EP3184898A1 - Brennkammer für eine gasturbine - Google Patents

Brennkammer für eine gasturbine Download PDF

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Publication number
EP3184898A1
EP3184898A1 EP15202500.3A EP15202500A EP3184898A1 EP 3184898 A1 EP3184898 A1 EP 3184898A1 EP 15202500 A EP15202500 A EP 15202500A EP 3184898 A1 EP3184898 A1 EP 3184898A1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
combustor
swirler
lip
pilot
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Withdrawn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
EP15202500.3A
Other languages
English (en)
French (fr)
Inventor
James Hird
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 EP15202500.3A priority Critical patent/EP3184898A1/de
Priority to CN201680075991.7A priority patent/CN108431503A/zh
Priority to US15/781,713 priority patent/US20180363904A1/en
Priority to PCT/EP2016/080488 priority patent/WO2017108454A1/en
Priority to EP16812715.7A priority patent/EP3394514A1/de
Publication of EP3184898A1 publication Critical patent/EP3184898A1/de
Withdrawn legal-status Critical Current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
    • 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
    • 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
    • F23R3/343—Pilot flames, i.e. fuel nozzles or injectors using only a very small proportion of the total fuel to insure continuous combustion

Definitions

  • the present invention relates to a combustor for a gas turbine.
  • a combustor generally comprises a main combustion chamber and a pre-combustion chamber, upstream of the main combustion chamber.
  • the pre-combustion chamber comprises a swirler section having a swirler through which a main fuel stream is provided.
  • the main fuel is mixed to a non-combustible gas flow comprising an oxidant, for example air.
  • the main fuel stream and the non-combustible gas flow are injected via the swirler into the pre-combustion chamber of the combustor in a generally tangential direction with respect to the centre axis of the combustor.
  • a pilot fuel is further injected in the pre-combustion chamber for controlling the combustor flame in which the main fuel in burned.
  • the pilot fuel is typically injected by a pilot burner, generally according a direction parallel to the centre axis of the combustor.
  • the pilot fuel is injected from the pilot burner into the pre-combustion chamber through a plurality of pilot fuel injectors arranged on the pilot burner surface, i.e. the surface separating the pilot burner from the pre-combustion chamber.
  • the main fuel and the pilot fuel is a gaseous fuel. Liquid fuel injection may also be provided in similar positions on the swirler and on the pilot burner.
  • the combustion of the pilot fuel is achieved through an oxidant, for example air, first being mixed together with the fuel in the pilot burner.
  • the injected pilot fuel generates a diffusion flame inside the pre-combustion chamber, close to pilot burner surface. This has the main drawback of increasing the local temperature at the pilot burner surface, with the consequence of reducing the life cycle of the pilot burner.
  • a combustor for a gas turbine comprises:
  • the combustor may be an annular-type or a can-type combustor.
  • the combustion chamber may have a cylindrical or oval shape.
  • the combustion chamber may comprise a main combustion chamber and a pre-combustion chamber with a swirler section.
  • the centre axis of the pre-combustion chamber may be a symmetry line of the pre-combustion chamber.
  • the swirler is mounted to the pre-combustion chamber and surrounds the pre-combustion chamber centre axis.
  • the inclined orientation of the lip guides the flow away from the pilot burner surface and towards a main combustion zone of the pre-combustion chamber.
  • the injection of oxidant gas through the feed passages enhance mixing of the oxidant gas with the pilot fuel from the pilot fuel injector.
  • temperature at the pilot burner surface is reduced, up to more acceptable values, which make life of the pilot burner longer.
  • an inclination angle is comprised between 30° and 60° has proved to be particularly advantageous.
  • the lip further comprises an external surface oriented towards the swirler for intercepting at least part of the flow of oxidant gas coming from the swirler, the feed passages being provided between the internal surface and the external surface.
  • the feed passages may be provided in plurality, regularly distributed around the centre axis.
  • the external surface may be provided with a plurality of turbulators for inducing turbulence in at least part of the flow of oxidant gas coming from the swirler.
  • the turbulators may comprise a plurality of protrusions extending orthogonally from the external surface and/or a plurality of channels having a depth extending from the external surface towards the internal surface.
  • the protrusions may comprise a circular rim concentric with the centre axis of the pre-combustion chamber.
  • the turbulators enhance the turbulence from the swirler oxidant gas to mix with the pilot fuel emerging from the lip towards the inside of the pre-combustion chamber. This produces premixed pilot for lowering temperatures and hence NOx emissions.
  • the internal surface and the external surface have a common trailing edge, at the end of the lip, where both the pilot fuel and the oxidant gas separate from the lip.
  • the trailing edge may have a circular profile around the centre axis of the pre-combustion chamber.
  • the trailing edge may have a waved profile.
  • the above described designs of the end portion of the lip improve turbulence and flow aerodynamics. Pressure loss may be also reduced.
  • the internal surface has an aerofoil shape.
  • this improves turbulence and flow aerodynamics and reduces pressure loss.
  • the lip is provided as an edge of a shroud of the pilot burner extending inside the pre-combustion chamber.
  • this allows to manufacture a pilot burner directly including a lip optimised for the present invention.
  • Fig. 1 shows an example of a gas turbine engine 10 in a sectional view.
  • the gas turbine engine 10 comprises, in flow series, an inlet 12, a compressor section 14, a burner section 16 and a turbine section 18 which are generally arranged in flow series and generally about and in the direction of a longitudinal or rotational axis 20.
  • the gas turbine engine 10 further comprises a shaft 22 which is rotatable about the rotational axis 20 and which extends longitudinally through the gas turbine engine 10.
  • the shaft 22 drivingly connects the turbine section 18 to the compressor section 14.
  • air 24 which is taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or burner section 16.
  • the burner section 16 comprises a burner plenum 26, one or more combustion chambers 28, each having a respective upstream pre-combustion chamber 101.
  • the burner section 16 further comprises at least one pilot burner 30 and a swirler section 31 fixed to each pre-combustion chamber 101.
  • the pre-combustion chambers 101, the combustion chambers 28, the pilot burners 30 and the swirler section 31 are located inside the burner plenum 26.
  • the compressed air passing through the compressor 14 enters a diffuser 32 and is discharged from the diffuser 32 into the burner plenum 26 from where a portion of the air enters the pilot burner 30 and is mixed with a gaseous or liquid pilot fuel.
  • the air/fuel mixture is then burned and the combustion gas 34 or working gas from the combustion is channelled through the combustion chamber 28 to the turbine section 18 via a transition duct 17.
  • a main flow of air/fuel mixture is further inserted in the pre-combustion chamber 101 through the swirler section 31, as better detailed in a following section of the present text.
  • the main fuel burns when mixing with the hot gasses in the pre-combustion chamber 101 and in the main combustor chamber 28.
  • This exemplary gas turbine engine 10 has a cannular combustor section arrangement, which is constituted by an annular array of combustor cans 19 each having a pilot burner 30 and a combustion chamber 28, the transition duct 17 having a generally circular inlet that interfaces with the combustor chamber 28 and an outlet in the form of an annular segment.
  • An annular array of transition duct outlets form an annulus for channelling the combustion gases to the turbine 18.
  • the turbine section 18 comprises a number of blade carrying discs 36 attached to the shaft 22.
  • two discs 36 each carry an annular array of turbine blades 38.
  • the number of blade carrying discs could be different, i.e. only one disc or more than two discs.
  • guiding vanes 40 which are fixed to a stator 42 of the gas turbine engine 10, are disposed between the stages of annular arrays of turbine blades 38. Between the exit of the combustion chamber 28 and the leading turbine blades 38 inlet guiding vanes 44 are provided and turn the flow of working gas onto the turbine blades 38. The combustion gas from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotate the shaft 22.
  • the guiding vanes 40, 44 serve to optimise the angle of the combustion or working gas on the turbine blades 38.
  • the turbine section 18 drives the compressor section 14.
  • the compressor section 14 comprises an axial series of vane stages 46 and rotor blade stages 48.
  • the rotor blade stages 48 comprise a rotor disc supporting an annular array of blades.
  • the compressor section 14 also comprises a casing 50 that surrounds the rotor stages and supports the vane stages 48.
  • the guide vane stages include an annular array of radially extending vanes that are mounted to the casing 50. The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point.
  • the casing 50 defines a radially outer surface 52 of the passage 56 of the compressor 14.
  • a radially inner surface 54 of the passage 56 is at least partly defined by a rotor drum 53 of the rotor which is partly defined by the annular array of blades 48.
  • the present invention is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine.
  • the present invention is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications.
  • upstream and downstream refer to the flow direction of the airflow and/or working gas flow through the engine unless otherwise stated.
  • the terms axial, radial and circumferential are made with reference to an axis 35 of the combustor.
  • Fig. 2 shows a combustor 100 for a gas turbine.
  • the combustor 100 has a centre axis 35 and comprises:
  • the swirler 103 is mounted on a peripheral wall 115 of the pre-combustion chamber 101, in such a way that the swirler 103 surrounds the pre-combustion chamber 101 in a circumferential direction with respect to the centre axis 35.
  • the swirler 103 comprises a bottom surface 104 which is orthogonal to the centre axis 35 and which forms a part of a slot 201 (see Fig. 3 ) through which, typically, an oxidant/fuel mixture flow F is injectable into the pre-combustion chamber 101.
  • the swirler 103 further comprises a cylindrical peripheral surface 119 having axis coincident with the combustor centre axis 35,
  • the swirler 103 comprises a plurality of slots 201 (twelve slots in the embodiment of figure 3 ).
  • Each slot 201 is formed by circumferentially spaced apart vanes 203 and the bottom surface 104.
  • Oxidant/fuel mixture which flows through the slots 201 is directed approximately tangentially with respect to the centre axis 35.
  • This orientation of the slots 201 induces a swirl movement, i.e. a movement according to a tangentially orientated direction around the centre axis 35, of the gasses inside the pre-combustion chamber 101.
  • Each slot 201 comprises a base fuel injector 107 which is arranged to the bottom surface 104 such that an air/fuel mixture is injectable into the slot 201 according to a main fuel injection direction which is orthogonal or inclined with respect to the bottom surface 104.
  • further side fuel injectors 202 may be provided for some of the slots 201 or for all of the slots 201 on the cylindrical peripheral surface 119 of the swirler 103.
  • two side fuel injectors 202 are provided for each of the slots 201.
  • the side fuel injectors 202 inject further fuel.
  • the further fuel may be mixed inside the slots 201 with the fuel which is injected by the base fuel injector 107 and with the oxidant.
  • Side fuel injectors 202 are in the form of holes, injecting further gaseous fuel.
  • atomizers or nozzles for liquid fuel injection are provided in the same slots 201, close to the trailing edges of the swirler vanes 203.
  • the combustor 100 further comprises a pilot burner 110, which comprises a burner face 111.
  • the burner face 111 is aligned or substantially parallel to the bottom surface 104.
  • the pilot burner 110 further comprises a cylindrical shroud 170, extending around the centre axis 35, for peripherally delimiting the pilot burner 110.
  • the pilot burner 110 comprises a plurality of pilot fuel injectors 112 which are arranged to the burner face 111 for injecting pilot fuel into the pre-combustion chamber 101.
  • twelve side pilot fuel injectors 112 regularly distributed 30 degrees apart circumferentially around the centre axis 35 are provided.
  • the pilot fuel injectors 112 are oriented substantially parallel to the centre axis 35.
  • the pilot fuel forms a separation layer and a flame front 105.
  • the circulation induced by the radial swirler 103 forms a central circular zone around the centre axis 35, inside of which the pilot fuel (i.e. the oxidant/fuel mixture) is burned. This central zone is called the reaction zone RZ.
  • the reaction zone RZ Around the central reaction zone RZ, the oxidant/fuel mixture is injected by the swirler 103.
  • the combustor 100 further includes a lip 150 extending from the pilot burner surface 111 in the pre-combustion chamber 101. In a circumferential direction, the lip 150 further extends around the centre axis 35. The lip 150 extends from a portion of the pilot burner surface 111 whose distance from the centre axis 35 of the pre-combustion chamber 101 is greater than the distance between the pilot fuel injectors 112 and the centre axis 35. With respect to the more internal portion of the pre-combustion chamber 101, identified as the portion around the centre axis 35, the lip 150 includes an internal surface 151 and an external surface 152.
  • the internal surface 151 is inclined towards the centre axis 35 and oriented towards the pilot fuel injectors 112 for intercepting at least part of the pilot fuel from the pilot fuel injectors 112.
  • the internal surface 151 is inclined of an inclination angle ⁇ comprised between 0 degrees and 90 degrees. More particularly, in the embodiments of Figs. 4 to 10 , the inclination angle ⁇ is comprised between 30 degrees and 60 degrees.
  • the external surface 152 is oriented towards the swirler 103 for intercepting at least part of the flow F coming from the swirler 103.
  • the lip 150 is integral with the pilot burner 110, being provided as an edge of the shroud 170, extending inside the pre-combustion 101. According to other embodiments of the present invention (not shown) the lip 150 is provided on pilot burner surface 111 or on the swirler 103.
  • the lip 150 further comprises a plurality of feed passages 155 provided between the internal surface 151 and the external surface 152, for connecting the internal surface 151 with the flow F coming from the swirler 103.
  • the feed passages 155 are regularly distributed around the centre axis 35.
  • the external surface 152 comprises a plurality of turbulators 160, 161, 162 for inducing turbulence in the flow F coming from the swirler 103.
  • the turbulators comprise a plurality of protrusions 160, 162 extending orthogonally from the external surface 152.
  • Some of the protrusions 160, 162 are constituted by a plurality of first protrusions 160, placed around the centre axis 35, at a same distance from the centre axis 35.
  • the first protrusions 160 have respective bases on the external surface 152, the bases having, for example, circular or rectangular shape.
  • the protrusions 160 are regularly distributed around the centre axis 35, at a fixed angular distance.
  • a further protrusion 162 is provided as a circular rim 162, concentric with the centre axis 35 of the pre-combustion chamber 101. With respect to the flow F coming from the swirler 103, the circular rim 162 is provided on the external surface 152, downstream of the first protrusions 160. According to other possible embodiments (not shown), the circular rim 162 is provided on the external surface 152, upstream of the first protrusions 160.
  • the turbulators comprise a plurality of channels 161 regularly distributed around the central axis Y.
  • Each channel 161 extends from the external surface 152 up the internal surface 151, in such a way that the channels 161 divide the lip 150 into a plurality of segments 158, each segment 158 being comprised between two consecutive channels 161.
  • the channels 161 are not completely extended from the external surface 152 up the internal surface 151, but are provided on the external surface 152 along a direction inclined of an inclination angle ⁇ with respect to the centre axis 35 and with a depth extending from the external surface 152 towards the internal surface 151.
  • any other array of the turbulators 160, 161, 162 may be arranged, each array being characterized by the type(s), number and distribution of the turbulators 160, 161, 162.
  • the internal surface 151 and the external surface 152 have a common trailing edge 156, at the end of the lip 150, where both the pilot fuel and the flow F separate from the lip 150.
  • the trailing edge 156 has a sharp circular profile around the centre axis 35.
  • the trailing edge 156 has a rounded profile in a section view (equivalent, for example to the view of Fig. 6 ) and a waved profile in a circumferential view, around the centre axis 35.
  • the trailing edge 156 is clipped, i.e. the lip 150 has, in a sectional plane including the centre axis 35, a trapezoidal shape, including a face 159, connecting the external surface 152 and the internal surface 151, at the end of the lip 150.
  • the lip 150 may be provided, in embodiments of the present invention, with an internal surface 151b having an aerofoil shape (dashed line of Fig. 6 ).
EP15202500.3A 2015-12-23 2015-12-23 Brennkammer für eine gasturbine Withdrawn EP3184898A1 (de)

Priority Applications (5)

Application Number Priority Date Filing Date Title
EP15202500.3A EP3184898A1 (de) 2015-12-23 2015-12-23 Brennkammer für eine gasturbine
CN201680075991.7A CN108431503A (zh) 2015-12-23 2016-12-09 用于燃气轮机的燃烧器
US15/781,713 US20180363904A1 (en) 2015-12-23 2016-12-09 Combustor for a gas turbine
PCT/EP2016/080488 WO2017108454A1 (en) 2015-12-23 2016-12-09 Combustor for a gas turbine
EP16812715.7A EP3394514A1 (de) 2015-12-23 2016-12-09 Brennkammer für eine gasturbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP15202500.3A EP3184898A1 (de) 2015-12-23 2015-12-23 Brennkammer für eine gasturbine

Publications (1)

Publication Number Publication Date
EP3184898A1 true EP3184898A1 (de) 2017-06-28

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Family Applications (2)

Application Number Title Priority Date Filing Date
EP15202500.3A Withdrawn EP3184898A1 (de) 2015-12-23 2015-12-23 Brennkammer für eine gasturbine
EP16812715.7A Withdrawn EP3394514A1 (de) 2015-12-23 2016-12-09 Brennkammer für eine gasturbine

Family Applications After (1)

Application Number Title Priority Date Filing Date
EP16812715.7A Withdrawn EP3394514A1 (de) 2015-12-23 2016-12-09 Brennkammer für eine gasturbine

Country Status (4)

Country Link
US (1) US20180363904A1 (de)
EP (2) EP3184898A1 (de)
CN (1) CN108431503A (de)
WO (1) WO2017108454A1 (de)

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP7207860B2 (ja) * 2018-04-09 2023-01-18 浜松ホトニクス株式会社 試料観察装置

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6151899A (en) * 1998-05-09 2000-11-28 Alstom Gas Turbines Limited Gas-turbine engine combustor
US20010027637A1 (en) * 1998-01-31 2001-10-11 Eric Roy Norster Gas-turbine engine combustion system
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
EP1835231A1 (de) * 2006-03-13 2007-09-19 Siemens Aktiengesellschaft Brenner für eine Turbinenbrennkammer und Verfahren zum Betrieb des Brenners

Family Cites Families (12)

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GB9023004D0 (en) * 1990-10-23 1990-12-05 Rolls Royce Plc A gas turbine engine combustion chamber and a method of operating a gas turbine engine combustion chamber
GB2328011A (en) * 1997-08-05 1999-02-10 Europ Gas Turbines Ltd Combustor for gas or liquid fuelled turbine
US7325402B2 (en) * 2004-08-04 2008-02-05 Siemens Power Generation, Inc. Pilot nozzle heat shield having connected tangs
EP1821035A1 (de) * 2006-02-15 2007-08-22 Siemens Aktiengesellschaft Gasturbinenbrenner und Verfahren zum Mischen von Brennstoff und Luft in einem Wirbelbereich eines Gasturbinenbrenners
EP1867925A1 (de) * 2006-06-12 2007-12-19 Siemens Aktiengesellschaft Brenner
GB2444737B (en) * 2006-12-13 2009-03-04 Siemens Ag Improvements in or relating to burners for a gas turbine engine
FR2911667B1 (fr) * 2007-01-23 2009-10-02 Snecma Sa Systeme d'injection de carburant a double injecteur.
DE102007043626A1 (de) * 2007-09-13 2009-03-19 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenmagerbrenner mit Kraftstoffdüse mit kontrollierter Kraftstoffinhomogenität
US8347630B2 (en) * 2008-09-03 2013-01-08 United Technologies Corp Air-blast fuel-injector with shield-cone upstream of fuel orifices
US9435537B2 (en) * 2010-11-30 2016-09-06 General Electric Company System and method for premixer wake and vortex filling for enhanced flame-holding resistance
US8931280B2 (en) * 2011-04-26 2015-01-13 General Electric Company Fully impingement cooled venturi with inbuilt resonator for reduced dynamics and better heat transfer capabilities
FR2996286B1 (fr) * 2012-09-28 2014-09-12 Snecma Dispositif d'injection pour une chambre de combustion de turbomachine

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
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
US20010027637A1 (en) * 1998-01-31 2001-10-11 Eric Roy Norster Gas-turbine engine combustion system
US6151899A (en) * 1998-05-09 2000-11-28 Alstom Gas Turbines Limited Gas-turbine engine combustor
EP1835231A1 (de) * 2006-03-13 2007-09-19 Siemens Aktiengesellschaft Brenner für eine Turbinenbrennkammer und Verfahren zum Betrieb des Brenners

Also Published As

Publication number Publication date
WO2017108454A1 (en) 2017-06-29
CN108431503A (zh) 2018-08-21
EP3394514A1 (de) 2018-10-31
US20180363904A1 (en) 2018-12-20

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