[특허]Gas turbine compressor with bleed path

Siemens Aktiengeselischaft

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특허 상세정보

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Gas turbine compressor with bleed path

특허상세정보
국가/구분	United States(US) Patent 등록
국제특허분류(IPC7판)	F02C-006/08 F02C-006/04
미국특허분류(USC)	060/782; 060/785
출원번호	US-0455509 (2012-04-25)
등록번호	US-9032738 (2015-05-19)
발명자 / 주소	Laurello, Vincent P. Tham, Kok-Mun
출원인 / 주소	Siemens Aktiengeselischaft
인용정보	피인용 횟수 : 1 인용 특허 : 11

초록
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A gas turbine engine includes a compressor for generating compressed air. The compressor includes a rotor defined by a plurality of axial disks including a first disk and a second disk. A first row of blades extends radially outwardly from the first disk, and a second row of blades extends radially outwardly from the second disk. A row of cantilevered vanes is located at an axial location between the first row of blades and the second row of blades. A bleed path extends at least partially through the second disk and includes an entrance at an axial locat...

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1. A gas turbine engine comprising a compressor for generating compressed air comprising: a rotor defined by a plurality of disks, wherein said disks comprise a first disk and a second disk;a first row of blades extending radially outwardly from said first disk;a second row of blades extending radially outwardly from said second disk;a row of cantilevered vanes at an axial location between said first row of blades and said second row of blades; anda bleed path extending at least partially through said second disk and comprising an entrance at an axial location between said first row of blades and at least a portion of said row of cantilevered vanes, said entrance communicating with a compressed air flowpath through said compressor;wherein said entrance comprises an annular first passage that defines a free vortex passage, said first passage is located at an outer horizontal surface of said first and second disks and is formed by a gap defined between an upstream boundary on said first disk and a downstream boundary on said second disk, both said upstream and downstream boundaries being at an axial location between said first row of blades and said row of cantilevered vanes. 2. The gas turbine engine as set forth in claim 1, wherein said bleed path further comprises at least one second passage having a first end communicating with said first passage, said second passage extending through said second disk and defining a forced vortex passage. 3. The gas turbine engine as set forth in claim 2, wherein said at least one second passage extends in radial and axial directions and further comprises a second end located downstream from said second row of blades. 4. The gas turbine engine as set forth in claim 2, wherein said at least one second passage comprises two or more second passages, each second passage comprising a first end communicating with said first passage. 5. The gas turbine engine as set forth in claim 4, wherein said two or more second passages are circumferentially spaced apart from one another. 6. The gas turbine engine as set forth in claim 2, wherein said second disk comprises a substantially vertical face, said first end of said at least one second passage being defined in said vertical face. 7. The gas turbine engine as set forth in claim 1, further comprising an inner passage in fluid communication with said bleed path through which compressed air from said compressor travels toward a turbine of the gas turbine engine. 8. A gas turbine engine comprising a compressor for generating compressed air comprising: a rotor defined by a plurality of disks, wherein said disks comprise a first disk and a second disk;a first row of blades extending radially outwardly from said first disk;a second row of blades extending radially outwardly from said second disk, said second disk including a substantially horizontal surface from which said second row of blades extends;a row of cantilevered vanes at an axial location between said first row of blades and said second row of blades;a bleed path extending at least partially through said second disk and comprising an entrance at an axial location between said first row of blades and at least a portion of said row of cantilevered vanes, said entrance communicating with a compressed air flowpath through said compressor; andwherein said bleed path comprises at least one passage having a first end defining said entrance of said bleed path and communicating with said compressed air flowpath, said at least one passage extending through said second disk, said first end of said at least one passage being defined in said horizontal surface at an axial location downstream from an upstream side of said second disk and upstream of said row of cantilevered vanes. 9. The gas turbine engine as set forth in claim 8, wherein said at least one passage extends in radial and axial directions. 10. The gas turbine engine as set forth in claim 8, wherein said at least one passage comprises two or more passages, each passage comprising a first end communicating with said compressed air flowpath. 11. The gas turbine engine as set forth in claim 10, wherein said two or more second passages are circumferentially spaced apart from one another. 12. A method of guiding bleed air from an air flowpath extending through a compressor in a gas turbine engine comprising: providing a rotor at least partially defined by a plurality of axial disks, the disks comprising a first disk from which a first row of blades extends and a second disk from which a second row of blades extends, the second disk being downstream from the first disk with respect to a direction of the air flowpath;providing a row of cantilevered vanes between the first row of blades and the second row of blades; andpassing bleed air from the air flowpath through a bleed path having an entrance at a location downstream from the first row of blades and upstream from at least a portion of the row of cantilevered vanes, at least a portion of the bleed path extending through the second disk;wherein said entrance is located at an outer horizontal surface of said second disk and is at an axial location downstream from a downstream side of said first disk and upstream of said row of cantilevered vanes. 13. The method as set forth in claim 12, wherein said passing bleed air comprises: passing the bleed air from the location downstream from the first row of blades and upstream from at least a portion of the row of cantilevered vanes through an annular first passage of the bleed path; andpassing the bleed air from the annular first passage through at least one second passage of the bleed path defined in the second disk. 14. The method as set forth in claim 13, wherein said passing the bleed air from the annular first passage through at least one second passage comprises passing the bleed air from the annular first passage through a plurality of circumferentially spaced apart second passages. 15. The method as set forth in claim 12, wherein said passing bleed air comprises passing the bleed air through at least one passage of the bleed path having a first end defining the entrance and communicating with the compressed air flowpath, the at least one passage extending in radial and axial directions through the second disk. 16. The method as set forth in claim 12, further comprising passing the bleed air from the bleed path into an inner passage in fluid communication the bleed path, the bleed air flowing through the inner passage toward a turbine of the gas turbine engine.

이 특허에 인용된 특허 (11)

이 특허를 인용한 특허 피인용횟수: 1

King, Jonathan M; Bolgar, Crispin D; Snowsill, Guy D. Gas turbine engine havinga rotatable off-take passage in a compressor section. USP2015089103281.

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Gas turbine compressor with bleed path

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Gas turbine compressor with bleed path

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