EP1642065B1 - Ensemble de brûleur pour une turbine à gaz et turbine à gaz - Google Patents

Ensemble de brûleur pour une turbine à gaz et turbine à gaz Download PDF

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Publication number
EP1642065B1
EP1642065B1 EP04740264.9A EP04740264A EP1642065B1 EP 1642065 B1 EP1642065 B1 EP 1642065B1 EP 04740264 A EP04740264 A EP 04740264A EP 1642065 B1 EP1642065 B1 EP 1642065B1
Authority
EP
European Patent Office
Prior art keywords
burner
gas turbine
stages
stage
combustion chamber
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.)
Expired - Lifetime
Application number
EP04740264.9A
Other languages
German (de)
English (en)
Other versions
EP1642065A1 (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
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Publication date
Application filed by Siemens AG 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

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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
    • F23R3/42—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 further 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 produced in the combustion chamber by the combustion of the fuel, under high pressure and at high temperature, is guided via a turbine unit arranged downstream of the combustion chamber, where it relaxes to perform 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 promoted to a particular extent. Furthermore, a gas turbine is to be specified, which is operable with particularly high operational safety.
  • this object is achieved according to the invention by a plurality of burner stages is arranged on the combustion chamber, which are defined in terms of the sum of the acoustic period of the respective burner stage, as the time period between an acoustic stimulus and the response of the respective Burning stage, and the respective delay time, defined as the period of time that requires a fluid element for the distance between the exit plane of the respective burner stage and the flame front, the burner stages from each other with regard to the acoustic impedance of their fuel supply, the acoustic impedance of their Air passage and / or the flame delay time or the injection delay time differ, wherein the burner stages are arranged with respect to the longitudinal direction of the gas turbine one behind the other.
  • the invention is based on the consideration that the burner unit can contribute in particular to the operational stability and safety of the gas turbine by possible sources of accident are consistently avoided.
  • thermoacoustically induced combustion instabilities which are not sufficiently limited by external damping mechanisms, can occur, especially in gas turbines designed for high power densities and combustion temperatures in a compact design as a possible source of accident can.
  • the burner unit comprising the combustion chamber should be designed appropriately with regard to its acoustic properties.
  • the design goal may be to suppress a coupling between the thermoacoustic response times of the burner flames and the acoustic natural frequencies of the combustion system, which could lead to the excitation of thermoacoustically induced combustion instabilities.
  • the burner unit should be designed in several stages with regard to the burners used. In this case, a plurality of burner stages are provided, each of which has in the manner of a conventional burner via a fuel gas supply, an air supply, optionally a premixing chamber and a burner outlet.
  • the burner stages should be designed appropriately in terms of their dimensions and the choice of their characteristic parameters. It is provided that the burner stages differ from each other in at least one of the features that characterize the respective acoustic 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 required for a fluid element from the burner exit to the flame front, also referred to as "flame delay time", or by the time a fuel-enriched fluid element of the injection site needed 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, ie, for example, in the alternating current technique 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. Between a pressure fluctuation and a resulting fluctuation in the flow velocity there is an amplitude ratio on the one hand and a phase difference on the other hand.
  • the phase difference expresses the extent to which the fluctuation of the flow velocity precedes or hinders the pressure fluctuation causing it, such that the acoustic impedance is, inter alia, a suitable measure for the time span between an acoustic excitation, for example an acoustic alternating pressure fluctuation, and the response of the respective burner stage this, ie a fluctuation of the exit velocity at the respective Brenneraustrittsebene is.
  • the multi-stage embodiment of the burner unit can be implemented in a particularly favorable manner by the burner stages are 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. Due to the design as an annular combustion chamber is also due to their rotational symmetry seen in the circumferential direction comparatively homogeneous temperature and flow distribution achievable.
  • the targeted adjustment of the acoustic properties of the burner stages can by suitable dimensioning and parameter selection, in particular with regard to the length of the fuel and / or air passage, so the distance between the fuel injection and the burner exit, and / or with respect to the length of the flow passages and volume sizes upstream of the fuel injection.
  • the burner stages are advantageously provided in each case with a number of throttle devices and / or with a number of resonator units.
  • the throttle devices may be designed in particular for the targeted generation of pressure losses in the interior of the burner stages, for example in their premixing chambers, wherein, for example, perforated plates with suitably dimensioned bore diameters may be provided as throttling device.
  • resonators may 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.
  • a reliable acoustic decoupling or detuning of the burner stages from each other they are advantageously designed such that the sum of the so-called acoustic period of each burner stage, so given by the acoustic impedances of the respective burner stage period between the acoustic stimulus and the response of the respective burner stage , And the so-called delay time, ie the period of time, which requires a fluid element for the distance between the exit plane of the respective burner stage and the flame front, different from each other.
  • the flame delay time is advantageously set via the specification of a suitably selected outlet speed at the respective burner outlet and / or via integrated swirl-generating means, wherein in particular the ratio of the size of the circumferential velocity component to the meridional velocity 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 by the invention are in particular that a consistent acoustic decoupling of the individual burner stages from each other can be achieved by the multi-stage design of the burner unit with burner stages, which differ in terms of their thermoacoustic properties suitable from each other.
  • 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 therefore particularly stable to pressure fluctuations in the combustion chamber, so that a gas turbine with such a burner unit has a particularly high operational stability.
  • the gas turbine 1 has a compressor 2 for combustion air, a burner unit 3 with a combustion chamber 4 and a turbine 6 for driving the compressor 2 and a Not shown generator or a working machine.
  • the turbine 6 and the compressor 2 are arranged on a common, also called turbine rotor turbine shaft 8, with which the generator or the working machine is connected, and which is rotatably mounted about its central axis 9.
  • the combustion chamber 4 is equipped with a number of burners 10 for the combustion of 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 blades 12 are arranged in a ring on the turbine shaft 8 and thus form a number of blade rows.
  • the turbine 6 comprises a number of fixed vanes 14, which are also fixed in a ring shape with the formation of rows of vanes on an inner casing 16 of the turbine 6.
  • the blades 12 serve to drive the turbine shaft 8 by momentum transfer from the turbine 6 flowing through the working medium M.
  • the vanes 14, however, serve to guide the flow of the working medium M between two seen in the flow direction of the working medium M consecutive blade rows or blade rings.
  • a successive pair of a ring of vanes 14 or a row of vanes and a ring of blades 12 or a blade row is also referred to as a turbine stage.
  • Each vane 14 has a platform 18, also referred to as a blade root 19, which is arranged to fix the respective vane 14 on the inner housing 16 of the turbine 6 as a wall element.
  • 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 analog Said manner attached to the turbine shaft 8 via a blade root 19, also referred to as a platform 18, wherein the blade root 19 each carries a profiled airfoil 20 extending along a blade axis.
  • each guide ring 21 on the inner housing 16 of the turbine 6 is arranged between the spaced-apart platforms 18 of the guide vanes 14 of two adjacent rows of guide vanes.
  • the outer surface of each guide ring 21 is also exposed to the hot, the turbine 6 flowing through the working medium M and spaced in the radial direction from the outer end 22 of the blade 12 opposite him through a gap.
  • the guide rings 21 arranged between adjacent rows of guide blades serve in particular as cover elements which protect the inner wall 16 or other housing installation parts from thermal overload by the hot working medium M flowing through the turbine 6.
  • the burner unit 3 which is in FIG. 2 is shown in a longitudinal section, executed in several stages, wherein seen at the combustion chamber designed as an annular combustion chamber 4 in the flow direction of the working medium M a plurality of burner stages 30 is arranged one behind the other.
  • Each burner stage 30 is in each case connected to a schematically indicated air supply or air passage 32 and to a not shown in detail, each opening in a number of inlet openings 34 fuel supply line.
  • each burner stage 30 is formed during operation of the gas turbine 1 in the interior of the combustion chamber 4, one of the respective burner stage 30 associated flame front 38.
  • three burner stages 30 are shown; but it can also be provided only two or four or more burner stages 30.
  • the burner stages 30 with respect to the acoustic impedance of their fuel supply, the acoustic impedance of their air passage 32 and / or their flame delay time designed differently from each other.
  • time constants can be derived from the acoustic impedances of the fuel supply 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, ie a fluctuation of the exit velocity on exit of the flow medium respective exit plane 36, play.
  • this time constant results in the total for the acoustic one Design of the respective burner stage 30 to be considered period.
  • the burner stages 30 are designed such that they differ from one another with regard to this characteristic period of time.
  • the dimensioning and parameterization of the burner stages 30 in particular 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 in the supply lines Exit velocity in the burner exit plane 36, stabilization (swirl stabilized or bluff body stabilized) and / or ratio of the size of the peripheral velocity component to the size of the meridional velocity component in the emerging from the respective burner stage 30 flow suitable selected.
  • a perforated plate in the embodiment, a perforated plate, and arranged in the flow passages upstream and downstream of the fuel gas injection, resonators, not shown.

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)

Claims (6)

  1. Unité (3) de brûleur pour une turbine (1) à gaz, comprenant une chambre de combustion (4) sur laquelle est disposée une pluralité d'étages (30) de brûleur, qui se distinguent les uns des autres en ce qui concerne la somme du laps de temps acoustique de l'étage (30) de brûleur respectif, défini comme étant le laps de temps entre une excitation acoustique et la réponse de l'étage (30) de brûleur respectif, et du temps de retard respectif, défini comme le laps de temps nécessaire à un élément fluide pour la distance entre le plan (36) de sortie de l'étage (30) de brûleur respectif et le front (38) de flamme, les étages (30) de brûleur se distinguant les uns des autres, en ce qui concerne l'impédance acoustique de leur conduit d'arrivée de combustible, l'impédance acoustique de leur passage d'air et/ou le temps de retard de flamme ou le temps de retard d'injection, caractérisée en ce que les étages (30) de brûleur sont disposés les uns derrière les autres, rapporté à la direction longitudinale de la turbine (1) à gaz.
  2. Unité (3) de brûleur suivant la revendication 1, dont la chambre de combustion (4) est constituée sous la forme d'une chambre de combustion annulaire.
  3. Unité (3) de brûleur suivant l'une des revendications 1 ou 2, dont les étages (30) de brûleur sont pourvus chacun d'un certain nombre de dispositifs (40) d'étranglement.
  4. Unité (3) de brûleur suivant l'une des revendications 1 à 3, dont les étages (30) de brûleur sont pourvus chacun d'un certain nombre d'unités de résonateur, qui débouchent de préférence dans le passage d'air et/ou dans le passage de combustible de chaque étage (30) de brûleur.
  5. Unité (3) de brûleur suivant l'une des revendications 1 à 4, dans les étages (30) de brûleur de laquelle le temps de retard de flamme est réglé par la prescription d'une vitesse de sortie à la sortie du brûleur et/ou par des moyens intégrés de production de tourbillonnement.
  6. Turbine (1) à gaz ayant un certain nombre d'unités (3) de brûleur suivant l'une des revendications 1 à 5.
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
PCT/EP2004/006851 WO2005003634A1 (fr) 2003-07-04 2004-06-24 Unite de combustion pour turbine a gaz et turbine a gaz
EP04740264.9A EP1642065B1 (fr) 2003-07-04 2004-06-24 Ensemble de brûleur pour une turbine à gaz et turbine à gaz

Publications (2)

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

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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
EP0925472B1 (fr) * 1996-09-16 2001-04-04 Siemens Aktiengesellschaft Procede pour la suppression des oscillations de combustion et dispositif pour la combustion d'un combustible avec de l'air
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
None *

Also Published As

Publication number Publication date
EP1493972A1 (fr) 2005-01-05
WO2005003634A1 (fr) 2005-01-13
EP1642065A1 (fr) 2006-04-05

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