Cool tip combustor

Abstract

A gas turbine engine is provided with a means for cooling tips of turbine blades in a hot turbine section of the engine. Inlet air holes are provided in a radially outer wall section of the combustor, just upstream of a turbine nozzle. Cooling air flows through these inlet air holes, into annulus regions protected from combustion gases, and then downstream along a radially outer wall of the turbine. The cooling air forms a film that cools the turbine blade tips in a localized manner that adds to total engine power output.

Having described the invention, what is claimed as novel and desired to be secured by Letters Patent of the United States is: 
     
       1. In a gas turbine engine having a compressor, a combustor, a turbine section with a turbine nozzle and turbine blades, all in serial flow relationship and disposed radially about an engine centerline, means for cooling tips of said turbine blades comprising: an inner wall section extending radially inwardly and in a downstream direction from a radially outer wall section of a downstream end portion of said combustor, said inner and outer wall sections defining a first annulus therebetween, said outer wall section including a plurality of first inlet air holes in flow communication with said first annulus, said first holes being sized for channeling a predetermined amount of high-pressure cooling air from said compressor, for forming a cooling air film that extends downstream from said first annulus and along a radially outer wall of said turbine section to cool said turbine blade tips.   
     
     
       2. The gas turbine engine recited in claim 1 wherein said first annulus and said inner wall section extend substantially to an inlet of said turbine nozzle for thereby preventing combustion of said cooling air in said combustor. 
     
     
       3. The gas turbine engine recited in claim 1 wherein said first inlet air holes include an intermediate row of inlet air holes extending circumferentially around said combustor and, additionally, a downstream row of inlet air holes extending circumferentially around said combustor. 
     
     
       4. In a gas turbine engine having a compressor, a combustor, a turbine section with a turbine nozzle and turbine blades, all in serial flow relationship and disposed about an engine centerline, means for providing a film of cooling air along a radially outer wall of said turbine section for cooling tips of said turbine blades, said means comprising: an upstream row of inlet air holes extending circumferentially around a radially outer wall section of said combustor, wherein compressor discharge air is directed into an annulus between a flange and an inner wall of said combustor,   an intermediate row of inlet air holes and a downstream row of inlet air holes, both rows extending circumferentially around said radially outer wall section of said combustor, wherein compressor discharge air is directed from said intermediate row and said downstream row into another annulus between said inner wall and said outer wall section of said combustor;   said annuli being open in a downstream direction, and said intermediate and downstream rows of inlet air holes being sized for collectively channeling a predetermined amount of said compressor discharge air for forming a film of cooling air that extends downstream along a radially outer wall of said turbine section thereby cooling tips of said turbine blades.   
     
     
       5. The gas turbine engine recited in claim 1 further including a flange extending from said outer wall section in a downstream direction, said flange being spaced radially inwardly from said inner wall section for defining a second annulus therebetween, said outer wall section including a plurality of second inlet air holes in flow communication with said second annulus, said second holes being sized for channeling a predetermined amount of said cooling air for protecting said inner wall section from combustion gases.