IPC분류정보
국가/구분 |
United States(US) Patent
등록
|
국제특허분류(IPC7판) |
|
출원번호 |
US-0926321
(2010-11-09)
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등록번호 |
US-8844260
(2014-09-30)
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발명자
/ 주소 |
- Beran, Martin
- Koranek, Michal
- Axelsson, Axel Lars-Uno Eugen
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출원인 / 주소 |
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대리인 / 주소 |
Finnegan, Henderson, Farabow, Garrett & Dunner, LLP
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인용정보 |
피인용 횟수 :
0 인용 특허 :
4 |
초록
▼
A low calorific value fuel-fired can combustor for a gas turbine include a generally cylindrical housing, and a generally cylindrical liner disposed coaxially within the housing to define with the housing a radial outer flow passage for combustion air, the liner also defining inner combustion and a
A low calorific value fuel-fired can combustor for a gas turbine include a generally cylindrical housing, and a generally cylindrical liner disposed coaxially within the housing to define with the housing a radial outer flow passage for combustion air, the liner also defining inner combustion and a dilution zone, the dilution zone being axially distant a closed housing end relative to the combustion zone. A nozzle assembly disposed at the closed housing end includes an air blast nozzle and surrounding swirl vanes. An impingement cooling sleeve coaxially disposed in the combustion air passage between the housing and the liner impingement cools the portion of the liner defining the combustion zone. The combustion liner has an L/D ratio of in the range 1≦L/D≦4, and a ratio of the combustion zone volume (m3) to heat energy flow rate Q (MJ/sec) in the range 0.0026≦V/Q≦0.018.
대표청구항
▼
1. A can combustor for burning fuels with low calorific values, the combustor comprising: a generally cylindrical housing having an interior, a longitudinal axis, an annular inlet for receiving compressed air at one longitudinal housing end with the other longitudinal housing end being closed;a gene
1. A can combustor for burning fuels with low calorific values, the combustor comprising: a generally cylindrical housing having an interior, a longitudinal axis, an annular inlet for receiving compressed air at one longitudinal housing end with the other longitudinal housing end being closed;a generally cylindrical combustor liner coaxially disposed in the housing interior, the liner and the housing defining a generally annular flow passage for the compressed air received through the housing inlet, an interior of the liner defining a combustion zone adjacent the closed housing end and a dilution zone distant the closed housing end;a fuel nozzle assembly disposed at the closed end, the nozzle assembly being supplied from a source of fuel having a calorific value of less than about 25 MJ/kg;an impingement cooling sleeve disposed in the compressed air passage surrounding the liner portion defining the combustion zone, the sleeve having a plurality of orifices sized and configured to impingement cool an outer surface of the liner portion with essentially all of the compressed air received at the housing inlet passing through the sleeve;a plurality of primary holes circumferentially disposed in the liner for introducing a first portion of the compressed air from a region downstream of the impingement cooling sleeve into the combustion zone;a plurality of dilution openings circumferentially disposed in the liner for introducing a second portion of the compressed air from the region downstream of the impingement cooling sleeve into the dilution zone,wherein at least part of a remainder portion of the compressed air from the region downstream of the impingement cooling sleeve is channeled through the fuel nozzle assembly for mixing with the supplied fuel to provide a fuel/air mixture directed into the combustion zone, andwherein the liner is sized to have an L/D ratio in the range 1.00≦L/D<4.00, where L is a liner length and D is a liner diameter. 2. The can combustor as in claim 1, wherein 1.5≦L/D<2.5. 3. The can combustor as in claim 1 wherein the first portion of compressed air is 5-15% of a total compressed air mass flow rate. 4. The can combustor as in claim 1, wherein the second portion of compressed air is 60-70% of a total compressed air mass flow rate. 5. The can combustor as in claim 1, wherein the fuel nozzle assembly includes an air blast nozzle, and wherein the nozzle assembly is configured to use a part of the remainder air portion of the compressed air to direct the fuel/air mixture into the combustion zone using a compressed air pressure differential between the region downstream of the impingement cooling sleeve and the combustion zone. 6. The can combustor as in claim 5, wherein the fuel nozzle assembly is disposed coaxially with the liner and includes swirl vanes distributed circumferentially about an exit of the nozzle assembly to induce swirling in the directed fuel/air mixture using another part of the remainder air portion. 7. The can combustor as in claim 1, wherein the fuel nozzle assembly and the liner are sized and configured to inject and burn liquid pyrolysis oil. 8. The can combustor as in claim 7, wherein the fuel nozzle assembly includes an air blast nozzle; wherein L/D is about 1.65. 9. The can combustor as in claim 7, wherein the source of fuel includes a light weight alcohol mixed with the pyrolysis oil for combustor operation at less than about 60% rated power. 10. The can combustor as in claim 1, wherein the primary holes have spout-shaped wall extensions into the combustion zone. 11. The can combustor as in claim 1, wherein a surface of the liner is coated with TBC to increase the inside surface temperature. 12. The can combustor as in claim 1, wherein the liner includes a tapered inlet portion adjacent an exit of the fuel nozzle assembly; wherein the liner further includes an entrance shield member coaxially disposed within, and spaced from, the tapered inlet liner portion; and wherein a plurality of impingement cooling orifices are provided in the tapered liner portion sized and directed to impingement cool the entrance shield member using compressed air from the sleeve downstream region. 13. The can combustor as in claim 1, wherein the liner includes a tapered transition portion disposed between the combustion zone and the dilution zone; wherein the liner further includes a transition shield member coaxially disposed within, and spaced from, the tapered transition liner portion; and wherein a plurality of impingement cooling orifices are provided in the tapered transition liner portion, the orifices being sized and directed to impingement cool the transition shield member using compressed air from the sleeve downstream region. 14. The can combustor as in claim 1, wherein the impingement cooling sleeve extends from a location on the liner downstream of the dilution ports to a location on the housing upstream of the combustion zone, relative to a flow direction of combustion gases. 15. A gas turbine engine having the can combustor of claim 1 operatively interconnected between an air compressor and a gas turbine. 16. A can combustor for burning a liquid fuel with a low calorific value, the combustor comprising: a generally cylindrical housing having an interior, longitudinal axis, an angular inlet for receiving compressed air at one longitudinal housing end with the other longitudinal housing end being closed;a generally cylindrical combustor liner and the housing defining a generally annular flow passage for the compressed air received through the housing inlet, an interior of the liner defining a combustion zone adjacent the closed housing end and a dilution zone distant the closed housing end;a fuel nozzle assembly disposed at the closed end, the nozzle assembly being supplied from a source of liquid fuel having a calorific value of less than about 25 MJ/kg, the nozzle assembly being configured to provide a fuel spray;an impingement cooling sleeve disposed in the compressed air passage surrounding the liner portion defining the combustion zone, the sleeve having a plurality of orifices sized and configured to impingement cool an outer surface of the liner portion with essentially all of the compressed air received at the housing inlet passing through the sleeve;a plurality of primary holes circumferentially disposed in the liner for introducing a first portion of the compressed air from a region downstream of the impingement cooling sleeve into the combustion zone;a plurality of dilution openings circumferentially disposed in the liner for introducing a second portion of the compressed air from the region downstream of the impingement cooling sleeve into the dilution zone,wherein at least part of a remainder portion of the compressed air from the region downstream of the impingement cooling screen is channeled through the fuel nozzle assembly for mixing with the fuel spray to provide a fuel/air mixture directed into the combustion zone,wherein the fuel nozzle assembly includes an air blast nozzle, and wherein the nozzle assembly is configured to use a part of the remainder air portion of the compressed air to direct the fuel/air mixture into the combustion zone using a compressed air pressure differential between the region downstream of the impingement cooling sleeve and the combustion zone,wherein the fuel nozzle assembly is disposed coaxially with the liner and includes swirl vanes distributed circumferentially about an exit of the nozzle assembly to induce swirling in the directed fuel/air mixture using another part of the remainder air portion,wherein the liner is sized to have an L/D ratio in the range 1.5≦L/D≦2.5 where L is a liner length and D is a liner diameter. 17. The can combustor as in claim 16 wherein the first portion of compressed air is 5-15% of a total compressed air mass flow rate. 18. The can combustor as in claim 16, wherein the second portion of compressed air is 60-70% of a total compressed air mass flow rate. 19. The can combustor as in claim 16, wherein the liquid fuel is pyrolysis oil having a calorific value of about 7 MJ/kg; wherein the L/D ratio is about 1.65. 20. A gas turbine engine having the can combustor as in claim 16 operatively interconnected between an air compressor and a gas turbine.
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