Gas turbine engine with radial diffuser and shortened mid section
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F01D-009/02
F01D-009/04
F02C-003/14
F23R-003/46
출원번호
US-0602422
(2012-09-04)
등록번호
US-9127554
(2015-09-08)
발명자
/ 주소
Charron, Richard C.
Montgomery, Matthew D.
출원인 / 주소
Siemens Energy, Inc.
인용정보
피인용 횟수 :
1인용 특허 :
3
초록▼
An industrial gas turbine engine (10), including: a can annular combustion assembly (80), having a plurality of discrete flow ducts configured to receive combustion gas from respective combustors (82) and deliver the combustion gas along a straight flow path at a speed and orientation appropriate fo
An industrial gas turbine engine (10), including: a can annular combustion assembly (80), having a plurality of discrete flow ducts configured to receive combustion gas from respective combustors (82) and deliver the combustion gas along a straight flow path at a speed and orientation appropriate for delivery directly onto the first row (56) of turbine blades (62); and a compressor diffuser (32) having a redirecting surface (130, 140) configured to receive an axial flow of compressed air and redirect the axial flow of compressed air radially outward.
대표청구항▼
1. An industrial gas turbine engine, comprising: a can annular combustion assembly, comprising a plurality of discrete flow ducts configured to receive combustion gas from respective combustors and deliver the combustion gas along a straight flow path at a speed and orientation appropriate for deliv
1. An industrial gas turbine engine, comprising: a can annular combustion assembly, comprising a plurality of discrete flow ducts configured to receive combustion gas from respective combustors and deliver the combustion gas along a straight flow path at a speed and orientation appropriate for delivery directly onto a first row of turbine blades without guide vanes, and an annular chamber configured to merge the plurality of discrete flow ducts into a single, annular flow duct defining an annular flow path immediately upstream of the first row of turbine blades; anda compressor diffuser positioned at a downstream end of an axial compressor and configured to receive an axial flow of compressed air from the axial compressor, the compressor diffuser comprising a diffuser radially outer wall, a diffuser radially inner wall, and a diffuser outlet there between, the diffuser radially inner wall comprising a redirecting surface configured to receive the axial flow of compressed air and redirect the axial flow of compressed air radially outward along a path that bypasses the annular chamber;wherein the axial flow of compressed air exiting the downstream end of the axial compressor comprises all compressed air exiting the downstream end of the axial compressor;wherein the redirecting surface extends in a radial direction, relative to a longitudinal axis of the industrial gas turbine engine, beyond a radially inner surface of the annular chamber, and into a plenum surrounding the can annular combustion assembly. 2. The industrial gas turbine engine of claim 1, wherein a combustion section length between a trailing edge of a last row of compressor airfoils and a leading edge of the first row of turbine blades is not more than 20% of an engine length between a leading edge of a first row of compressor airfoils and a trailing edge of a last row of turbine airfoils. 3. The industrial gas turbine engine of claim 1, wherein an engine length between a leading edge of a first row of compressor airfoils and a trailing edge of a last row of turbine airfoils is at least 5 meters and a combustion section length between a trailing edge of a last row of compressor airfoils and a leading edge of the first row of turbine blades is less than 1 meter. 4. The industrial gas turbine engine of claim 1, wherein an engine length between a leading edge of a first row of compressor airfoils and a trailing edge of a last row of turbine airfoils is at least 6 meters and a combustion section length between a trailing edge of a last row of compressor airfoils and a leading edge of the first row of turbine blades is less than 1.2 meters. 5. The industrial gas turbine engine of claim 2, wherein an engine output of the industrial gas turbine engine is less than 75 MW. 6. The industrial gas turbine engine of claim 1, wherein an engine output of the industrial gas turbine engine is less than 75 MW. 7. The industrial gas turbine engine of claim 5, further comprising a rotor shaft supported by hydrodynamic bearings. 8. The industrial gas turbine engine of claim 1, wherein the compressor diffuser axially overlaps the can annular combustion assembly. 9. The industrial gas turbine engine of claim 1, wherein the redirecting surface is curved and wherein the curved redirecting surface redirects the axial flow of compressed air more than 90 degrees. 10. The industrial gas turbine engine of claim 1, wherein the redirecting surface is curved and wherein the curved redirecting surface redirects the axial flow of compressed air in a direction essentially parallel to a combustor can longitudinal axis. 11. The industrial gas turbine engine of claim 1, wherein the compressor diffuser redirects compressed air into top hats enclosing the combustors. 12. The industrial gas turbine engine of claim 1, wherein the redirecting surface comprises a conically diverging surface. 13. The industrial gas turbine engine of claim 1, wherein the redirecting surface comprises an arcuately diverging surface, and wherein the diffuser outer wall comprises an arcuately diverging surface. 14. An industrial gas turbine engine, comprising: a can annular combustion assembly comprising a plurality of discrete and straight flow ducts configured to receive combustion gas from respective combustors, and an annular chamber configured to merge the flow ducts and deliver the combustion gas directly onto a first row of turbine blades, wherein the can annular combustion assembly accelerates and orients the combustion gas without guide vanes; anda radial diffuser positioned at a downstream end of an axial compressor and configured to receive an axial flow of compressed air from the axial compressor, the radial diffuser comprising a diffuser radially outer wall, a diffuser radially inner wall, and a diffuser outlet there between, the diffuser radially inner wall configured to receive the axial flow of compressed air exiting the axial compressor and redirect it radially outward of the annular chamber via a redirecting surface, wherein a downstream end of the redirecting surface is disposed to extend in a radial direction, with respect to a longitudinal axis of the industrial gas turbine engine, beyond a radially inner surface of the annular chamber;wherein the axial flow of compressed air exiting the downstream end of the axial compressor comprises all compressed air exiting the downstream end of the axial compressor. 15. The industrial gas turbine engine of claim 14, wherein the can annular combustion assembly comprises an acceleration geometry for each flow duct of the plurality of discrete and straight flow ducts that is configured to accelerate the combustion gas to approximately 0.8 mach. 16. The industrial gas turbine engine of claim 14, wherein the redirecting surface comprises an arcuate surface that diverges radially outward toward a downstream end of the radial diffuser, wherein the arcuate surface directs the compressed air into top hats surrounding combustor inlets. 17. The industrial gas turbine engine of claim 16, the arcuate surface comprising a conically diverging surface. 18. The industrial gas turbine engine of claim 14, wherein the industrial gas turbine engine is rated for a maximum output of at least 75 MW. 19. The industrial gas turbine engine of claim 1, wherein each flow duct of the plurality of flow ducts traverses the redirecting surface. 20. The industrial gas turbine engine of claim 1, wherein the redirecting surface is disposed between the diffuser outlet and the annular chamber.
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이 특허에 인용된 특허 (3)
Bancalari, Eduardo; Wilson, Jody; Little, David A.; Fadok, Joseph, Combustion transition duct providing stage 1 tangential turning for turbine engines.
McCarty William L. (West Chester FL) Wescott Kermit R. (Winter Springs FL), Gas turbine control system having maximum instantaneous load-pickup limiter.
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