Transitional region for a combustion chamber of a gas turbine
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02C-007/24
F23R-003/04
F01D-009/02
F23R-003/00
출원번호
US-0226020
(2011-09-06)
등록번호
US-9097118
(2015-08-04)
우선권정보
EP-10175744 (2010-09-08)
발명자
/ 주소
Schnieder, Martin
Krückels, Jörg
Rüdel, Uwe
Appel, Christoph
Benz, Urs
출원인 / 주소
ALSTOM TECHNOLOGY LTD.
대리인 / 주소
Buchanan Ingersoll & Rooney PC
인용정보
피인용 횟수 :
0인용 특허 :
3
초록▼
A gas turbine including a combustion chamber and a first row of guide vanes, arranged essentially directly downstream thereof, of a turbine. The outer and/or inner limitation of the combustion chamber defined by at least one outer and/or inner heat shield, mounted on at least one combustion chamber
A gas turbine including a combustion chamber and a first row of guide vanes, arranged essentially directly downstream thereof, of a turbine. The outer and/or inner limitation of the combustion chamber defined by at least one outer and/or inner heat shield, mounted on at least one combustion chamber structure arranged radially outside and/or inside. The hot gases flow path in the region of the guide vane row being restricted radially on the outside and/or inside by an outer and/or inner vane platform, mounted at least indirectly on at least one turbine carrier. A minimal gap size directly upstream of the first row of guide vanes is achieved by mounting at least indirectly on the turbine carrier at least one mini heat shield, arranged upstream of the first row of guide vanes and essentially adjacent the vane platform and in the flow direction between the heat shield and the vane platform.
대표청구항▼
1. A gas turbine (22) comprising: at least one combustion chamber (9) and a first row of guide vanes (2) arranged directly downstream of the combustion chamber, wherein at least one of radially outer and radially inner limitations of the combustion chamber (9) being defined by a respective outer or
1. A gas turbine (22) comprising: at least one combustion chamber (9) and a first row of guide vanes (2) arranged directly downstream of the combustion chamber, wherein at least one of radially outer and radially inner limitations of the combustion chamber (9) being defined by a respective outer or inner heat shield (7), which is mounted on at least one combustion chamber structure (6) arranged at least one of radially outside and radially inside, respectively, of the at least one combustion chamber structure; a flow path of hot gases (10) in a region of the first row of guide vanes (2) being restricted radially on the outside and/or radially on the inside by an outer and/or inner vane platform (3), which is mounted at least indirectly on at least one turbine carrier (4);at least one mini heat shield (13) is mounted at least indirectly on the at least one turbine carrier (4), arranged upstream of the first row of guide vanes (2) and essentially adjacent the outer and/or inner vane platform (3), and in the flow path direction (10) between the respective outer and/or inner heat shield (7) and the outer and/or inner vane platform (3), and forming a flow wall therebetween in the form of a shape adapted to the flow, with an upstream gap (17) disposed between the at least one mini heat shield (13) and the respective outer and/or inner heat shield (7), wherein the upstream gap (17) extends directly between and separates the turbine carrier (4, 18) and the combustion chamber structure (6). 2. The gas turbine (22) as claimed in claim 1, wherein the at least one mini heat shield includes a plurality of mini heat shields and at least one of the mini heat shields (13) is mounted on an extension (4′) of the at least one turbine carrier (4) extending upstream with respect to the direction of flow (10) of the hot gases. 3. The gas turbine (22) as claimed in claim 1, wherein the at least one mini heat shield includes a plurality of mini heat shields and at least one of the mini heat shields (13) is mounted on at least one additional turbine carrier element (18), arranged upstream of, and mounted, on the at least one turbine carrier (4). 4. The gas turbine (22) as claimed in claim 1, wherein a plurality of mini heat shields (13) are arranged around a circumference of a wall of the combustion chamber (9) having essentially axially running gaps (24) between them, or gaps running essentially in the main direction of flow (10), while a device that applies a cooling air stream to said gaps is provided, and/or seals are provided in these gaps (24). 5. The gas turbine (22) as claimed in claim 1, further comprising a peripheral gap (23) between the outer and/or inner vane platform (3) and the at least one mini heat shield (13), which has a gap width (d), in an axial direction, in the range of 1-5 mm, and a device that applies a cooling air stream to said peripheral gap (23). 6. The gas turbine (22) as claimed in claim 1, further comprising a peripheral gap (23) between the outer and/or inner vane platform (3) and the at least one mini heat shield (13), which has a gap width (d), in an axial direction, in the range of 2-4 mm, and a device that applies a cooling air stream to said peripheral gap (23). 7. The gas turbine (22) as claimed in claim 1, wherein the at least one mini heat shield (13) has, in an axial direction (25), a length in the range of 5-500 mm. 8. The gas turbine (22) as claimed in claim 1, wherein the at least one mini heat shield (13) has, in an axial direction (25), a length in the range of 10-350 mm. 9. The gas turbine (22) as claimed in claim 1, wherein the upstream gap (17) is arranged at a point at which a wall of the combustion chamber is arranged conically tapering, and a gap size (d′) thereof in a radial direction (26) is in the range of 0.1-20 mm, and/or a gap size (d″) thereof in an axial direction (25) is in the range of 0.1-20 mm. 10. The gas turbine (22) as claimed in claim 1, wherein the upstream gap (17) is arranged at a point at which a wall of the combustion chamber is arranged conically tapering, and a gap size (d′) thereof in a radial direction (26) is in the range of 0.5-20 mm, and/or the gap size (d″) thereof in an axial direction (25) is in a range of 0.5-20 mm. 11. The gas turbine (22) as claimed in claim 1, further comprising a device that applies cooling air to the upstream gap (17), and in particular a cavity (20) arranged behind the upstream gap; and at least one step element, arranged in an entry region of said cavity, which reduces a width of the cavity by at least 10% in at least one step running essentially perpendicularly to the direction of flow of the hot gas in the cavity (20), said element being formed peripherally with respect to the axis of the turbine. 12. The gas turbine (22) as claimed in claim 1, wherein the number of combustion chamber heat shields (7) is an integral multiple of a number of the at least one mini heat shield (13) or the number of the at least one mini heat shield (13) is an integral multiple of the number of combustion chamber heat shields (7). 13. The gas turbine (22) as claimed in claim 1, wherein a number of the at least one mini heat shield (13) is an integral multiple of the number of guide vanes (2) of the first row of guide vanes (13) or the number of guide vanes (2) of the first row of guide vanes is an integral multiple of the number of the at least one mini heat shield (13). 14. The gas turbine (22) as claimed in claim 2, wherein the mini heat shields (13) are formed as individual heat accumulation elements that are adapted to a form of flow and are mounted by way of positively and/or non-positively connecting and/or material-bonder on the turbine carrier (4′) and/or on an additional turbine carrier element (18) mounted on the turbine carrier (4). 15. The gas turbine (22) as claimed in claim 14, wherein at least on a side facing the hot gases flowing in the combustion chamber, the mini heat shields (13) comprise a thermal barrier coating. 16. The gas turbine (22) as claimed in claim 1, wherein the upstream gap (17) extends on a rear side, facing away from the combustion chamber (9), into a cavity (20), the cavity (20) extends essentially in a radial direction (25), in the form of a peripheral gap (20′) running essentially around an axis of the gas turbine (22) and forming a cylindrical enclosure. 17. The gas turbine as claimed in claim 1, wherein a contour of an annular space formed by the at least one mini heat shield (13) is not circular over its entire axial extent, but instead slight protuberances are provided locally, extending out from a circular shape normal to the direction of the flow of hot gas (10), in order to locally increase or reduce static pressure, thereby achieving an overall improvement in uniformity of static pressure distribution in a circumferential direction in a region of the upstream gap (17). 18. The gas turbine as claimed in claim 1, wherein the upstream gap (17) goes over on a rear side, facing away from the combustion chamber (9), into a cavity (20), the cavity (20) extends essentially in a radial direction (26), in the form of a peripheral gap (20″) running essentially around the axis of the gas turbine (22) and forming a circular disk.
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이 특허에 인용된 특허 (3)
Lauck Lawrence J. (South Windsor CT), Gas turbine engine having an improved transition duct support.
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