[특허]Hot gas path component cooling for hybrid pulse detonation combustion systems

General Electric Company

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특허 상세정보

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Hot gas path component cooling for hybrid pulse detonation combustion systems

특허상세정보
국가/구분	United States(US) Patent 등록
국제특허분류(IPC7판)	F02C-006/08 F02C-005/02 F02C-007/18 F23C-015/00 F02C-009/18
미국특허분류(USC)	060/782; 060/039.76; 060/039.38; 060/247; 060/248; 060/806
출원번호	US-0974085 (2010-12-21)
등록번호	US-8978387 (2015-03-17)
발명자 / 주소	Tangirala, Venkat Eswarlu Joshi, Narendra Digamber Rasheed, Adam Brzek, Brian Gene Hofer, Douglas Carl Lavertu, Thomas Michael Ma, Fuhua
출원인 / 주소	General Electric Company
대리인 / 주소	McCarthy, Robert M.
인용정보	피인용 횟수 : 0 인용 특허 : 15

초록

The flow through the core of a hybrid pulse detonation combustion system is passed through a compressor and then separated into a primary flow, that passes directly to the combustor, and a bypass flow, which is routed to a portion of the system to be used to cool components of the system. The bypass flow is routed to a nozzle of the pulse detonation combustor. The flow is then passed back into the primary flow through the core downstream of where it was extracted.

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1. A gas turbine comprising: a pulse detonation combustor, the combustor comprising a combustion tube and a nozzle disposed downstream of the combustion tube;a compressor in fluid communication with the combustor and disposed upstream of the combustor;a turbine in fluid communication with the combustor and disposed downstream of the combustor;a bypass in fluid communication with the compressor, a booster pump, an external surface of the nozzle, a plurality of fins disposed on an inner surface of the nozzle, and an exterior surface of the combustion tube of the pulse detonation combustor, the booster pump disposed downstream of the compressor and upstream of the external surface of the nozzle; andan internal passage in each of the plurality of fins fluidly connecting the bypass to the inner surface of the nozzle through openings in each of the plurality of fins;wherein a portion of a primary flow from the compressor flows to the booster pump as a bypass flow, the booster pump increases the pressure of the bypass flow allowing the increased pressure bypass flow to flow in a downstream direction to the external surface of the nozzle and then in an upstream direction along the external surface of the combustion tube and reintegrated with the primary flow from the compressor upstream of the combustion tube, at the external surface of the nozzle a portion of the increased pressure bypass flow passes into the internal passage in each of the plurality of fins and out through the openings to mix with an exhaust flow from the combustor. 2. A gas turbine as in claim 1, wherein the fins increase the surface area of the nozzle. 3. A gas turbine as in claim 1, further comprising external fins disposed on the external surface of the nozzle, the increased pressure bypass flow passes over the external fins and the external surface of the nozzle to provide enhanced convective cooling of the nozzle. 4. A gas turbine as in claim 1, wherein the openings in each of the plurality of fins provide film cooling of the fins. 5. A gas turbine as in claim 1, wherein the bypass flow through the bypass is equal to about 10 percent of an inlet flow entering the compressor. 6. A gas turbine as in claim 1 further comprising a chiller configured to reduce the temperature of the bypass flow through the bypass, the chiller disposed along the bypass downstream of the compressor and upstream of the external surface of the nozzle. 7. A gas turbine as in claim 1, wherein the bypass is further in fluid communication with the turbine. 8. A gas turbine as in claim 7, wherein the bypass flow through the bypass is equal to about 12 percent of an inlet flow entering the compressor. 9. A cooling system for a pulse detonation combustor having a combustion tube and a nozzle disposed downstream of the combustion tube, the cooling system comprising; a compressor in fluid communication with the pulse detonation combustor and disposed upstream of the pulse detonation combustor;a bypass in fluid communication with the compressor, a booster pump, an external surface of the nozzle, a plurality of fins disposed on an inner surface of the nozzle, and an exterior surface of the combustion tube of the pulse detonation combustor, the booster pump disposed downstream of the compressor and upstream of the external surface of the nozzle; andan internal passage in each of the plurality of fins fluidly connecting the bypass to the inner surface of the nozzle through openings in each of the plurality of fins;wherein a portion of a primary flow from the compressor flows to the booster pump as a bypass flow, the booster pump increases the pressure of the bypass flow allowing the increased pressure bypass flow to flow in a downstream direction to the external surface of the nozzle and then in an upstream direction along the external surface of the combustion tube and reintegrated with the primary flow from the compressor upstream of the combustion tube, at the external surface of the nozzle a portion of the increased pressure bypass flow passes into the internal passage in each of the plurality of fins and out through the openings to mix with an exhaust flow from the combustor. 10. A cooling system as in claim 9, wherein the fins increase the surface area of the nozzle. 11. A cooling system as in claim 9, further comprising external fins disposed on the external surface of the nozzle, the increased pressure bypass flow passes over the external fins and the external surface of the nozzle to provide enhanced convective cooling of the nozzle. 12. A cooling system as in claim 9, wherein the openings in each of the plurality of fins provide film cooling of the fins. 13. A cooling system as in claim 9, wherein the bypass flow through the bypass is equal to about 10 percent of an inlet flow entering the compressor. 14. A cooling system as in claim 9 further comprising a chiller configured to reduce the temperature of the bypass flow through the bypass, the chiller disposed along the bypass downstream of the compressor and upstream of the external surface of the nozzle. 15. A gas turbine comprising: a pulse detonation combustor, the combustor comprising a combustion tube and a nozzle disposed downstream of the combustion tube;a compressor comprising a plurality of sequential stages disposed upstream of the combustor;a primary flow diverted from the compressor at a stage other than a furthest downstream stage of the plurality of sequential stages, the primary flow routed to the combustor;a turbine in fluid communication with the combustor and disposed downstream of the combustor;a bypass in fluid communication with the further downstream stage of the compressor, an external surface of the nozzle, a plurality of fins disposed on an inner surface of the nozzle, and an exterior surface of the combustion tube of the pulse detonation combustor; andan internal passage in each of the plurality of fins fluidly connecting the bypass to the inner surface of the nozzle through openings in each of the plurality of fins;wherein a remaining flow from the stage of the compressor other than the furthest downstream stage of the compressor flows to the furthest downstream stage of the compressor as a bypass flow, the furthest downstream stage of the compressor increases the pressure of the bypass flow allowing the increased pressure bypass flow to flow in a downstream direction through the bypass to the external surface of the nozzle and then in an upstream direction along the external surface of the combustion tube and reintegrate with the primary flow upstream of the combustion tube, at the external surface of the nozzle a portion of the increased pressure bypass flow passes into the internal passage in each of the plurality of fins and out through the openings to mix with an exhaust flow from the combustor.

이 특허에 인용된 특허 (15)

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Hot gas path component cooling for hybrid pulse detonation combustion systems

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특허 상세정보

Hot gas path component cooling for hybrid pulse detonation combustion systems

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