EP2728258A1 - Gas Turbine - Google Patents
Gas Turbine Download PDFInfo
- Publication number
- EP2728258A1 EP2728258A1 EP12191049.1A EP12191049A EP2728258A1 EP 2728258 A1 EP2728258 A1 EP 2728258A1 EP 12191049 A EP12191049 A EP 12191049A EP 2728258 A1 EP2728258 A1 EP 2728258A1
- Authority
- EP
- European Patent Office
- Prior art keywords
- combustion chamber
- gas turbine
- mixing section
- combustor
- downstream
- 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.)
- Withdrawn
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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
- F23R3/002—Wall structures
-
- 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
- F23R3/346—Feeding into different combustion zones for staged combustion
-
- 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
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/03341—Sequential combustion chambers or burners
Definitions
- the present invention relates to the technology of gas turbines. It refers to a gas turbine according to the preamble of claim 1.
- Fig. 1 shows in a perspective view a stationary gas turbine of the applicant of the well-known type GT26.
- the gas turbine 10 of Fig. 1 comprises a rotor 11 surrounded by a casing, a compressor 13, which compresses air entering the gas turbine 10 through an air intake 12, a ring of first burners 14 with a first fuel supply 14a, a first (high pressure) turbine 16, a ring of second burners 15 with a second fuel supply 15a, a second (low pressure) turbine 16', and an exhaust gas outlet 17 for releasing the exhaust gas either into a stack or a heat recovery steam generator of an associated water/steam cycle in case of a combined cycle power plant CCPP.
- a gas turbine of the type shown in Fig. 1 is called to have a sequential combustion, due to the sequential arrangement of a first and a second combustion chamber with first and second burners.
- FIG.2 A typical design of the second burner/combustion chamber (second combustor) arrangement, which is known in the art (see for example document EP 2 169 314 A2 or EP 2 423 599 A2 ) is illustrated in Fig.2 .
- the second (SEV) combustor of Fig. 2 comprises a second burner/injector 15 axially connected to a combustion chamber 18.
- the hot gas flow entering the burner section (from the left side in Fig. 2 ; see arrow) is fed with fuel by means of fuel supply 15a via the injector (e.g. a lance) through injection holes 15b, flows along a mixing section 32, and leaves the burner section at its exit to expand into the combustion chamber 18.
- the interface between the burner section 15, 32 and combustion chamber 18 is characterized by a sudden cross-sectional area change (i.e. backward facing step) comprising a perpendicular front panel 19 extending from the exit of the burner section to inner liner 21 and outer liner 20.
- the mechanical interface is at the front panel seal, separating the burner front panel 19 from the liners 20 and 21.
- the transition between burner section 15, 32 and combustion chamber 18 is located at the periphery of the burner front panel 19.
- the front panel 19 is flat and a large portion of the combustion chamber 18 is therefore occupied by recirculation bubbles 22.
- the leakages were minimised with seals between front panel 19 and liners 20, 21.
- they are completely bypassing the flame and the decreased turbine inlet temperature due to bypassing air needs to be compensated with higher flame temperature.
- the leakages may also locally quench the flame and increase the part load CO emissions.
- the gas turbine according to the invention comprises a compressor, a first combustor downstream of said compressor, a first turbine, a second combustor downstream of said first turbine, and a second turbine downstream of said second combustor, whereby said second combustor comprises a plurality of burners having a mixing section, which opens downstream into a wider combustion chamber. It is characterized in that the transition between the mixing section and the combustion chamber is continuous.
- the transition between the mixing section and the combustion chamber is streamlined.
- the mixing section and the combustion chamber are assembled from at least two parts, which are joined at a split plane.
- said split plane is located at the exit of said mixing section.
- said split plane is located upstream of the exit of said mixing section.
- said burners inject fuel in an injection plane perpendicular to the flow direction within said mixing section, and that said split plane coincides with said injection plane.
- said combustion chamber has an annular layout.
- said combustion chamber has a can layout.
- a basic idea of the present invention is to minimize the second (SEV) burner pressure drop and associated emissions by means of a smoother transition between burner section and combustion chamber.
- An additional measure is to move the split plane and hence the position of the associated leakages upstream.
- the transition between burner mixing section 32 and combustion chamber 24 of a gas turbine 30 with sequential combustion is done with a smoothed, especially streamlined, geometry of the outer liner 25 and inner liner 26.
- the split between burner 15, mixing section 32 and combustion chamber 24 can be located at the mixing section 32 exit (split plane 29), so that leakages 27 are entering into the flame and participate in the fuel/air reaction.
- Both measures can be used separately or in combination. Both measures can be applied to combustion chambers with annular or can layout.
- the split could be even positioned at the injection plane 31, with the leakages 28 being mixed with the injected fuel, which is particularly suitable for a can layout.
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
Abstract
A gas turbine (30) comprises a compressor, a first combustor downstream of said compressor, a first turbine, a second combustor (15, 15a) downstream of said first turbine, and a second turbine (16) downstream of said second combustor (15, 15a), whereby said second combustor (15, 15a) comprises a plurality of burners (15) having a mixing section (32), which opens downstream into a wider combustion chamber (24).
The emissions of the gas turbine are improved by providing a continuous transition between the mixing section (32) and the combustion chamber (24).
Description
- The present invention relates to the technology of gas turbines. It refers to a gas turbine according to the preamble of claim 1.
-
Fig. 1 shows in a perspective view a stationary gas turbine of the applicant of the well-known type GT26. - The
gas turbine 10 ofFig. 1 comprises arotor 11 surrounded by a casing, acompressor 13, which compresses air entering thegas turbine 10 through anair intake 12, a ring offirst burners 14 with afirst fuel supply 14a, a first (high pressure)turbine 16, a ring ofsecond burners 15 with asecond fuel supply 15a, a second (low pressure) turbine 16', and anexhaust gas outlet 17 for releasing the exhaust gas either into a stack or a heat recovery steam generator of an associated water/steam cycle in case of a combined cycle power plant CCPP. - A gas turbine of the type shown in
Fig. 1 is called to have a sequential combustion, due to the sequential arrangement of a first and a second combustion chamber with first and second burners. - A typical design of the second burner/combustion chamber (second combustor) arrangement, which is known in the art (see for example document
EP 2 169 314 A2 orEP 2 423 599 A2 ) is illustrated inFig.2 . - The second (SEV) combustor of
Fig. 2 comprises a second burner/injector 15 axially connected to acombustion chamber 18. In particular, the hot gas flow entering the burner section (from the left side inFig. 2 ; see arrow) is fed with fuel by means offuel supply 15a via the injector (e.g. a lance) throughinjection holes 15b, flows along amixing section 32, and leaves the burner section at its exit to expand into thecombustion chamber 18. - The interface between the
burner section combustion chamber 18 is characterized by a sudden cross-sectional area change (i.e. backward facing step) comprising aperpendicular front panel 19 extending from the exit of the burner section toinner liner 21 andouter liner 20. The mechanical interface is at the front panel seal, separating theburner front panel 19 from theliners - The transition between
burner section combustion chamber 18 is located at the periphery of theburner front panel 19. Thefront panel 19 is flat and a large portion of thecombustion chamber 18 is therefore occupied byrecirculation bubbles 22. - The two main drawbacks of this current concept are:
- (1) The sudden expansion (backward facing step) causes a large pressure drop due to the large recirculation zones associated with it.
- (2) The
leakages 23 at the burner section/liner interface and between adjacent burners are entering into thecombustion chamber 18 in therecirculation bubbles 22, without participating in combustion, i.e. the leakage air is bypassing the flame, which needs to be at higher temperature to fulfil the target combustion chamber exit temperature. Higher flame temperature increases NOx emissions. At part load, CO emissions may increase because of local quenching of the flame. - Thus, although the transition between the second (SEV) burner and combustion chamber with its large backward facing step is a key element for stabilising the flame, it has nevertheless a negative impact on the SEV pressure drop and emissions associated to the
leakages 23 between theburner front panel 19 andliners - In the past, it has been proposed to design the burner with a diffuser in its mixing section to minimise the front panel area and therefore the pressure drop.
- According to another proposal the leakages were minimised with seals between
front panel 19 andliners - It is an object of the present invention to provide a gas turbine with sequential combustion, which avoids the drawbacks of the prior art second burner/combustion chamber arrangement and has reduced emissions.
- This object is obtained by a gas turbine according to claim 1.
- The gas turbine according to the invention comprises a compressor, a first combustor downstream of said compressor, a first turbine, a second combustor downstream of said first turbine, and a second turbine downstream of said second combustor, whereby said second combustor comprises a plurality of burners having a mixing section, which opens downstream into a wider combustion chamber. It is characterized in that the transition between the mixing section and the combustion chamber is continuous.
- According to an embodiment of the invention the transition between the mixing section and the combustion chamber is streamlined.
- According to another embodiment of the invention the mixing section and the combustion chamber are assembled from at least two parts, which are joined at a split plane.
- Specifically, said split plane is located at the exit of said mixing section.
- Specifically, said split plane is located upstream of the exit of said mixing section.
- More specifically, said burners inject fuel in an injection plane perpendicular to the flow direction within said mixing section, and that said split plane coincides with said injection plane.
- According to a further embodiment of the invention said combustion chamber has an annular layout.
- According to just another embodiment of the invention said combustion chamber has a can layout.
- The present invention is now to be explained more closely by means of different embodiments and with reference to the attached drawings.
- Fig. 1
- shows in a perspective view a stationary (or industrial) gas turbine with sequential combustion;
- Fig. 2
- shows a sectional view of a secondary burner/combustion chamber arrangement according to the state of the art;
- Fig. 3
- shows a sectional view of a secondary burner/combustion chamber arrangement according to one embodiment of the invention; and
- Fig. 4
- shows a sectional view of a secondary burner/combustion chamber arrangement according to another embodiment of the invention.
- A basic idea of the present invention is to minimize the second (SEV) burner pressure drop and associated emissions by means of a smoother transition between burner section and combustion chamber. An additional measure is to move the split plane and hence the position of the associated leakages upstream.
- According to
Fig. 3 , the transition betweenburner mixing section 32 andcombustion chamber 24 of agas turbine 30 with sequential combustion is done with a smoothed, especially streamlined, geometry of theouter liner 25 andinner liner 26. - In addition, the split between
burner 15, mixingsection 32 andcombustion chamber 24 can be located at themixing section 32 exit (split plane 29), so thatleakages 27 are entering into the flame and participate in the fuel/air reaction. - However, there are two measures that may be separated:
- (1) The upstream shift of leakages (by shifting the
split plane 29 upstream); and - (2) The smoothing of the liner shapes of the inner and/or outer liner (26, 25).
- Both measures can be used separately or in combination. Both measures can be applied to combustion chambers with annular or can layout.
- According to another embodiment of the invention, as shown for a gas turbine 30' in
Fig. 4 , the split could be even positioned at theinjection plane 31, with theleakages 28 being mixed with the injected fuel, which is particularly suitable for a can layout. -
- 10,30,30'
- gas turbine
- 11
- rotor
- 12
- air intake
- 13
- compressor
- 14,15
- burner
- 14a,15a
- fuel supply
- 16,16'
- turbine
- 17
- exhaust gas outlet
- 18,24
- combustion chamber
- 19
- front panel
- 20,25
- outer liner
- 21,26
- inner liner
- 22
- recirculation bubble
- 23,27,28
- leakage
- 29
- split plane
- 31
- injection plane (split plane)
- 32
- mixing section
Claims (8)
- Gas turbine (10, 30, 30'), comprising a compressor (13), a first combustor (14, 14a) downstream of said compressor (13), a first turbine (16), a second combustor (15, 15a) downstream of said first turbine (16), and a second turbine (16') downstream of said second combustor (15, 15a), whereby said second combustor (15, 15a) comprises a plurality of burners (15) having a mixing section (32), which opens downstream into a wider combustion chamber (18, 24), characterized in that the transition between the mixing section (32) and the combustion chamber (24) is continuous.
- Gas turbine according to claim 1, characterized in that the transition between the mixing section (32) and the combustion chamber (24) is streamlined.
- Gas turbine according to claim 1, characterized in that the mixing section (32) and the combustion chamber (24) are assembled from at least two parts, which are joined at a split plane (29, 31).
- Gas turbine according to claim 3, characterized in that said split plane (29) is located at the exit of said mixing section (32).
- Gas turbine according to claim 3, characterized in that said split plane (31) is located upstream of the exit of said mixing section (32).
- Gas turbine according to claim 5, characterized in that said burners (15) inject fuel in an injection plane (31) perpendicular to the flow direction within said mixing section (32), and that said split plane coincides with said injection plane (31).
- Gas turbine according to one of the claims 1-6, characterized in that said combustion chamber (24) has an annular layout.
- Gas turbine according to one of the claims 1-6, characterized in that said combustion chamber (24) has a can layout.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP12191049.1A EP2728258A1 (en) | 2012-11-02 | 2012-11-02 | Gas Turbine |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP12191049.1A EP2728258A1 (en) | 2012-11-02 | 2012-11-02 | Gas Turbine |
Publications (1)
Publication Number | Publication Date |
---|---|
EP2728258A1 true EP2728258A1 (en) | 2014-05-07 |
Family
ID=47216089
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP12191049.1A Withdrawn EP2728258A1 (en) | 2012-11-02 | 2012-11-02 | Gas Turbine |
Country Status (1)
Country | Link |
---|---|
EP (1) | EP2728258A1 (en) |
Cited By (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US11156164B2 (en) | 2019-05-21 | 2021-10-26 | General Electric Company | System and method for high frequency accoustic dampers with caps |
US11174792B2 (en) | 2019-05-21 | 2021-11-16 | General Electric Company | System and method for high frequency acoustic dampers with baffles |
Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0623786A1 (en) * | 1993-04-08 | 1994-11-09 | ABB Management AG | Combustion chamber |
US20020187448A1 (en) * | 2001-06-09 | 2002-12-12 | Adnan Eroglu | Burner system |
EP2169314A2 (en) | 2008-09-30 | 2010-03-31 | Alstom Technology Ltd | A method of reducing emissions for a sequential combustion gas turbine and combustor for such a gas turbine |
US20110030375A1 (en) * | 2009-08-04 | 2011-02-10 | General Electric Company | Aerodynamic pylon fuel injector system for combustors |
EP2423599A2 (en) | 2010-08-27 | 2012-02-29 | Alstom Technology Ltd | Method for operating a burner arrangement and burner arrangement for implementing the method |
-
2012
- 2012-11-02 EP EP12191049.1A patent/EP2728258A1/en not_active Withdrawn
Patent Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0623786A1 (en) * | 1993-04-08 | 1994-11-09 | ABB Management AG | Combustion chamber |
US20020187448A1 (en) * | 2001-06-09 | 2002-12-12 | Adnan Eroglu | Burner system |
EP2169314A2 (en) | 2008-09-30 | 2010-03-31 | Alstom Technology Ltd | A method of reducing emissions for a sequential combustion gas turbine and combustor for such a gas turbine |
US20110030375A1 (en) * | 2009-08-04 | 2011-02-10 | General Electric Company | Aerodynamic pylon fuel injector system for combustors |
EP2423599A2 (en) | 2010-08-27 | 2012-02-29 | Alstom Technology Ltd | Method for operating a burner arrangement and burner arrangement for implementing the method |
Cited By (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US11156164B2 (en) | 2019-05-21 | 2021-10-26 | General Electric Company | System and method for high frequency accoustic dampers with caps |
US11174792B2 (en) | 2019-05-21 | 2021-11-16 | General Electric Company | System and method for high frequency acoustic dampers with baffles |
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