EP2808610A1 - Chambre de combustion d'une turbine à gaz avec injection tangentielle comme injection pauvre tardive - Google Patents

Chambre de combustion d'une turbine à gaz avec injection tangentielle comme injection pauvre tardive Download PDF

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Publication number
EP2808610A1
EP2808610A1 EP13170045.2A EP13170045A EP2808610A1 EP 2808610 A1 EP2808610 A1 EP 2808610A1 EP 13170045 A EP13170045 A EP 13170045A EP 2808610 A1 EP2808610 A1 EP 2808610A1
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EP
European Patent Office
Prior art keywords
combustion chamber
injector
combustor
fuel
main flow
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Withdrawn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
EP13170045.2A
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German (de)
English (en)
Inventor
Christian Dr. Beck
Olga Deiss
Werner Krebs
Bernhard Dr. Wegner
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
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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 EP13170045.2A priority Critical patent/EP2808610A1/fr
Publication of EP2808610A1 publication Critical patent/EP2808610A1/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Withdrawn legal-status Critical Current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
    • F23R3/286—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
    • F23R3/346—Feeding into different combustion zones for staged combustion
    • 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 combustion chamber, in particular a tube combustion chamber, a gas turbine and a method for operating a combustion chamber and a gas turbine.
  • Modern gas turbines should meet the requirements in terms of pollutant emissions and environmental friendliness in a wide operating range. The fulfillment of these requirements depends essentially on the combustion system used in the gas turbine.
  • NOx nitrogen oxides
  • To reduce emissions of nitrogen oxides (NOx) lean premix is used.
  • high turbine inlet temperatures are sought to achieve a high efficiency, which are associated with high flame temperatures.
  • the aforementioned premixed flames are susceptible to thermoacoustic instabilities due to the high thermal power density and the NOx emissions increase exponentially with increasing flame temperature.
  • axial staging consists of a conventional burner that fires a primary combustion zone. This primary zone can in turn be internally graded like conventional burners and covers the load range up to today's firing temperatures. Downstream of the primary zone is followed by a secondary combustion zone. In this additional fuel is injected through an axially offset from the primary zone stage. This is then burned in a diffusion-like regime.
  • the fuel may be diluted with inert components (steam, nitrogen, carbon dioxide) to greatly lower the stoichiometric combustion temperature, thereby suppressing NOx formation.
  • inert components steam, nitrogen, carbon dioxide
  • the US 2011/0067402 A1 discloses a gas turbine with a combustion chamber having a dual stage combustion concept.
  • the combustor includes a combustor head end having a burner assembly, a combustor exit and a combustor wall, the combustor wall extending from the combustor head end to the combustor exit, and a primary zone and a secondary zone.
  • the secondary zone is located in the main flow direction of the hot gas downstream of the primary zone.
  • injectors opening into the secondary zone are arranged, which form a second axial stage of the combustion system.
  • a second object is to provide a corresponding gas turbine.
  • a third object of the present invention is to provide an advantageous method for operating a combustion chamber or a gas turbine comprising a combustion chamber, which makes it possible to reduce emissions of nitrogen oxides and / or to reduce CO emissions.
  • the first object is achieved by a combustion chamber according to claim 1.
  • the second object is achieved by a gas turbine according to claim 12.
  • the third object is achieved by a method according to claim 13.
  • the dependent claims contain further advantageous embodiments of the invention.
  • the combustion chamber according to the invention comprises a longitudinal axis, a combustion chamber head end and a combustion chamber exit.
  • the combustor further includes a combustor wall extending from the combustor head end to the combustor exit.
  • the combustion chamber according to the invention comprises a primary zone and a secondary zone.
  • the secondary zone is arranged in the main flow direction of the hot gas downstream of the primary zone.
  • the combustion chamber comprises at least one injector for introducing a fuel-air mixture into the secondary zone.
  • the injector is disposed in the secondary zone on the combustion chamber wall.
  • the injector also includes an outlet opening into the secondary zone or an injection opening with an inflow direction.
  • the inflow direction has a component in the circumferential direction of the combustion chamber.
  • the combustor wall includes an outer surface and at least one injector is at least partially helically disposed with respect to the longitudinal axis of the combustor along the outer surface.
  • the term "along the surface” also includes a first course of the injector spaced from the surface, so that the injector is in direct contact with the combustion chamber wall only in a section passing through the combustion chamber wall.
  • helical is also understood to be helical, for example helices of decreasing diameter, wherein the diameter of the helix may be essentially circular or elliptical.
  • the fact that the injector is at least partially helically disposed with respect to the longitudinal axis of the combustion chamber along the outer surface also includes such injectors that at least partially follow a portion of such a spiral or helical path.
  • a number of corresponding injectors are arranged on the combustion chamber wall in the region of the secondary zone.
  • the combustion chamber is preferably a tube combustion chamber. At least one burner may be arranged at the end of the combustion chamber.
  • the primary zone is determined by the area in which the fuel supplied via the burner is primarily burned within the combustion chamber.
  • the secondary zone is characterized by the fact that in it the hot gas generated in the primary zone is further burned out as completely as possible.
  • the secondary zone can in principle be arranged at any desired position between the primary zone and the combustion chamber exit.
  • the airborne axial stage itself has several advantages. By premixing fuel and air outside the combustion chamber as with conventional burner technology, the resulting peak temperatures and thus NOx emissions can be reduced. Lower residence times in the secondary zone and turbine entry continue to result in lower overall NOx emissions. In addition, no additional media are needed, but an operation takes place only with the originating from the compressor outlet air, which are treated with fuel in the axial stage to a mixture. Therefore, the resulting system is robust and stable available.
  • the Airborne axial stage therefore equally serves to expand the operating range of the combustion system to lower and higher loads.
  • the combustion chamber wall may comprise an outer surface. At least one injector may be disposed at least partially along the outer surface. This has the advantage that good utilization of the available installation space around the combustion chamber results in a large premix length in the injectors despite the compact design. Furthermore, emissions are reduced and thermoacoustic tuning is possible by adapting deadtime elements in flame transfer functions.
  • the present invention also has the following special advantages: Due to the helical arrangement, a long mixing length can be achieved in the injectors or scoops despite their compact design.
  • the swirl generation provides for the generation of additional gradients and shear layers and thus for a better mixing with the main flow.
  • a smoother turbine entry profile reduces emissions.
  • a simple and inexpensive construction of the guide vanes of the first turbine stage (TLe 1) is made possible.
  • the present invention opens up great potential for saving cooling air and possibly saving potential by dispensing with the vanes of the first turbine stage (TLe 1).
  • the at least one injector may include an output having a central axis.
  • the central axis can include an angle ⁇ 1 between 0 ° and 180 ° with the main flow direction in the combustion chamber at the position of the respective injector.
  • the fuel-air mixture can be introduced both in the opposite direction and in the main flow direction in the secondary zone.
  • the angle ⁇ 1 between the center axis of the injector outlet and the main flow direction may be greater than 45 ° and less than 90 °, preferably less than 70 °. As a result, a good mixing with the main flow while generating a twist is achieved.
  • the center axis of the injector outlet determines the inflow direction of the injected fuel-air mixture into the combustion chamber.
  • the main flow direction of the hot gas in the combustion chamber is determined in particular by the burner axis and the geometry of the combustion chamber.
  • the main flow direction may be in the form of a curved curve extending from the burner to the combustion chamber exit.
  • the at least one injector may comprise an output with a central axis, which may include an angle ⁇ 2 between 0 ° and 180 ° with the longitudinal axis of the combustion chamber.
  • the angle ⁇ 2 is between 0 ° and 90 °, preferably between 20 ° and 70 °, which corresponds to an inflow in the main flow direction.
  • the angle ⁇ 2 is greater than 45 ° and less than 90 ° or less than 70 ° in order to achieve a favorable mixing of the introduced fuel-air mixture with the main flow with simultaneous swirl generation.
  • the output of the injector may be arranged with respect to the main flow direction such that a line radial to the main flow direction intersects the center axis of the injector in the region of its output at an angle ⁇ 1 .
  • the output of the injector can be arranged with respect to the longitudinal axis of the combustion chamber such that a radial line to the longitudinal axis intersects the center axis of the injector in the region of its output at an angle ⁇ 2 , the angles ⁇ 1 and ⁇ 2 respectively in the region between 0 ° and 90 °, advantageously between 20 ° and 70 ° or between 45 ° and 90 ° or between 45 ° and 70 °.
  • the combustion chamber may have a radial direction with respect to the main flow direction.
  • the inflow direction and / or the center axis of the injector can be perpendicular to the main flow direction an angle ⁇ 1 with the radial direction with respect to the main flow direction between 0 ° and 90 °, advantageously between 20 ° and 70 ° or between 45 ° and 90 ° or between 45 ° and Include 70 °.
  • This means that the inflow direction and / or the center axis of the injector are arranged perpendicular to the main flow direction or has a tangential component with respect to the main flow direction.
  • the tangential component with respect to the main flow direction is described by the angle ⁇ 1 . According to the invention, this tangential component has a component pointing in the circumferential direction.
  • the combustion chamber may comprise a radial direction with respect to the longitudinal axis.
  • the inflow direction and / or the center axis of the injector can be perpendicular to the burner axis an angle ⁇ 2 with the radial direction with respect to the longitudinal axis between 0 ° and 90 °, advantageously between 20 ° and 70 °, preferably between 45 ° and 90 ° or between 45 ° and 70 °.
  • This means that the inflow direction and / or the center axis of the injector has a component which is perpendicular to the longitudinal axis and described by the angle ⁇ 2 , which at the same time describes a tangential component with respect to the longitudinal axis. According to the invention, this tangential component has a component pointing in the circumferential direction.
  • the combustion chamber may include a fuel distributor.
  • the fuel distributor may be connected to at least one nozzle which is arranged in the at least one injector.
  • the fuel distributor may be connected to a plurality of obliquely arranged nozzles and fuel distribute to these.
  • the fuel is mixed with air and then injected into the combustion chamber with a component pointing in the circumferential direction of the combustion chamber.
  • the fuel distributor is an annular fuel distributor, which is arranged, for example, annularly around the combustion chamber wall.
  • the fuel distributor may be disposed on the outer surface of the combustion chamber wall.
  • the fuel distributor may be disposed in the axial direction to (downstream of) the at least one injector and upstream of (upstream of) the combustor exit along the outer combustor surface.
  • the at least one injector may be disposed in the axial direction to (downstream of) the fuel rail and upstream of (upstream of) the combustor exit along the outer combustor surface.
  • the injectors may comprise an area arranged outside the combustion chamber, which extends at least partially in or counter to the main flow direction along the outer combustion chamber surface.
  • the injector may comprise a flow channel for supplying air.
  • a fuel nozzle can be arranged in the interior of the injector (scoop). The fuel nozzle may be at least partially surrounded by the flow channel for supplying air.
  • a number of injectors are arranged circumferentially on the combustion chamber wall.
  • the number of injectors can be distributed uniformly along the circumference of the combustion chamber wall.
  • a liner region may adjoin the primary zone in the main flow direction, followed by a transition region to the combustion chamber outlet.
  • at least one injector can be arranged in the liner area.
  • the gas turbine according to the invention comprises a combustion chamber described above. It has the same characteristics and advantages as the combustion chamber described above.
  • the inventive method for operating a combustion chamber previously described or for operating a gas turbine described above is characterized in that a fuel-air mixture is introduced through the at least one injector in the secondary zone of the combustion chamber so that the inflow direction is a component in the circumferential direction of the combustion chamber having.
  • the method according to the invention has the same advantages as the combustion chamber according to the invention described above. In particular, with the aid of the introduced into the secondary zone fuel-air mixture improved mixing of the main flow and a reduction of emissions by a more uniform turbine inlet profile can be achieved. Incidentally, reference is made to the advantages mentioned in connection with the combustion chamber according to the invention.
  • FIG. 1 shows by way of example a gas turbine 100 in a longitudinal partial section.
  • the gas turbine 100 has inside a rotatably mounted about a rotation axis 102 rotor 103 with a shaft 101, which is also referred to as a turbine runner.
  • an intake housing 104 a compressor 105, for example a toroidal combustion chamber 110, in particular annular combustion chamber, with a plurality Coaxially arranged burners 107, a turbine 108 and the exhaust housing 109th
  • a compressor 105 for example a toroidal combustion chamber 110, in particular annular combustion chamber, with a plurality Coaxially arranged burners 107, a turbine 108 and the exhaust housing 109th
  • the annular combustion chamber 110 communicates with an annular annular hot gas channel 111, for example.
  • annular annular hot gas channel 111 for example.
  • turbine stages 112 connected in series form the turbine 108.
  • Each turbine stage 112 is formed, for example, from two blade rings.
  • the hot gas channel 111 of a row of guide vanes 115 is followed by a row 125 formed of rotor blades 120.
  • the guide vanes 130 are fastened to an inner housing 138 of a stator 143, whereas the moving blades 120 of a row 125 are attached to the rotor 103 by means of a turbine disk 133, for example.
  • air 105 is sucked in and compressed by the compressor 105 through the intake housing 104.
  • the compressed air provided at the turbine-side end of the compressor 105 is supplied to the burners 107 where it is mixed with a fuel.
  • the mixture is then burned to form the working fluid 113 in the combustion chamber 110.
  • the working medium 113 flows along the hot gas channel 111 past the guide vanes 130 and the rotor blades 120.
  • the working medium 113 expands in a pulse-transmitting manner so that the rotor blades 120 drive the rotor 103 and drive the machine coupled to it.
  • the components exposed to the hot working medium 113 are subject to thermal loads during operation of the gas turbine 100.
  • the vanes 130 and blades 120 of the As seen in the flow direction of the working medium 113 first turbine stage 112 are next to the annular combustion chamber 110 lining heat shield elements most thermally stressed.
  • iron, nickel or cobalt-based superalloys are used as the material for the components, in particular for the turbine blades 120, 130 and components of the combustion chamber 110.
  • the vane 130 has a guide vane foot (not shown here) facing the inner casing 138 of the turbine 108 and a vane head opposite the vane root.
  • the vane head faces the rotor 103 and fixed to a mounting ring 140 of the stator 143.
  • the FIG. 2 schematically shows a combustion chamber 110 of a gas turbine.
  • the combustion chamber 110 is designed, for example, as a so-called annular combustion chamber, in which a plurality of burners 107 arranged around a rotation axis 102 in the circumferential direction open into a common combustion chamber space 154, which generate flames 156.
  • the combustion chamber 110 is configured in its entirety as an annular structure, which is positioned around the axis of rotation 102 around.
  • the combustion chamber 110 is designed for a comparatively high temperature of the working medium M of about 1000 ° C to 1600 ° C.
  • the combustion chamber wall 153 is on their the working medium M facing side provided with an inner lining formed of heat shield elements 155.
  • the heat shield elements 155 are then, for example, hollow and possibly still have cooling holes (not shown) which open into the combustion chamber space 154.
  • FIG. 3 schematically shows a part of a combustion chamber according to the invention in a partially perspective and partially sectioned view.
  • the combustion chamber comprises a combustion chamber wall 1 and a combustion chamber outlet 6.
  • the main flow direction of the hot gas in the combustion chamber during operation of the combustion chamber is indicated by an arrow 3.
  • the combustion chamber further comprises a primary zone 4, in which the fuel introduced from the burner into the combustion chamber is burned.
  • a secondary zone 5 adjoins the primary zone in the direction of flow 3.
  • the hot gas from the primary zone 4 is further burned off. This is done by additionally introducing a fuel-air mixture 14 in the secondary zone 5 by means of injectors. 8
  • the injectors 8 comprise an air supply 13 and an outlet 9 opening into the combustion chamber. Furthermore, a fuel nozzle 10 is arranged in the interior of each injector 8. The fuel nozzle 10 is connected to a fuel distributor 11, preferably an annular fuel distributor 11. With the help of the fuel nozzle 10, fuel is injected into the interior of the injector 8 and in this way generates a fuel-air mixture in the interior of the injector 8. The fuel-air mixture thus produced is then through the injector or the injection port 9 into the combustion chamber injected in the region of the secondary zone 5. According to the invention with at least one extending in the circumferential direction of the combustion chamber component.
  • a liner region 7 and a transition region 25 which in the FIG. 3 are each designed as separate components.
  • a liner region 7 and a transition region 25 which in the FIG. 3 are each designed as separate components.
  • Between the primary zone 4 and the liner area 7 at least one sealing ring 12 is arranged. Furthermore, at least one sealing ring 12 is also arranged between the liner region 7 and the transitional component 25.
  • the injectors 8 are connected to the liner area 7.
  • the injector or injection ports 9 open in the region of the liner region 7 in the secondary zone 5 of the combustion chamber.
  • FIG. 4 shows a section of the already in the FIG. 3 partially shown combustion chamber in perspective and sectional view.
  • a fuel supply 15 is shown, which supplies the fuel distributor 11 with fuel.
  • the FIG. 5 schematically shows the arrangement of the fuel nozzle within the injector in a perspective and partially sectioned view.
  • the fuel nozzle 10 is connected to the annular fuel distributor 11 and is supplied with fuel from this.
  • the injector includes an inlet 33 having an air inlet 9.
  • the tip of the fuel nozzle 10 includes fuel injection ports 36.
  • the fuel nozzle 10 is partially disposed within the injector 8.
  • the tip of the fuel nozzle 10 including the fuel injection openings 36 in the region of the central axis 2 of the injector 8 within the injector 8 is arranged. Through the fuel injection openings 36, fuel in the direction 37 perpendicular to the central axis 2 of the injector 8 in the region of the air supply 9 in the injector. 8 injected. In this case, a fuel-air mixture is generated, which is passed through the injector 8 in the flow direction 35 to the combustion chamber.
  • FIG. 6 schematically shows a part of the combustion chamber according to the invention in the region of the secondary zone in a perspective view.
  • the liner region 7 and the transition region 25 is integrally designed as a coherent transition element.
  • the Indian FIG. 6 The area of the combustion chamber shown comprises a central axis 34, which runs parallel to the main flow direction 3 of the hot gas mixture flowing through the combustion chamber, and an outer surface 32. On the outer surface 32, an annular fuel distributor 11 is arranged. This is connected to a number of fuel nozzles 10, as in connection with the FIGS. 3 to 5 described, feed fuel to a number of injectors 8.
  • the injectors 8 are arranged uniformly along the circumference of the combustion chamber on the outer surface 32.
  • the injectors 8 and the fuel nozzles 10 with respect to the central axis 34 and the main flow direction 3 are arranged spirally along the outer surface 32.
  • an introduction of the fuel-air mixture generated in the injectors is achieved with a pointing in the circumferential direction of the combustion chamber component of the inflow into the combustion chamber. In this way, a swirl of the combustion chamber flowing through the hot gas mixture is generated within the combustion chamber.
  • FIG. 7 schematically shows a plan view of a section through the combustion chamber according to the invention in the region of the injectors 8 in a partially perspective view.
  • the respective inflow direction of the fuel-air mixture from the injectors 8 into the combustion chamber is identified by the reference numerals 23.
  • FIG. 8 schematically shows the center axis 2 of the injector 8 and the inflow direction 23 with respect to the main flow direction 3 of the hot gas in the combustion chamber or with respect to the longitudinal axis 34 of the combustion chamber.
  • the main flow direction 3 is shown schematically in the form of an axis.
  • the center axis of the injector 2 or the inflow direction 23 of the fuel-air mixture flowing from the injector outlet 9 into the combustion chamber encloses an angle ⁇ 1 with the main flow direction 3 or an angle ⁇ 2 with the longitudinal axis 34 of the combustion chamber.
  • the angles ⁇ 1 can basically assume values between 0 ° and 180 °, for example between 20 ° and 70 °, preferably between 45 ° and 70 °.
  • FIG. 9 schematically shows the tangential component of the inflow or the component of the inflow direction in the circumferential direction of the combustion chamber.
  • This is in the FIG. 9 schematically a section through a portion of the combustion chamber perpendicular to the main flow direction 3 or alternatively shown perpendicular to the longitudinal axis 34.
  • a radial line to the main flow direction 3, which intersects the central axis 2 of the injector 8 in the region of its output 9 is indicated by the reference numeral 19.
  • a radial line to the longitudinal axis 34 of the combustion chamber, which intersects the central axis 2 of the injector 8 in the region of its output 9, is also identified by the reference numeral 19.
  • the radial line 19 has a right angle to the main flow direction 3 and / or a right angle to the longitudinal axis 34.
  • the respective radial direction or radial line 19 encloses an angle ⁇ 1 with the inflow direction 23 or with the center axis of the injector 2, if the radial line 19 relates to the main flow direction 3. If the radial line 19 refers to the longitudinal axis 34, so includes the radial line 19 with the inflow direction 23 and the central axis 2 of the injector 8 an angle ⁇ 2 a.
  • the angles ⁇ 1 and ⁇ 2 can be between 0 ° and 90 °, preferably between 20 ° and 70 °, for example between 45 ° and 70 °.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
EP13170045.2A 2013-05-31 2013-05-31 Chambre de combustion d'une turbine à gaz avec injection tangentielle comme injection pauvre tardive Withdrawn EP2808610A1 (fr)

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EP13170045.2A EP2808610A1 (fr) 2013-05-31 2013-05-31 Chambre de combustion d'une turbine à gaz avec injection tangentielle comme injection pauvre tardive

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Application Number Priority Date Filing Date Title
EP13170045.2A EP2808610A1 (fr) 2013-05-31 2013-05-31 Chambre de combustion d'une turbine à gaz avec injection tangentielle comme injection pauvre tardive

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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2017018982A1 (fr) * 2015-07-24 2017-02-02 Siemens Aktiengesellschaft Conduit de transition de turbine à gaz à injection pauvre tardive présentant un temps réduit de séjour de combustion
CN115234943A (zh) * 2022-06-30 2022-10-25 北京航空航天大学 中心分级与轴向分级耦合式燃烧室
EP4187153A1 (fr) * 2021-11-24 2023-05-31 Raytheon Technologies Corporation Chambre de combustion de moteur à turbine à gaz avec conduit(s) de carburant intégré(s)

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JPS61105029A (ja) * 1984-10-29 1986-05-23 Kawasaki Heavy Ind Ltd 予混合型ガスタ−ビン燃焼器
DE4232383A1 (de) 1992-09-26 1994-03-31 Asea Brown Boveri Gasturbogruppe
US6047550A (en) 1996-05-02 2000-04-11 General Electric Co. Premixing dry low NOx emissions combustor with lean direct injection of gas fuel
US6418725B1 (en) 1994-02-24 2002-07-16 Kabushiki Kaisha Toshiba Gas turbine staged control method
US6868676B1 (en) 2002-12-20 2005-03-22 General Electric Company Turbine containing system and an injector therefor
US20060156735A1 (en) * 2005-01-15 2006-07-20 Siemens Westinghouse Power Corporation Gas turbine combustor
DE102006053679A1 (de) 2005-11-15 2007-05-24 General Electric Co. Niedrigemissionsbrennkammer und Betriebsverfahren
US20090084082A1 (en) 2007-09-14 2009-04-02 Siemens Power Generation, Inc. Apparatus and Method for Controlling the Secondary Injection of Fuel
US20110067402A1 (en) 2009-09-24 2011-03-24 Wiebe David J Fuel Nozzle Assembly for Use in a Combustor of a Gas Turbine Engine

Patent Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS61105029A (ja) * 1984-10-29 1986-05-23 Kawasaki Heavy Ind Ltd 予混合型ガスタ−ビン燃焼器
DE4232383A1 (de) 1992-09-26 1994-03-31 Asea Brown Boveri Gasturbogruppe
US6418725B1 (en) 1994-02-24 2002-07-16 Kabushiki Kaisha Toshiba Gas turbine staged control method
US6047550A (en) 1996-05-02 2000-04-11 General Electric Co. Premixing dry low NOx emissions combustor with lean direct injection of gas fuel
US6192688B1 (en) 1996-05-02 2001-02-27 General Electric Co. Premixing dry low nox emissions combustor with lean direct injection of gas fule
US6868676B1 (en) 2002-12-20 2005-03-22 General Electric Company Turbine containing system and an injector therefor
US20060156735A1 (en) * 2005-01-15 2006-07-20 Siemens Westinghouse Power Corporation Gas turbine combustor
DE102006053679A1 (de) 2005-11-15 2007-05-24 General Electric Co. Niedrigemissionsbrennkammer und Betriebsverfahren
US20090084082A1 (en) 2007-09-14 2009-04-02 Siemens Power Generation, Inc. Apparatus and Method for Controlling the Secondary Injection of Fuel
US20110067402A1 (en) 2009-09-24 2011-03-24 Wiebe David J Fuel Nozzle Assembly for Use in a Combustor of a Gas Turbine Engine

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2017018982A1 (fr) * 2015-07-24 2017-02-02 Siemens Aktiengesellschaft Conduit de transition de turbine à gaz à injection pauvre tardive présentant un temps réduit de séjour de combustion
CN107923621A (zh) * 2015-07-24 2018-04-17 西门子公司 具有减少的燃烧停留时间的带延迟稀薄喷射的燃气涡轮过渡管道
CN107923621B (zh) * 2015-07-24 2020-03-10 西门子公司 具有减少的燃烧停留时间的带延迟稀薄喷射的燃气涡轮过渡管道
EP4187153A1 (fr) * 2021-11-24 2023-05-31 Raytheon Technologies Corporation Chambre de combustion de moteur à turbine à gaz avec conduit(s) de carburant intégré(s)
US11808455B2 (en) 2021-11-24 2023-11-07 Rtx Corporation Gas turbine engine combustor with integral fuel conduit(s)
CN115234943A (zh) * 2022-06-30 2022-10-25 北京航空航天大学 中心分级与轴向分级耦合式燃烧室

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