EP2639505A1 - Système de combustion de turbine à gaz et procédé de stabilisation de la flamme dans un tel système - Google Patents

Système de combustion de turbine à gaz et procédé de stabilisation de la flamme dans un tel système Download PDF

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Publication number
EP2639505A1
EP2639505A1 EP12159203.4A EP12159203A EP2639505A1 EP 2639505 A1 EP2639505 A1 EP 2639505A1 EP 12159203 A EP12159203 A EP 12159203A EP 2639505 A1 EP2639505 A1 EP 2639505A1
Authority
EP
European Patent Office
Prior art keywords
radial
inflow swirler
gas turbine
combustion system
fluid
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Withdrawn
Application number
EP12159203.4A
Other languages
German (de)
English (en)
Inventor
Suresh Sadasivuni
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
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 filed Critical Siemens AG
Priority to EP12159203.4A priority Critical patent/EP2639505A1/fr
Priority to PCT/EP2012/074412 priority patent/WO2013135324A1/fr
Priority to US14/382,314 priority patent/US20150033752A1/en
Priority to EP12798270.0A priority patent/EP2825823B1/fr
Publication of EP2639505A1 publication Critical patent/EP2639505A1/fr
Withdrawn legal-status Critical Current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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air
    • F23R3/12—Air inlet arrangements for primary air inducing a vortex
    • F23R3/14—Air inlet arrangements for primary air inducing a vortex by using swirl vanes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN  A CARRIER GAS OR AIR 
    • F23C7/00—Combustion apparatus characterised by arrangements for air supply
    • F23C7/002—Combustion apparatus characterised by arrangements for air supply the air being submitted to a rotary or spinning motion
    • F23C7/004—Combustion apparatus characterised by arrangements for air supply the air being submitted to a rotary or spinning motion using vanes
    • 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
    • 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
    • F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN  A CARRIER GAS OR AIR 
    • F23C2900/00—Special features of, or arrangements for combustion apparatus using fluid fuels or solid fuels suspended in air; Combustion processes therefor
    • F23C2900/07001—Air swirling vanes incorporating fuel injectors
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23D—BURNERS
    • F23D2900/00—Special features of, or arrangements for burners using fluid fuels or solid fuels suspended in a carrier gas
    • F23D2900/14—Special features of gas burners
    • F23D2900/14021—Premixing burners with swirling or vortices creating means for fuel or air

Definitions

  • the present invention relates to a gas turbine combustion system and to flame stabilisation in a gas turbine combustion system.
  • the invention relates to flame stabilisation in swirl stabilized diffusion flames.
  • US 6,311,496 B1 describes a gas turbine combustion system with two radial inflow swirlers that are successively used by the airstream.
  • the first objective is achieved by a gas turbine combustion system as claimed in claim 1.
  • the second objective is achieved by a method of flame stabilisation in a gas turbine combustion system as claimed in claim 9.
  • the depending claims contain further developments of the invention.
  • An inventive gas turbine combustion system comprises a central axis and a radial direction with respect to said central axis, a first radial inflow swirler and a second radial inflow swirler.
  • the first radial inflow swirler has radial outer intake openings located at a radial outer circumference of the first radial inflow swirler.
  • the radial outer intake openings of the first radial inflow swirler are refered to as first radial outer intake openings in the following.
  • the first radial inflow swirler has outlet openings located at a radial inner circumference of the first radial inflow swirler. These outlet openings are referred to as first radial inner outlet openings in the following.
  • Flow passages named first flow passages in the following, extend from the first radial outer intake openings to the first radial inner outlet openings. Each first flow passage includes a first angle with respect to the radial direction.
  • the gas turbine combustion system further comprises a second radial inflow swirler having radial outer intake openings which are located at a radial outer circumference of the second radial inflow swirler and which are referred to as second radial outer intake openings in the following.
  • the second radial inflow swirler has radial inner outlet openings, which are referred to as second radial inner outlet openings in the following and which are located at a radial inner circumference of the second radial inflow swirler.
  • Flow passages named second flow passages in the following, extend from the second radial outer intake openings to the second radial inner outlet openings.
  • Each second flow passage includes an angle with respect to the radial direction. This angle is referred to as a second angle in the following.
  • the number of second flow passages may be identical to the number of first flow passages.
  • the radial outer circumference of the second radial inflow swirler has a diameter that is at least slightly smaller than the diameter of the radial inner circumference of the first radial inflow swirler, and the second radial inflow swirler is located coaxially with and radially inside the first radial inflow swirler.
  • the first angle has a different sign than the second angle with respect to the radial direction.
  • the second radial inflow swirler produces a swirl counterrotating with respect to the swirl generated by the first radial inflow swirler.
  • the counterrotation produced by the two swirlers leads to a more uniform mixing of an oxidant, like in particular the oxygen in the air, and fuel and to a stable flame which has the advantages of lesser flameouts, a more distributed mixing of fuel and the oxidant, a better control of the combustion burner, lesser hotspots and a lower heat load across the metal surfaces like, for example, the combustor walls.
  • the first angle and the second angle may have the same absolute value so that they only differ in their orientation with respect to the radial direction.
  • fuel injection openings are located in the second radial inflow swirler and are open towards the second flow passages. More preferably, the fuel injection openings are located inside the second flow passages, in particular in the radial outer half of the second flow passages, preferably in the outer third of the second flow passages.
  • a radial gap may be present between the radial inner circumference of the first radial inflow swirler and the radial outer circumference of the second radial inflow swirler.
  • the flow cross section of the second flow passages may be smaller than the flow cross section of the first flow passages since part of the fluid can be introduced into a combustion chamber through the radial gap while another part will be introduced into the combustion chamber through the second radial inflow swirler.
  • a method of flame stabilisation in a gas turbine combustion system is provided.
  • a fluid flows along a flow path with a radial component from a fluid inlet to a fluid outlet.
  • the fluid is a fluid that comprises an oxidant
  • a fuel is mixed with the fluid that comprises an oxidant so as to transform the fluid into a mixture comprising fuel and the oxidant.
  • air is used as the fluid (that comprises oxygen as the oxidant) the fluid is transformed into a fuel/air mixture.
  • a first swirl with a first rotational direction is introduced into the flowing fluid in a radial upstream section of the flow path.
  • a second swirl with a second rotational direction is introduced into at least a portion of the fluid in a radial downstream section of the flow path.
  • the second rotational direction represents a counterrotation with respect to the first rotational direction.
  • the inventive method is particularly effective in improving flame stability and uniform mixing of fuel and oxidant if fuel is introduced into the fluid where the second swirl is generated.
  • the fuel is introduced into the fluid at a location where generation of the second swirl begins.
  • no second swirl is introduced into a portion of the fluid.
  • inventive combustion system will be described with respect to Figures 1 and 2 in the context of a combustor arrangement including an inventive combustion system.
  • inventive combustion system is adapted for performing the inventive method of flame stabilisation in a gas turbine combustion system which will also be described with respect to Figures 1 and 2 .
  • Figure 1 shows part of a combustor arrangement in a sectional view.
  • the combustor arrangement comprises a combustion chamber 3 and a combustion system 1 that is connected to a combustion chamber 3 via a small pre-chamber 5.
  • the pre-chamber is sometimes also called transition section and may be part of the combustion system 1 like in the present embodiment.
  • the pre-chamber 5 may as well be a part of the combustion chamber 3 or a distinct part that is neither part of the combustion system 1 nor of the combustion chamber 3.
  • the combustion system 1 comprises a first radial inflow swirler 7 that, shows rotational symmetry with respect to a central combustor axis A.
  • the first radial inflow swirler is equipped with a number of vanes 9 that are distributed along the circumferential direction of the swirler 7 and are spaced apart from each other.
  • Flow passages 11 are formed between neighbouring vanes 9.
  • Each flow passage 11 extends from a first radial outer intake opening 13 located at a radial outer circumference of the swirler 7 to a first radial inner outlet opening 15 located at a radial inner circumference of the swirler 7.
  • the flow passages 11 of the first swirler 7 are angled with respect to the radial direction of the swirler with a first angle ⁇ so that a swirl is imparted to a fluid flowing through the flow channel 11.
  • the combustion system 1 further comprises a second radial inflow swirler 17 that, like the first radial inflow swirler, shows radial symmetry.
  • the second radial inflow swirler 17 has an outer circumference the diameter of which is smaller than the inner circumference of the first radial inflow swirler 11.
  • the second radial inflow swirler 17 is located inside an opening formed by the inner circumference of the first radial inflow swirler 7 so that a fluid that exists the outlet openings 15 of the first radial inflow swirler 7 is directed towards the second radial inflow swirler 17.
  • the second radial inflow swirler 17 comprises a number of vanes 19 that are distributed in circumferential direction of the swirler such that second flow passages 21 are formed between them.
  • Each second flow passages 21, i.e. each flow passage of the second radial inflow swirler 17, extends from a second radial outer intake opening 23 located at the radial outer circumference of the second swirler to a second radial inner outlet opening 25, i.e. an outlet opening of the second swirler 17 that is located at the inner circumference of the second radial inflow swirler 17.
  • the flow channels 21 of the second radial inflow swirler 17 include an angle with the radial direction (denominated ⁇ in Figure 2 ) which has, in the present embodiment, the same absolute value as the angle of the flow channels 11 of the first radial inflow swirler 7 but a different sign.
  • the flow channels 11 of the first radial inflow swirler 7 impart a clockwise swirl to a flowing fluid
  • the flow channels 21 of the second radial inflow swirler 17 impart a counter-clockwise swirl to a fluid flowing therethrough, or vice versa.
  • Both swirlers 7, 17 are mounted to a base plate 31 such that they are arranged coaxially with each other and with respect to the combustor axis A. Moreover, in the present embodiment they are arranged such that a radial gap 27 is formed between the inner circumference of the first radial inflow swirler 7 and the outer circumference of the second radial inflow swirler 17.
  • Fuel channels 33 extend through the base plate 31 and lead to fuel opening 29 in the flow passages 21 of the second radial inflow swirler 7.
  • the fuel openings 29 are located in the outer half of the second flow passages 21, preferably in the outer third of the second flow passages 21, and more preferably in the outer fourth of the second flow passages 21.
  • the first radial inflow swirler 7 is surrounded by a flow channel 35 which allows feeding a fluid, in particular air or any other suitable fluid that comprises an oxidant, to the intake openings 13 of the first radial inflow swirler.
  • the intake openings 23 of the second radial inflow swirler generate turbulences in the flow channel sections adjoining the intake openings 15.
  • the turbulences are generated due to a reversal in rotation direction that is necessary for the air to enter the flow passages 21 of the second swirler 17.
  • the turbulence are highest in a flow passage zone adjoining the intake openings 23 of the flow passages.
  • a fuel gas like, for example, syngas or coke oven gas (COG) is introduced into the turbulent airstreams in the second flow passages 21 through the fuel holes 29.
  • the strong turbulence leads to a highly uniform mixing of fuel and air until the fuel/air mixture leaves the second flow channels 21 through the second outlet openings 25.
  • Due to the angle ⁇ the second flow passages 21 include with the radial direction a second swirl (indicate by arrow 39) with a counter-clockwise rotation is imparted to the fuel/air mixture flowing through the second flow passages 21.
  • a further effect of giving the angle of the flow channels of the first and second swirlers a different sign with respect to the radial direction is that the fuel/air mixture has a different direction of rotation than the air entering the pre-chamber 5 through the gap 27 that is present between both swirlers 7, 17 in the described embodiment.
  • the air rotating clockwise in the present embodiment can form an envelop around the fuel/air mixture rotating counter-clockwise in the present embodiment which makes it more difficult for fuel/air mixture to reach the wall of the pre-chamber 5 and the combustion chamber 3, thereby reducing heat load across the metal surface of the combustor wall.
  • a further advantage is that turbulences are formed where the counter-clockwise swirling fuel/air mixture is in contact with the clockwise swirling air, which turbulences lead to a more distributed mixing of fuel and air.
  • the mentioned effects contribute to leading to less flameouts and less hotspots, in particular with use of H 2 containing gases like syngas or COG. In the end, this leads to a better controllable combustion burner.
  • the present invention has been described with respect to a specific embodiment to describe a method of improve mixing of gas and air and to stabilise the flame by using the concepts of swirl strength in diffusion flames to anchor it in a stabile way.
  • counterrotating swirls are used to improve mixing and stabilising of the flame.
  • the invention shall not be restricted to the specific embodiment described with respect to the figures, since deviations thereform are possible.
  • both swirlers have the same number of flow passages the second wirler could have a higher or lower number of flow passages than the first swirler.
  • the flow passages of both swirlers are angled by the same absolute value with respect to the radial direction but with a different sign.
  • a further possible deviation from the embodiment described with respect to the figures is the number of fuel opening that are present in each flow passage of the second swirler. While in the described embodiment only one fuel openings is present in each flow passage a higher number of fuel openings may be present as well. Moreover, the fuel openings do not need to be present in the base plate. Alternatively or additionally, fuel openings could be located in the sidewalls of the vanes. Since the location of the fuel openings is closely related to the geometry of the swirler and the fuel to be used the exact position of the fuel openings may depend on the concrete design of the first and second radial inflow swirler and/or on the intended use of the combustion system.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
EP12159203.4A 2012-03-13 2012-03-13 Système de combustion de turbine à gaz et procédé de stabilisation de la flamme dans un tel système Withdrawn EP2639505A1 (fr)

Priority Applications (4)

Application Number Priority Date Filing Date Title
EP12159203.4A EP2639505A1 (fr) 2012-03-13 2012-03-13 Système de combustion de turbine à gaz et procédé de stabilisation de la flamme dans un tel système
PCT/EP2012/074412 WO2013135324A1 (fr) 2012-03-13 2012-12-05 Système de combustion de turbine à gaz et procédé de stabilisation des flammes dans un tel système
US14/382,314 US20150033752A1 (en) 2012-03-13 2012-12-05 Gas turbine combustion system and method of flame stabilization in such a system
EP12798270.0A EP2825823B1 (fr) 2012-03-13 2012-12-05 Système de combustion de turbine à gaz et procédé de stabilisation de la flamme dans un tel système

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP12159203.4A EP2639505A1 (fr) 2012-03-13 2012-03-13 Système de combustion de turbine à gaz et procédé de stabilisation de la flamme dans un tel système

Publications (1)

Publication Number Publication Date
EP2639505A1 true EP2639505A1 (fr) 2013-09-18

Family

ID=47324135

Family Applications (2)

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EP12159203.4A Withdrawn EP2639505A1 (fr) 2012-03-13 2012-03-13 Système de combustion de turbine à gaz et procédé de stabilisation de la flamme dans un tel système
EP12798270.0A Not-in-force EP2825823B1 (fr) 2012-03-13 2012-12-05 Système de combustion de turbine à gaz et procédé de stabilisation de la flamme dans un tel système

Family Applications After (1)

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EP12798270.0A Not-in-force EP2825823B1 (fr) 2012-03-13 2012-12-05 Système de combustion de turbine à gaz et procédé de stabilisation de la flamme dans un tel système

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US (1) US20150033752A1 (fr)
EP (2) EP2639505A1 (fr)
WO (1) WO2013135324A1 (fr)

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SG186083A1 (en) * 2010-07-02 2013-01-30 Exxonmobil Upstream Res Co Low emission triple-cycle power generation systems and methods
US20150285502A1 (en) * 2014-04-08 2015-10-08 General Electric Company Fuel nozzle shroud and method of manufacturing the shroud
EP3882547A1 (fr) * 2020-03-20 2021-09-22 Primetals Technologies Germany GmbH Tube de brûleur, module de tube de brûleur et unité de brûleur

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EP0939275A2 (fr) * 1997-12-30 1999-09-01 United Technologies Corporation Injecteur de combustible et dispositif de guidage de l'injecteur pour une turbine à gaz
WO2000049337A1 (fr) * 1999-02-16 2000-08-24 Siemens Aktiengesellschaft Systeme de bruleurs et procede permettant de le faire fonctionner
US6253555B1 (en) 1998-08-21 2001-07-03 Rolls-Royce Plc Combustion chamber comprising mixing ducts with fuel injectors varying in number and cross-sectional area
US6311496B1 (en) 1997-12-19 2001-11-06 Alstom Gas Turbines Limited Gas turbine fuel/air mixing arrangement with outer and inner radial inflow swirlers
US20050257530A1 (en) * 2004-05-21 2005-11-24 Honeywell International Inc. Fuel-air mixing apparatus for reducing gas turbine combustor exhaust emissions
EP2192347A1 (fr) 2008-11-26 2010-06-02 Siemens Aktiengesellschaft Dispositif de tourbillonnement double

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EP0660038A2 (fr) * 1993-12-23 1995-06-28 ROLLS-ROYCE plc Dispositif d'injection de carburant
US6311496B1 (en) 1997-12-19 2001-11-06 Alstom Gas Turbines Limited Gas turbine fuel/air mixing arrangement with outer and inner radial inflow swirlers
EP0939275A2 (fr) * 1997-12-30 1999-09-01 United Technologies Corporation Injecteur de combustible et dispositif de guidage de l'injecteur pour une turbine à gaz
US6253555B1 (en) 1998-08-21 2001-07-03 Rolls-Royce Plc Combustion chamber comprising mixing ducts with fuel injectors varying in number and cross-sectional area
WO2000049337A1 (fr) * 1999-02-16 2000-08-24 Siemens Aktiengesellschaft Systeme de bruleurs et procede permettant de le faire fonctionner
US20050257530A1 (en) * 2004-05-21 2005-11-24 Honeywell International Inc. Fuel-air mixing apparatus for reducing gas turbine combustor exhaust emissions
EP2192347A1 (fr) 2008-11-26 2010-06-02 Siemens Aktiengesellschaft Dispositif de tourbillonnement double

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Also Published As

Publication number Publication date
EP2825823B1 (fr) 2016-03-23
WO2013135324A1 (fr) 2013-09-19
US20150033752A1 (en) 2015-02-05
EP2825823A1 (fr) 2015-01-21

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