Turbine shroud segment transpiration cooling with individual cast inlet and outlet cavities

Abstract

A shroud segment of a turbine shroud of a gas turbine engine comprises a platform with front and rear legs. The platform defines a plurality of axially extending holes with individual inlets on an outer surface of the platform for transpiration cooling of the platform of the turbine shroud segment.

1 - 14 . (canceled) 
   
   
       15 . A gas turbine engine cooling arrangement for serially cooling a turbine static shroud and a turbine vane outer shroud, the cooling arrangement comprising a plurality of passages extending through a turbine static shroud platform substantially in a direction of a turbine rotating axis for directing cooling air through the turbine static shroud platform to cool the turbine shroud, the passages having respective outlets positioned and configured to direct air exhausted from the passages to impinge substantially all of a circumferential leading end of the turbine vane outer shroud 
   
   
       16 . The cooling arrangement as claimed in  claim 15  wherein each of the passages extends through a major section of, an entire axial length of the shroud platform. 
   
   
       17 . The cooling arrangement as claimed in  claim 16  wherein each of the passages is in fluid communication with a cavity defined between front and rear legs of the turbine shroud segment. 
   
   
       18 . The cooling arrangement as claimed in  claim 17  wherein each of the passages discharges the cooling air at a trailing end of the shroud platform for impingement, on the turbine vane outer shroud. 
   
   
       19 . The cooling arrangement as claimed in  claim 18  wherein the passages comprise a plurality of substantially axial straight holes. 
   
   
       20 . A gas turbine engine assembly defining a portion of an outer wall of a gas path, the assembly comprising a turbine static shroud having a platform, and a turbine vane outer shroud disposed immediately downstream of the turbine static shroud, the turbine vane outer shroud having a plurality of vanes extending radially inwardly from the turbine vane outer shroud, the turbine static shroud defining a plurality of passages extending substantially axially in a direction of a turbine rotating axis through the platform of the turbine static shroud, the passages having outlets on a trailing end of the platform which substantially radially align with a leading edge of the turbine vane outer shroud, the outlets configured for discharging air exhausted from the passages to impinge on the turbine vane outer shroud leading edge. 
   
   
       21 . The gas turbine engine structure as claimed in  claim 20  wherein the passages comprise a plurality of holes having at least one inlet thereof on an outer surface of the platform. 
   
   
       22 . The gas turbine engine structure as claimed in  claim 21  wherein the at least one inlet of the holes is located in a position close to and downstream a front leg of the shroud. 
   
   
       23 . The gas turbine engine structure as claimed in  claim 21  wherein the holes extend in a substantially straight direction over a major section of the entire axial length of the platform. 
   
   
       24 . A method of reusing turbine shroud cooling air for impingement cooling on a downstream turbine vane outer shroud, the method comprising steps of:
 (a) directing cooling air within and through a platform of a shroud segment of a turbine shroud for cooling the turbine shroud; and   (b) using the cooling air of step (a) to form a plurality of substantially straight cooling air streams axially towards a leading end of the turbine vane outer shroud for impingement cooling on the turbine vane outer shroud.   
   
   
       25 . The method as claimed in  claim 24  wherein steps (a) and (b) are conducted substantially simultaneously. 
   
   
       26 . The method as claimed in  claim 25  wherein steps (a) and (b) are practiced by directing the cooling air through a plurality of substantially axial and straight passages extending within the platform of the shroud segment of the turbine shroud to deliver the substantially straight cooling, air streams with a high velocity thereof. 
   
   
       27 . The method as claimed in  claim 26  wherein the cooling air is directed from a cavity defined between front and rear legs of the shroud segment, into the substantially axial and straight passages, and is discharged therefrom at a trailing end of the platform. 
   
   
       28 . The method as claimed in  claim 24  further comprising a step (c) of discharging the cooling air into a gas path upon the impingement cooling thereof on the turbine vane outer shroud. 
   
   
       29 . The method as claimed in  claim 24  comprising directing the substantially straight cooling air streams in a manner for impingement cooling on a substantially entire circumference of the leading end of the turbine vane outer shroud.