EP2242955A1 - Turbine à gaz à chambre de combustion annulaire - Google Patents

Turbine à gaz à chambre de combustion annulaire

Info

Publication number
EP2242955A1
EP2242955A1 EP09712985A EP09712985A EP2242955A1 EP 2242955 A1 EP2242955 A1 EP 2242955A1 EP 09712985 A EP09712985 A EP 09712985A EP 09712985 A EP09712985 A EP 09712985A EP 2242955 A1 EP2242955 A1 EP 2242955A1
Authority
EP
European Patent Office
Prior art keywords
connecting elements
machine according
thermal machine
outer shell
halves
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.)
Granted
Application number
EP09712985A
Other languages
German (de)
English (en)
Other versions
EP2242955B1 (fr
Inventor
Remigi Tschuor
Russell Bond Jones
Nilze Isabel Seda-Maurell
Marion Oneil Duggans
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.)
GE Vernova GmbH
Original Assignee
Alstom Technology 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 Alstom Technology AG filed Critical Alstom Technology AG
Publication of EP2242955A1 publication Critical patent/EP2242955A1/fr
Application granted granted Critical
Publication of EP2242955B1 publication Critical patent/EP2242955B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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
    • F23R3/002—Wall structures
    • 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
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/23—Manufacture essentially without removing material by permanently joining parts together
    • F05D2230/232—Manufacture essentially without removing material by permanently joining parts together by welding
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/30—Retaining components in desired mutual position
    • F05D2260/36—Retaining components in desired mutual position by a form fit connection, e.g. by interlocking
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/94—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF]
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/94—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF]
    • F05D2260/941—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF] particularly aimed at mechanical or thermal stress reduction
    • 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
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00017—Assembling combustion chamber liners or subparts
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49229—Prime mover or fluid pump making

Definitions

  • the present invention relates to the field of thermal machines. It relates to a thermal machine according to the preamble of claim 1 and a method for mounting such a thermal machine.
  • IGT industrial gas turbines
  • annular combustion chamber Usually smaller IGTs are designed as so-called "Can Annular Combustors.”
  • the combustion chamber is bounded by the sidewalls and the entrance and exit planes of the hot gas
  • FIGs 1 and 2 Such a gas turbine is shown in Figures 1 and 2.
  • the in 1 and 2 has a turbine housing 1 1 in which a rotor 12 rotating about an axis 27 is accommodated, and on the right side a compressor 17 for compressing combustion and cooling air is formed on the rotor 12
  • a turbine 13 is arranged on the left-hand side and the compressor 17 compresses air which flows into a plenum 14.
  • annular combustion chamber 15 is arranged concentrically with the axis 27, which is closed on the inlet side by a front plate 19 cooled by front plate cooling air 20 and on the output side via a hot gas channel 25 with the input of the turbine 13 is in communication.
  • burners 16 are arranged in a ring, which are designed for example as a premix burner, as they emerge preferably from EP-A1 -321 809 or EP-AI-704 657, and inject a fuel-air mixture into the combustion chamber 15 .
  • the cited publications and the further developments derived therefrom form an integral part of this application.
  • the resulting during the combustion of the mixture hot air stream 26 passes through the hot gas channel 25 in the turbine 13 and is relaxed there under work.
  • the combustion chamber 15 with the hot gas duct 25 is surrounded at the outside by a distance from an outer and inner cooling jacket 21 and 31, which are fastened by means of fastening elements 24 to the combustion chamber 15, 25 and between each and the combustion chamber 15, 25 respectively an outer and form inner cooling channel 22 and 32 respectively.
  • the cooling channels 22, 32 flows in the opposite direction to the hot gas flow 26 cooling air on the walls of the combustion chamber 15, 25 along a combustion chamber hood 18 and from there into the burner 16 and front plate cooling air 20 directly into the combustion chamber 15th
  • the side walls of the combustion chamber 15, 25 are carried out either as shell elements or as solid shells (outer shell 23, inner shell 33).
  • solid shells the necessity of a parting plane (29 in FIG. 4 ff.) Arises due to the assembly, which makes it possible to remove an upper half of the shell 23, 33 (the upper part), for example the gas turbine rotor 12 to assemble or disassemble.
  • the parting plane 29 accordingly has two parting plane welding seams, which are located at the height of the machine axis 27 using the example of the gas turbine constructed by the applicant.
  • the side walls in the region of the parting planes 29 have a reduced strength and service life.
  • TBC Thermal Barrier Coating Thermal Barrier Coating
  • the thermally very heavily loaded outer and inner shells 23 and 33 act on the four parting planes (29 and others) with high compressive and tensile stresses.
  • the required service life of outer and inner shells 23 and 33 is typically two so-called service intervals (service intervals / service cycles). An operating interval describes the time between (re-) commissioning of the combustion chamber and reconditioning of the components. Both shells, the outer and inner shell 23, 33, often begin to break at the beginning and end of the parting plane welding seams during operation.
  • an additional mechanical positive connection is provided for receiving tensile and shear forces acting on the parting planes on the parting planes.
  • a preferred embodiment of the invention is characterized in that as an additional mechanical positive connection in each case a extending over the parting plane connecting element is provided in the form of a bridge that the outer shell and inner shell at the entrance and / or exit of the combustion chamber have a flange that the connecting elements the outside of the flange are arranged, that the flange on the outside has a circumferential groove, and that the connecting elements are inserted into the groove.
  • the connecting elements can be releasably connected to the two halves of the outer shell or inner shell.
  • the connecting elements with the two halves of the outer shell or inner shell are then releasably connected by screws or bolts.
  • connecting elements can also be materially connected, in particular welded, to the two halves of the outer shell or inner shell.
  • Another embodiment of the invention is characterized in that the groove and the connecting elements are designed such that the connecting elements are held by positive engagement in the groove.
  • the connecting elements have first means for improving the mechanical integrity, wherein throat-shaped incisions are preferably provided as means for improving the mechanical integrity, preferably at the ends.
  • Another embodiment is characterized in that the connecting elements have second means for improving the mountability, wherein as a means for improving the mountability is preferably provided on the top of a cam.
  • Another embodiment is characterized in that the connecting elements have third means for improving the cooling of the connecting elements.
  • the connecting elements fourth means for forming cooling channels between the connecting element and the flange, wherein as a means for forming cooling channels preferably on the bottom of a wavy base is provided.
  • An embodiment of the inventive method is characterized in that the connecting element is loosely inserted in the first step in the upper half and welded in the final position with the two halves.
  • Another embodiment is characterized in that the connecting element is inserted into the upper half in its final position in the first step and secured with screws or bolts, and that in the third step, the upper half positioned with simultaneous retraction of the connecting element on the lower half becomes.
  • FIG. 1 shows a longitudinal section through a cooled annular combustion chamber of a
  • FIG. 2 shows in detail the annular combustion chamber from FIG. 1 with the cooling shirts fastened to the outside;
  • FIG. 4 shows a detail of the halves of the outer shell which abut on the dividing plane together with a bolted bridge arranged on the flange according to a preferred embodiment of the invention
  • Fig. 5 viewed the detail of Figure 4 from another direction.
  • FIG. 6 shows a first partial step in the assembly of the bridge according to FIG. 4;
  • FIG. 7 shows different views of a bridge according to FIG. 4 in different subfigures (a), (b) and (c);
  • FIG. 8 shows a detail of the halves of the outer shell which abut one another at the dividing plane with a welded bridge arranged on the flange according to another preferred embodiment of the invention;
  • Fig. 9 viewed from the section of Figure 8 from another direction.
  • FIG. 10 in different sub-figures (a), (b) and (c) different views of a bridge according to FIG. 8, and
  • FIG. 1 1 in two sub-figures (a) and (b) different views of a provided with additional coolant bridge similar to FIG. 10th
  • An essential feature of the inventive idea is an additional, mechanical form fit of the parting plane welding seams between the half shells of the outer shell and / or inner shell of an annular combustion chamber (note: all the following explanations and illustrations relate to the outer shell, but also apply correspondingly to an inner shell ).
  • a bridge is used as an additional connecting element on both sides of the parting plane, preferably in a respective already existing flange.
  • this bridge may or may not be designed to continue to permit or permit cooling of the flange portion.
  • the structural design is generally subject to the following principles:
  • the bridges should come as close as possible to the "cold" shell outer wall, so that no further, unnecessarily high leverage forces are generated •
  • the bridges can be welded, clamped or bolted with positive locking. Cooling air can be used to cool the underside of the bridges in the immediate vicinity of the thermally stressed shell structure to cause increased transfer of stresses away from the parting line weld through the bridge.
  • the bridge is used on one side, in the upper part of the outer shell, in a flanged groove.
  • the two shells are stacked in the gas turbine (GT) and the bridge is pushed into position or beaten (a cam or a nose on the outer diameter of the bridge can serve as a starting point for a mandrel or hammer.)
  • GT gas turbine
  • the bridge is pushed into position or beaten (a cam or a nose on the outer diameter of the bridge can serve as a starting point for a mandrel or hammer.)
  • the bridge is welded at its top to the flange.
  • the geometric design of the flange and the bridge itself allows preferably the cooling air to flow through the flange under the bridge over - and thus to ensure the conditions for a convective cooling.
  • the bridge is then inserted on one side, in the upper part (in the upper half) of the outer shell, into the flanged groove and positioned with bolts at its destination.
  • the two half shells are placed one above the other in the gas turbine and the bridge is retracted into the lower half shell.
  • the bridge can also be secured in the lower half shell (by bolts and / or screws). For better accessibility when welding the parting line, the bridge can also be removed and reused at any time.
  • FIGS. 4 to 11 The two above-mentioned alternatives (welded or bolted bridge) are to be explained below using the exemplary embodiments of FIGS. 4 to 11.
  • the shells 23, 33 of the annular combustion chamber 15, 25 are preferably provided with flanges on the burner-side and the turbine-side end, which flanges are used for connecting be used between the combustion chamber and adjacent components.
  • Fig. 3 shows an example in longitudinal section of the turbine end of the outer shell 23 of the combustion chamber 15, 25 of FIG. 1 with the attached flange 28.
  • the flange 28 has on the outside of a groove 34, which the mechanical Relieving the parting plane welding seams provided bridges.
  • FIGS. 4 and 5 seen from different angles, the halves 23a, 23b of the outer shell 23 which collide with the dividing plane 29 are reproduced in a section with a bolted bridge 30 arranged on the flange 34 according to a preferred exemplary embodiment of the invention.
  • the bridge 30 itself is shown in various views in Figs. 7a to 7c.
  • the bridge 30 is in the form of an elongate planar strip of rectangular cross section having the slightly curved shape of a circular arc segment.
  • the length of the bridge 30 is selected so that on both sides of the parting plane 29 with sufficient distance two mounting holes 36 can be attached, which serve the screw / Verbolzung the bridge 30 with the two welded half-shells 23 a, 23 b.
  • a connecting element 40 for a discharge arrangement with a welded bridge, a connecting element 40 according to FIG. 8-10 or 11 is preferably used.
  • the bridge 40 is adapted in its cross-sectional contour (FIG. 10 b) to the cross-sectional contour of the flange groove 34 such that the bridge 40 can be inserted into the groove 34 in a form-fitting manner and engages with a foot strip 37 in an undercut in the groove 34.
  • a transversely projecting cam 39 is provided in the middle, at which when driving the bridge 40 into the groove 34 with a striking tool can be recognized.
  • a wave-shaped base 38 is formed (Fig.
  • Throat-shaped cuts 41, 42 are advantageously arranged at the ends of the bridge 40, which are partially incorporated on one side (FIG. 10c) or as a cross (FIG. 11).
  • the radii of curvature of the incisions may vary.
  • novel, positive-locking connecting elements which act as "structural bridges for the combustion chamber shell parting plane" ensure significantly improved force transmission at the ends of the parting plane.
  • the bridges (40) can have throat-shaped incisions (41, 42) at their ends for improved mechanical integrity-improved transfer of force flux, breaking of the force peaks;
  • the cuts in the bridge can be partially integrated on one side or as a cross;
  • the radii of the cuts shown in FIG. 10) can vary;
  • the wall thicknesses of the two shown bridges (30, 40) can vary;
  • the bridges can be supplemented with turbulence ribs on the cooling air side to increase the cooling efficiency
  • the bridges may have a cam (39) at the top for ease of mounting, in order to achieve a simplified hammerability

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne une machine thermique, en particulier une turbine à gaz, qui comprend une chambre de combustion annulaire délimitée vers l'extérieur par une coque extérieure et une coque intérieure. La coque extérieure et la coque intérieure sont respectivement divisées, dans un plan de séparation (29), en une moitié supérieure (23a) et une moitié inférieure (23b), qui sont soudées entre elles dans le plan de séparation (29). On obtient une stabilité mécanique et une longévité accrue de la chambre de combustion par le fait qu'un engagement mécanique supplémentaire par complémentarité de forme (30) est prévue au niveau des plans de séparation (29) afin d'absorber les forces de traction et de cisaillement agissant sur les plans de séparation (29).
EP09712985.2A 2008-02-20 2009-02-12 Turbine à gaz à chambre de combustion annulaire et procédé d'assemblage Active EP2242955B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
CH2452008 2008-02-20
PCT/EP2009/051644 WO2009103658A1 (fr) 2008-02-20 2009-02-12 Turbine à gaz à chambre de combustion annulaire

Publications (2)

Publication Number Publication Date
EP2242955A1 true EP2242955A1 (fr) 2010-10-27
EP2242955B1 EP2242955B1 (fr) 2018-10-17

Family

ID=39735175

Family Applications (1)

Application Number Title Priority Date Filing Date
EP09712985.2A Active EP2242955B1 (fr) 2008-02-20 2009-02-12 Turbine à gaz à chambre de combustion annulaire et procédé d'assemblage

Country Status (5)

Country Link
US (1) US20110113785A1 (fr)
EP (1) EP2242955B1 (fr)
AU (1) AU2009216857B2 (fr)
MY (1) MY158901A (fr)
WO (1) WO2009103658A1 (fr)

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US8549861B2 (en) 2009-01-07 2013-10-08 General Electric Company Method and apparatus to enhance transition duct cooling in a gas turbine engine
EP2309099B1 (fr) 2009-09-30 2015-04-29 Siemens Aktiengesellschaft Conduit de transition
EP2852735B1 (fr) * 2011-10-24 2016-04-27 Alstom Technology Ltd Turbine à gaz
US9915428B2 (en) * 2014-08-20 2018-03-13 Mitsubishi Hitachi Power Systems, Ltd. Cylinder of combustor, method of manufacturing of cylinder of combustor, and pressure vessel
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
US10655853B2 (en) 2016-11-10 2020-05-19 United Technologies Corporation Combustor liner panel with non-linear circumferential edge for a gas turbine engine combustor
US10830433B2 (en) 2016-11-10 2020-11-10 Raytheon Technologies Corporation Axial non-linear interface for combustor liner panels in a gas turbine combustor
US10935235B2 (en) * 2016-11-10 2021-03-02 Raytheon Technologies Corporation Non-planar combustor liner panel for a gas turbine engine combustor
US10935236B2 (en) * 2016-11-10 2021-03-02 Raytheon Technologies Corporation Non-planar combustor liner panel for a gas turbine engine combustor
US11359810B2 (en) * 2017-12-22 2022-06-14 Raytheon Technologies Corporation Apparatus and method for mitigating particulate accumulation on a component of a gas turbine
US10697634B2 (en) * 2018-03-07 2020-06-30 General Electric Company Inner cooling shroud for transition zone of annular combustor liner

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CH342794A (fr) * 1954-10-11 1959-11-30 Canadian Patents Dev Dispositif de combustion de section annulaire pour installation à turbine à gaz
US3031844A (en) * 1960-08-12 1962-05-01 William A Tomolonius Split combustion liner
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DE1626025A1 (de) * 1967-01-12 1970-08-20 Daimler Benz Ag Grossen Temperaturunterschieden ausgesetztes Bauteil
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DE4435266A1 (de) * 1994-10-01 1996-04-04 Abb Management Ag Brenner
JP3831638B2 (ja) * 2001-08-09 2006-10-11 三菱重工業株式会社 板状体接合方法、接合体、ガスタービン燃焼器用の尾筒、及び、ガスタービン燃焼器
US7036316B2 (en) * 2003-10-17 2006-05-02 General Electric Company Methods and apparatus for cooling turbine engine combustor exit temperatures
GB2434199B (en) * 2006-01-14 2011-01-05 Alstom Technology Ltd Combustor liner with heat shield
US7681403B2 (en) * 2006-04-13 2010-03-23 General Electric Company Forward sleeve retainer plate and method

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Also Published As

Publication number Publication date
MY158901A (en) 2016-11-30
WO2009103658A1 (fr) 2009-08-27
AU2009216857B2 (en) 2014-01-16
AU2009216857A1 (en) 2009-08-27
US20110113785A1 (en) 2011-05-19
EP2242955B1 (fr) 2018-10-17

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