US20180187563A1 - 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
US20180187563A1
US20180187563A1 US15/739,819 US201515739819A US2018187563A1 US 20180187563 A1 US20180187563 A1 US 20180187563A1 US 201515739819 A US201515739819 A US 201515739819A US 2018187563 A1 US2018187563 A1 US 2018187563A1
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Prior art keywords
flow
accelerating
combustion
cone
accelerating structure
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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.)
Abandoned
Application number
US15/739,819
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English (en)
Inventor
Walter Ray Laster
Juan Enrique Portillo Bilbao
Timothy A. Fox
Grant L. Powers
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Siemens AG
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Siemens AG
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Assigned to SIEMENS ENERGY, INC. reassignment SIEMENS ENERGY, INC. ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: PORTILLO BILBAO, JUAN ENRIQUE, LASTER, WALTER RAY, POWERS, Grant L.
Assigned to SIEMENS CANADA LIMITED reassignment SIEMENS CANADA LIMITED ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: FOX, TIMOTHY A.
Assigned to SIEMENS ENERGY, INC. reassignment SIEMENS ENERGY, INC. ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: SIEMENS CANADA LIMITED
Assigned to SIEMENS AKTIENGESELLSCHAFT reassignment SIEMENS AKTIENGESELLSCHAFT ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: SIEMENS ENERGY, INC.
Publication of US20180187563A1 publication Critical patent/US20180187563A1/en
Abandoned legal-status Critical Current

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    • 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.
  • a flow-accelerating structure such as a flow-accelerating cone
  • 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 Ili& 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
  • 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 vanes 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
  • 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 IEP 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 arrangements for affixing the fuel injectors to the flow-accelerating cone since such mechanical design and/or arrangements can be readily tailored based on the needs of a given application.
  • plot 44 of static temperature as a function of flow speed between the cone inlet and the cone outlet in the context of flow-accelerating cone 51 , as shown in FIG. 3 .
  • a portion 46 of plot 44 corresponds to flattened portion 60 of cone 51 , where, although the flow speed may be constant over flattened portion 60 , such flow speed would be lower compared to the static temperature at cone inlet 26 .
  • injectors 64 may be disposed to provide jet in cross-flow injection, as schematically illustrated in FIG. 5
  • injectors 66 may be positioned normal to a wall 62 of the flow-accelerating cone, as schematically illustrated in FIG. 6 , where arrow 68 schematically represents flow direction.
  • injector angles relative to the flow direction other than those illustrated in FIGS. 5 and 6 , and thus 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.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
US15/739,819 2015-07-24 2015-07-24 Gas turbine transition duct with late lean injection having reduced combustion residence time Abandoned US20180187563A1 (en)

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)

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US20180187563A1 true US20180187563A1 (en) 2018-07-05

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US15/739,819 Abandoned US20180187563A1 (en) 2015-07-24 2015-07-24 Gas turbine transition duct with late lean injection having reduced combustion residence time

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

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US20180252410A1 (en) * 2017-03-02 2018-09-06 General Electric Company Combustor for Use in a Turbine Engine
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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US20180245792A1 (en) * 2017-02-24 2018-08-30 General Electric Company Combustion System with Axially Staged Fuel Injection
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
US10816203B2 (en) 2017-12-11 2020-10-27 General Electric Company Thimble assemblies for introducing a cross-flow into a secondary combustion zone
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

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US20100170216A1 (en) * 2009-01-07 2010-07-08 General Electric Company Late lean injection system configuration
US20110067402A1 (en) * 2009-09-24 2011-03-24 Wiebe David J Fuel Nozzle Assembly for Use in a Combustor of a Gas Turbine Engine
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EP2660519A1 (en) * 2012-04-30 2013-11-06 General Electric Company Transition duct with late lean injection for a gas turbine
US20140260279A1 (en) * 2013-03-18 2014-09-18 General Electric Company Hot gas path duct for a combustor of a gas turbine
US20140260273A1 (en) * 2013-03-18 2014-09-18 General Electric Company Continuous combustion liner for a combustor of a gas turbine
US20140260272A1 (en) * 2013-03-18 2014-09-18 General Electric Company System for providing fuel to a combustor
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US20150052905A1 (en) * 2013-08-20 2015-02-26 General Electric Company Pulse Width Modulation for Control of Late Lean Liquid Injection Velocity
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US8689559B2 (en) * 2009-03-30 2014-04-08 General Electric Company Secondary combustion system for reducing the level of emissions generated by a turbomachine
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Publication number Priority date Publication date Assignee Title
US20180252410A1 (en) * 2017-03-02 2018-09-06 General Electric Company Combustor for Use in a Turbine Engine
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
US11612938B2 (en) 2018-12-07 2023-03-28 Raytheon Technologies Corporation Engine article with integral liner and nozzle
US12053821B2 (en) 2018-12-07 2024-08-06 Rtx Corporation Engine article with integral liner and nozzle

Also Published As

Publication number Publication date
CN107923621B (zh) 2020-03-10
WO2017018982A1 (en) 2017-02-02
CN107923621A (zh) 2018-04-17
JP6584634B2 (ja) 2019-10-02
EP3325887A1 (en) 2018-05-30
JP2018526603A (ja) 2018-09-13

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