WO2017018982A1 - Gas turbine transition duct with late lean injection having reduced combustion residence time - Google Patents

Gas turbine transition duct with late lean injection having reduced combustion residence time Download PDF

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Publication number
WO2017018982A1
WO2017018982A1 PCT/US2015/041948 US2015041948W WO2017018982A1 WO 2017018982 A1 WO2017018982 A1 WO 2017018982A1 US 2015041948 W US2015041948 W US 2015041948W WO 2017018982 A1 WO2017018982 A1 WO 2017018982A1
Authority
WO
WIPO (PCT)
Prior art keywords
flow
combustion
cone
accelerating
turbine engine
Prior art date
Application number
PCT/US2015/041948
Other languages
English (en)
French (fr)
Inventor
Walter Ray Laster
Juan Enrique Portillo Bilbao
Timothy A. Fox
Grant L. POWERS
Original Assignee
Siemens Aktiengesellschaft
Siemens Energy, Inc.
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 Aktiengesellschaft, Siemens Energy, Inc. filed Critical Siemens Aktiengesellschaft
Priority to JP2018503579A priority Critical patent/JP6584634B2/ja
Priority to CN201580081825.3A priority patent/CN107923621B/zh
Priority to EP15747907.2A priority patent/EP3325887A1/en
Priority to PCT/US2015/041948 priority patent/WO2017018982A1/en
Priority to US15/739,819 priority patent/US20180187563A1/en
Publication of WO2017018982A1 publication Critical patent/WO2017018982A1/en

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • 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
    • F23R3/28—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
    • F23R3/425—Combustion chambers comprising a tangential or helicoidal arrangement of the flame tubes
    • 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

  • Disclosed embodiments are generally related to combustion turbine engines, such as gas turbine engines and, more particularly, to a combustion system having a reduced combustion residence time.
  • DCS distributed combustion system
  • FIG. 1 is a fragmentary schematic representation of one non-limiting embodiment of a ducting arrangement with fuel injectors disposed at a location in a flow-accelerating structure, such as a flow-accelerating cone, characterized by a relatively lower static temperature and a reduced combustion residence time, each of which is conducive to reduce NOx emissions at the high firing temperatures of a combustion turbine engine.
  • FIG, 2 illustrate non-limiting plots of decreasing static temperatures as a function of increasing flow speed between the cone inlet and the cone outlet in the flow-accelerating cone shown in FIG. 1.
  • FIGs. 3 and 4 illustrate further non-limiting embodiments of ducting arrangements with fuel injectors disposed at respective flow-accelerating cones.
  • FIG. 5 is a schematic of a fuel injector, which in one non-limiting embodiment may be arranged to provide jet in cross-flow injection.
  • FIG. 6 is a schematic of a fuel injector, which in another non-limiting embodiment may be arranged without providing jet in cross-flow injection DETAILED DESCRIPTION
  • the inventors of the present invention have recognized synergies that result from an innovative integration of what up to the present invention have been perceived as seemingly independent combustor design approaches, such as may involve a distributed combustion system (DCS) approach, and an advanced ducting approach in the combustor system of a combustion turbine engine, such as a gas turbine engine.
  • DCS distributed combustion system
  • a combustion turbine engine such as a gas turbine engine.
  • FIG. 1 is a fragmentary schematic representation of an advanced ducting arrangement 10 in one non-limiting embodiment of a combustor system of a combustion turbine engine, such as a gas turbine engine.
  • a plurality of flow paths 12 blends smoothly into a single, annular chamber 14.
  • each flow path 12 may be configured to deliver combustion gases formed in a respective combustor to a turbine section of the engine without a need of a first stage of flow-directing va es in the turbine section of the engine.
  • each flow path 12 includes a cone 16 and an integrated exit piece (IEP) 18.
  • each cone 16 has a cone inlet 26 having a circular cross section and configured to receive the combustion gases from a combustor outlet (not shown). The cross-sectional profile of cone 16 narrows toward a cone outlet 28 that is associated with an IEP inlet 30 in fluid communication with each other.
  • cone 16 Based on the narrowing cross-sectional profile of cone 16, as the flow travels from cone inlet 26 to cone outlet 28, the flow of combustion gases is accelerated to a relatively high subsonic Mach (M) number, such as without limitation may comprise a range from approximately 0.3 M to approximately a 0.8 M, and thus cone 16 may be generally conceptualized as a non-limiting embodiment of a flow- accelerating structure. Accordingly, the combustion gases may flow through cone 16 with an increasing flow speed, and as a result, this flow of combustion gases can experience a decreasing static temperature in cone 16.
  • M subsonic Mach
  • FIG. 2 illustrates a non-limiting plot 40 of decreasing static temperature as a function of increasing flow speed between the cone inlet and the cone outlet in cone 16, as illustrated in FIG. 1.
  • FIG. 2 further illustrates a plot 42 of total temperature, which is essentially independent of the increasing flow speed between the cone inlet and the cone outlet.
  • FIG. 1 illustrates a single injector 32, as may comprise an assembly of an air scoop and a fuel nozzle, in connection with each of the cones illustrated in FIG. 1; it will be appreciated, however, that multiple injectors may be circumferentially distributed in each cone 16.
  • FIG. 3 illustrates another non- limiting embodiment of a ducting arrangement 50 where a flow-accelerating cone 51 may be made up of two or more interconnected cone sections, in lieu of a single-piece flow- accelerating cone, as described above.
  • a first cone section 52 may be arranged to receive the combustion gases from a combustor outlet 54, and a second cone section 56, affixed at one end to first cone section 52, may be arranged to supply the combustion gases to a corresponding lEP inlet 58.
  • cone sections 52, 54 may each include a respective flattened portion 60 defining a non-varying cross sectional profile where the injectors 32 may be located.
  • a respective manifold 34 (e.g., a ring manifold) is fluidly coupled to the fuel injectors 32.
  • manifold 34 may be affixed (e.g., bolted) between respective interconnecting flanges 33, 35. It will be appreciated that aspects of the present invention are not limited to any specific configuration regarding the mechanical design of the flow-accelerating cone; or regarding mechanical
  • injectors 64 may ⁇ be disposed to provide jet in cross-flow injection, as schematically illustrated in FIG.
  • injectors 66 may be positioned normal to a wall 62 of the flow- accelerating cone, as schematically illustrated in FIG. 6, where arrow 68
  • FIG. 5 schematically represents flow direction. It will be appreciated that one can use injector angles relative to the flow direction other than those illustrated in FIGs. 5 and
  • aspects of the present invention are not limited to injector angles normal to the flow or normal to the wall. That is, aspects of the present invention are not limited to any particular modality of injectors or to any particular injector angle relative to the flow direction.
  • disclosed embodiments are expected to be conducive to a combustion system capable of realizing approximately a 65% combined cycle efficiency or greater in a gas turbine engine.
  • Disclosed embodiments are also expected to realize a combustion system capable of maintaining stable operation at turbine inlet temperatures of approximately 1700°C and higher while maintaining a relatively low level of NOx emissions, and acceptable temperatures in components of the engine without an increase in cooling air consumption.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
PCT/US2015/041948 2015-07-24 2015-07-24 Gas turbine transition duct with late lean injection having reduced combustion residence time WO2017018982A1 (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
JP2018503579A JP6584634B2 (ja) 2015-07-24 2015-07-24 燃焼滞留時間が短縮された遅延希薄噴射を有するガスタービントランジションダクト
CN201580081825.3A CN107923621B (zh) 2015-07-24 2015-07-24 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
EP15747907.2A EP3325887A1 (en) 2015-07-24 2015-07-24 Gas turbine transition duct with late lean injection having reduced combustion residence time
PCT/US2015/041948 WO2017018982A1 (en) 2015-07-24 2015-07-24 Gas turbine transition duct with late lean injection having reduced combustion residence time
US15/739,819 US20180187563A1 (en) 2015-07-24 2015-07-24 Gas turbine transition duct with late lean injection having reduced combustion residence time

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/US2015/041948 WO2017018982A1 (en) 2015-07-24 2015-07-24 Gas turbine transition duct with late lean injection having reduced combustion residence time

Publications (1)

Publication Number Publication Date
WO2017018982A1 true WO2017018982A1 (en) 2017-02-02

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ID=53785745

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2015/041948 WO2017018982A1 (en) 2015-07-24 2015-07-24 Gas turbine transition duct with late lean injection having reduced combustion residence time

Country Status (5)

Country Link
US (1) US20180187563A1 (zh)
EP (1) EP3325887A1 (zh)
JP (1) JP6584634B2 (zh)
CN (1) CN107923621B (zh)
WO (1) WO2017018982A1 (zh)

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2020508431A (ja) * 2017-02-24 2020-03-19 ゼネラル・エレクトリック・カンパニイ 軸方向多段型燃料噴射を備える燃焼システム
US10816203B2 (en) 2017-12-11 2020-10-27 General Electric Company Thimble assemblies for introducing a cross-flow into a secondary combustion zone
US11137144B2 (en) 2017-12-11 2021-10-05 General Electric Company Axial fuel staging system for gas turbine combustors
US11156164B2 (en) 2019-05-21 2021-10-26 General Electric Company System and method for high frequency accoustic dampers with caps
US11174792B2 (en) 2019-05-21 2021-11-16 General Electric Company System and method for high frequency acoustic dampers with baffles
US11187415B2 (en) 2017-12-11 2021-11-30 General Electric Company Fuel injection assemblies for axial fuel staging in gas turbine combustors

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US10823418B2 (en) * 2017-03-02 2020-11-03 General Electric Company Gas turbine engine combustor comprising air inlet tubes arranged around the combustor
US11248789B2 (en) * 2018-12-07 2022-02-15 Raytheon Technologies Corporation Gas turbine engine with integral combustion liner and turbine nozzle

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EP2808610A1 (de) * 2013-05-31 2014-12-03 Siemens Aktiengesellschaft Gasturbinen-Brennkammer mit Tangentialeindüsung als späte Mager-Einspritzung

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Publication number Priority date Publication date Assignee Title
JP2020508431A (ja) * 2017-02-24 2020-03-19 ゼネラル・エレクトリック・カンパニイ 軸方向多段型燃料噴射を備える燃焼システム
JP7051884B2 (ja) 2017-02-24 2022-04-11 ゼネラル・エレクトリック・カンパニイ 軸方向多段型燃料噴射を備える燃焼システム
US10816203B2 (en) 2017-12-11 2020-10-27 General Electric Company Thimble assemblies for introducing a cross-flow into a secondary combustion zone
US11137144B2 (en) 2017-12-11 2021-10-05 General Electric Company Axial fuel staging system for gas turbine combustors
US11187415B2 (en) 2017-12-11 2021-11-30 General Electric Company Fuel injection assemblies for axial fuel staging in gas turbine combustors
US11156164B2 (en) 2019-05-21 2021-10-26 General Electric Company System and method for high frequency accoustic dampers with caps
US11174792B2 (en) 2019-05-21 2021-11-16 General Electric Company System and method for high frequency acoustic dampers with baffles

Also Published As

Publication number Publication date
US20180187563A1 (en) 2018-07-05
CN107923621A (zh) 2018-04-17
JP2018526603A (ja) 2018-09-13
EP3325887A1 (en) 2018-05-30
JP6584634B2 (ja) 2019-10-02
CN107923621B (zh) 2020-03-10

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