US20110197591A1 - Axially staged premixed combustion chamber - Google Patents
Axially staged premixed combustion chamber Download PDFInfo
- Publication number
- US20110197591A1 US20110197591A1 US13/025,307 US201113025307A US2011197591A1 US 20110197591 A1 US20110197591 A1 US 20110197591A1 US 201113025307 A US201113025307 A US 201113025307A US 2011197591 A1 US2011197591 A1 US 2011197591A1
- Authority
- US
- United States
- Prior art keywords
- combustion chamber
- nozzles
- air
- center nozzle
- combustor
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Abandoned
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Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/46—Combustion chambers comprising an annular arrangement of several essentially tubular flame tubes within a common annular casing or within individual casings
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/286—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply having fuel-air premixing devices
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/34—Feeding into different combustion zones
- F23R3/343—Pilot flames, i.e. fuel nozzles or injectors using only a very small proportion of the total fuel to insure continuous combustion
Definitions
- This invention relates to gas turbine technology and, more specifically, to an axially-staged gas turbine combustor nozzle configuration that promotes enhanced CO burn-off.
- combustor exit temperature must be kept relatively high in order to ensure CO burn-off to meet required emission levels for CO.
- the customer In order to keep the combustor exit temperature high enough to maintain low CO levels under at low- or no-load conditions, the customer must either shut the turbine down or keep the turbine “on-line”, even during periods of low power requirements, thus increasing the amount of fuel consumed.
- the invention in a first exemplary but nonlimiting embodiment, relates to a combustor for a gas turbine comprising a plurality of radially outer nozzles arranged in a substantially annular array, each of the radially outer nozzles having an outlet end located to supply fuel and/or air to a first combustion chamber; at least a center nozzle having an outlet end located axially upstream of the outlet ends of the radially outer nozzles, configured and arranged to supply fuel and air to a second combustion chamber axially upstream of the first combustion chamber, the second combustion chamber opening into the first combustion chamber and having a length sufficient to maintain a center nozzle flame confined to the second combustion chamber.
- the invention in another exemplary aspect, relates to a combustor for a gas turbine comprising a plurality of nozzles arranged in a substantially annular array, each of the nozzles having an outlet end located to supply fuel and/or air to a first combustion chamber; a center nozzle and at least one of the plurality of nozzles having outlet ends located axially upstream of the outlet ends of remaining ones of the plurality of nozzles, configured and arranged to supply fuel and air to a second combustion chamber axially upstream of the first combustion chamber, the second primary combustion chamber opening into the first combustion chamber and having a length sufficient to maintain a center nozzle flame and a flame of the at least one of the plurality of nozzles confined to the second combustion chamber.
- the invention provides a method of operating a gas turbine having at least one combustor supplied with fuel and/or air through a plurality of nozzles including an outer array of nozzles surrounding a center nozzle, the method comprising (a) at no- or low-load conditions, supplying fuel and air to the center nozzle and air only to the outer array of nozzles while isolating a flame generated by the center nozzle from air flowing through the outer array of nozzles; and (b) at higher load conditions, supplying a fuel/air mixture through both the outer array of nozzles and the center nozzle such that flames generated by the outer array of nozzles are maintained in a first combustion chamber and a flame generated by the center nozzle is maintained in a second combustion chamber upstream of the first combustion chamber.
- FIG. 1 is a cross section through a gas turbine combustor in accordance with a first exemplary but nonlimiting embodiment of the invention
- FIG. 2 is a partial, enlarged perspective view of the combustor shown in FIG. 1 ;
- FIG. 3 is a partially sectioned perspective view of the combustor shown in FIG. 2 ;
- FIG. 4 is a schematic diagram of a combustor configuration in accordance with another exemplary but nonlimiting embodiment.
- FIGS. 1-3 a gas turbine combustor 10 in accordance with an exemplary but non-limiting embodiment of the invention is illustrated.
- the combustor 10 is typically combined with several other similar combustors arranged in an annular array about the gas turbine casing, each combustor supplying combustion gases to the first stage of the turbine.
- Each combustor 10 is supplied with air from a compressor (not shown). The compressor air is reverse-flowed (as indicated by the flow arrows), into an annular passage 12 located between the radially inner and axially-aligned transition piece 14 and combustion chamber liner 16 on the one hand, and the radially outer, axially-aligned flow sleeves 18 and 20 on the other.
- the compressor air flows into the passage 12 through impingement cooling holes 22 , 24 in the respective flow sleeves 18 and 20 , thus also providing cooling to the transition piece and combustor liner, before reversing flow at the inlet end of the combustor.
- the air will flow into air injectors associated with each of a plurality of six radially outer nozzles 26 and a center nozzle 28 (the number of nozzles in the combustor typically varies between 6 and 8 ) where it pre-mixes with fuel supplied to the nozzles via the combustor end cover 30 .
- the air/fuel mixture from the radially outer nozzles 28 is injected into the burning zone, or main combustion chamber 32 .
- Ignition is achieved by spark plugs (not shown) in conjunction with crossfire tubes (also not shown) that connect adjacent combustors.
- Hot combustion gases flow from the combustion chamber 32 into the transition piece 14 and then to the first stage of the gas turbine, represented by a single nozzle blade 34 .
- the combustor as described is generally well known, and the invention here relates to the location of the center nozzle 28 relative to the radially outer nozzles 26 and 30 , and to the establishment of a second (or primary) combustion chamber 36 upstream of the first (or main) combustion chamber 32 .
- the center nozzle 28 is recessed in an upstream direction (relative to a flow direction of combustion gases from left to right in the various FIGS.).
- the center nozzle 28 is located axially behind the outlets of the radially outer surrounding nozzles 26 .
- a combustor cap 38 supports the outlet ends of the outer nozzles, but is configured and mounted so as to allow compressor air to flow between the cap and the casing wall 40 ( FIG. 1 ).
- a substantially cylindrical tubular member 42 extends rearwardly from the cap 38 to the outlet end of the center nozzle 28 , thus forming the primary combustion chamber 36 which opens into the main combustion chamber 32 at the forwardmost plate 44 of the cap 38 .
- the length of the chamber 36 is determined so as to be sufficient to allow complete combustion of CO while protecting the center nozzle flame from the surrounding cold air flowing into the main chamber 32 via the radially outer nozzles 26 .
- Fuel is supplied to the radially outer nozzle tubes (two shown at 46 ( FIG. 1 ) and to the center nozzle tube 48 through the end cover 30 as noted above, while air is supplied to the radially outer 26 nozzles at conventionally-configured premix swirler inlets 50 (two shown in FIG. 3 ), and to the center nozzle 28 through premix swirler inlet via apertures 52 in the radial vane 54 .
- the radially outer nozzles 26 are then brought on board, with fuel supplied to the radially outer nozzles mixing with combustion air supplied by the compressor as described above.
- the combustion flames associated with the outer nozzles 26 are anchored downstream of the primary combustion chamber 36 , within the main combustion chamber 32 .
- the radially outer nozzles 26 may be “lit” or ignited simultaneously, or in some predetermined sequence (or simultaneously in groups of 2 or three, for example) as dictated by combustion optimization for specific combustor applications.
- the center nozzle flame remains anchored in the primary combustion chamber 36 while the outer nozzle flames remain anchored in the main combustion chamber 32 , downstream of the primary combustion chamber 36 .
- the tubular member 42 defining the primary combustion chamber 36 is exposed directly to the center nozzle flame, it must be cooled by any suitable means such as, for example, application of a thermal barrier coating, impingement cooling, the addition of turbulators, or any combination of the above.
- phi is an equivalence ratio defined as the ratio of the actual fuel/air ratio to the stoichiometric value.
- Typical phi values range from 0.50 to 0.65.
- the flame in the center nozzle 28 may be extinguished for a relatively short time, and then resupplied with fuel such that the flame re-ignites (and is maintained) downstream of the primary combustion chamber 36 .
- the temperature of the tubular member 42 will be cooler, and the mixing zone for the fuel and air supplied to the center nozzle 28 is extended, resulting in better mixing and in lower NOx emissions.
- more than one nozzle can be protected from the cold air flowing through the surrounding or adjacent nozzles at FSNL.
- a center nozzle and one or two other nozzles in the outer array could be recessed in the same manner as described above in connection with the center nozzle 28 .
- the one or two additional nozzles could be located in a single oblong, oval or other shape combustion chamber, i.e., the chamber shape would be dictated by the number and location of the recessed nozzles.
- FIG. 4 shows a center nozzle 128 and one of a surrounding array of radially outer nozzles 126 are recessed within a second combustion chamber 136 defined by an oblong tubular member 142 .
- This developed multi-stage combustor is thus capable of isolating fueled nozzles (for example, the center nozzle 28 ) reacting flames from excessively cold surrounding air exiting adjacent unfueled nozzles (for example, the radially outer nozzles 26 at part-or no-load regimes by establishing a combustion zone in a recessed combustion chamber (the primary combustion chamber 36 ) for complete CO burnout at the end of that chamber.
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
RU2010105138 | 2010-02-16 | ||
RU2010105138/06A RU2534189C2 (ru) | 2010-02-16 | 2010-02-16 | Камера сгорания для газовой турбины(варианты) и способ эксплуатации газовой турбины |
Publications (1)
Publication Number | Publication Date |
---|---|
US20110197591A1 true US20110197591A1 (en) | 2011-08-18 |
Family
ID=44317404
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US13/025,307 Abandoned US20110197591A1 (en) | 2010-02-16 | 2011-02-11 | Axially staged premixed combustion chamber |
Country Status (6)
Country | Link |
---|---|
US (1) | US20110197591A1 (ru) |
JP (1) | JP5775319B2 (ru) |
CN (1) | CN102192508B (ru) |
CH (1) | CH702737B1 (ru) |
DE (1) | DE102011000589A1 (ru) |
RU (1) | RU2534189C2 (ru) |
Cited By (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20100180603A1 (en) * | 2009-01-16 | 2010-07-22 | General Electric Company | Fuel nozzle for a turbomachine |
US20140338339A1 (en) * | 2013-03-12 | 2014-11-20 | General Electric Company | System and method having multi-tube fuel nozzle with multiple fuel injectors |
US9528444B2 (en) | 2013-03-12 | 2016-12-27 | General Electric Company | System having multi-tube fuel nozzle with floating arrangement of mixing tubes |
US9534787B2 (en) | 2013-03-12 | 2017-01-03 | General Electric Company | Micromixing cap assembly |
US9651259B2 (en) | 2013-03-12 | 2017-05-16 | General Electric Company | Multi-injector micromixing system |
US9650959B2 (en) | 2013-03-12 | 2017-05-16 | General Electric Company | Fuel-air mixing system with mixing chambers of various lengths for gas turbine system |
US9671112B2 (en) | 2013-03-12 | 2017-06-06 | General Electric Company | Air diffuser for a head end of a combustor |
US9765973B2 (en) | 2013-03-12 | 2017-09-19 | General Electric Company | System and method for tube level air flow conditioning |
WO2019012559A1 (en) * | 2017-07-12 | 2019-01-17 | Bharat Forge Limited | ADDITIVE FABRICATION PROCESS FOR COMBUSTION CHAMBER |
Families Citing this family (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20130081397A1 (en) * | 2011-10-04 | 2013-04-04 | Brandon Taylor Overby | Forward casing with a circumferential sloped surface and a combustor assembly including same |
US9404655B2 (en) * | 2012-01-20 | 2016-08-02 | General Electric Company | Process of fabricating a fuel nozzle assembly |
AU2013219140B2 (en) | 2012-08-24 | 2015-10-08 | Ansaldo Energia Switzerland AG | Method for mixing a dilution air in a sequential combustion system of a gas turbine |
EP2796789B1 (en) | 2013-04-26 | 2017-03-01 | General Electric Technology GmbH | Can combustor for a can-annular combustor arrangement in a gas turbine |
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US4505666A (en) * | 1981-09-28 | 1985-03-19 | John Zink Company | Staged fuel and air for low NOx burner |
US4677822A (en) * | 1985-02-22 | 1987-07-07 | Hitachi, Ltd. | Gas turbine combustor |
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US4805411A (en) * | 1986-12-09 | 1989-02-21 | Bbc Brown Boveri Ag | Combustion chamber for gas turbine |
US5054280A (en) * | 1988-08-08 | 1991-10-08 | Hitachi, Ltd. | Gas turbine combustor and method of running the same |
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-
2010
- 2010-02-16 RU RU2010105138/06A patent/RU2534189C2/ru not_active IP Right Cessation
-
2011
- 2011-02-09 DE DE102011000589A patent/DE102011000589A1/de not_active Withdrawn
- 2011-02-11 US US13/025,307 patent/US20110197591A1/en not_active Abandoned
- 2011-02-14 CH CH00257/11A patent/CH702737B1/de not_active IP Right Cessation
- 2011-02-15 JP JP2011029253A patent/JP5775319B2/ja not_active Expired - Fee Related
- 2011-02-16 CN CN201110043127.8A patent/CN102192508B/zh not_active Expired - Fee Related
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US7631499B2 (en) * | 2006-08-03 | 2009-12-15 | Siemens Energy, Inc. | Axially staged combustion system for a gas turbine engine |
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US7578130B1 (en) * | 2008-05-20 | 2009-08-25 | General Electric Company | Methods and systems for combustion dynamics reduction |
Cited By (11)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20100180603A1 (en) * | 2009-01-16 | 2010-07-22 | General Electric Company | Fuel nozzle for a turbomachine |
US8161750B2 (en) * | 2009-01-16 | 2012-04-24 | General Electric Company | Fuel nozzle for a turbomachine |
US20140338339A1 (en) * | 2013-03-12 | 2014-11-20 | General Electric Company | System and method having multi-tube fuel nozzle with multiple fuel injectors |
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US9534787B2 (en) | 2013-03-12 | 2017-01-03 | General Electric Company | Micromixing cap assembly |
US9651259B2 (en) | 2013-03-12 | 2017-05-16 | General Electric Company | Multi-injector micromixing system |
US9650959B2 (en) | 2013-03-12 | 2017-05-16 | General Electric Company | Fuel-air mixing system with mixing chambers of various lengths for gas turbine system |
US9671112B2 (en) | 2013-03-12 | 2017-06-06 | General Electric Company | Air diffuser for a head end of a combustor |
US9759425B2 (en) * | 2013-03-12 | 2017-09-12 | General Electric Company | System and method having multi-tube fuel nozzle with multiple fuel injectors |
US9765973B2 (en) | 2013-03-12 | 2017-09-19 | General Electric Company | System and method for tube level air flow conditioning |
WO2019012559A1 (en) * | 2017-07-12 | 2019-01-17 | Bharat Forge Limited | ADDITIVE FABRICATION PROCESS FOR COMBUSTION CHAMBER |
Also Published As
Publication number | Publication date |
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CN102192508A (zh) | 2011-09-21 |
DE102011000589A1 (de) | 2011-08-18 |
JP5775319B2 (ja) | 2015-09-09 |
JP2011169575A (ja) | 2011-09-01 |
CH702737A2 (de) | 2011-08-31 |
CH702737B1 (de) | 2016-02-15 |
CN102192508B (zh) | 2015-11-25 |
RU2534189C2 (ru) | 2014-11-27 |
RU2010105138A (ru) | 2011-08-27 |
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