EP2383518A2 - Tangentialbrennkammer - Google Patents

Tangentialbrennkammer Download PDF

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Publication number
EP2383518A2
EP2383518A2 EP11163440A EP11163440A EP2383518A2 EP 2383518 A2 EP2383518 A2 EP 2383518A2 EP 11163440 A EP11163440 A EP 11163440A EP 11163440 A EP11163440 A EP 11163440A EP 2383518 A2 EP2383518 A2 EP 2383518A2
Authority
EP
European Patent Office
Prior art keywords
combustion
transition pieces
combustion section
products
combustor cans
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Withdrawn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
EP11163440A
Other languages
English (en)
French (fr)
Other versions
EP2383518A3 (de
Inventor
David Richard Johns
Craig Steven Humanchuk
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 EP2383518A2 publication Critical patent/EP2383518A2/de
Publication of EP2383518A3 publication Critical patent/EP2383518A3/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Withdrawn legal-status Critical Current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
    • F23R3/42—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
    • F23R3/58—Cyclone or vortex type combustion chambers

Definitions

  • the invention relates to a gas turbine and, more particularly, to a combustion section for a gas turbine including structure for turning air flow into the turbine.
  • gas flow is exhausted from the combustor and routed by the transition duct to first stage vanes and blades (rotating buckets).
  • first stage vanes and blades rotating buckets
  • the function of the first stage vanes is to accelerate and turn the flow in a circumferential direction so that the predominant flow direction of the gas flow leaving the trailing edges of the vanes is angled in the circumferential or tangential direction relative to the longitudinal direction.
  • This turned flow thus has a longitudinal component and a circumferential component.
  • the flow angle can be substantial, in the range of 40 degrees to 85 degrees measured from the longitudinal axis.
  • first stage vanes to accelerate and turn the longitudinal gas flow in the circumferential direction presents several challenges.
  • the vanes and the associated vane support structure must have high strength characteristics to withstand the forces generated in changing the direction of hot, high pressure gas flow over a substantial angle in a relatively short distance.
  • the temperature of the gas flow and the heat generated by this turning process also require a vane cooling system.
  • the forces and heat involved can crack and otherwise damage the vanes and associated support structure.
  • the first stage vanes and the associated support structure and cooling systems have developed into a complex system that can be expensive to manufacture, install, and, in the event of damage, repair and replace.
  • First stage nozzles add approximately 1.5% total pressure drop, reducing the total pressure ratio of the machine. Performance is closely tied to pressure ratio. Another major draw on engine performance is parasitic flows. The minimum compressor mass flow in current state of the art air-cooled gas turbines represents a major reduction in overall turbine performance. This air flow could be reduced or effectively eliminated by (1) much smaller wetted surface area of the turning/throttling vanes (if turning/throttling vanes are required), (2) possible use of closed circuit cooling of the transition piece vanes, and (3) possible use of air that cools the vanes then passes through the combustor.
  • a combustion section for a gas turbine including rotating buckets that are driven by products of combustion from the combustion section.
  • the combustion section includes a casing defining a chamber, a plurality of combustor cans disposed in the casing and oriented in an annular pattern, and a plurality of transition pieces one each coupled with each of the combustor cans.
  • the transition pieces direct the products of combustion from the combustor cans into contact with the rotating buckets.
  • Each of the transition pieces is angled in two planes to effect turning of the products of combustion and to shorten the gas turbine.
  • transition pieces of the combustion section are angled tangentially relative to the annular pattern and axially in a flow direction.
  • a method of directing products of combustion into contact with rotating buckets of a gas turbine includes the steps of orienting a plurality of combustor cans in an annular pattern in a combustor casing, providing a plurality of transition pieces each one coupled with one of the combustor cans, and angling the transition pieces tangentially relative to the axial flow direction.
  • FIG. 1 illustrates a typical combustor for a gas turbine, which includes a compressor, a plurality of combustors, and a turbine.
  • the compressor pressurizes inlet air, which is then reverse flowed to the combustor where it is used to cool the combustor and to provide air to the combustion process.
  • the combustor 10 includes a liner 12 that defines a combustion zone and a transition piece 14 that connects the outlet end of the combustor with an inlet end of the turbine to deliver products of combustion to the turbine.
  • the interface between the combustion transition piece 14 and the turbine first stage nozzle requires the use of seals to reduce leakages into the gas path.
  • FIG. 2 is an end view of a combustion section 30 for the gas turbine.
  • the combustion section 30 includes a casing 32 defming a chamber, and a plurality of combustor cans 34 disposed in the casing and oriented in an annular pattern as shown.
  • a plurality of transition pieces 36 one each coupled with each of the combustor cans 34, serves to direct products of combustion from the combustor cans 34 into contact with rotating buckets of the turbine.
  • Each of the transition pieces 36 is angled in two planes to effect turning of the products of combustion and to shorten the gas turbine.
  • the transition pieces 36 are angled tangentially relative to the annular pattern of combustor cans 34 (in an X-Y plane, with the Z axis into the page in FIG. 2 ). Additionally, the transition pieces are angled axially in a flow direction (i.e., toward the turbine buckets) as shown in FIG. 3 (in the Y-Z plane, with the X axis into the page in FIG. 3 ). The angles are determined to effect a proper angle of incidence to extract work from the rotating buckets. These angles vary on turbine design.
  • This structure offers benefits to gas turbine performance and configuration.
  • the flange-to-flange length can be shortened, the pressure loss across the first stage nozzle can be drastically reduced, and the first stage nozzle cooling and leakage flows can be drastically reduced.
  • Angling of the combustor in the tangential direction allows the combustor to provide all or a portion of the turning effect that is typically applied in the first stage nozzle of the turbine. As a consequence, much lower pressure drop will occur in the process of turning the air flow into the first stage bucket.
  • the transition piece includes throttling vanes 38.
  • the throttling vanes may be formed by contouring the sides of the transition piece. The throttling vanes would further accelerate and straighten the flow to bring the air to ideal conditions to extract work in the first stage bucket.
  • ends of the transition pieces 36 may be turned, possibly up to 30° or more, to give the required overall turning.
  • the turning/throttling vanes (if required) will have much smaller wetted surface area than current first stage nozzles. As a result, less heat transfer will be required to keep them within acceptable material temperatures. Moreover, by integrating the combustor transition piece and the first stage turning vanes, the opportunity to use air that eventually passes through the combustor to cool the nozzle is facilitated. Since the combustor transition piece will take the place of the first stage nozzle, there will be no interference leakages between these two pieces as in conventional designs.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Combustion Of Fluid Fuel (AREA)
EP11163440.8A 2010-04-27 2011-04-21 Tangentialbrennkammer Withdrawn EP2383518A3 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US12/768,385 US20110259015A1 (en) 2010-04-27 2010-04-27 Tangential Combustor

Publications (2)

Publication Number Publication Date
EP2383518A2 true EP2383518A2 (de) 2011-11-02
EP2383518A3 EP2383518A3 (de) 2014-11-19

Family

ID=44280768

Family Applications (1)

Application Number Title Priority Date Filing Date
EP11163440.8A Withdrawn EP2383518A3 (de) 2010-04-27 2011-04-21 Tangentialbrennkammer

Country Status (4)

Country Link
US (1) US20110259015A1 (de)
EP (1) EP2383518A3 (de)
JP (1) JP2011232022A (de)
CN (1) CN102252345A (de)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2660519A1 (de) * 2012-04-30 2013-11-06 General Electric Company Überleitkanal mit später Mager-Einspritzung für eine Gasturbine
EP2738355A1 (de) * 2012-11-30 2014-06-04 General Electric Company Gasturbinentriebwerk und zugehöriges Verfahren
EP2636847A3 (de) * 2012-03-09 2015-10-14 General Electric Company Vorrichtung und System zum Leiten von Heißgas
WO2017023330A1 (en) * 2015-08-06 2017-02-09 Siemens Aktiengesellschaft A ducting arrangement for directing combustion gas

Families Citing this family (32)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8978388B2 (en) 2011-06-03 2015-03-17 General Electric Company Load member for transition duct in turbine system
US8448450B2 (en) 2011-07-05 2013-05-28 General Electric Company Support assembly for transition duct in turbine system
US8650852B2 (en) 2011-07-05 2014-02-18 General Electric Company Support assembly for transition duct in turbine system
US8974179B2 (en) 2011-11-09 2015-03-10 General Electric Company Convolution seal for transition duct in turbine system
US8701415B2 (en) 2011-11-09 2014-04-22 General Electric Company Flexible metallic seal for transition duct in turbine system
US8459041B2 (en) 2011-11-09 2013-06-11 General Electric Company Leaf seal for transition duct in turbine system
US20130239585A1 (en) * 2012-03-14 2013-09-19 Jay A. Morrison Tangential flow duct with full annular exit component
US9145778B2 (en) * 2012-04-03 2015-09-29 General Electric Company Combustor with non-circular head end
US20130283802A1 (en) * 2012-04-27 2013-10-31 General Electric Company Combustor
US9038394B2 (en) 2012-04-30 2015-05-26 General Electric Company Convolution seal for transition duct in turbine system
US20130283817A1 (en) * 2012-04-30 2013-10-31 General Electric Company Flexible seal for transition duct in turbine system
US9021783B2 (en) * 2012-10-12 2015-05-05 United Technologies Corporation Pulse detonation engine having a scroll ejector attenuator
US8707673B1 (en) 2013-01-04 2014-04-29 General Electric Company Articulated transition duct in turbomachine
US9322335B2 (en) * 2013-03-15 2016-04-26 Siemens Energy, Inc. Gas turbine combustor exit piece with hinged connections
US9080447B2 (en) 2013-03-21 2015-07-14 General Electric Company Transition duct with divided upstream and downstream portions
US9134029B2 (en) * 2013-09-12 2015-09-15 Siemens Energy, Inc. Radial midframe baffle for can-annular combustor arrangement having tangentially oriented combustor cans
US9458732B2 (en) * 2013-10-25 2016-10-04 General Electric Company Transition duct assembly with modified trailing edge in turbine system
US9309774B2 (en) 2014-01-15 2016-04-12 Siemens Energy, Inc. Assembly for directing combustion gas
US9404421B2 (en) 2014-01-23 2016-08-02 Siemens Energy, Inc. Structural support bracket for gas flow path
US9593853B2 (en) 2014-02-20 2017-03-14 Siemens Energy, Inc. Gas flow path for a gas turbine engine
EP3204615B1 (de) 2014-10-07 2018-11-28 Siemens Energy, Inc. Anordnung für ein verbrennungsgasturbinentriebwerk
CN104654357B (zh) * 2015-01-23 2020-01-03 北京华清燃气轮机与煤气化联合循环工程技术有限公司 燃气轮机燃烧室
CN104776449A (zh) * 2015-01-23 2015-07-15 北京华清燃气轮机与煤气化联合循环工程技术有限公司 热通道补燃燃烧室
CN104595926A (zh) * 2015-01-23 2015-05-06 北京华清燃气轮机与煤气化联合循环工程技术有限公司 热通道部件一体燃烧室
WO2017023331A1 (en) * 2015-08-06 2017-02-09 Siemens Aktiengesellschaft Ducting arrangement for directing combustion gas
US9810434B2 (en) * 2016-01-21 2017-11-07 Siemens Energy, Inc. Transition duct system with arcuate ceramic liner for delivering hot-temperature gases in a combustion turbine engine
US10260360B2 (en) * 2016-03-24 2019-04-16 General Electric Company Transition duct assembly
US10260752B2 (en) 2016-03-24 2019-04-16 General Electric Company Transition duct assembly with late injection features
US10260424B2 (en) 2016-03-24 2019-04-16 General Electric Company Transition duct assembly with late injection features
US10227883B2 (en) 2016-03-24 2019-03-12 General Electric Company Transition duct assembly
US10145251B2 (en) 2016-03-24 2018-12-04 General Electric Company Transition duct assembly
CN115355528A (zh) * 2022-09-02 2022-11-18 中航通飞华南飞机工业有限公司 一种涡桨发动机斜流式燃烧室

Family Cites Families (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US972642A (en) * 1909-07-21 1910-10-11 William A Reed Gas-turbine engine.
US2851853A (en) * 1953-12-28 1958-09-16 Thomas E Quick Thrust augmentation means for jet propulsion engines
US2945672A (en) * 1956-10-05 1960-07-19 Marquardt Corp Gas turbine unit
US3238718A (en) * 1964-01-30 1966-03-08 Boeing Co Gas turbine engine
US3657884A (en) * 1970-11-20 1972-04-25 Westinghouse Electric Corp Trans-nozzle steam injection gas turbine
US4016718A (en) * 1975-07-21 1977-04-12 United Technologies Corporation Gas turbine engine having an improved transition duct support
GB2300909B (en) * 1995-05-18 1998-09-30 Europ Gas Turbines Ltd A gas turbine gas duct arrangement
EP1284391A1 (de) * 2001-08-14 2003-02-19 Siemens Aktiengesellschaft Brennkammeranordnung für Gasturbinen
US6840048B2 (en) * 2002-09-26 2005-01-11 General Electric Company Dynamically uncoupled can combustor
US7373772B2 (en) * 2004-03-17 2008-05-20 General Electric Company Turbine combustor transition piece having dilution holes
US7721547B2 (en) * 2005-06-27 2010-05-25 Siemens Energy, Inc. Combustion transition duct providing stage 1 tangential turning for turbine engines
EP1975373A1 (de) * 2007-03-06 2008-10-01 Siemens Aktiengesellschaft Leitschaufelkanalelement für eine Leitschaufelanordnung eines Gasturbinentriebwerks
US20090266047A1 (en) * 2007-11-15 2009-10-29 General Electric Company Multi-tube, can-annular pulse detonation combustor based engine with tangentially and longitudinally angled pulse detonation combustors
FR2927951B1 (fr) * 2008-02-27 2011-08-19 Snecma Ensemble diffuseur-redresseur pour une turbomachine
US8113003B2 (en) * 2008-08-12 2012-02-14 Siemens Energy, Inc. Transition with a linear flow path for use in a gas turbine engine

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
None

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2636847A3 (de) * 2012-03-09 2015-10-14 General Electric Company Vorrichtung und System zum Leiten von Heißgas
EP2660519A1 (de) * 2012-04-30 2013-11-06 General Electric Company Überleitkanal mit später Mager-Einspritzung für eine Gasturbine
US9133722B2 (en) 2012-04-30 2015-09-15 General Electric Company Transition duct with late injection in turbine system
EP2738355A1 (de) * 2012-11-30 2014-06-04 General Electric Company Gasturbinentriebwerk und zugehöriges Verfahren
US9551492B2 (en) 2012-11-30 2017-01-24 General Electric Company Gas turbine engine system and an associated method thereof
WO2017023330A1 (en) * 2015-08-06 2017-02-09 Siemens Aktiengesellschaft A ducting arrangement for directing combustion gas

Also Published As

Publication number Publication date
CN102252345A (zh) 2011-11-23
JP2011232022A (ja) 2011-11-17
EP2383518A3 (de) 2014-11-19
US20110259015A1 (en) 2011-10-27

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