초록 ▼

A method of cooling a shroud ring in a turbine section of gas turbine engine includes 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, and impinging cooling air on to a

A method of cooling a shroud ring in a turbine section of gas turbine engine includes 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, and impinging cooling air on to an outer surface of the shroud ring. More cooling air is impinged onto regions which correspond to the high temperature regions on the shroud ring than to regions corresponding to the lower temperature regions of the shroud ring.

대표청구항 ▼

1. 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 surfa

1. 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; andimpinging 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. 2. The method as defined in claim 1, further comprising determining an amount of cooling air flow required to reduce the temperature in said high temperature regions such that an axial temperature distribution across the shroud ring is substantially uniform. 3. The method as defined in claim 1, further comprising determining a desired distribution of the cooling air flow on the turbine shroud ring by clocking the targeted locations for cooling to the circumferential positions of the identified local high temperature regions. 4. The method as defined in claim 1, wherein the step of providing more impingement cooling air to the targeted locations further comprising providing a plurality of said impingement cooling holes in the turbine support case surrounding said shroud ring, and grouping said impingement cooing holes in greater concentrations in areas of the turbine support case which circumferentially align with the targeted locations on the shroud ring. 5. The method as defined in claim 1, further comprising selecting an engine running condition during which the circumferential temperature distribution is determined. 6. The method as defined in claim 1, further comprising determining a desired distribution of the cooling air flow which permits impinging more cooling air onto the regions corresponding to said high temperature regions on the shroud ring. 7. The method as defined in claim 6, further comprising targeting said regions corresponding to said high temperature regions by circumferentially aligning said cooling air flow to the high temperature regions on the shroud ring. 8. The method as defined in claim 1, wherein identifying the high temperature regions includes locating an axial coordinate of a center of each of the identified local high temperature regions at identified circumferential positions of the high temperature regions. 9. The method as defined in claim 1, further comprising selecting an amount of cooling air flow such that an axial temperature distribution across the shroud ring is substantially uniform. 10. An impingement cooling system for cooling a static annular turbine shroud ring located downstream of a combustor in a gas turbine engine and surrounding a turbine rotor, the turbine shroud ring including a plurality of individual turbine shroud segments circumferentially arranged to form the annular turbine shroud ring, the combustor including a plurality of circumferentially spaced apart fuel nozzles, the turbine shroud ring being exposed to an annular hot gas flow produced from said combustor, the system comprising a turbine support casing which surrounds the turbine shroud ring, the turbine support casing having a plurality of impingement cooling holes defined therethrough which direct cooling air from a pressurized air source onto a radially outer surface of the turbine shroud ring for impingement cooling of the turbine shroud ring, said impingement cooling holes in the turbine support casing being arranged in an annular band about a circumference of the turbine support casing assembly and including alternating first and second groups of holes, the first groups of holes in use providing more impingement cooling air onto the turbine shroud ring than the second groups of holes, the first groups of holes being provided in varying density, with regions of higher density circumferentially aligned with discrete high temperature regions distributed in use circumferentially around the component, the high temperature regions of the component correspond to periodic high temperature regions in the hot gas flow resulting from the circumferentially spaced apart locations of the fuel nozzles in the combustor, and wherein said impingement cooling holes in the turbine support casing are arranged such that the impingement cooling of the turbine shroud segments forming the turbine shroud ring is non-uniform for all segments and at least one of said discrete high temperature regions extends across one or more of said turbine shroud segments. 11. The system as defined in claim 10, wherein a greater concentration of the first groups of holes are disposed in the casing at a greater concentration than the second groups of holes. 12. The system as defined in claim 10, wherein the annular band of the impingement cooling holes includes a first row of holes and a second row of holes downstream from the first row of holes. 13. The system as defined in claim 12, wherein holes of the first row of holes are circumferentially aligned with each other and holes of the second row of holes are circumferentially aligned with each other. 14. The system as defined in claim 13, wherein the holes of the second row of holes are axially offset from those of the first row of holes. 15. A gas turbine engine comprising: a compressor, a combustor and a turbine serially connected to one another in flow communication, the turbine section including an annular turbine shroud ring for surrounding a stage of turbine blades, the turbine shroud ring including a plurality of individual turbine shroud segments circumferentially arranged to form the annular turbine shroud ring, the turbine shroud ring being concentrically mounted within a supporting turbine support case, the turbine support case having a plurality of impingement cooling holes extending between an inner and an outer surface of the turbine support case and being arranged in a circumferentially extending band, the impingement cooling holes being fed with cooling air from a source disposed outside the turbine support case and directing said cooling air through the turbine support case and onto targeted locations on an outer surface of the turbine shroud ring for impingement cooling of said targeted locations, said impingement cooling holes being arranged in a configuration having first zones of holes which provide more cooling air through the turbine support case and second zones of holes which provide less cooling air through the turbine support case than the first zones, the first and second zones alternating about the circumference of the band, the first zones being circumferentially aligned in the turbine support case to correspond to identified circumferentially spaced high temperature regions of the turbine shroud ring which correspond to circumferentially spaced apart locations of the fuel nozzles in the combustor, and wherein said impingement cooling holes in the turbine support case are arranged such that the impingement cooling of the individual turbine shroud segments forming the turbine shroud ring is non-uniform for all segments and at least one of said high temperature regions cooled by the first zone of holes extends across one or more of said individual turbine shroud segments.