EP3159609A1 - Chambre de combustion pour turbine à gaz - Google Patents

Chambre de combustion pour turbine à gaz Download PDF

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Publication number
EP3159609A1
EP3159609A1 EP15190773.0A EP15190773A EP3159609A1 EP 3159609 A1 EP3159609 A1 EP 3159609A1 EP 15190773 A EP15190773 A EP 15190773A EP 3159609 A1 EP3159609 A1 EP 3159609A1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
combustor
swirler
mixing
centre axis
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Withdrawn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
EP15190773.0A
Other languages
German (de)
English (en)
Inventor
Nishant Parsania
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 EP15190773.0A priority Critical patent/EP3159609A1/fr
Priority to US15/766,054 priority patent/US20180299129A1/en
Priority to PCT/EP2016/073597 priority patent/WO2017067775A1/fr
Priority to EP16775712.9A priority patent/EP3365604A1/fr
Publication of EP3159609A1 publication Critical patent/EP3159609A1/fr
Withdrawn legal-status Critical Current

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    • 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
    • F23R3/283—Attaching or cooling of fuel injecting means including supports for fuel injectors, stems, or lances
    • 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
    • F23R3/06—Arrangement of apertures along the flame tube
    • 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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/46—Combustion chambers comprising an annular arrangement of several essentially tubular flame tubes within a common annular casing or within individual casings

Definitions

  • the present invention relates to a combustor for a gas turbine and to a method for operating a combustor for a gas turbine.
  • a combustor comprise a main combustion chamber and a pre-combustion chamber, upstream 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 stream is injected via the swirler into the combustor in a generally tangential direction with respect to the centre axis of the combustor.
  • a pilot fuel is further injected typically by a pilot burner, generally according a direction parallel to the centre axis of the combustor, wherein the pilot fuel is used for controlling the combustor flame in which the main fuel in burned.
  • 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 injected pilot fuel generates a predefined flame shape inside the combustor, in particular inside the pre-combustion chamber.
  • the injected main fuel stream and the pilot fuel stream may be a liquid fuel or gaseous fuel.
  • the combustion is achieved by a substantially non-combustible gas flow comprising an oxidant, for example air, first being mixed together with the fuel in the pilot burner.
  • a high fuel concentration in the mixture of the gas e.g. air and fuel inside the combustion chamber at the centre of the pilot burner close to the pilot burner surface, may occur due to a back-circulation of the injected gas and fuel. This has the effect of increasing temperatures at the pilot burner surface, hence increasing also NOx emissions.
  • the temperature at the centre of the pilot burner close to the pilot burner surface can be reduced according to the combustor described in WO 2013/120558 A1 , which includes a swirler design having a plurality of main fuel injectors inclined with respect to the centre axis of the combustor.
  • the latter solution cannot yet be considered optimal, as it is also observed that, independently from the inclination of the main fuel injectors, high fuel concentration pockets form in the outer periphery of the combustor pre-chamber. Like at the pilot burner surface, temperatures within these pockets increase, hence causing also NOx emissions to increase.
  • a combustor for a gas turbine comprises:
  • the swirler surrounds the pre-combustion chamber in a circumferential direction with respect to the centre axis.
  • the burner plenum surrounds the pre-combustion chamber in a circumferential direction with respect to the centre axis.
  • the swirler comprises a plurality of slots through which an oxidant/fuel mixture is injectable into the pre-combustion chamber, at least one slot comprising at least a fuel injector through which the mixture is injectable into the slot.
  • the peripheral wall of the pre-combustion chamber comprises a plurality of mixing holes downstream of the swirler for letting a portion of the oxidant gas in burner plenum to flow from the burner plenum to the pre-combustion chamber.
  • 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 combustor comprises:
  • the swirler surrounds the pre-combustion chamber in a circumferential direction with respect to the centre axis.
  • the burner plenum surrounds the pre-combustion chamber in a circumferential direction with respect to the centre axis.
  • the swirler comprises a plurality of slots through which an oxidant/fuel mixture is injectable into the pre-combustion chamber, at least one slot comprising at least a fuel injector through which the mixture is injectable into the slot. According to the method, a portion of the oxidant gas in burner plenum is let to flow from the burner plenum to the pre-combustion chamber, downstream of the swirler.
  • the mixing holes are inclined of a first mixing angle ⁇ with respect to centre axis of the combustor.
  • the orientation of the mixing holes may be towards the swirler or towards a combustion chamber of the combustor.
  • the value of the first mixing angle ⁇ may be chosen in such a way be comprised between 30° and 60°. More particularly, the value of the first mixing angle ⁇ may be approximately 45°.
  • the inclination of the mixing holes with respect to the centre axis of the combustor provides better mixing of the oxidant/fuel mixture in the pre-combustion chamber, thus lowering NOx emissions.
  • the mixing holes are inclined of a second mixing angle ⁇ with respect to a radial direction of the combustor, in a transversal plane orthogonal to the centre axis of the combustor.
  • the mixing holes are inclined towards the negative direction of a circumferential swirl flowing inside the pre-combustion chamber.
  • the value of the second mixing angle ⁇ may be chosen in such a way be comprised between 30° and 60°. More particularly, the value of the second mixing angle ⁇ may be approximately 45°.
  • the inclination of the mixing holes with respect to a radial direction in the transversal plane of the combustor provides better mixing of the oxidant/fuel mixture in the pre-combustion chamber, thus lowering NOx emissions.
  • the axis of the mixing holes lies on a plane which is inclined of a third mixing angle ⁇ with respect to centre axis Y of the pre-combustion chamber.
  • the value of the third mixing angle ⁇ may be chosen in such a way be comprised between 30° and 60°. More particularly, the value of the third mixing angle ⁇ may be approximately 45°.
  • the inclination of the mixing holes with respect to the three main symmetry direction of the combustor (one axial and two radial directions) provides better mixing of the oxidant/fuel mixture in the pre-combustion chamber, thus lowering NOx emissions.
  • the combustor further comprises at least a base fuel injector arranged to a bottom surface of the swirler, the base fuel injector being inclined of an angle comprised between 30° and 60° with respect to the centre axis of the combustor.
  • the base fuel injector being inclined of an angle comprised between 30° and 60° with respect to the centre axis of the combustor.
  • 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 combustor 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 chamber 28.
  • This exemplary gas turbine engine 10 has a cannular combustor section arrangement 16, 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.
  • Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operations conditions.
  • 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 the rotational axis 20 of the engine.
  • Fig. 2 shows a combustor 100 for a gas turbine.
  • the combustor 100 has a centre axis Y and comprises:
  • the pre-combustion chamber 101, the swirler 103 and the combustion chamber 109 are all axially symmetric around the centre axis Y. With respect to the centre axis Y, the pre-combustion chamber 101 has a smaller diameter than the combustion chamber 109.
  • the pre-combustion chamber 101 and the combustion chamber 109 are adjacent to one another along the centre axis Y are in fluid communication through an exit plane 209 of pre-combustion chamber 101, downstream of which the combustion chamber 109 extends up to the transition duct 17.
  • the combustion chamber 109 is conventional and therefore not described in further detail.
  • 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 Y.
  • the swirler 103 comprises a bottom surface 104 which is orthogonal to the centre axis Y and which forms a part of a slot 201 (see Fig. 3 ) through which typically an oxidant/fuel mixture is injectable into the pre-combustion chamber 101.
  • the swirler 103 further comprises a cylindrical peripheral surface 119 having axis coincident with the centre axis Y.
  • 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 Y.
  • This orientation of the slots 201 induces a swirl movement, i.e. a movement according to a tangentially orientated direction W around the centre axis Y (see Fig. 6 ), 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 to the bottom surface 104.
  • the base fuel injector 107 is arranged to the bottom surface 104 such that the main fuel injection direction inclined of an angle between 0° and 90° or, more particularly between 30° and 60° with respect to the bottom surface 104 or the centre angle Y.
  • the base fuel injector 107 is arranged to the bottom surface 104 such that the main fuel injection direction inclined of approximately 45° with respect to the bottom surface 104 and the centre angle Y.
  • 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 locations.
  • a pilot burner 110 which comprises a burner face 111 is mounted immediately upstream to the swirler 103 and to the pre-combustion chamber 101.
  • the burner face 111 is aligned or substantially parallel to the bottom surface 104.
  • a pilot burner 110 comprises a plurality pilot fuel injector 112 which are arranged to the burner face 111 for injecting pilot fuel into the pre-combustion chamber 101.
  • twelve side pilot fuel injector 112 regularly distributed 30 degree apart circumferentially around the centre axis Y are provided.
  • Pilot fuel is injected through the pilot fuel injectors 112 basically along the axial direction with respect to the centre axis Y.
  • the pilot fuel forms a separation layer and a flame front 105.
  • the pilot fuel injectors 112 may be located along a circumferential direction to the pilot burner face 111 such that the injected pilot fuel forms a central circular zone inside of which the fuel (i.e. the oxidant/fuel mixture) is burned. This central zone may be called the recirculation zone RZ.
  • the recirculation zone RZ i.e. between the peripheral wall 115 of the pre-combustion chamber 101 and the separation layer generated by the pilot fuel, the oxidant/fuel mixture is injected by the swirler 103.
  • the fuel is injected into the slots 201 of the swirler 103 by the base fuel injectors 107 and the side fuel injectors 202 and then the fuel enters the pre-combustion chamber 101, where it is guided by the pilot fuel along the axial direction of the centre axis Y.
  • the pilot fuel stream is weakened and the pre-combustion products of the oxidant/fuel mixture flows abruptly back towards the pilot burner face 111. This causes a hot spot to be located near the burner face 111 in the central recirculation zone RZ due to the backflow of the ignitioned oxidant /fuel mixture.
  • the base fuel injectors 107 are inclined, i.e. not perpendicular to the bottom surface 104, the temperature and extension of the hot spot can be significantly reduced, as known from WO 2013/120558 A1 .
  • a plurality of mixing holes 120 provided on the peripheral wall 115 of the pre-combustion chamber 101, downstream of the swirler 103.
  • twelve mixing holes 120 regularly distributed around the centre axis Y are provided.
  • the mixing holes 120 let a portion of the oxidant gas in burner plenum 26 to flow directly from the burner plenum 26 to the pre-combustion chamber 101.
  • the air from the burner plenum 26 to the pre-combustion chamber 101 improves mixing of the air/fuel mixture in the pre-combustion chamber 101, in particular fuel from side fuel injectors 202.
  • the mixing of the air/fuel mixture can be furtherly improved by controlling the velocity of the pilot fuel stream to be sufficient high, in order to prevent backflow of the ignitioned oxidant /fuel mixture.
  • the mixing holes 120 are inclined of a first mixing angle ⁇ with respect to longitudinal axis Y, the value of the first mixing angle ⁇ being lower than 90°.
  • the value of the first mixing angle ⁇ is particularly comprised between 30° and 60°. Even more particularly the value of the first mixing angle ⁇ may be approximately 45°.
  • value of first mixing angle ⁇ may be positive or negative, i.e. the mixing holes 120 may inclined towards the swirler 103 ( Fig. 2 ) or towards the combustion chamber 109 ( Fig. 4 ).
  • the mixing holes 120 are inclined of a second mixing angle ⁇ with respect to a radial direction Z of the combustor 100, orthogonal to the centre axis Y.
  • the value of the second mixing angle ⁇ is lower than 90°.
  • the value of the second mixing angle ⁇ is particularly comprised between 30° and 60°. Even more particularly the value of the second mixing angle ⁇ may be approximately 45°.
  • the mixing holes 120 are inclined in the transversal plane XZ towards the negative direction of a circumferential swirl flowing inside the pre-combustion chamber 101, i.e. if the direction W of the swirl inside the pre-combustion chamber 101 is directed clockwards, the mixing holes 120 are directed in order to insert air in the pre-combustion chamber 101 according to a counter-clockward direction.
  • the mixing holes 120 are inclined of a third mixing angle ⁇ with respect to centre axis Y of the pre-combustion chamber 101, the value of the third mixing angle ⁇ being lower than 90°.
  • the value of the third mixing angle ⁇ is particularly comprised between 30° and 60°. Even more particularly the value of the third mixing angle ⁇ may be approximately 45°.
  • the orientation of the axis Z1 of a mixing hole 120 is shown with reference to the main orthogonal axis X, Y and Z of the combustor 100.
  • the mixing angles ⁇ , ⁇ and ⁇ of the projections of the axis Z1 on the three mutually orthogonal planes YZ, XZ and XY, respectively, are also shown.
  • results which can be obtained with the present invention are shown.
  • standard deviation 301 of the mixture fraction in five consecutive position 101a-e along the combustor pre-chamber 101, the first position 101a being adjacent to the swirler 103 and the last position 101e being coincident with the exit plane 209 are shown.
  • the standard deviation curve 301 is compared to a second and a third standard deviation curves 302, 303 in two respective known combustors, not provided with mixing holes 120.
  • the second standard deviation 302 is obtained in a combustor with orthogonal base fuel injector 107.
  • the third standard deviation 303 is obtained in a combustor with base fuel injector 107 inclined of 45° with respect to the centre axis Y.
  • the standard deviation 302 is higher than the standard deviation 303, while at the last position 101e the standard deviation 302 is lower than the standard deviation 303.
  • An ideal curve 301 is obtainable with the combustor according to the present invention.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
EP15190773.0A 2015-10-21 2015-10-21 Chambre de combustion pour turbine à gaz Withdrawn EP3159609A1 (fr)

Priority Applications (4)

Application Number Priority Date Filing Date Title
EP15190773.0A EP3159609A1 (fr) 2015-10-21 2015-10-21 Chambre de combustion pour turbine à gaz
US15/766,054 US20180299129A1 (en) 2015-10-21 2016-10-04 Combustor for a gas turbine
PCT/EP2016/073597 WO2017067775A1 (fr) 2015-10-21 2016-10-04 Chambre de combustion pour turbine à gaz
EP16775712.9A EP3365604A1 (fr) 2015-10-21 2016-10-04 Chambre de combustion pour turbine à gaz

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP15190773.0A EP3159609A1 (fr) 2015-10-21 2015-10-21 Chambre de combustion pour turbine à gaz

Publications (1)

Publication Number Publication Date
EP3159609A1 true EP3159609A1 (fr) 2017-04-26

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Application Number Title Priority Date Filing Date
EP15190773.0A Withdrawn EP3159609A1 (fr) 2015-10-21 2015-10-21 Chambre de combustion pour turbine à gaz
EP16775712.9A Withdrawn EP3365604A1 (fr) 2015-10-21 2016-10-04 Chambre de combustion pour turbine à gaz

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Application Number Title Priority Date Filing Date
EP16775712.9A Withdrawn EP3365604A1 (fr) 2015-10-21 2016-10-04 Chambre de combustion pour turbine à gaz

Country Status (3)

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US (1) US20180299129A1 (fr)
EP (2) EP3159609A1 (fr)
WO (1) WO2017067775A1 (fr)

Cited By (1)

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CN111288490A (zh) * 2020-03-23 2020-06-16 上海电力大学 一种分散凸台处高温回流区的燃烧室装置

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20220364511A1 (en) * 2021-05-11 2022-11-17 General Electric Company Integral fuel-nozzle and mixer with angled jet-in-crossflow fuel injection

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WO2007060216A1 (fr) * 2005-11-26 2007-05-31 Siemens Aktiengesellschaft Appareil de combustion
EP1975506A1 (fr) * 2007-03-30 2008-10-01 Siemens Aktiengesellschaft Pré-chambre de combustion
EP2405200A1 (fr) * 2010-07-05 2012-01-11 Siemens Aktiengesellschaft Appareil de combustion et moteur de turbine à gaz
WO2013120558A1 (fr) 2012-02-15 2013-08-22 Siemens Aktiengesellschaft Injection inclinée de carburant dans une fente de coupelle de turbulence

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DE50313028D1 (de) * 2002-05-02 2010-10-14 Alstom Technology Ltd Katalytischer Brenner
US6775983B2 (en) * 2002-05-21 2004-08-17 General Electric Company Flow control device for a combustor
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US8640463B2 (en) * 2011-06-28 2014-02-04 United Technologies Corporation Swirler for gas turbine engine fuel injector
FR2979416B1 (fr) * 2011-08-26 2013-09-20 Turbomeca Paroi de chambre de combustion
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Publication number Priority date Publication date Assignee Title
WO2007060216A1 (fr) * 2005-11-26 2007-05-31 Siemens Aktiengesellschaft Appareil de combustion
EP1975506A1 (fr) * 2007-03-30 2008-10-01 Siemens Aktiengesellschaft Pré-chambre de combustion
EP2405200A1 (fr) * 2010-07-05 2012-01-11 Siemens Aktiengesellschaft Appareil de combustion et moteur de turbine à gaz
WO2013120558A1 (fr) 2012-02-15 2013-08-22 Siemens Aktiengesellschaft Injection inclinée de carburant dans une fente de coupelle de turbulence

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN111288490A (zh) * 2020-03-23 2020-06-16 上海电力大学 一种分散凸台处高温回流区的燃烧室装置

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Publication number Publication date
WO2017067775A1 (fr) 2017-04-27
US20180299129A1 (en) 2018-10-18
EP3365604A1 (fr) 2018-08-29

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