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EP2921778A1 - Chambre de combustion d'une turbine à gaz - Google Patents

Chambre de combustion d'une turbine à gaz Download PDF

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Publication number
EP2921778A1
EP2921778A1 EP15158426.5A EP15158426A EP2921778A1 EP 2921778 A1 EP2921778 A1 EP 2921778A1 EP 15158426 A EP15158426 A EP 15158426A EP 2921778 A1 EP2921778 A1 EP 2921778A1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
groove
shingle
base plate
wall
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Withdrawn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
EP15158426.5A
Other languages
German (de)
English (en)
Inventor
Carsten Dr.-Ing. Clemen
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.)
Rolls Royce Deutschland Ltd and Co KG
Original Assignee
Rolls Royce Deutschland Ltd and Co KG
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 Rolls Royce Deutschland Ltd and Co KG filed Critical Rolls Royce Deutschland Ltd and Co KG
Publication of EP2921778A1 publication Critical patent/EP2921778A1/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Withdrawn legal-status Critical Current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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • 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
    • F23R3/60—Support structures; Attaching or mounting means
    • 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

Definitions

  • the invention relates to a combustion chamber of a gas turbine with an outer combustion chamber wall and at least one shingle mounted thereon, and with a base plate.
  • combustion chamber shingles are fastened to a supporting structure of the combustion chamber outer wall, also referred to as a liner.
  • the combustion chamber shingles have a large number of effusion cooling holes on the side facing the combustion chamber. These effusion cooling holes serve to cool the shingle from the high temperatures in the combustion chamber.
  • the combustion chamber shingle has at least one mixing air hole, through which air is passed from the space surrounding the combustion chamber (annular channel / annulus) into the combustion chamber in order to cool the combustion gases and to abase the combustion. In particular, this results in a reduction of NOx formation in the combustion chamber.
  • the shingles are often provided with a ceramic coating, which acts as an insulating layer against the high temperatures in the combustion chamber.
  • the invention has for its object to provide a combustion chamber of a gas turbine, which avoids the disadvantages of the prior art with a simple structure and simple, cost manufacturability, in particular by means of additive manufacturing process and enables a reliable construction.
  • the shingle extends over the entire length of the combustion chamber and is mounted at its front and at its rear end in each case in a groove.
  • the groove is formed at the front end on the base plate of the combustion chamber, while the groove is provided at the rear end to the outer combustion chamber wall.
  • the inventive solution it is thus possible to dispense completely with the threaded bolt. Rather, the shingle is stored as an entire component only at its front and at its rear end. This allows a simple, inexpensive manufacture of the shingle.
  • the production of the combustion chamber according to the invention can be carried out in a simple manner, that the shingle is inserted with its rear end portion in the groove of the outer combustion chamber wall. Subsequently, the outer combustion chamber wall is mounted together with the shingles on the base plate by the front end portion of the shingle is inserted into the groove of the base plate. Subsequently, the outer combustion chamber wall is welded to the base plate.
  • the shingle according to the invention can be produced inexpensively by means of additive processes in vertical production.
  • the shingle thus has at its front and its rear end in each case a spring, which is inserted into the respective groove.
  • Both the spring and the groove may extend over the entire circumference or be formed segmented.
  • the gas turbine engine 110 is a generalized example of a turbomachine, in which the invention can be applied.
  • the engine 110 is formed in a conventional manner and comprises in succession an air inlet 111, a fan 112 circulating in a housing, a medium pressure compressor 113, a high pressure compressor 114, a combustion chamber 115, a high pressure turbine 116, a medium pressure turbine 117 and a low pressure turbine 118 and a Exhaust nozzle 119, which are all arranged around a central engine center axis 101.
  • the intermediate pressure compressor 113 and the high pressure compressor 114 each include a plurality of stages, each of which includes a circumferentially extending array of fixed stationary vanes 120, commonly referred to as stator vanes, that radially inwardly from the engine casing 121 into an annular flow passage through the compressors 113, 114 protrude.
  • the compressors further include an array of compressor blades 122 projecting radially outward from a rotatable drum or disc 125 coupled to hubs 126 of high pressure turbine 116 and intermediate pressure turbine 117, respectively.
  • the turbine sections 116, 117, 118 have similar stages, comprising an array of fixed vanes 123 projecting radially inward from the housing 121 into the annular flow passage through the turbines 116, 117, 118, and a downstream array of turbine blades 124 projecting outwardly from a rotatable hub 126.
  • the compressor drum or compressor disk 125 and the vanes 122 disposed thereon and the turbine rotor hub 126 and the Turbine blades 124 disposed thereon rotate about the engine centerline 101 during operation.
  • the Fig. 2 shows a simplified enlarged view of a known from the prior art combustion chamber 1.
  • This includes a heat shield 2, a combustion chamber head 3 and a combustion chamber seal 4.
  • an outer combustion chamber wall 9 is provided, in which Zumischlöcher 5 are formed. The presentation of effusion holes and impingement cooling holes has been omitted for clarity.
  • the combustion chamber wall 9 is supported by means of combustion chamber suspensions 10 and combustion chamber flanges 11, as known from the prior art.
  • shingles 8 are arranged, which are integrally provided with bolts 6 and by means of nuts 7, which pass through holes in the combustion chamber wall 9, are secured.
  • the combustion chamber wall 9 is connected at its front end region with a base plate 12, usually welded.
  • Fig. 3 and 4 show the inventive design of the combustion chamber. The same parts are provided with the same reference numerals.
  • the base plate 12 connected to the combustion chamber head 3 has a groove 15 at its circumference.
  • a spring 16 can be inserted, which is formed at the front end of the shingle 8.
  • the shingle 8 extends over the entire length of the combustion chamber and also has a spring 16 at its rear end. This is also inserted into a groove 15, which is formed at the rear end portion of the combustion chamber wall 9.
  • cooling air is introduced through impingement cooling holes 19.
  • the cooling air flows through effusion holes 20 through the shingle 8 to cool it.
  • the Fig. 3 shows a state in which the shingle 8 is inserted by means of its rear spring 16 in the groove of the combustion chamber wall 9. To secure a temporary locking pin 22 may serve. The combustion chamber wall 9 is then pushed together with the shingle 8 on the base plate 12. Then the combustion chamber wall 9 can be welded to the combustion chamber head 3. Thus, the results in Fig. 4 shown completed condition.
  • the reference numeral 13 shows a weld 13 between the combustion chamber head 3 and the combustion chamber wall 9. The weld 13 is attached to a in Fig. 3 shown welding surface 14 is formed.
  • FIGS. 5 and 6 This is, as mentioned, inserted by means of its spring 16 in the groove 15 of the combustion chamber wall 9. For cooling this area additional cooling holes 21 may be provided.

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)
EP15158426.5A 2014-03-11 2015-03-10 Chambre de combustion d'une turbine à gaz Withdrawn EP2921778A1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DE102014204466.9A DE102014204466A1 (de) 2014-03-11 2014-03-11 Brennkammer einer Gasturbine

Publications (1)

Publication Number Publication Date
EP2921778A1 true EP2921778A1 (fr) 2015-09-23

Family

ID=52633149

Family Applications (1)

Application Number Title Priority Date Filing Date
EP15158426.5A Withdrawn EP2921778A1 (fr) 2014-03-11 2015-03-10 Chambre de combustion d'une turbine à gaz

Country Status (3)

Country Link
US (1) US9447973B2 (fr)
EP (1) EP2921778A1 (fr)
DE (1) DE102014204466A1 (fr)

Families Citing this family (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102013007443A1 (de) * 2013-04-30 2014-10-30 Rolls-Royce Deutschland Ltd & Co Kg Brennerdichtung für Gasturbinen-Brennkammerkopf und Hitzeschild
EP3236155B1 (fr) 2016-04-22 2020-05-06 Rolls-Royce plc Chambre de combustion à paroi segmentée
EP3306199B1 (fr) * 2016-10-06 2020-12-30 Ansaldo Energia Switzerland AG Dispositif de combustion pour un moteur à turbine à gaz et moteur de turbine à gaz intégrant ledit dispositif de combustion
US10393381B2 (en) 2017-01-27 2019-08-27 General Electric Company Unitary flow path structure
US10253643B2 (en) 2017-02-07 2019-04-09 General Electric Company Airfoil fluid curtain to mitigate or prevent flow path leakage
US10385776B2 (en) * 2017-02-23 2019-08-20 General Electric Company Methods for assembling a unitary flow path structure
US10378373B2 (en) 2017-02-23 2019-08-13 General Electric Company Flow path assembly with airfoils inserted through flow path boundary
US10247019B2 (en) 2017-02-23 2019-04-02 General Electric Company Methods and features for positioning a flow path inner boundary within a flow path assembly
US10253641B2 (en) 2017-02-23 2019-04-09 General Electric Company Methods and assemblies for attaching airfoils within a flow path
US10385731B2 (en) 2017-06-12 2019-08-20 General Electric Company CTE matching hanger support for CMC structures
US11402100B2 (en) * 2018-11-15 2022-08-02 Pratt & Whitney Canada Corp. Ring assembly for double-skin combustor liner

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2263733A (en) * 1992-01-28 1993-08-04 Snecma Turbomachine with removable combustion chamber.
EP0927992A1 (fr) 1997-07-17 1999-07-07 Sony Corporation Support d'enregistrement magnetique et dispositif d'enregistrement/de reproduction magnetique le comprenant
US6145319A (en) 1998-07-16 2000-11-14 General Electric Company Transitional multihole combustion liner
US6314739B1 (en) * 2000-01-13 2001-11-13 General Electric Company Brazeless combustor dome assembly
DE10214570A1 (de) 2002-04-02 2004-01-15 Rolls-Royce Deutschland Ltd & Co Kg Mischluftloch in Gasturbinenbrennkammer mit Brennkammerschindeln
EP1491823A1 (fr) * 2003-06-27 2004-12-29 General Electric Company Chambre de combustion pour une turbine à gaz montée en raccord d'angle
EP2604926A1 (fr) * 2011-12-16 2013-06-19 General Electric Company Système permettant d'intégrer des chicanes pour un refroidissement amélioré de chemises CMC

Family Cites Families (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3031844A (en) * 1960-08-12 1962-05-01 William A Tomolonius Split combustion liner
US4628694A (en) * 1983-12-19 1986-12-16 General Electric Company Fabricated liner article and method
US4686823A (en) * 1986-04-28 1987-08-18 United Technologies Corporation Sliding joint for an annular combustor
JP2597800B2 (ja) * 1992-06-12 1997-04-09 ゼネラル・エレクトリック・カンパニイ ガスタービンエンジン用燃焼器
US5291732A (en) * 1993-02-08 1994-03-08 General Electric Company Combustor liner support assembly
US6397603B1 (en) * 2000-05-05 2002-06-04 The United States Of America As Represented By The Secretary Of The Air Force Conbustor having a ceramic matrix composite liner
US6401447B1 (en) * 2000-11-08 2002-06-11 Allison Advanced Development Company Combustor apparatus for a gas turbine engine
US6513330B1 (en) * 2000-11-08 2003-02-04 Allison Advanced Development Company Diffuser for a gas turbine engine
JP2004524479A (ja) * 2001-04-27 2004-08-12 シーメンス　アクチエンゲゼルシヤフト 特にガスタービンの燃焼室
EP1486730A1 (fr) * 2003-06-11 2004-12-15 Siemens Aktiengesellschaft Elément de bouclier thermique
WO2005108869A1 (fr) * 2004-05-05 2005-11-17 Alstom Technology Ltd Chambre de combustion pour une turbine a gaz
GB2432902B (en) * 2005-12-03 2011-01-12 Alstom Technology Ltd Gas turbine sub-assemblies
US20090090110A1 (en) * 2007-10-04 2009-04-09 Honeywell International, Inc. Faceted dome assemblies for gas turbine engine combustors

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2263733A (en) * 1992-01-28 1993-08-04 Snecma Turbomachine with removable combustion chamber.
EP0927992A1 (fr) 1997-07-17 1999-07-07 Sony Corporation Support d'enregistrement magnetique et dispositif d'enregistrement/de reproduction magnetique le comprenant
US6145319A (en) 1998-07-16 2000-11-14 General Electric Company Transitional multihole combustion liner
US6314739B1 (en) * 2000-01-13 2001-11-13 General Electric Company Brazeless combustor dome assembly
DE10214570A1 (de) 2002-04-02 2004-01-15 Rolls-Royce Deutschland Ltd & Co Kg Mischluftloch in Gasturbinenbrennkammer mit Brennkammerschindeln
EP1491823A1 (fr) * 2003-06-27 2004-12-29 General Electric Company Chambre de combustion pour une turbine à gaz montée en raccord d'angle
EP2604926A1 (fr) * 2011-12-16 2013-06-19 General Electric Company Système permettant d'intégrer des chicanes pour un refroidissement amélioré de chemises CMC

Also Published As

Publication number Publication date
DE102014204466A1 (de) 2015-10-01
US20150260405A1 (en) 2015-09-17
US9447973B2 (en) 2016-09-20

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