Cooling air temperature reduction using nozzles
IPC분류정보
국가/구분 |
United States(US) Patent
등록
|
국제특허분류(IPC7판) |
|
출원번호 |
US-0077475
(2013-11-12)
|
등록번호 |
US-9644539
(2017-05-09)
|
발명자
/ 주소 |
- Heinrich, Chad W
- Holland, Stephen Erick
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출원인 / 주소 |
|
인용정보 |
피인용 횟수 :
0 인용 특허 :
6 |
초록
▼
A converging-diverging nozzle that has particular application for providing a cooling air flow to ring segments in a gas turbine engine. The engine includes a turbine section that receives a hot working gas. The turbine section includes at least one row of vanes, at least one row of blades and a plu
A converging-diverging nozzle that has particular application for providing a cooling air flow to ring segments in a gas turbine engine. The engine includes a turbine section that receives a hot working gas. The turbine section includes at least one row of vanes, at least one row of blades and a plurality of ring segments forming at least one ring. The ring segments and the vanes are mounted to a vane carrier, where the vane carrier includes a cooling flow channel for each of the ring segments that receives an air flow to cool the ring segments. A plug is provided in each channel and has an internal bore shaped to define the converging-diverging nozzle through which the air flows so as to create a supersonic flow that reduces the temperature of the air and thus provides more cooling for the same amount of air flow.
대표청구항
▼
1. A gas turbine engine comprising: a compressor section being operable to produce a compressed gas;a combustion section in fluid communication with the compressor section that receives the compressed gas, said combustion section mixing the compressed gas with a fuel and combusting the compressed ga
1. A gas turbine engine comprising: a compressor section being operable to produce a compressed gas;a combustion section in fluid communication with the compressor section that receives the compressed gas, said combustion section mixing the compressed gas with a fuel and combusting the compressed gas and fuel mixture to produce a hot working fluid; anda turbine section in fluid communication with the combustion section, said turbine section receiving the hot working fluid to produce mechanical power, said turbine section including at least one row of vanes, at least one row of blades and a plurality of ring segments forming at least one ring, wherein the ring segments provide a sealing structure at an end of the blades, said ring segments and said at least one row of vanes being mounted to a vane carrier, said vane carrier including a ring cooling channel for each ring segment that receives a portion of the compressed gas to cool the ring segments, said turbine section further including a plug positioned in each ring cooling channel and having an internal flow channel, said internal flow channel defining a converging-diverging nozzle through which the compressed gas flows so as to create a supersonic flow that reduces the temperature of the compressed gas. 2. The gas turbine engine according to claim 1 wherein the internal flow channel includes a wide end section coupled to a converging section. 3. The gas turbine engine according to claim 2 wherein the internal flow channel includes a narrow throat section coupled to the converging section at one end and a diverging section at an opposite end. 4. The gas turbine engine according to claim 3 wherein the internal flow channel includes a wide output end section at an output end of the nozzle coupled to the diverging section where the supersonic flow occurs. 5. The gas turbine engine according to claim 1 wherein the turbine section includes four rows of vanes, four rows of blades and four rows of rings. 6. The gas turbine engine according to claim 5 wherein the ring cooling channels are only provided for the first three rows of rings. 7. The gas turbine engine according to claim 1 wherein the compressed gas is air. 8. A gas turbine engine comprising a turbine section receiving a hot working fluid, said turbine section including at least one row of vanes, at least one row of blades and a plurality of ring segments forming at least one ring, wherein the ring segments provide a sealing structure at an end of the blades, said turbine section further including at least one cooling flow channel providing a flow of cooling air to one or more of the vanes, the blades or the ring segments, said cooling flow channel including a converging-diverging nozzle through which the cooling air flows, said nozzle being configured to receive a subsonic flow of the cooling air and to create a supersonic flow of the cooling air that reduces the temperature of the cooling air. 9. The gas turbine engine according to claim 8 wherein the cooling flow channel includes a wide end section coupled to a converging section. 10. The gas turbine engine according to claim 9 wherein the cooling flow channel includes a narrow throat section coupled to the converging section at one end and a diverging section at an opposite end. 11. The gas turbine engine according to claim 10 wherein the cooling flow channel includes a wide output end section at an output end of the nozzle coupled to the diverging section where the supersonic flow occurs. 12. The gas turbine engine according to claim 8 wherein the turbine section includes four rows of vanes, four rows of blades and four rows of rings. 13. The gas turbine engine according to claim 8 wherein the at least one cooling flow channel provides a flow of cooling air to the ring segments. 14. The gas turbine engine according to claim 13 wherein the nozzle is provided within a plug positioned within the cooling flow channel. 15. A gas turbine engine comprising: a compressor section being operable to produce a compressed gas;a combustion section in fluid communication with the compressor section that receives the compressed gas, said combustion section mixing the compressed gas with a fuel and combusting the compressed gas and fuel mixture to produce a hot working fluid; anda turbine section in fluid communication with the combustion section, said turbine section receiving the hot working fluid to produce mechanical power, said turbine section including at least one row of vanes, at least one row of blades and a plurality of ring segments forming at least one ring, wherein the ring segments provide a sealing structure at an end of the blades, said ring segments and said at least one row of vanes being mounted to a vane carrier, said vane carrier including a ring cooling channel for each ring segment that receives a portion of the compressed gas to cool the ring segments, said turbine section further including a plug positioned in each ring cooling channel and having an internal flow channel, said internal flow channel defining a converging-diverging nozzle including a wide end section coupled to a converging section, a narrow throat section coupled to the converging section at one end and a diverging section at an opposite end and a wide output end section at an output end of the nozzle coupled to the diverging section, wherein the compressed air flows at supersonic speeds in the diverging section that reduces the temperature of the compressed gas. 16. The gas turbine engine according to claim 15 wherein the turbine section includes four rows of vanes, four rows of blades and four rows of rings. 17. The gas turbine engine according to claim 15 wherein the ring cooling channels are only provided for the first three rows of rings.
이 특허에 인용된 특허 (6)
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Eskesen John H. (Schenectady NY) Leibowitz Herman M. (Schenectady NY), Bar for sealing the gap between adjacent shroud plates in liquid-cooled gas turbine.
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Giri, Sheo Narain, Cooling hole exits for a turbine bucket tip shroud.
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Liotta Gary C. ; Acquaviva Paul J., Dual cooled shroud.
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Stroud David (Bristol GB2) Corfe Arthur G. (Bristol GB2) Towill Jonathan P. W. (Cardiff GB7) Cooper Brian G. (Derby GB2), Film cooled components.
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Hoffmann, Juergen; Rofka, Stefan; Waelchli, Rene; Dittmann, Rolf, Gas turbine set.
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Rice Ivan G. (P.O. Box 233 Spring TX 77373), Integrated gas/steam nozzle.
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