EP2927591A1 - Anneau de refroidissement et brûleur de turbine à gaz dotée d'un tel anneau de refroidissement - Google Patents

Anneau de refroidissement et brûleur de turbine à gaz dotée d'un tel anneau de refroidissement Download PDF

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Publication number
EP2927591A1
EP2927591A1 EP14162655.6A EP14162655A EP2927591A1 EP 2927591 A1 EP2927591 A1 EP 2927591A1 EP 14162655 A EP14162655 A EP 14162655A EP 2927591 A1 EP2927591 A1 EP 2927591A1
Authority
EP
European Patent Office
Prior art keywords
cooling
cooling ring
ring
gas turbine
combustion chamber
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Withdrawn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
EP14162655.6A
Other languages
German (de)
English (en)
Inventor
Olga Deiss
Julian Timmermann
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.)
Siemens AG
Original Assignee
Siemens 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 Siemens AG filed Critical Siemens AG
Priority to EP14162655.6A priority Critical patent/EP2927591A1/fr
Publication of EP2927591A1 publication Critical patent/EP2927591A1/fr
Withdrawn legal-status Critical Current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
    • F23R3/005—Combined with pressure or heat exchangers
    • 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
    • F23R3/42—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
    • 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
    • F23R2900/00012—Details of sealing 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
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03043—Convection cooled combustion chamber walls with means for guiding the cooling air flow

Definitions

  • the present invention relates to a cooling ring with a plurality of cooling channels, each extending from one end face to the opposite end face of the cooling ring therethrough, and a gas turbine burner with such a cooling ring.
  • Gas turbine assemblies are known in the art in various configurations. They include a compressor, multiple gas turbine burners and a turbine. During operation, ambient air is compressed using the compressor and supplied to the gas turbine combustors, where the compressed air is mixed with fuel and the mixture is burned to produce combustion gases. The combustion gases exit the respective combustors of the gas turbine combustors and are routed via transition ducts, each connected to the combustion chambers of the gas turbine combustors, to the turbine, the blades of which are rotationally driven by the combustion gases.
  • the connection between a combustor of a gas turbine combustor and an associated transition duct is normally made by inserting the exhaust end of the combustor into the inlet end of the transition duct during assembly of the gas turbine combustor.
  • spring elements are provided at the outlet end of the transition channel, which protrude radially outwardly from the outer surface of the combustion chamber and engage during insertion with the inner periphery of the inlet end of the transitional channel clamped.
  • the combustion chambers are for the most part directly supplied with or flowed around by the compressor discharge air, so that sufficient cooling is present. Only in the transition region between the combustion chambers and the associated transitional channels, this is not due to the spring elements arranged there the case, which requires additional cooling.
  • the cooling ring may for example be provided as a separate component and welded to the free end of the combustion chamber, so that it continues in one piece the combustion chamber.
  • the cooling ring is provided along its circumference with a plurality of cooling channels in the form of bores, which extend in each case from one end face to the opposite end face of the cooling ring in the axial direction therethrough and are flowed through during operation of the Verêtrendluft.
  • improved cooling is achieved.
  • the holes only cool at certain points, creating stresses in the ring during operation, which is not desirable.
  • the production of long axial bores is relatively expensive.
  • the present invention provides a cooling ring of the type mentioned above, which is characterized in that the cooling ring is formed of at least two concentric ring elements whose mutually facing annular surfaces touch each other and are interconnected, wherein at least one of the mutually facing annular surfaces the cooling channels having defining grooves.
  • the cooling channels can be made simple and inexpensive.
  • the cooling channels can be formed using a cutter on one or both of the ring elements.
  • the shape and the course of each cooling channel are almost freely selectable, which is why the cooling channels effectively to provide even cooling, thereby reducing the stresses occurring in the cooling ring during operation.
  • the grooves extend in a straight line from one end face to the opposite end face of the cooling ring.
  • the grooves extend meander-shaped, which is associated with longer, the cooling-promoting cooling channels.
  • each cooling channel Preferably, the inlet and the outlet of each cooling channel are circumferentially offset from one another by a predetermined amount. Also by this measure, improved cooling is achieved by longer cooling channels.
  • the mutually facing annular surfaces of the ring elements are integrally connected to one another.
  • the ring elements can be fastened to one another by means of soldering, friction welding or the like. Such a cohesive connection prevents the cooling air from escaping from the cooling channels.
  • the present invention further provides a cooling ring with a plurality of cooling channels, each extending from one end face to the opposite end side through this, in particular a cooling ring of the type described above, which is characterized in that the inner peripheral surface and / or the outer peripheral surface of the cooling ring is provided with cooling fins.
  • a cooling ring with a plurality of cooling channels, each extending from one end face to the opposite end side through this, in particular a cooling ring of the type described above, which is characterized in that the inner peripheral surface and / or the outer peripheral surface of the cooling ring is provided with cooling fins.
  • the cooling fins each extend between the two end faces.
  • the cooling ribs extend in a straight line.
  • the cooling ribs extend meander-shaped, which leads to an extension of the cooling ribs and thus to a better cooling.
  • each fin is circumferentially offset by a predetermined amount from each other, which is also conducive to the achievement of improved cooling.
  • the present invention provides a gas turbine combustor having a combustion chamber, a cooling ring of the invention provided at the outlet end of the combustion chamber, and spring members disposed along the outer circumference of the cooling ring and projecting radially outward therefrom.
  • the cooling ring defines the outlet end of the combustion chamber and receives the spring elements.
  • the cooling ring may be welded to the free end of the combustion chamber and continue in one piece.
  • the cooling ring is arranged between the outer surface of the combustion chamber and the spring elements and receives the spring elements. In other words, in this variant, the cooling ring is thus pushed onto the free end of the combustion chamber.
  • the cooling ring is materially connected to the combustion chamber, such as by means of a soldered or welded connection.
  • FIG. 1 shows a portion of a gas turbine assembly in which a gas turbine combustor 1 is inserted into a housing 2 of the gas turbine assembly.
  • the gas turbine burner 1 is connected via a flange 3 with a connecting housing 4, which in turn is screwed to the housing 2.
  • the flange 3 can also be fastened directly to the housing 2, that is, the connection housing 4 can be dispensed with.
  • the gas turbine combustor 1 has a tubular combustion chamber 5, the outlet end of which is connected to an inlet end of a tubular transition channel 6 positioned in the housing 2 of the gas turbine arrangement, which is held on the housing 2 via an adjusting and fixing device 7.
  • a cooling ring 8 provided at the outlet end of the combustion chamber 5 is inserted into the inlet end of the transitional passage 6.
  • On the outer circumference of the cooling ring 8 are fixed and radially outwardly projecting spring elements 9 fixedly secured to the inner circumference of the inlet end of the transition channel 6 in such a clamping manner that the outlet end of the combustion chamber 5 is fixed within the inlet end of the transition channel 6.
  • On the outer circumference of the cooling ring 8, a stop ring 10 is further positioned, which has an L-shaped cross-section and limits the radial inward movement of the spring elements 9 during the insertion of the cooling ring 8 in the inlet end of the transition channel 6.
  • the cooling ring 8 is welded to the free end of the combustion chamber 5, so that the cooling ring 8 continues the combustion chamber 5 and defines the outlet end thereof.
  • the cooling ring 8 can also be pushed onto the free end of the combustion chamber 5 and fixed thereto, even if this is not shown here.
  • the cooling ring 8 comprises a plurality of cooling channels 11, each extending from one end face 12 to the opposite end face 13 of the cooling ring 8 therethrough.
  • the cooling ring 8 is formed from two concentric ring elements 14 and 15, the mutually facing annular surfaces touching each other and are materially connected to each other, for example by means of a solder or Reibsch waititati.
  • the cooling channels 11 are formed by grooves which are formed in the outer peripheral surface of the inner ring member 14, wherein the grooves are covered by the outer ring member 15.
  • the grooves may be incorporated in the inner surface of the outer ring member 15 or in both ring members 14 and 15, although not shown herein.
  • the production of the grooves can be done for example by using a milling cutter or the like.
  • the cooling channels 11 have a uniform rectangular cross-section.
  • the distances between the respective cooling channels 11 have a uniform dimension 2a, and the distances between the cooling channels 11 and the inner surface of the inner ring element 14 and the distances between the cooling channels 11 and the outer surface of the outer ring element 15 each have a uniform dimension a, which results in a very even cooling of the cooling ring 8 during operation.
  • the cooling channels 11 may extend between the two end faces 12 and 13 of the cooling ring 8 meandering.
  • Both in the Figures 5 and 6 illustrateddekanal losses advantagen are characterized in that their cooling channels 11 have a greater cooling channel length compared to extending in the axial direction of the cooling channels, which goes hand in hand with improved cooling.
  • FIG. 7 shows a cooling ring 16 according to a second embodiment of the present invention, instead of the cooling ring 8 shown in Figure 1 at the outlet end of the combustion chamber 5 can be arranged.
  • the cooling ring 16 includes cooling channels 17 in the form of bores, which extend in each case from one end face to the opposite end face of the cooling ring 16 in the axial direction therethrough.
  • cooling ribs 18 are provided on the outer circumference of the cooling ring, which point radially outward starting from the cooling channels 17.
  • the positions and the diameters of the cooling channels 17 as well as the positions and the shapes of the cooling fins 18 are each chosen such that the cooling channels 17 to the inner surface and the outer surface of the cooling ring 16 each have approximately a uniform dimension b.
  • the cooling ribs 18 extend in the present case in a straight line in the axial direction along the outer surface of the cooling ring 16 from one end face to the other. It should be understood, however, that the cooling fins 18 may alternatively be disposed on the inner surface of the cooling ring 16 and / or both on the outer surface and on the inner surface of the cooling ring 16.
  • gas turbine assembly are mixed in the gas turbine combustor 1 compressed ambient air and fuel, burned in the combustion chamber 5 and the thus generated combustion gases introduced into the transition channel 6 and fed through this turbine, where the combustion gases drive the rotor of the turbine.
  • the combustion chamber 5 is for the most part directly flowed against or flowed around by the compressor discharge air, so that sufficient cooling is present in the areas directly flowed around or flowed around.
  • the cooling ring 8, 16 provides the necessary cooling, the cooling channels 11, 17 of which are traversed by the compressor end air. Thanks to the arrangement and design of the cooling channels and / or cooling fins a very uniform and sufficient cooling is achieved with a simple and inexpensive to produceméringetter.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP14162655.6A 2014-03-31 2014-03-31 Anneau de refroidissement et brûleur de turbine à gaz dotée d'un tel anneau de refroidissement Withdrawn EP2927591A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP14162655.6A EP2927591A1 (fr) 2014-03-31 2014-03-31 Anneau de refroidissement et brûleur de turbine à gaz dotée d'un tel anneau de refroidissement

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP14162655.6A EP2927591A1 (fr) 2014-03-31 2014-03-31 Anneau de refroidissement et brûleur de turbine à gaz dotée d'un tel anneau de refroidissement

Publications (1)

Publication Number Publication Date
EP2927591A1 true EP2927591A1 (fr) 2015-10-07

Family

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Application Number Title Priority Date Filing Date
EP14162655.6A Withdrawn EP2927591A1 (fr) 2014-03-31 2014-03-31 Anneau de refroidissement et brûleur de turbine à gaz dotée d'un tel anneau de refroidissement

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EP (1) EP2927591A1 (fr)

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3915619A (en) * 1972-03-27 1975-10-28 Phillips Petroleum Co Gas turbine combustors and method of operation
WO2000032920A1 (fr) * 1998-11-27 2000-06-08 Volvo Aero Corporation Structure tuyere pour cols de tuyeres avec paroi refroidie
US6640538B1 (en) * 1999-01-18 2003-11-04 Astrium Gmbh Combustion chamber cooling structure for a rocket engine
FR2883929A1 (fr) * 2005-04-04 2006-10-06 United Technologies Corp Caracteristiques d'augmentation du transfert thermique pour une chambre a combustion a paroi tubulaire
DE102008062486A1 (de) * 2007-12-28 2009-07-02 Showa Denko K.K. Doppelwandrohr-Wärmetauscher
US20100064693A1 (en) * 2008-09-15 2010-03-18 Koenig Michael H Combustor assembly comprising a combustor device, a transition duct and a flow conditioner
US20110247341A1 (en) * 2010-04-09 2011-10-13 General Electric Company Combustor liner helical cooling apparatus
US20120247111A1 (en) * 2011-03-29 2012-10-04 Narcus Andrew R Turbine combustion system liner

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3915619A (en) * 1972-03-27 1975-10-28 Phillips Petroleum Co Gas turbine combustors and method of operation
WO2000032920A1 (fr) * 1998-11-27 2000-06-08 Volvo Aero Corporation Structure tuyere pour cols de tuyeres avec paroi refroidie
US6640538B1 (en) * 1999-01-18 2003-11-04 Astrium Gmbh Combustion chamber cooling structure for a rocket engine
FR2883929A1 (fr) * 2005-04-04 2006-10-06 United Technologies Corp Caracteristiques d'augmentation du transfert thermique pour une chambre a combustion a paroi tubulaire
DE102008062486A1 (de) * 2007-12-28 2009-07-02 Showa Denko K.K. Doppelwandrohr-Wärmetauscher
US20100064693A1 (en) * 2008-09-15 2010-03-18 Koenig Michael H Combustor assembly comprising a combustor device, a transition duct and a flow conditioner
US20110247341A1 (en) * 2010-04-09 2011-10-13 General Electric Company Combustor liner helical cooling apparatus
US20120247111A1 (en) * 2011-03-29 2012-10-04 Narcus Andrew R Turbine combustion system liner

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