Patent ID: 8182199

Claim:
A method of cooling a turbine shroud ring within a surrounding turbine support case in a turbine section of a gas turbine engine, the turbine shroud ring including a plurality of individual turbine shroud segments circumferentially arranged to form the turbine shroud ring, at least an inner surface of the shroud ring being exposed to an annular hot gas flow produced from a combustor of the gas turbine engine having a plurality of circumferentially spaced apart fuel nozzles, the method comprising: identifying a series of alternating high temperature regions and lower temperature regions of a circumferential temperature distribution about the inner surface of the shroud ring, the high temperature regions of the circumferential temperature distribution corresponding to circumferential locations of the fuel nozzles in the combustor; and impinging cooling air on to an outer surface of the turbine shroud ring using impingement cooling holes in the turbine support case, including the step of impinging more cooling air to regions corresponding to said high temperature regions on the shroud ring than to regions corresponding to said lower temperature regions of the shroud ring, and wherein impingement cooling of the turbine shroud segments forming the turbine shroud ring is non-uniform for all segments.