Cooling Holes For Gas Turbine Combustor Having A Non-Uniform Diameter Therethrough

Abstract

A gas turbine combustor liner, including a shell having a first end adjacent to an upstream end of the combustor and a second end adjacent to a downstream end of the combustor, where the shell also has a hot side, a cold side, and a centerline axis therethrough. A plurality of small, closely-spaced film cooling holes are formed in the shell through which air flows for providing a cooling film along the hot side of the shell. Each cooling hole has a non-uniform diameter as it extends through the shell. In particular, each cooling hole includes a first opening located at the cold side of the shell having a first diameter and a second opening located at the hot side of the shell having a second diameter, wherein the second diameter of the second opening is larger than the first diameter of the first opening. It is preferred that the shape of each cooling hole be substantially frusto-conical.

1 . A gas turbine combustor liner, comprising:
 (a) a shell having a first end adjacent to an upstream end of said combustor and a second end adjacent to a downstream end of said combustor, said shell also having a hot side, a cold side, and a centerline axis therethrough; and,   (b) a plurality of small, closely-spaced film cooling holes formed in said shell through which air flows for providing a cooling film along said hot side of said shell, said cooling holes having a non-uniform diameter through said shell.   
   
   
       2 . The combustor liner of  claim 1 , each of said cooling holes further comprising:
 (a) a first opening located at said cold side of said shell having a first diameter; and,   (b) a second opening located at said hot side of said shell having a second diameter;   
     wherein said second diameter of said second opening is larger than said first diameter of said first opening. 
   
   
       3 . The combustor liner of  claim 2 , wherein a ratio of said second diameter for said second opening to said first diameter for said first opening is approximately 3.0-5.0. 
   
   
       4 . The combustor liner of  claim 2 , wherein spacing between adjacent first openings of said cooling holes is greater than spacing between adjacent second openings of said cooling holes. 
   
   
       5 . The combustor liner of  claim 2 , wherein spacing between adjacent first openings of said cooling holes is approximately 3.0-6.0 times said first diameter. 
   
   
       6 . The combustor liner of  claim 2 , wherein spacing between said adjacent second openings of said cooling holes is approximately 0.2-0.7 times said second diameter. 
   
   
       7 . The combustor liner of  claim 1 , wherein a diameter of each cooling hole grows progressively larger from said cold side of said shell to said hot side of said shell. 
   
   
       8 . The combustor liner of  claim 7 , wherein each said cooling hole has a substantially frusto-conical shape. 
   
   
       9 . The combustor liner of  claim 7 , wherein each said cooling hole has an angle of diffusion with respect to an axis therethrough of approximately 1° to approximately 15°. 
   
   
       10 . The combustor liner of  claim 7 , wherein each said cooling hole has an angle of diffusion with respect to an axis therethrough of approximately 3° to approximately 10°. 
   
   
       11 . The combustor liner of  claim 7 , wherein each said cooling hole has an angle of diffusion with respect to an axis therethrough of approximately 5° to approximately 9°. 
   
   
       12 . The combustor liner of  claim 1 , wherein an axis through each said cooling hole is oriented at an axial angle to said centerline axis in a range of approximately 15° to approximately 35°. 
   
   
       13 . The combustor liner of  claim 1 , wherein an axis through each said cooling hole is oriented at a circumferential angle to said centerline axis in a range of approximately 30° to approximately 60°. 
   
   
       14 . The combustor liner of  claim 1 , wherein a ratio of a jet velocity of air at said hot side of said shell to a jet velocity of air at said cold side of said shell is approximately 0.25-0.50. 
   
   
       15 . The combustor liner of  claim 1 , each said cooling hole further comprising:
 (a) a first portion having a substantially uniform diameter through said liner; and,   (b) a second portion having a non-uniform diameter through said liner.   
   
   
       16 . The combustor liner of  claim 15 , wherein said first portion of said cooling hole is located adjacent said cold side of said shell. 
   
   
       17 . The combustor liner of  claim 15 , wherein said first portion of said cooling hole is located adjacent said hot side of said shell. 
   
   
       18 . The combustor liner of  claim 1 , wherein said cooling holes are provided over essentially an entire length of said shell. 
   
   
       19 . The combustor liner of  claim 1 , wherein said cooling holes are provided at certain designated locations of said shell. 
   
   
       20 . The combustor liner of  claim 1 , wherein said cooling holes are provided at an upstream portion of said shell. 
   
   
       21 . The combustor liner of  claim 1 , wherein said cooling holes are provided in said shell immediately downstream of an air-fuel mixer for said combustor. 
   
   
       22 . A method of forming a cooling hole in a liner of a gas turbine engine combustor, wherein said cooling hole has a non-uniform diameter therethrough, comprising the following steps:
 (a) forming a first portion of said cooling hole from a hot side of said liner, wherein said first portion is substantially conical in shape and extends substantially through said liner; and,   (b) forming a second portion of said cooling hole from said first portion of said cooling hole to a cold side of said liner, wherein said second portion is substantially uniform in diameter.   
   
   
       23 . The method of  claim 22 , wherein said first portion of said cooling hole has a diameter which progressively decreases from said hot side of said liner.