초록 ▼

The present invention provides an apparatus and method for reducing the pressure loss of air prior to entering a combustion system, such that, for a known combustion system having a predetermined pressure loss, the resulting fluid entering the turbine has a higher supply pressure that will result in

The present invention provides an apparatus and method for reducing the pressure loss of air prior to entering a combustion system, such that, for a known combustion system having a predetermined pressure loss, the resulting fluid entering the turbine has a higher supply pressure that will result in more efficient turbine and increased engine output. Significant enhancements include the addition of a plurality of deflector assemblies to direct the air from a compressor outlet towards an exposed single-wall transition duct to provide direct cooling to a first panel of the transition duct.

대표청구항 ▼

What we claim is: 1. A gas turbine engine having reduced pressure drop and a lower heat rate, said gas turbine engine comprising: an axial compressor coupled to an axially extending shaft said compressor having a compressor inlet and a compressor outlet; an inner compressor discharge case positione

What we claim is: 1. A gas turbine engine having reduced pressure drop and a lower heat rate, said gas turbine engine comprising: an axial compressor coupled to an axially extending shaft said compressor having a compressor inlet and a compressor outlet; an inner compressor discharge case positioned proximate said compressor outlet for receiving air from said compressor, said compressor discharge case having a compressor discharge end, a turbine inlet end opposite said compressor discharge end, and a radially outer surface extending therebetween, and means for directing said air from said compressor outlet away from said shaft said means attached to said radially outer surface; a plurality of combustors arranged in an annular array about said shaft and fixed to said compressor discharge case, each of said combustors comprising: an outer case; a flow sleeve positioned radially within said outer case; a combustion liner positioned radially within said flow sleeve and having a liner inlet and liner outlet; an end cover fixed to said outer case, said end cover including a plurality of fuel nozzles for injecting fuel into said combustion liner proximate said liner inlet; an exposed single-wall transition duct in fluid communication with said combustion liner, said transition duct comprising a first panel and a second panel, said first panel fixed to said second panel thereby forming a duct having an inner wall, an outer wall, a thickness therebetween, a generally cylindrical duct inlet, and a generally rectangular duct outlet; a turbine coupled to said axially extending shaft for driving said axial compressor; and, wherein said air from said compressor outlet is directed towards said first panel of said transition duct to provide direct cooling to said first panel of said transition duct; said means for directing said air away from said shaft comprises a plurality of deflector assemblies; each of said deflector assemblies comprises: an inner deflector fixed to said radially outer surface of said inner compressor discharge case and extending generally radially outward, said inner deflector having a first inner deflector end, a second inner deflector end, and a deflector surface extending therebetween, said deflector surface including a first inner deflector portion extending from said first inner deflector end and a second inner deflector portion extending from said second inner deflector end; and, wherein said inner deflector is positioned axially such that said first inner deflector end is located adjacent but disconnected from said compressor discharge end, and said second inner deflector end is located adjacent but disconnected from said duct outlet. 2. The gas turbine engine of claim 1 wherein said plurality of deflector assemblies comprises fourteen assemblies. 3. The gas turbine engine of claim 1 wherein each of said deflector assemblies is located between compressor discharge struts. 4. The gas turbine engine of claim 1 wherein said first inner deflector portion is substantially parallel to said shaft and said second inner deflector portion is at an angle a, of between 10 degrees and 70 degrees to said shaft. 5. The gas turbine engine of claim 1 wherein each of said deflector assemblies further comprises an outer vane fixed to a second portion of said compressor discharge case, said outer vane is positioned axially proximate said inner deflector portion and radially outward thereof to form a deflector channel therebetween, said deflector channel having a channel inlet and a channel outlet. 6. The gas turbine engine of claim 5 wherein said deflector channel expands from said channel inlet to said channel outlet.