초록

A gas turbine engine combustor liner having a plurality of holes defined therein for directing air into the combustion chamber. The plurality of holes provide a greater cooling air flow in fuel nozzle regions than in other areas of the combustor liner.

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The invention claimed is: 1. A gas turbine engine reverse-flow combustor comprising a liner enclosing a combustion chamber, the liner including a dome portion at an upstream end thereof and two annular liner walls extending downstream from and circumscribing said dome portion, the two annular liner

The invention claimed is: 1. A gas turbine engine reverse-flow combustor comprising a liner enclosing a combustion chamber, the liner including a dome portion at an upstream end thereof and two annular liner walls extending downstream from and circumscribing said dome portion, the two annular liner walls being spaced apart such that the dome circumscribed thereby and disposed there between is annular, the dome portion having defined therein a plurality of openings each adapted to receive a fuel nozzle, the openings defined in said dome being positioned substantially equidistant between the two annular liner walls, at least one of said liner walls having a plurality of holes defined therein to form an annular cooling band extending around said liner wall immediately downstream of said dome portion for directing cooling air into the combustion chamber, said plurality of holes within said annular cooling band including a first set of cooling holes disposed within circumferentially spaced regions aligned with said openings and located downstream therefrom and a second set of cooling holes disposed outside said regions, wherein said regions having said first set of cooling holes provide a greater cooling air flow there through than similarly sized areas of said combustor liner having said second set of cooling holes therein. 2. The combustor as defined in claim 1, wherein said regions substantially correspond in shape to a fuel spray cone adapted to be ejected into said combustion chamber from said fuel nozzles. 3. The combustor as defined in claim 2, wherein said regions define a wedge shaped area extending from a narrow upstream end to a wider downstream end. 4. The combustor as defined in claim 1, wherein said first set of cooling holes are defined within said region in a spacing density greater than that of said second set of cooling holes. 5. The combustor as defined in claim 1, wherein each hole of said first set of cooling holes defines a larger cross-sectional opening than that of said second set of cooling holes. 6. The combustor as defined in claim 3, wherein said narrow upstream end is proximate said dome portion. 7. The combustor as defined in claim 3, wherein said wedge shaped areas are substantially trapezoidal. 8. A gas turbine engine reverse-flow combustor comprising an annular liner enclosing a combustion chamber, the liner having an annular dome portion and two liner wall portions spaced apart such that the dome portion is circumscribed thereby and disposed therebetween, the dome portion having defined therein a plurality of openings each adapted to receive a fuel nozzle for directing fuel into the combustion chamber in a spray cone, the openings being positioned in the dome portion substantially equidistant between the two liner wall portions, at least one of the liner wall portions having means for directing cooling air into the combustion chamber, said means providing more cooling air in regions corresponding substantially in shape to said spray cone and located downstream of each opening in alignment therewith. 9. The combustor as defined in claim 8, wherein said means include a plurality of cooling holes, said plurality of holes including first cooling holes disposed within said regions and second cooling holes disposed outside said regions, wherein said first cooling holes provide a greater cooling air flow there through than similarly sized areas of said liner having said second cooling holes therein. 10. The combustor as defined in claim 9, wherein said first cooling holes within said regions are disposed in a spacing density greater than that of said second cooling holes. 11. The combustor as defined in claim 9, wherein each of said first cooling holes defines a larger cross-sectional opening than that of said second cooling holes. 12. The combustor as defined in claim 9, wherein said plurality of holes define an annular cooling band extending around said at least one of the liner walls portions immediately downstream from the dome portion of said liner, said annular cooling band having said regions circumferentially spaced throughout, and said second cooling holes being defined within said annular cooling band between each said regions. 13. An annular combustor for a gas turbine engine comprising: combustor walls including an inner liner and an outer liner spaced apart to define at least a portion of a combustion chamber there between; a plurality of fuel nozzles disposed in openings defined in an annular dome end of the combustor between said inner and outer liners for injecting a spray cone of fuel mixture into the combustion chamber; and a plurality of cooling apertures defined though at least one of said inner and outer liners for delivering pressurized cooling air surrounding said combustor into said combustion chamber, said plurality of cooling apertures defining an annular cooling band extending around said at least one of said inner and outer liners immediately downstream from said annular dome end of the combustor having said fuel nozzles, said cooling apertures in said annular cooling band being provided in one of two spacing densities, thereby providing circumferentially adjacent first and second regions of cooling apertures within the annular cooling band, a boundary between the first and second regions being defined by a change in spacing density of said cooling apertures, the cooling apertures being disposed in multiple rows and having a first spacing density in the first regions of said annular cooling band, each of the first regions being centered on one of said fuel nozzles and the boundary of the first regions in use substantially corresponding in shape to the spray cone of said fuel nozzles, said cooling apertures being disposed in multiple rows and having a second spacing density in the second regions of said annular cooling band, said annular cooling band having said first regions circumferentially spaced throughout and said second regions being all of a remaining portion of the band disposed between each of said first regions, and wherein said first spacing density is greater than said second spacing density. 14. The combustor as defined in claim 13, wherein said first regions substantially correspond in shape to the fuel spray cone injected into the combustion chamber by said fuel nozzles. 15. The combustor as defined in claim 14, wherein said first regions define a wedge shaped area extending from a narrow upstream end to a wider downstream end. 16. The combustor as defined in claim 15, wherein said wedge shaped areas are substantially trapezoidal. 17. The combustor as defined in claim 15, wherein said narrow upstream end is proximate said dome end of the combustor having said fuel nozzles. 18. The combustor of claim 13 wherein the first regions each provide a greater cooling air flow there through than a respective ones of said second region.