EP1642065A1 - Unite de combustion pour turbine a gaz et turbine a gaz - Google Patents

Unite de combustion pour turbine a gaz et turbine a gaz

Info

Publication number
EP1642065A1
EP1642065A1 EP04740264A EP04740264A EP1642065A1 EP 1642065 A1 EP1642065 A1 EP 1642065A1 EP 04740264 A EP04740264 A EP 04740264A EP 04740264 A EP04740264 A EP 04740264A EP 1642065 A1 EP1642065 A1 EP 1642065A1
Authority
EP
European Patent Office
Prior art keywords
burner
gas turbine
stages
combustion chamber
burner unit
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.)
Granted
Application number
EP04740264A
Other languages
German (de)
English (en)
Other versions
EP1642065B1 (fr
Inventor
Werner Krebs
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Siemens AG
Original Assignee
Siemens AG
Siemens Corp
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Siemens AG, Siemens Corp filed Critical Siemens AG
Priority to EP04740264.9A priority Critical patent/EP1642065B1/fr
Publication of EP1642065A1 publication Critical patent/EP1642065A1/fr
Application granted granted Critical
Publication of EP1642065B1 publication Critical patent/EP1642065B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23M—CASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M20/00—Details of combustion chambers, not otherwise provided for, e.g. means for storing heat from flames
    • F23M20/005—Noise absorbing means
    • 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
    • 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
    • F23R2900/00014—Reducing thermo-acoustic vibrations by passive means, e.g. by Helmholtz resonators

Definitions

  • the invention relates to a burner unit for a gas turbine with a combustion chamber. It also relates to a gas turbine with a number of such burner units.
  • Gas turbines are used in many areas to drive generators or work machines.
  • the energy content of a fuel is used to generate a rotational movement of a turbine shaft.
  • the fuel is burned in a combustion chamber, compressed air being supplied by an air compressor.
  • the working medium generated in the combustion chamber by the combustion of the fuel and under high pressure and at high temperature is guided over a turbine unit downstream of the combustion chamber, where it relaxes while performing work.
  • the invention is therefore based on the object of specifying a burner unit for a gas turbine of the type mentioned above, with which the operational safety and stability of the gas turbine is particularly promoted. Furthermore, a gas turbine is to be specified that can be operated with a particularly high level of operational safety.
  • this object is achieved according to the invention in that a plurality of burner stages are arranged on the combustion chamber, which differ from one another with regard to the acoustic impedance of their fuel feed line, the acoustic impedance of their air passage and / or the flame delay time or the injection delay time.
  • the invention is based on the consideration that the
  • Burner unit can make a particular contribution to the operational stability and safety of the gas turbine by consistently avoiding possible sources of accidents.
  • thermo-acoustically induced combustion instabilities can occur, which cannot be adequately limited by external damping mechanisms.
  • the burner unit comprising the combustion chamber should be suitably designed with regard to its acoustic properties.
  • the burner unit should be designed in several stages with regard to the burners used. A plurality of burner stages are provided, each of which has a fuel gas supply, an air supply, possibly a premixing chamber and a burner outlet, in the manner of a conventional burner.
  • the burner stages should be suitably designed with regard to their dimensioning and the selection of their characteristic parameters. It is provided that the burner stages differ from one another in at least one of the features that distinguish the respective acoustic Characterize the response times of the burner stages to a pressure fluctuation in the combustion chamber, namely the thermoacoustic properties of the fuel supply, characterized by the acoustic impedance of the fuel supply, the thermoacoustic properties of the air supply, characterized by the acoustic impedance of the air passage, and the delay time of the flame, characterized by the time which a fluid element requires from the burner outlet to the flame front, also known as "flame delay time", or by the time that a fluid enriched with fuel ement of the injection on site up to the flame front, also referred to as "injection delay time".
  • the impedance generally expresses the relationship between a force excitation and a movement resulting therefrom, that is to say, for example, in AC technology between the electric field and the resulting current density.
  • the acoustic impedance thus reflects the ratio of a pressure fluctuation to the resulting flow velocity of a medium. There is an amplitude ratio on the one hand and a phase difference on the other hand between a pressure fluctuation and a resulting fluctuation in the flow velocity.
  • the phase difference expresses the extent to which the fluctuation in the flow velocity leads or lags behind the pressure fluctuation causing it, so that the acoustic impedance is, among other things, a suitable measure for the time period between acoustic excitation, for example an acoustic alternating pressure fluctuation, and the response of the respective burner stage on this, so one
  • the multi-stage configuration of the burner unit can be implemented in a particularly favorable manner in that the burner stages are advantageously arranged one behind the other with respect to the longitudinal direction of the gas turbine.
  • Combustion chamber of the burner unit is advantageously designed as an annular combustion chamber.
  • the design as an annular combustion chamber also makes it possible to achieve a relatively homogeneous temperature and flow distribution in the circumferential direction due to its rotational symmetry.
  • the targeted setting of the acoustic properties of the burner stages can be achieved by suitable dimensioning and parameter selection, in particular with regard to the length of the fuel and / or air passage, that is to say the distance between the fuel gas injection and the burner outlet, and / or with regard to the length of the flow passages and volume sizes upstream of the Fuel injection can be set.
  • the burner stages are advantageously each provided with a number of throttle devices and / or with a number of resonator units.
  • the throttle devices can in particular be designed for the targeted generation of pressure losses in the interior of the burner stages, for example in their premixing chambers, where as
  • Throttle device for example perforated plates with appropriately dimensioned bore diameters can be provided.
  • resonators can be used, preferably in flow passages upstream or downstream of the fuel gas injection, advantageously in such a way that they open into the air passage and / or into the fuel passage of the respective burner stage.
  • acoustic decoupling or detuning of the burner stages from one another they are advantageously designed such that the sum of the so-called acoustic time period of each burner stage, i.e. the time period given by the acoustic impedances of the respective burner stage, between the acoustic excitation and the response of the respective burner stage , and the so-called delay time, that is to say the period of time that a fluid element requires for the distance between the exit plane of the respective burner stage and the flame front.
  • the flame delay time is advantageously set via the specification of a suitably selected exit speed at the respective burner outlet and / or via integrated swirl generating means, in particular the ratio of the size of the peripheral speed component to the meridional speed component of the flow medium flowing out of the respective burner stage is used.
  • the stated object is achieved by designing its burner unit as a burner unit of the aforementioned type.
  • the advantages achieved with the invention consist in particular in that a consistent acoustic decoupling of the individual burner stages from one another can be achieved due to the multi-stage design of the burner unit with burner stages which differ suitably in terms of their thermoacoustic properties.
  • the possible excitation of thermoacoustically induced combustion instabilities in the combustion system of the gas turbine can be kept particularly low.
  • a burner unit designed in this way is thus particularly stable with respect to pressure fluctuations in the combustion chamber, so that a gas turbine with such a burner unit has particularly high operational stability.
  • FIG. 2 shows a longitudinal section of a burner unit of the gas turbine according to FIG. 1.
  • the gas turbine 1 according to FIG. 1 has a compressor 2 for
  • Combustion air Combustion air, a burner unit 3 with a combustion chamber 4 and a turbine 6 for driving the compressor 2 and a Nes not shown generator or a work machine.
  • the turbine 6 and the compressor 2 are arranged on a common turbine shaft 8, also referred to as a turbine rotor, to which the generator or the working machine is also connected, and which is rotatably mounted about its central axis 9.
  • the combustion chamber 4 is equipped with a number of burners 10 for burning a liquid or gaseous fuel. It is also provided on its inner wall with heat shield elements, not shown.
  • the turbine 6 has a number of rotatable blades 12 connected to the turbine shaft 8.
  • the rotor blades 12 are arranged in a ring on the turbine shaft 8 and thus form a number of rows of rotor blades.
  • the turbine 6 comprises a number of stationary guide vanes 14, which are also attached to an inner casing 16 of the turbine 6 in a ring shape, with the formation of rows of guide vanes.
  • the blades 12 serve to drive the turbine shaft 8 by means of a pulse transmission from the working medium M flowing through the turbine 6.
  • the guide blades 14, serve to guide the flow of the working medium M between two successive rows of blades or rotating blade rings seen in the flow direction of the working medium M.
  • a successive pair of a ring of guide vanes 14 or a row of guide vanes and a ring of vanes 12 or a row of rotor blades is also referred to as a turbine stage.
  • Each guide vane 14 has a platform 18, also designated as a blade root 19, which is arranged as a wall element for fixing the respective guide vane 14 to the inner housing 16 of the turbine 6.
  • the platform 18 is a thermally comparatively heavily loaded component, which forms the outer boundary of a heating gas channel for the working medium M flowing through the turbine 6.
  • Each blade 12 is in a- similarly attached to the turbine shaft 8 via a blade root 19, also referred to as platform 18, the blade root 19 each carrying a profiled blade 20 extending along a blade axis.
  • each guide ring 21 is arranged on the inner casing 16 of the turbine 6.
  • the outer surface of each guide ring 21 is likewise exposed to the hot working medium M flowing through the turbine 6 and is spaced in the radial direction from the outer end 22 of the rotor blade 12 lying opposite it by a gap.
  • the guide rings 21 arranged between adjacent guide vane rows serve in particular as cover elements which protect the inner wall 16 or other housing installation parts against thermal overloading by the hot working medium M flowing through the turbine 6.
  • each burner stage 30 is in each case connected to a schematically indicated air supply or air passage 32 and to a fuel supply line, not shown in more detail, each opening into a number of inflow openings 34.
  • a flame front 38 assigned to the respective burner stage 30 is formed in the interior of the combustion chamber 4 during operation of the gas turbine 1.
  • three burner stages 30 are shown; however, only two or even four or more burner stages 30 can also be provided.
  • the burner stages 30 are configured differently with respect to the acoustic impedance of their fuel supply line, the acoustic impedance of their air passage 32 and / or their flame delay time .
  • time constants can be derived from the acoustic impedances of the fuel feed line and the air passage 32, which are the time span between a pressure fluctuation occurring in the interior of the combustion chamber 4 and the subsequent reaction of the respective burner stage 30, that is to say a fluctuation in the exit speed at the exit of the flow medium from the respective exit plane 36, playback.
  • the flame delay time that is the period of time that a fluid element in the designed operating state of the gas turbine 1 requires to get from the outlet plane 36 of the respective burner stage 30 to the associated flame front 38
  • the overall time constant for this results in the acoustic period of time to be taken into account for the respective burner stage 30.
  • the burner stages 30 are designed in such a way that they differ from one another with regard to this characteristic time period.
  • the parameters, length of the fuel and / or air passage, length of the flow passages and volume sizes upstream of the fuel injection, pressure losses are in particular in the dimensioning and parameterization of the burner stages 30 the feed lines, outlet speed in the burner outlet plane 36, type of stabilization (swirl stabilized or bluff body stabilized) and / or ratio of the size of the peripheral speed component to the size of the meridional speed component in the flow emerging from the respective burner stage 30. suitably chosen.
  • the first burner stage 30, seen in the flow direction of the working medium M is equipped with an integrated throttle device 40, in the exemplary embodiment a perforated plate, and with resonators, not shown, which are arranged in the flow passages upstream and downstream of the fuel gas injection.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Pre-Mixing And Non-Premixing Gas Burner (AREA)

Abstract

L'invention concerne une unité de combustion (3) destinée à une turbine à gaz (1), cette unité comprenant une chambre de combustion (4). Selon l'invention, cette unité de combustion doit favoriser particulièrement la sécurité de fonctionnement et la stabilité de la turbine à gaz. A cet effet, une pluralité d'étages de combustion (30) sont disposés sur la chambre de combustion (4), ces étages se différenciant du point de vue de l'impédance acoustique de leur alimentation en combustible, de l'impédance acoustique de leur passage d'air et/ou du délai d'allumage ou du délai d'injection.
EP04740264.9A 2003-07-04 2004-06-24 Ensemble de brûleur pour une turbine à gaz et turbine à gaz Expired - Lifetime EP1642065B1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP04740264.9A EP1642065B1 (fr) 2003-07-04 2004-06-24 Ensemble de brûleur pour une turbine à gaz et turbine à gaz

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
EP03015214A EP1493972A1 (fr) 2003-07-04 2003-07-04 Ensemble de brûleur pour une turbine à gaz et turbine à gaz
EP04740264.9A EP1642065B1 (fr) 2003-07-04 2004-06-24 Ensemble de brûleur pour une turbine à gaz et turbine à gaz
PCT/EP2004/006851 WO2005003634A1 (fr) 2003-07-04 2004-06-24 Unite de combustion pour turbine a gaz et turbine a gaz

Publications (2)

Publication Number Publication Date
EP1642065A1 true EP1642065A1 (fr) 2006-04-05
EP1642065B1 EP1642065B1 (fr) 2018-04-25

Family

ID=33427129

Family Applications (2)

Application Number Title Priority Date Filing Date
EP03015214A Withdrawn EP1493972A1 (fr) 2003-07-04 2003-07-04 Ensemble de brûleur pour une turbine à gaz et turbine à gaz
EP04740264.9A Expired - Lifetime EP1642065B1 (fr) 2003-07-04 2004-06-24 Ensemble de brûleur pour une turbine à gaz et turbine à gaz

Family Applications Before (1)

Application Number Title Priority Date Filing Date
EP03015214A Withdrawn EP1493972A1 (fr) 2003-07-04 2003-07-04 Ensemble de brûleur pour une turbine à gaz et turbine à gaz

Country Status (2)

Country Link
EP (2) EP1493972A1 (fr)
WO (1) WO2005003634A1 (fr)

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7886539B2 (en) * 2007-09-14 2011-02-15 Siemens Energy, Inc. Multi-stage axial combustion system
EP2587158A1 (fr) * 2011-10-31 2013-05-01 Siemens Aktiengesellschaft Chambre de combustion pour une turbine à gaz et agencement de brûleur
DE102017207487A1 (de) 2017-05-04 2018-11-08 Siemens Aktiengesellschaft Brennkammer

Family Cites Families (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4122674A (en) * 1976-12-27 1978-10-31 The Boeing Company Apparatus for suppressing combustion noise within gas turbine engines
US5943866A (en) * 1994-10-03 1999-08-31 General Electric Company Dynamically uncoupled low NOx combustor having multiple premixers with axial staging
RU2186298C2 (ru) * 1996-09-16 2002-07-27 Сименс Акциенгезелльшафт Способ и устройство для сжигания топлива с воздухом
SE9802707L (sv) * 1998-08-11 2000-02-12 Abb Ab Brännkammaranordning och förfarande för att reducera inverkan av akustiska trycksvängningar i en brännkammaranordning
DE19939235B4 (de) * 1999-08-18 2012-03-29 Alstom Verfahren zum Erzeugen von heissen Gasen in einer Verbrennungseinrichtung sowie Verbrennungseinrichtung zur Durchführung des Verfahrens
DE19948674B4 (de) * 1999-10-08 2012-04-12 Alstom Verbrennungseinrichtung, insbesondere für den Antrieb von Gasturbinen
GB0019533D0 (en) * 2000-08-10 2000-09-27 Rolls Royce Plc A combustion chamber
US6622487B2 (en) * 2001-01-16 2003-09-23 Rolls-Royce Plc Fluid flow control valve
DE10164097A1 (de) * 2001-12-24 2003-07-03 Alstom Switzerland Ltd Vormischbrenner mit hoher Flammenstabilität
DE10164099A1 (de) * 2001-12-24 2003-07-03 Alstom Switzerland Ltd Brenner mit gestufter Brennstoffeinspritzung
EP1342953A1 (fr) * 2002-03-07 2003-09-10 Siemens Aktiengesellschaft Turbine à gaz

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
See references of WO2005003634A1 *

Also Published As

Publication number Publication date
EP1493972A1 (fr) 2005-01-05
EP1642065B1 (fr) 2018-04-25
WO2005003634A1 (fr) 2005-01-13

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