustor of  claim 11  wherein the first passage and/or the turning passage comprise a fuel injection means and a second swirling device; 
     
     
         13 . The combustor of  claim 11  wherein the first swirling device and/or the second swirling device are axial or radial swirlers. 
     
     
         14 . A gas turbine comprising the combustor according  claim 9 . 
     
     
         15 . A method for operating the gas turbine combustor according to  claim 10 , the method comprising:
 supplying a first flow of air into the pilot channel;   supplying a first stream of fuel into the pilot channel to mix with the first flow of air, and feeding the resulting first mixture into the combustion zone for providing pilot flame;   supplying a second flow of air into the first passage; supplying a second stream of fuel into the first passage or turning passage to mix with the second flow of air, and feeding the resulting second mixture into the combustion zone for providing a main flame;   wherein the first mixture and second mixture are guided along the inner wall of the liner and form a central recirculation zone in the center of the combustion zone.