EP2629018A2 - System für späte Magergemischeinspritzung - Google Patents

System für späte Magergemischeinspritzung Download PDF

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Publication number
EP2629018A2
EP2629018A2 EP13154946.1A EP13154946A EP2629018A2 EP 2629018 A2 EP2629018 A2 EP 2629018A2 EP 13154946 A EP13154946 A EP 13154946A EP 2629018 A2 EP2629018 A2 EP 2629018A2
Authority
EP
European Patent Office
Prior art keywords
injection system
late lean
lean injection
fuel
duct
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Withdrawn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
EP13154946.1A
Other languages
English (en)
French (fr)
Inventor
Patrick Benedict Melton
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.)
General Electric Co
Original Assignee
General Electric Co
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 General Electric Co filed Critical General Electric Co
Publication of EP2629018A2 publication Critical patent/EP2629018A2/de
Withdrawn legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements

Definitions

  • the subject matter disclosed herein relates to turbines, and more particularly to late lean injection systems.
  • a late lean injection system includes at least one fuel injector disposed proximate a combustion zone. Also included is at least one guide for directing an airflow from a region proximate a compressor discharge exit to the at least one fuel injector.
  • a late lean injection system includes a transition duct defining a transition interior, the transition duct having an end adapted for connection to a first turbine zone, and an opposite end. Also included is a sleeve spaced radially outward of the transition duct and extending circumferentially around the transition duct. Further included is at least one fuel injector configured to inject fuel into the transition interior. Yet further included is at least one guide for directing an airflow to the at least one fuel injector.
  • a late lean injection system includes a transition duct having an upstream end and a downstream end. Also included is a liner duct disposed proximate the upstream end of the transition duct. Further included is a flowsleeve spaced radially outward of the liner duct and extending circumferentially around the liner duct. Yet further included is at least one fuel injector disposed proximate at least one of the transition duct and the liner duct. Also included is at least one guide for directing an airflow from a region proximate a compressor discharge exit to the at least one fuel injector.
  • the late lean injection system 10 includes a transition piece assembly 11 of a gas turbine system that is operably connected between a combustor (not labeled) and a first turbine stage (not illustrated) and includes an interior 21 defined by a transition duct 12.
  • the transition duct 12 carries hot combustion gases from the combustor, which is typically upstream of the transition duct 12, to an inlet of the turbine. At least a portion of the transition duct 12 may be surroundingly enclosed by an impingement sleeve 14 that is spaced radially outward of the transition duct 12. Upstream of the transition piece assembly 11 is a liner duct 16.
  • the interior region of the liner duct 16 and the transition duct 12 comprises a combustion zone, wherein combustion of the hot gases occurs and is directed toward the turbine. At least a portion of the liner duct 16 is surroundingly enclosed by a flowsleeve 17 that is spaced radially outward of the liner duct 16.
  • a compressor discharge casing 32 is illustrated and includes a compressor discharge exit 34.
  • the combustor of the gas turbine is late lean injection (LLI) compatible.
  • LLI compatible combustor is any combustor with either an exit temperature that exceeds 2500°F or handles fuels with components that are more reactive than methane with a hot side residence time greater than 10 milliseconds (ms).
  • the late lean injection system 10 of a second embodiment is illustrated.
  • the late lean injection system 10 of the second embodiment is similar to that of the first embodiment, however, does not include an impingement sleeve 14 that surroundingly encloses the transition duct 12.
  • the late lean injection system 10 of a third embodiment comprises merely a single duct, that being the transition duct 12 that extends upstream to a region that included the liner duct 16 in the first and second embodiments. Furthermore, a single sleeve, referred to generally as a sleeve 19 surroundingly encloses the transition duct 12 at a location radially outward of the transition duct 12.
  • a plurality of fuel injectors 18 are each integrated with or structurally supported by a plurality of housings that extend radially into at least one of the transition duct 12 or the liner duct 16.
  • the plurality of fuel injectors 18 extend through the respective duct, i.e., the transition duct 12 or the liner duct 16, to varying depths. That is, the fuel injectors 18 are each configured to supply a second fuel (i.e., LLI fuel) to the combustion zone through fuel injection in a direction that is generally transverse to a predominant flow direction through the transition duct 12 and/or the liner duct 16.
  • a second fuel i.e., LLI fuel
  • the plurality of fuel injectors 18 may be disposed proximate the transition duct 12 or the liner duct 16, in spite of the illustrated embodiments showing disposal of the plurality of fuel injectors 18 disposed in connection with only one of the transition duct 12 and the liner duct 16. Furthermore, the plurality of fuel injectors 18 may be disposed in connection with both the transition duct 12 and the liner duct 16. The plurality of fuel injectors 18 may be disposed in a single axial circumferential stage that includes multiple currently operating fuel injectors 18 respectively disposed around a circumference of a single axial location of the transition duct 12 and/or the liner duct 16.
  • the plurality of fuel injectors 18 may be situated in a single axial stage, multiple axial stages, or multiple axial circumferential stages.
  • a single axial stage includes a currently operating single fuel injector 18.
  • a multiple axial stage includes multiple currently operating fuel injectors 18 that are respectively disposed at multiple axial locations.
  • a multiple axial circumferential stage includes multiple currently operating fuel injectors 18, which are disposed around a circumference of the transition duct 12 and/or the liner duct 16 at multiple axial locations thereof.
  • Airflow from a compressor enters into a compressor discharge casing 32.
  • a high pressure dynamic airflow 20 exits the compressor discharge casing 32 proximate a compressor discharge exit 34 and rushes downstream toward the transition duct 12 and/or the liner duct 16 to locations proximate the fuel injectors 18.
  • the impingement sleeve 14 and/or the flowsleeve 17, or the transition duct 12 in the case of the embodiment illustrated in FIG. 2 includes one or more guides 22 to redirect the high pressure dynamic airflow 20 into the fuel injectors 18.
  • the guides 22 are in the form of scoops that are positioned to correspond to the fuel injectors 18. Based on this correspondence to the fuel injectors 18, the guides may be disposed in a single axial circumferential stage, a single axial stage, a multiple axial stage, or a multiple axial circumferential stage, as is the case with the fuel injectors 18.
  • the impingement sleeve 14 and/or the flowsleeve 17 include apertures 24 that correspond to the fuel injectors 18 and the guides 22 are positioned proximate the apertures 24.
  • a typical scoop can either fully or partially surround each aperture 24 or partially or fully cover the aperture 24 and be generally part-spherical in shape.
  • the scoop may be in the shape of a half cylinder with or without a top.
  • the guides 22 may take the form of various other shapes that provide a similar functionality, specifically harnessing of the high pressure dynamic airflow 20.
  • the guides 22 may be disposed radially inward of the impingement sleeve 14 and/or the flowsleeve 17 and may be in direct connection with the plurality of fuel injectors 18 in embodiments where a sleeve is not present.
  • the guides 22 may be attached individually proximate the impingement sleeve 14 and/or flowsleeve 17, or the transition duct 12 in the case of the embodiment illustrated in FIG. 2 , so as to direct the compressor discharge air radially inboard, through the guides 22, apertures 24, into the fuel injectors 18, and projecting into the transition duct 12 and/or the liner duct 16.
  • the airflow 20 is quickly turned and redirected inboard. Such a redirection may lead to turning vortices within the airflow, thereby hindering the flow into the fuel injector 18.
  • each guide 22 may include one or more straightening vanes 26 proximate a bend 28 in the guide 22.
  • FIG. 5 a penetration profile of the mixed airflow and LLI fuel is illustrated.
  • the harnessing of the high pressure dynamic airflow 20 allows deeper penetration of the late lean injection into the combustion zone.
  • airflow is channeled toward the fuel injectors 18 by the guides 22 that project out into the high pressure dynamic airflow 20 passing the impingement sleeve 14 and/or the flowsleeve 17 of the transition duct 12 and/or the liner duct 16.
  • the guides 22, by a combination of stagnation and redirection, catch air that would previously have passed the apertures 24 aligned with the fuel injectors 18 due to the lack of static pressure differential to drive the flow through them, and directs the airflow 20 inward into the fuel injectors 18 to mix with LLI fuel, and into the transition duct 12 and/or the liner duct 16, thus producing deeper penetration into the combustion zone.
  • the guides 122 are substantially longer than the above-described guides 22 in the form of scoops or the like.
  • the guides 122 function similarly to guides 22, such that high pressure dynamic airflow 120 is directed from a compressor discharge exit 133 to one or more fuel injectors 118.
  • the guides 122 include a first end 130 disposed proximate the compressor discharge exit 133 and a second end 132 disposed proximate an aperture 124 of an impingement sleeve 114 and/or a flowsleeve 117, where the aperture 124 is relatively aligned with an inlet 134 of each fuel injector 118.
  • the guides 122 function as passages that take the high pressure dynamic airflow 120 to the fuel injectors 118.
  • Each guide 122 may be mounted to numerous components within the gas turbine assembly including, but not limited to, a compressor discharge casing 131 or various other combustion hardware components.
  • the contour of the guides 122 as they extend from the first end 130 to the second end 132 may vary based on the specific application of use.
  • One typical contour comprises a substantially elongated straight portion 136 extending from a region proximate the first end 130 and a bend portion 128 that functions to transition the airflow 120 from the substantially elongated straight portion 136 to the inlet 134 of the fuel injector 118.
  • a bend portion 128 may impose turning vortices on the airflow.
  • the bend portion 128 may include one or more straightening vanes 126.
  • FIG. 7 a penetration profile of the mixed airflow and LLI fuel is illustrated.
  • the harnessing of the high pressure dynamic airflow 120 allows deeper penetration of the late lean injection into the combustion zone.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Fuel-Injection Apparatus (AREA)
EP13154946.1A 2012-02-16 2013-02-12 System für späte Magergemischeinspritzung Withdrawn EP2629018A2 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/398,630 US20130213046A1 (en) 2012-02-16 2012-02-16 Late lean injection system

Publications (1)

Publication Number Publication Date
EP2629018A2 true EP2629018A2 (de) 2013-08-21

Family

ID=47709980

Family Applications (1)

Application Number Title Priority Date Filing Date
EP13154946.1A Withdrawn EP2629018A2 (de) 2012-02-16 2013-02-12 System für späte Magergemischeinspritzung

Country Status (5)

Country Link
US (1) US20130213046A1 (de)
EP (1) EP2629018A2 (de)
JP (1) JP2013167435A (de)
CN (1) CN103256630A (de)
RU (1) RU2013106577A (de)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9803555B2 (en) 2014-04-23 2017-10-31 General Electric Company Fuel delivery system with moveably attached fuel tube

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US9551492B2 (en) * 2012-11-30 2017-01-24 General Electric Company Gas turbine engine system and an associated method thereof
EP2789915A1 (de) * 2013-04-10 2014-10-15 Alstom Technology Ltd Verfahren zum Betrieb einer Brennkammer und Brennkammer
US20160047317A1 (en) * 2014-08-14 2016-02-18 General Electric Company Fuel injector assemblies in combustion turbine engines
US10054314B2 (en) 2015-12-17 2018-08-21 General Electric Company Slotted injector for axial fuel staging
US10976052B2 (en) 2017-10-25 2021-04-13 General Electric Company Volute trapped vortex combustor assembly
US10976053B2 (en) 2017-10-25 2021-04-13 General Electric Company Involute trapped vortex combustor assembly
US11181269B2 (en) 2018-11-15 2021-11-23 General Electric Company Involute trapped vortex combustor assembly

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US4226083A (en) * 1978-01-19 1980-10-07 United Technologies Corporation Method and apparatus for reducing nitrous oxide emissions from combustors
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JP2950720B2 (ja) * 1994-02-24 1999-09-20 株式会社東芝 ガスタービン燃焼装置およびその燃焼制御方法
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Publication number Priority date Publication date Assignee Title
US9803555B2 (en) 2014-04-23 2017-10-31 General Electric Company Fuel delivery system with moveably attached fuel tube

Also Published As

Publication number Publication date
JP2013167435A (ja) 2013-08-29
US20130213046A1 (en) 2013-08-22
RU2013106577A (ru) 2014-08-20
CN103256630A (zh) 2013-08-21

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