EP1989486A1 - Gas turbine burner and method of operating a gas turbine burner - Google Patents
Gas turbine burner and method of operating a gas turbine burnerInfo
- Publication number
- EP1989486A1 EP1989486A1 EP07712256A EP07712256A EP1989486A1 EP 1989486 A1 EP1989486 A1 EP 1989486A1 EP 07712256 A EP07712256 A EP 07712256A EP 07712256 A EP07712256 A EP 07712256A EP 1989486 A1 EP1989486 A1 EP 1989486A1
- Authority
- EP
- European Patent Office
- Prior art keywords
- gas
- fuel
- turbine burner
- combustion
- combustion exhaust
- 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.)
- Ceased
Links
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
- F23R3/28—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
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/08—Plants including a gas turbine driving a compressor or a ducted fan with supplementary heating of the working fluid; Control thereof
- F02K3/10—Plants including a gas turbine driving a compressor or a ducted fan with supplementary heating of the working fluid; Control thereof by after-burners
-
- 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
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/16—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
- F23R3/18—Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants
- F23R3/24—Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants of the fluid-screen type
Definitions
- the invention is directed to a gas turbine combustor having a combustion zone for combusting a mixture of combustion gas added with fuel gas and a fuel mixing arrangement having a fuel nozzle for injecting the fuel gas into the combustion exhaust gas.
- the invention is based on a method for operating a gas turbine burner with a combustion zone in which a mixture of combustion gas mixed with fuel gas is burned, wherein the fuel gas is injected with a fuel nozzle into the combustion exhaust gas.
- the object of the invention is, in particular, to provide a gas turbine burner and a method for operating a gas turbine burner, in which a low-emission combustion can be ensured.
- the fuel ignites after 0.3 ms or less, so that the fuel can mix little with the combustion exhaust gas. This creates an unfavorable diffusion flame that leads to unacceptable NOx emissions. It is called a diffusion flame when a flame without air premix burns.
- the oxygen required for combustion as well as all other air components diffuse over the edge of the flame the flame into it, which is why the flame to the flame kernel is supplied with less and less oxygen and the fuel burns therefore slower.
- the reference system may be the quiescent combustion chamber, in particular when the combustion exhaust gas to be injected is flowing slowly, so that its speed is negligible. If the hot gas to be injected is also in rapid motion, the reference system can be the reference system moving with the combustion exhaust gas surrounding the jet. Then, the speed at which the fuel gas is injected into the combustion exhaust gas is advantageously based on the reference system moving with the combustion exhaust gas.
- the speed of sound is expediently to be seen here as the speed of sound of the unburnt fuel-containing fuel mixture emerging from the nozzle-also referred to below simply as fuel gas-which depends on the temperature and the pressure of the fuel Fuel gas.
- the fuel gas can thus be injected into the combustion exhaust gas with a jet at a speed which is at least as great as 0.2 times the speed of sound in the fuel gas.
- the injection rate can be measured, for example, in the center of the beam or averaged over the entire or part of the beam cross section.
- the gas turbine burner is expediently an afterburner or Wiederaufsammlungverbrennungssystems or part of such.
- the fuel gas suitably contains a proportion of fuel sufficient to enrich the combustion exhaust gas with a predetermined temperature so that it ignites itself.
- fuel all usable in gas turbines fuels can be used, for example
- the achievable by the high injection rate principle of delaying the combustion by a high shear gradient is characterized by a high degree of independence from the fuel used.
- the gas turbine combustor comprises a primary combustion chamber, wherein the combustion zone is arranged in an exhaust gas stream downstream of the primary combustion chamber and the
- the fuel gas can be injected into the combustion exhaust gas without recirculation of the combustion exhaust gas being necessary, whereby a stable jet of injection with a high shear gradient can be achieved.
- the fuel mixing arrangement is designed to inject the fuel gas (4) into the combustion exhaust gas (6) at at least 0.4 times the speed of sound. In general, the faster and harder the beam, the larger the area in which the value of the shear gradient is above the critical value.
- the fuel mixing arrangement is designed to inject the fuel gas into the combustion exhaust gas at a rate less than 0.9 times the speed of sound in the combustion exhaust gas, a satisfactory balance can be achieved between higher speed demands on the one hand and low cost fuel injection arrangements on the other become.
- the fuel mixing arrangement comprises a premixing unit for premixing the fuel gas with oxygen-containing gas
- lean lean combustion of low pollutant concentration in the combustion products can be achieved.
- the mixed product of the premix is the fuel gas that is injected into the exhaust gas.
- the premixing unit is designed to premix the fuel gas with the oxygen-containing gas so that a ratio between the number of fuel molecules to the number of oxygen molecules is between 0.2 and 10.
- the lean combustion can be achieved even at jet speeds in the lower part of the speed range according to the invention, if the premixing unit is designed to Pre-mix the fuel gas with the oxygen-containing gas such that a ratio between the number of fuel molecules to the number of oxygen molecules is less than 1.0.
- inert material can be admixed with the fuel, it also being expedient to take account of the abovementioned ratios, but now with inert material instead of the oxygen-containing gas.
- Suitable inert material is particularly water vapor, CO 2 or
- the number particle amount of the inert substance may be up to ten times the amount of the fuel.
- the fuel can also be injected as fuel gas without admixing oxygen-containing gas or inert material.
- the delay in autoignition can be ensured if a shear gradient in an edge region of the jet in an area in front of the nozzle exit, ie downstream of the nozzle exit, is above a critical shear gradient for autoignition.
- a length of the region in front of the nozzle exit, in which the shear gradient lies above the critical shear gradient for the self-ignition is at least 10 cm long.
- the length of the region depends on the velocities of the jet and the combustion exhaust gas and is particularly advantageously chosen so that the auto-ignition delays by at least 1 ms.
- the jet can be realized in a particularly simple manner.
- the ratio of the pressure difference between the jet pressure and the pressure of the combustion exhaust gas to the pressure of the combustion exhaust gas equal to the ratio of the velocity of the jet and the speed of sound in the combustion exhaust gas.
- injection jet of fuel gas at least one inner jet of fuel-containing gas and a
- Inner jet surrounding outer jet of cooling gas wherein the cooling gas has a lower temperature than the combustion exhaust gas, a particularly effective premix can be achieved because the auto-ignition is further delayed by the cooling gas by the achievement of the autoignition temperature is delayed.
- the critical value of the shear gradient is temperature dependent so that it is reduced by the addition of cooling gas. This can eventually lead to an enlargement of the premixing zone, in which the shear gradient lies above the critical value dependent on the local temperature.
- Effective cooling can be achieved when the temperature of the cooling gas is between 200 0 C and 400 0 C.
- the velocity of the outer jet of cooling gas is equal to the velocity of the inner jet, the hardness of the jet edge does not decrease due to the additional outer jet, so that a large shear gradient can be achieved.
- the advantage of the combustion delay can be further increased if the velocity of the outer jet of cooling gas is greater than the velocity of the inner jet. An even higher shear gradient can be achieved between the outer jet and the environment than with only the inner jet and the environment, which further delays the combustion.
- Cooling gas is smaller than the velocity of the inner jet, the outer beam can be generated in a cost-effective manner without complex compressors and nozzles. If the cooling gas contains fuel, a homogeneous fuel concentration in the flame zone can be achieved.
- a cost-effective implementation of the gas turbine burner can be achieved by the cooling gas consists at least substantially of air.
- the advantages of the invention are particularly useful because of the particularly rapid auto-ignition in this temperature range, when the temperature of the combustion exhaust gas between 900 0 C and 1600 0 C.
- the object directed to the method is achieved by a method for operating a gas turbine of the type mentioned, in which, according to the invention, the fuel gas is injected into the combustion exhaust gas at at least 0.2 times the speed of sound. It can be achieved for the reasons mentioned above a low-emission combustion.
- FIG. 1 shows a gas turbine burner with a secondary combustion zone according to a first exemplary embodiment of the invention
- FIG. 3 shows a lance designed as a fuel nozzle of a Wiederauf rempliverbrennungssystems according to another alternative embodiment of the invention.
- FIG. 1 shows a reheat combustion system 2 for a gas turbine plant with a gas turbine burner 4 with a secondary combustion zone 6 in which a mixture of combustion exhaust gas 10 mixed with fuel gas 8 is combusted.
- the combustion exhaust 10 is from a primary combustion chamber 12 of the gas turbine plant located upstream of the combustion zone 6, separated from the combustion zone 6 by a turbine stage 14 of the gas turbine whose blades 16 are driven by the combustion exhaust gases 10 from the combustion chamber 12.
- the secondary combustion zone 2 is substantially annular and rotationally symmetric to a rotational axis of the turbine stage 14, not shown.
- the combustion exhaust gas 10 flowing into the secondary combustion zone 6 has a temperature which is between 900 0 C and 1600 0 C.
- a combustion precursor upstream of the secondary combustion zone 2 in a common combustion chamber is possible.
- the reheat combustion system 2 includes a fuel mixing assembly 18 having a fuel nozzle 20 through which the fuel gas 8 is disposed in a direction relative to the fuel gas
- Rotation axis of the turbine stage 14 is introduced with a radially inwardly directed direction component in the axially into the secondary combustion zone 2 inflowing combustion exhaust gas 10.
- the fuel mixing arrangement 18 is designed by strong compressors and the nozzle geometry to inject the fuel gas 8 into the combustion exhaust gas 10 in the pulsed and fast injection jet 22.
- the speed of the injection jet 22 can be flexibly adapted to the detected state, depending on sensor signals which contain parameters for a state of the reheat combustion system 2, by not one here shown control unit of
- Fuel mixing arrangement 18 sets.
- the speed is at least in an operating mode in which high-shear combustion is performed, in the range between 0.4 times and 0.9 times the speed of sound in the combustion exhaust gas 10.
- the control unit may increase the speed depending on the pressure and determine the temperature of the combustion exhaust gas 10 or drive a fixed speed of the injection jet 22, which exceeds the corresponding minimum speed of 0.4 times the speed of sound in any case at all temperatures and pressures occurring.
- Fuel mixing arrangement 8 the fuel gas 4 at a speed in the combustion exhaust gas 6, which is between 0.6 and 0.8 times the speed of sound in the combustion exhaust gas 10.
- the fuel nozzle 20 is designed in this embodiment as a subsonic nozzle, so that the
- the fuel gas 8 can inject maximum with a speed in the combustion exhaust gas 10, which corresponds to 0.9 times the speed of sound in the combustion exhaust gas 10.
- fuel mixing arrangement 18 comprises a premixing unit 24, shown only schematically here, for premixing the fuel gas 8 with oxygen-containing gas or an inert substance.
- the premixing unit 24 may premix the fuel gas 8 in a variably adjustable mixing ratio with the corresponding gas.
- the range of possible mixing ratios ie the possible ratios between the number of fuel molecules to the number of In particular, oxygen molecules range between 0.2 and 2.0.
- control unit operates the premixing unit 24 to premix the fuel gas 8 with the oxygen-containing gas in such a ratio that the ratio between the number of fuel molecules and the number of oxygen molecules is less than 1.0.
- the speed of the injection jet 22 is so great that a shear gradient in an edge region 26 of the pulsed jet 12 in a region in front of a nozzle exit 28 is above a critical shear gradient for the self-ignition.
- a length of the region in front of the nozzle outlet 28, in which the shear gradient lies above the critical shear gradient for the self-ignition is at least 10 cm.
- the pressure of the combustion exhaust gas 6 from the primary combustion zone in the secondary combustion zone 2 is about 20 bar, and the pressure of the fuel gas 4 is 30 bar.
- the injection jet 22 consists of fuel gas 8 consisting of an inner jet 30 of fuel-containing gas and an outer jet 32 of cooling gas surrounding the inner jet 30.
- the temperature of the cooling gas is between 200 0 C and 600 0 C, so that the cooling gas has a lower temperature than that
- Combustion exhaust gas 10 which flows from the primary combustion zone into the secondary combustion zone 6.
- fuel gas is combusted in the primary combustion chamber 12 and the hot combustion exhaust gases 10 flow through the turbine stage 14 into the secondary combustion zone 6.
- the fuel gas 8 is injected into the combustion exhaust gas 10 in a jet 12 at a rate at least is as large as 0.2 times the speed of sound in the combustion exhaust gas 10.
- the speed of the outer jet 32 of cooling gas is equal to the speed of the inner jet 30, so that between the inner beam 30 and the outer beam 32 no shear gradient arises.
- the large shear gradient then arises in the edge region 26 at the transition between the outer edge of the outer jet 32 and the combustion exhaust gas 10 surrounding the entire injection jet 22.
- the velocity of the outer jet 32 of cooling gas is less than the velocity of the inner jet 30.
- the cooling gas consists at least substantially of inert material, such as nitrogen, CO 2 or water vapor, wherein the Brennscherinmischan extract 18 can mix the cooling gas in an adjustable ratio fuel to homogenize the flame.
- inert material such as nitrogen, CO 2 or water vapor
- the Brennscherinmischan extract 18 can mix the cooling gas in an adjustable ratio fuel to homogenize the flame.
- air in or as a cooling gas is conceivable
- FIG. 2 shows a fuel nozzle 34 of an alternative reheat combustion system.
- the fuel nozzle 34 comprises an inner tube 36 and an inner tube 36 concentrically unbending outer tube 38, which projects beyond the inner tube 36 in the flow direction and in a front mixing region 40 has a conically tapered cross section, the at a circular outlet opening 42 of the fuel nozzle 34th ends.
- In the inner tube 36 pure fuel or at least a highly fuel-containing gas is guided, while in the space between the inner tube 36 and the outer tube 38, an oxygen-rich sheath flow is performed, which leads air in a preferred embodiment.
- the highly fuel-containing gas and the oxygen-containing jacket stream mix to form the premixed fuel gas
- the fuel gas 8 is accelerated, since the averaged over the beam profile velocity is substantially inversely proportional to the cross-sectional area.
- the premixed fuel gas 8 is finally introduced into an injection jet 22 in the secondary combustion zone 6.
- FIG. 3 shows an alternative reheat combustion system 44, different from that shown in FIGS.
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Gas Burners (AREA)
Abstract
Description
Claims
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
DE102006009562 | 2006-02-28 | ||
PCT/EP2007/051597 WO2007099046A1 (en) | 2006-02-28 | 2007-02-20 | Gas turbine burner and method of operating a gas turbine burner |
Publications (1)
Publication Number | Publication Date |
---|---|
EP1989486A1 true EP1989486A1 (en) | 2008-11-12 |
Family
ID=38009771
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP07712256A Ceased EP1989486A1 (en) | 2006-02-28 | 2007-02-20 | Gas turbine burner and method of operating a gas turbine burner |
Country Status (6)
Country | Link |
---|---|
US (1) | US20100043440A1 (en) |
EP (1) | EP1989486A1 (en) |
JP (1) | JP4776697B2 (en) |
CN (1) | CN101395428B (en) |
RU (1) | RU2406034C2 (en) |
WO (1) | WO2007099046A1 (en) |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
WO2020200568A1 (en) | 2019-04-03 | 2020-10-08 | Siemens Aktiengesellschaft | Heat-shield tile having a damping function |
Families Citing this family (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US9528439B2 (en) * | 2013-03-15 | 2016-12-27 | General Electric Company | Systems and apparatus relating to downstream fuel and air injection in gas turbines |
US10222066B2 (en) * | 2016-05-26 | 2019-03-05 | Siemens Energy, Inc. | Ducting arrangement with injector assemblies arranged in an expanding cross-sectional area of a downstream combustion stage in a gas turbine engine |
US11156156B2 (en) | 2018-10-04 | 2021-10-26 | Raytheon Technologies Corporation | Gas turbine engine with a unitary structure and method for manufacturing the same |
US11174792B2 (en) | 2019-05-21 | 2021-11-16 | General Electric Company | System and method for high frequency acoustic dampers with baffles |
US11156164B2 (en) | 2019-05-21 | 2021-10-26 | General Electric Company | System and method for high frequency accoustic dampers with caps |
Family Cites Families (23)
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FR1217843A (en) * | 1958-12-10 | 1960-05-05 | Snecma | Hot fuel combustion or post-combustion burner |
DE1235670B (en) * | 1962-11-06 | 1967-03-02 | Deutsche Forsch Luft Raumfahrt | Device for flame stabilization in constant pressure combustion chambers |
DE1800611A1 (en) * | 1968-10-02 | 1970-05-27 | Hertel Dr Ing Heinrich | Arrangement for injecting fuel into an air stream flowing past an injection nozzle at supersonic speed |
DE1926728B1 (en) * | 1969-05-24 | 1971-03-25 | Messerschmitt Boelkow Blohm | Combustion chamber for jet engines, especially for rocket ramjet engines |
GB1283827A (en) * | 1970-09-26 | 1972-08-02 | Rolls Royce | Improvements in or relating to combustion apparatus |
FR2392231A1 (en) * | 1977-05-23 | 1978-12-22 | Inst Francais Du Petrole | GAS TURBINE WITH A COMBUSTION CHAMBER BETWEEN THE STAGES OF THE TURBINE |
US4255777A (en) * | 1977-11-21 | 1981-03-10 | Exxon Research & Engineering Co. | Electrostatic atomizing device |
US4581675A (en) * | 1980-09-02 | 1986-04-08 | Exxon Research And Engineering Co. | Electrostatic atomizing device |
US4683541A (en) * | 1985-03-13 | 1987-07-28 | David Constant V | Rotary fluidized bed combustion system |
US4821512A (en) * | 1987-05-05 | 1989-04-18 | United Technologies Corporation | Piloting igniter for supersonic combustor |
US4793305A (en) * | 1987-07-16 | 1988-12-27 | Dresser Industries, Inc. | High turbulence combustion chamber for turbocharged lean burn gaseous fueled engine |
US4896501A (en) * | 1987-10-22 | 1990-01-30 | Faulkner Robie L | Turbojet engine with sonic injection afterburner |
US5070690A (en) * | 1989-04-26 | 1991-12-10 | General Electric Company | Means and method for reducing differential pressure loading in an augmented gas turbine engine |
US4991774A (en) * | 1989-08-24 | 1991-02-12 | Charged Injection Corporation | Electrostatic injector using vapor and mist insulation |
US5093602A (en) * | 1989-11-17 | 1992-03-03 | Charged Injection Corporation | Methods and apparatus for dispersing a fluent material utilizing an electron beam |
US5341640A (en) * | 1993-03-30 | 1994-08-30 | Faulkner Robie L | Turbojet engine with afterburner and thrust augmentation ejectors |
US5515681A (en) * | 1993-05-26 | 1996-05-14 | Simmonds Precision Engine Systems | Commonly housed electrostatic fuel atomizer and igniter apparatus for combustors |
CH688899A5 (en) * | 1994-05-26 | 1998-05-15 | Asea Brown Boveri | A method for controlling a gas turbine group. |
JPH08193716A (en) * | 1995-01-17 | 1996-07-30 | Hitachi Ltd | Gas turbine combustion device |
US6112512A (en) * | 1997-08-05 | 2000-09-05 | Lockheed Martin Corporation | Method and apparatus of pulsed injection for improved nozzle flow control |
GB2390150A (en) * | 2002-06-26 | 2003-12-31 | Alstom | Reheat combustion system for a gas turbine including an accoustic screen |
US6883302B2 (en) * | 2002-12-20 | 2005-04-26 | General Electric Company | Methods and apparatus for generating gas turbine engine thrust with a pulse detonation thrust augmenter |
FR2858661B1 (en) * | 2003-08-05 | 2005-10-07 | Snecma Moteurs | POST-COMBUSTION DEVICE |
-
2007
- 2007-02-20 WO PCT/EP2007/051597 patent/WO2007099046A1/en active Application Filing
- 2007-02-20 CN CN2007800071274A patent/CN101395428B/en not_active Expired - Fee Related
- 2007-02-20 US US12/224,482 patent/US20100043440A1/en not_active Abandoned
- 2007-02-20 JP JP2008556749A patent/JP4776697B2/en not_active Expired - Fee Related
- 2007-02-20 EP EP07712256A patent/EP1989486A1/en not_active Ceased
- 2007-02-20 RU RU2008138545/06A patent/RU2406034C2/en not_active IP Right Cessation
Non-Patent Citations (2)
Title |
---|
None * |
See also references of WO2007099046A1 * |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
WO2020200568A1 (en) | 2019-04-03 | 2020-10-08 | Siemens Aktiengesellschaft | Heat-shield tile having a damping function |
Also Published As
Publication number | Publication date |
---|---|
US20100043440A1 (en) | 2010-02-25 |
CN101395428A (en) | 2009-03-25 |
CN101395428B (en) | 2010-12-08 |
JP4776697B2 (en) | 2011-09-21 |
RU2008138545A (en) | 2010-04-10 |
WO2007099046A1 (en) | 2007-09-07 |
RU2406034C2 (en) | 2010-12-10 |
JP2009528503A (en) | 2009-08-06 |
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