Combustor arrangement including flow control vanes
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F23R-003/04
F23R-003/42
F23R-003/54
출원번호
US-0506055
(2014-09-05)
등록번호
US-9982893
(2018-05-29)
국제출원번호
PCT/US2014/054214
(2014-09-05)
국제공개번호
WO2016/036381
(2016-03-10)
발명자
/ 주소
Rodriguez, Jose L.
Golsen, Matthew J.
출원인 / 주소
SIEMENS ENERGY, INC.
인용정보
피인용 횟수 :
0인용 특허 :
8
초록▼
A combustor assembly (17) including guide vanes (44) located between an inner cylinder (24) and a flow sleeve (25). Each guide vane (44) includes a circumferentially angled flow directing portion (60) adjacent to a leading edge (46). The leading edge (46) of at least one guide vane (44) can be locat
A combustor assembly (17) including guide vanes (44) located between an inner cylinder (24) and a flow sleeve (25). Each guide vane (44) includes a circumferentially angled flow directing portion (60) adjacent to a leading edge (46). The leading edge (46) of at least one guide vane (44) can be located radially inward along the longitudinal axis (54) relative to the leading edge (46) of at least one other of the guide vanes (44). The length of the guide vanes (44) may vary, and the circumferential spacing between a first pair of the guide vanes (44) can be different from a spacing between a second pair of the guide vanes (44).
대표청구항▼
1. A can annular gas turbine engine having a gas delivery structure for delivering gases from a plurality of combustors each having a longitudinal axis tangentially arranged with respect to a circumference defined by an annular chamber that extends circumferentially about a gas turbine engine axis f
1. A can annular gas turbine engine having a gas delivery structure for delivering gases from a plurality of combustors each having a longitudinal axis tangentially arranged with respect to a circumference defined by an annular chamber that extends circumferentially about a gas turbine engine axis for delivering the gas flow to a first row of blades, a gas flow path formed by a duct arrangement between a respective combustor of the plurality of combustors and the annular chamber for conveying gases in a downstream direction from the respective combustor to the first row of turbine blades, a combustor arrangement comprising: the respective combustor having an inner cylinder surrounding a combustion zone of the combustor, the inner cylinder defining the longitudinal axis of the respective combustor that is tangentially arranged with respect to the circumference defined by the annular chamber;a cone-shaped section having an inlet end receiving the gas flow from the inner cylinder, wherein the cone section defines a decreasing flow area in the downstream direction between the inlet end and an outlet end of the cone-shaped section;a flow sleeve surrounding the inner cylinder;an annular space defined between the inner cylinder and the flow sleeve and defining an air flow path having an annular air inlet defined along the inner cylinder radially inward along the longitudinal axis of a junction with respect to the gas turbine engine axis between the inner cylinder and the cone section; anda plurality of guide vanes located in circumferentially spaced relation to each other in the annular space, spanning between the flow sleeve and the inner cylinder, the guide vanes each having a length dimension defined between a radially inner leading edge and a radially outer trailing edge, wherein the leading edges of the guide vanes are located along the longitudinal axis radially inward with respect to the gas turbine engine axis from the junction between the inner cylinder and the cone section. 2. The combustor arrangement of claim 1, wherein the leading edge of at least one guide vane is located radially inward along the longitudinal axis relative to the leading edge of at least one other of the guide vanes with respect to the gas turbine engine axis. 3. The combustor arrangement of claim 2, wherein air flowing into the annular space has a circumferential swirl flow direction around a circumference of the inner cylinder, and the at least one guide vane is located in an upstream direction of the swirl flow from the at least one other of the guide vanes. 4. The combustor arrangement of claim 3, wherein at least two of the guide vanes have leading edges located radially inward along the longitudinal axis with respect to the gas turbine engine axis relative to at least two other of the guide vanes. 5. The combustor arrangement of claim 4, wherein the at least two of the guide vanes have trailing edges located radially inward along the longitudinal axis with respect to the gas turbine engine axis relative to the trailing edges of the at least two other guide vanes. 6. The combustor arrangement of claim 4, wherein a first guide vane of the plurality of guide vanes has a first leading edge which is located along the longitudinal axis radially inward of the leading edges of all the other guide vanes, and has a length that is less than half the length of the other guide vanes. 7. The combustor arrangement of claim 4, wherein a first guide vane is circumferentially located at an upstream-most location and is spaced from an adjacent guide vane a distance that is less than a circumferential spacing between other adjacent guide vanes. 8. The combustor arrangement of claim 4, wherein each of the guide vanes include a circumferentially angled flow directing portion and an angle of each of the circumferentially angled flow directing portions of the at least two guide vanes, as measured relative to the longitudinal axis, is greater than an angle of each of the circumferentially angled flow directing portions of the at least two other of the guide vanes. 9. The combustor arrangement of claim 1, wherein all of the guide vanes are located in a circumferential area that is less than 180 degrees around the circumference of the inner cylinder. 10. The combustor arrangement of claim 1, wherein a radially inner end of the flow sleeve is located radially inward along the longitudinal axis with respect to the gas turbine engine axis from the junction between the inner cylinder and the cone section. 11. The combustor arrangement of claim 1, wherein one or more of the guide vanes includes a straight main body aligned parallel to the longitudinal axis of the combustor and extending from the trailing edge to a radially inner intermediate location, and a circumferentially angled flow directing portion extending from the intermediate location to the leading edge. 12. A can annular gas turbine engine having a gas delivery structure for delivering gases from a plurality of combustors each having a longitudinal axis tangentially arranged with respect to a circumference defined by an annular chamber that extends circumferentially about a gas turbine engine axis for delivering the gas flow to a first row of blades, a gas flow path formed by a duct arrangement between a respective combustor of the plurality of combustors and the annular chamber for conveying gases in a downstream direction from the respective combustor to the first row of turbine blades, a combustor arrangement comprising: the respective combustor having an inner cylinder surrounding a combustion zone of the combustor, the inner cylinder defining the longitudinal axis of the respective combustor that is tangentially arranged with respect to the circumference defined by the annular chamber;a cone-shaped section having an inlet end receiving the gas flow from the inner cylinder, wherein the cone section defines a decreasing flow area in the downstream direction between the inlet end and an outlet end of the cone-shaped section;a flow sleeve surrounding the inner cylinder;an annular space defined between the inner cylinder and the flow sleeve and defining an air flow path having an annular air inlet defined along the inner cylinder radially inward along the longitudinal axis of a junction with respect to the gas turbine engine axis between the inner cylinder and the cone section; anda plurality of guide vanes located in circumferentially spaced relation to each other in the annular space, spanning between the flow sleeve and the inner cylinder, the guide vanes each having a length dimension defined between a radially inner leading edge and a radially outer trailing edge, wherein the leading edge of at least one guide vane is located radially inward along the longitudinal axis relative to the leading edge of at least one other of the guide vanes with respect to the gas turbine engine axis. 13. The combustor arrangement of claim 12, wherein air flowing into the annular space has a circumferential swirl flow direction around a circumference of the inner cylinder, and the at least one guide vane is located in an upstream direction of the swirl flow from the at least one other of the guide vanes. 14. The combustor arrangement of claim 13, wherein a first pair of the guide vanes have leading edges located radially inward along the longitudinal axis relative to a second pair of the guide vanes, and the first pair of guide vanes is located in the upstream direction of the swirl flow from the second pair of guide vanes. 15. The combustor arrangement of claim 14, wherein a first, upstream-most guide vane of the first pair of guide vanes has a length that is less than the length of the other guide vanes. 16. The combustor arrangement of claim 14, wherein each of the guide vanes include a circumferentially angled flow directing portion and an angle of the circumferentially angled flow directing portions of each of the guide vanes in the first pair of guide vanes, as measured relative to the longitudinal axis, is greater than an angle of each of the circumferentially angled flow directing portions of the guide vanes in the second pair of guide vanes. 17. A can annular gas turbine engine having a gas delivery structure for delivering gases from a plurality of combustors each having a longitudinal axis tangentially arranged with respect to a circumference defined by an annular chamber that extends circumferentially about a gas turbine engine axis for delivering the gas flow to a first row of blades, a gas flow path formed by a duct arrangement between a respective combustor of the plurality of combustors and the annular chamber for conveying gases in a downstream direction from the respective combustor to the first row of turbine blades, a combustor arrangement comprising: the respective combustor having an inner cylinder surrounding a combustion zone of the combustor, the inner cylinder defining the longitudinal axis of the respective combustor that is tangentially arranged with respect to the circumference defined by the annular chamber;a cone-shaped section having an inlet end receiving the gas flow from the inner cylinder, wherein the cone section defines a decreasing flow area in the downstream direction between the inlet end and an outlet end of the cone-shaped section;a flow sleeve surrounding the inner cylinder;an annular space defined between the inner cylinder and the flow sleeve and defining an air flow path having an annular air inlet defined along the inner cylinder radially inward along the longitudinal axis of a junction with respect to the gas turbine engine axis between the inner cylinder and the cone section;a plurality of guide vanes located in circumferentially spaced relation to each other in the annular space, spanning between the flow sleeve and the inner cylinder, the guide vanes each having a length dimension defined between a radially inner leading edge and a radially outer trailing edge, each guide vane includes a circumferentially angled flow directing portion; andwherein air flowing into the annular space has a circumferential swirl flow direction around a circumference of the inner cylinder, and a first pair of guide vanes are circumferentially spaced apart a distance that is less than a circumferential spacing between a second pair of guide vanes adjacent to the first pair of guide vanes, the first pair of guide vanes being located in an upstream direction of the swirl flow relative to the second pair of guide vanes. 18. The combustor arrangement of claim 17, wherein leading edges of the first pair of guide vanes is located radially inward along the longitudinal axis from leading edges of the second pair of guide vanes with respect to the gas turbine engine axis. 19. The combustor arrangement of claim 18, wherein a first, upstream-most guide vane of the first pair of guide vanes has a length that is less than the length of the other guide vanes.
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이 특허에 인용된 특허 (8)
Martling, Vincent C.; Xiao, Zhenhua, Airflow distribution to a low emissions combustor.
Charron, Richard C.; Nordlund, Raymond S.; Morrison, Jay A.; Campbell, Ernie B.; Pierce, Daniel J.; Montgomery, Matthew D.; Wilson, Jody W., Assembly for directing combustion gas.
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