EP1983265A2 - Paroi de chambre de combustion de turbine à gaz - Google Patents

Paroi de chambre de combustion de turbine à gaz Download PDF

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Publication number
EP1983265A2
EP1983265A2 EP08007322A EP08007322A EP1983265A2 EP 1983265 A2 EP1983265 A2 EP 1983265A2 EP 08007322 A EP08007322 A EP 08007322A EP 08007322 A EP08007322 A EP 08007322A EP 1983265 A2 EP1983265 A2 EP 1983265A2
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
gas turbine
chamber wall
wall according
cooling holes
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Withdrawn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
EP08007322A
Other languages
German (de)
English (en)
Other versions
EP1983265A3 (fr
Inventor
Miklós Dr.-Ing. Gerendás
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 EP1983265A2 publication Critical patent/EP1983265A2/fr
Publication of EP1983265A3 publication Critical patent/EP1983265A3/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Withdrawn legal-status Critical Current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
    • F23R3/002—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
    • F23R3/007—Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
    • 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
    • F23R2900/03041—Effusion cooled combustion chamber walls or domes
    • 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
    • F23R2900/03042—Film cooled combustion chamber walls or domes
    • 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
    • F23R2900/03044—Impingement cooled combustion chamber walls or subassemblies

Definitions

  • the invention relates to a gas turbine combustion chamber wall according to the features of the preamble of claim 1.
  • the GB 9 106 085 A and the WO 92/16798 A describe the construction of a gas turbine combustion chamber by metallic shingles fastened with studs, which by the combination of impingement and effusion cooling leads to a quite effective cooling effect and thus allows the reduction of the cooling air consumption.
  • the pressure loss which exists over the wall, distributed to two throttle points, the shingle support and the shingle itself. To avoid edge leakage usually the greater proportion of the pressure loss is generated via the shingle support, so that the cooling air has less cause, at the Pass the effusion shingles.
  • the GB 2 087 065 A describes an impingement cooling configuration with a shingled shingled piping, with each individual impingement cooling jet being protected from cross flow by an upstream pin or rib on the shingle. Furthermore, the pins or ribs increase the area available for heat transfer.
  • the GB 2 360 086 A describes an impingement cooling configuration with hexagonal ribs and partly additional prisms centrally located within the hexagonal ribs to increase the heat transfer.
  • the GB 9 106 085 A uses only a flat surface as the target of impingement cooling. An attachment of ribs would bring little except the simple increase in the area, since the ribs, such as in the GB 2 360 086 A are shown to require an overflow to take effect. Due to the congruence of impingement cooling air supply and removal of the air through the effusion bores, however, there is no appreciable Speed at the overflow of the ribs. In part, the pressure difference across the shingle is reduced by the torch swirl so that no effective flow through the effusion holes takes place more or even threatens hot gas burglary in the impingement cooling chamber of the shingle.
  • Film cooling is the most effective way to reduce the wall temperature, as the component is protected by the insulating cooling film from transferring heat from the hot gas, instead of removing heat that has already been injected by other methods afterwards.
  • the GB 2 087 065 A and the GB 2 360 086 A contain no technical teaching to renew the cooling film on the hot gas side within the extension of the shingle.
  • the shingle must be made so short in each case in the flow direction that the cooling film produced by the upstream shingle over the entire length of the shingle carries. This forces a multitude of shingles along the combustion chamber wall and does not allow to cover this distance with a single shingle.
  • the invention has for its object to provide a gas turbine combustor wall of the type mentioned, which with a simple structure and easier, cost-effective manufacturability which has a high cooling efficiency and good damping.
  • impact-cooled shingles having a surface structure, e.g. by hexagonal ribs or other polygonal shapes, wherein the spent air is discharged through effusion holes from the baffle cooling gap so that the extension of baffle cooling holes for air supply, and the effusion hole field for air discharge are not congruent.
  • the area which is equipped with a surface structure, can cover the entire shingle, or only an optimized area, in which a considerable overflow of the surface structure takes place and thus increases the noticeable heat transfer.
  • the displacement may be provided in the circumferential direction or in the axial direction or any combination thereof.
  • the hexagonal ribs may be filled with a prism so that the tip of the prism is at or above the level of the ribs.
  • the surface structure may be formed of tri-, tetra- or other polygonal cells.
  • the surface structure may also consist of circular or droplet-shaped depressions, whereby here too a shift between impact field field, surface structure area and effusion hole field in the axial or / and circumferential direction is decisive. If impingement cooling holes are present in the area of the surface structure, then the impact cooling jets strike the shingle substantially in the middle of the polygonal cell or at the lowest point of the circular or drop-shaped depression.
  • the impingement cooling holes may vary in diameter in the axial and / or circumferential direction, as well as the effusion holes and the dimensions of the surface structure.
  • the impingement cooling holes are substantially perpendicular to the impingement cooling surface, but the effusion holes are at a shallow angle to the hot gas side surface in the range of 10-45 degrees, advantageously in the range of 15-30 degrees.
  • the effusion holes may be purely axially aligned or form a circumferential angle.
  • the effusion hole pattern can be oriented on the surface structure.
  • a defined overflow of the ribs or depressions results in maximizing the rib effect while at the same time minimizing the impairment of the impingement cooling by the transverse flow.
  • the shingles temperature is lowered, thus extending the life of the component.
  • the Fig. 1 shows a schematic representation of a cross section of a gas turbine combustor according to the prior art.
  • compressor outlet blades 1 and a combustion chamber outer housing 2 and a combustion chamber inner housing 3 are shown schematically.
  • the reference numeral 4 denotes a burner with arm and head
  • the reference numeral 5 denotes a combustion chamber head, which is followed by a multilayer combustion chamber wall 6, from which the flow is directed to turbine inlet blades 7.
  • the Fig. 2 shows an embodiment according to the prior art, as for example from the WO 92/16798 A is already known.
  • a combustion chamber wall 9 (shingle support) is shown, in which a plurality of inflow bores 8 (impingement cooling holes) are formed, through which cooling air from the compressor exit air 12 is introduced into a gap 14 between a shingle 10 and the combustion chamber wall 9.
  • the shingle 10 is secured by stud bolts 15 and fastening nuts 16.
  • the shingle comprises several effusion cooling holes 11.
  • the Fig. 3 shows a first embodiment of the combustion chamber wall according to the invention.
  • FIG. 3 shows, in a schematic plan view, the displacement of the region 17 of the impingement cooling holes 8 and the region 18 of the effusion cooling holes 11 and 23. It can be seen that between the regions 17 and 18 with a partial overlap the region of the surface structure 20 is arranged, wherein the individual elements of FIG Surface structure are indicated schematically by the reference numeral 22.
  • the Fig. 5 shows a further modification in an analogous view Fig. 4 with only partially overlapping areas (area 17 for the impingement cooling holes 8, area 18 for the Effusion cooling holes 11 and area 20 for the surface structure 22).
  • Reference numeral 21 schematically shows the projection of an impingement cooling hole 8 in the combustion chamber wall 9 (shingle support) onto the shingle 10.
  • the Fig. 6 shows a schematic side view (cross section) of different embodiments of the surface structure 19, 22.
  • a rib 24 is provided with a rectangular cross-section and a rib 25 with a trapezoidal cross-section.
  • the surface structure 19 may comprise circular depressions 26 as well as drop-shaped depressions 27 (see also FIGS Fig. 7 ).
  • Reference numeral 30 schematically represents a prismatic elevation (prism). The prism may be lower than the ribs 24, 25, higher than the ribs 24, 25 or the same height as the ribs 24, 25.
  • the Fig. 7 shows a schematic plan view, analog Fig. 6 , a further embodiment variant, from which square cells 28 and hexagonal cells 29 result, which may also be provided with a prism 30.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Ceramic Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP08007322A 2007-04-17 2008-04-14 Paroi de chambre de combustion de turbine à gaz Withdrawn EP1983265A3 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DE102007018061A DE102007018061A1 (de) 2007-04-17 2007-04-17 Gasturbinenbrennkammerwand

Publications (2)

Publication Number Publication Date
EP1983265A2 true EP1983265A2 (fr) 2008-10-22
EP1983265A3 EP1983265A3 (fr) 2011-04-27

Family

ID=39522222

Family Applications (1)

Application Number Title Priority Date Filing Date
EP08007322A Withdrawn EP1983265A3 (fr) 2007-04-17 2008-04-14 Paroi de chambre de combustion de turbine à gaz

Country Status (3)

Country Link
US (1) US8099961B2 (fr)
EP (1) EP1983265A3 (fr)
DE (1) DE102007018061A1 (fr)

Cited By (11)

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Publication number Priority date Publication date Assignee Title
CN101858256A (zh) * 2009-04-13 2010-10-13 通用电气公司 组合型对流/泻流冷却的一件式筒形燃烧器
EP2489836A1 (fr) * 2011-02-21 2012-08-22 Karlsruher Institut für Technologie Composant pouvant être refroidi
EP2685170A1 (fr) 2012-07-10 2014-01-15 Alstom Technology Ltd Structure de paroi refroidie pour les parties chaudes d'une turbine à gaz et procédé de fabrication d'une telle structure
WO2014055887A2 (fr) 2012-10-04 2014-04-10 United Technologies Corporation Chemise de chambre de combustion de turbine à gaz
WO2014105269A2 (fr) 2012-12-19 2014-07-03 United Technologies Corporation Diffuseur pour système d'alimentation
DE102013003444A1 (de) * 2013-02-26 2014-09-11 Rolls-Royce Deutschland Ltd & Co Kg Prall-effusionsgekühlte Schindel einer Gasturbinenbrennkammer mit verlängerten Effusionsbohrungen
WO2014201249A1 (fr) 2013-06-14 2014-12-18 United Technologies Corporation Panneau de chemisage de chambre de combustion à géométrie ondulée de moteur à turbine à gaz
WO2015065579A1 (fr) 2013-11-04 2015-05-07 United Technologies Corporation Ensemble paroi pour moteur à turbine à gaz comprenant un rail décalé
EP3009744A1 (fr) * 2014-10-13 2016-04-20 Rolls-Royce plc Élément de garniture pour une chambre de combustion et procédé associé
EP2384392B1 (fr) 2009-01-30 2017-05-31 Ansaldo Energia IP UK Limited Élément structural refroidi pour turbine à gaz
US10408452B2 (en) 2015-10-16 2019-09-10 Rolls-Royce Plc Array of effusion holes in a dual wall combustor

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DE102008026463A1 (de) * 2008-06-03 2009-12-10 E.On Ruhrgas Ag Verbrennungseinrichtung für eine Gasturbinenanlage
US20100037620A1 (en) * 2008-08-15 2010-02-18 General Electric Company, Schenectady Impingement and effusion cooled combustor component
US8438856B2 (en) 2009-03-02 2013-05-14 General Electric Company Effusion cooled one-piece can combustor
CH703657A1 (de) * 2010-08-27 2012-02-29 Alstom Technology Ltd Verfahren zum betrieb einer brenneranordnung sowie brenneranordnung zur durchführung des verfahrens.
FR2972027B1 (fr) * 2011-02-25 2013-03-29 Snecma Chambre annulaire de combustion de turbomachine comprenant des orifices de dilution ameliores
JP5696566B2 (ja) * 2011-03-31 2015-04-08 株式会社Ihi ガスタービンエンジン用燃焼器及びガスタービンエンジン
DE102011007562A1 (de) * 2011-04-18 2012-10-18 Man Diesel & Turbo Se Brennkammergehäuse und damit ausgerüstete Gasturbine
EP2559854A1 (fr) * 2011-08-18 2013-02-20 Siemens Aktiengesellschaft Composant refroidissable de l'intérieur pour une turbine à gaz dotée d'au moins un canal de refroidissement
GB201116608D0 (en) * 2011-09-27 2011-11-09 Rolls Royce Plc A method of operating a combustion chamber
DE102011114928A1 (de) * 2011-10-06 2013-04-11 Lufthansa Technik Ag Brennkammer für eine Gasturbine
JP5821550B2 (ja) * 2011-11-10 2015-11-24 株式会社Ihi 燃焼器ライナ
US9151173B2 (en) * 2011-12-15 2015-10-06 General Electric Company Use of multi-faceted impingement openings for increasing heat transfer characteristics on gas turbine components
US20130180252A1 (en) * 2012-01-18 2013-07-18 General Electric Company Combustor assembly with impingement sleeve holes and turbulators
US9052111B2 (en) 2012-06-22 2015-06-09 United Technologies Corporation Turbine engine combustor wall with non-uniform distribution of effusion apertures
DE102012016493A1 (de) 2012-08-21 2014-02-27 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammer mit prallgekühlten Bolzen der Brennkammerschindeln
US9869279B2 (en) * 2012-11-02 2018-01-16 General Electric Company System and method for a multi-wall turbine combustor
DE102012025375A1 (de) * 2012-12-27 2014-07-17 Rolls-Royce Deutschland Ltd & Co Kg Verfahren zur Anordnung von Prallkühllöchern und Effusionslöchern in einer Brennkammerwand einer Gasturbine
US9939154B2 (en) * 2013-02-14 2018-04-10 United Technologies Corporation Combustor liners with U-shaped cooling channels
CA2904200A1 (fr) 2013-03-05 2014-09-12 Rolls-Royce Corporation Tuile de chambre de combustion a effusion, convexion, impact a double paroi
US20160370008A1 (en) * 2013-06-14 2016-12-22 United Technologies Corporation Conductive panel surface cooling augmentation for gas turbine engine combustor
EP3039347B1 (fr) * 2013-08-30 2019-10-23 United Technologies Corporation Ensemble paroi de turbine à gaz doté de zones de contour d'enveloppe de support
US20160169515A1 (en) * 2013-09-10 2016-06-16 United Technologies Corporation Edge cooling for combustor panels
US9644843B2 (en) * 2013-10-08 2017-05-09 Pratt & Whitney Canada Corp. Combustor heat-shield cooling via integrated channel
US10690348B2 (en) * 2013-11-04 2020-06-23 Raytheon Technologies Corporation Turbine engine combustor heat shield with one or more cooling elements
DE102013223258A1 (de) * 2013-11-14 2015-06-03 Rolls-Royce Deutschland Ltd & Co Kg Brennkammerhitzeabschirmelement einer Gasturbine
WO2015077755A1 (fr) * 2013-11-25 2015-05-28 United Technologies Corporation Structure à multiples parois refroidie par film ayant une ou plusieurs indentations
US10344979B2 (en) * 2014-01-30 2019-07-09 United Technologies Corporation Cooling flow for leading panel in a gas turbine engine combustor
WO2015117137A1 (fr) * 2014-02-03 2015-08-06 United Technologies Corporation Refroidissement par film d'air d'une paroi de chambre de combustion d'un moteur à turbine
US9410702B2 (en) 2014-02-10 2016-08-09 Honeywell International Inc. Gas turbine engine combustors with effusion and impingement cooling and methods for manufacturing the same using additive manufacturing techniques
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US10746403B2 (en) * 2014-12-12 2020-08-18 Raytheon Technologies Corporation Cooled wall assembly for a combustor and method of design
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US10876730B2 (en) * 2016-02-25 2020-12-29 Pratt & Whitney Canada Corp. Combustor primary zone cooling flow scheme
DE102016224632A1 (de) * 2016-12-09 2018-06-14 Rolls-Royce Deutschland Ltd & Co Kg Plattenförmiges Bauteil einer Gasturbine sowie Verfahren zu dessen Herstellung
US10731562B2 (en) 2017-07-17 2020-08-04 Raytheon Technologies Corporation Combustor panel standoffs with cooling holes
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US10823414B2 (en) * 2018-03-19 2020-11-03 Raytheon Technologies Corporation Hooded entrance to effusion holes
US11306918B2 (en) * 2018-11-02 2022-04-19 Chromalloy Gas Turbine Llc Turbulator geometry for a combustion liner
US11415320B2 (en) 2019-01-04 2022-08-16 Raytheon Technologies Corporation Combustor cooling panel with flow guide
JP7234006B2 (ja) * 2019-03-29 2023-03-07 三菱重工業株式会社 高温部品及び高温部品の製造方法
US11112114B2 (en) * 2019-07-23 2021-09-07 Raytheon Technologies Corporation Combustor panels for gas turbine engines

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Cited By (19)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2384392B1 (fr) 2009-01-30 2017-05-31 Ansaldo Energia IP UK Limited Élément structural refroidi pour turbine à gaz
CN101858256A (zh) * 2009-04-13 2010-10-13 通用电气公司 组合型对流/泻流冷却的一件式筒形燃烧器
EP2489836A1 (fr) * 2011-02-21 2012-08-22 Karlsruher Institut für Technologie Composant pouvant être refroidi
EP2685170A1 (fr) 2012-07-10 2014-01-15 Alstom Technology Ltd Structure de paroi refroidie pour les parties chaudes d'une turbine à gaz et procédé de fabrication d'une telle structure
EP2904236A4 (fr) * 2012-10-04 2015-12-09 United Technologies Corp Chemise de chambre de combustion de turbine à gaz
WO2014055887A2 (fr) 2012-10-04 2014-04-10 United Technologies Corporation Chemise de chambre de combustion de turbine à gaz
WO2014105269A2 (fr) 2012-12-19 2014-07-03 United Technologies Corporation Diffuseur pour système d'alimentation
US9476429B2 (en) 2012-12-19 2016-10-25 United Technologies Corporation Flow feed diffuser
EP2935868A4 (fr) * 2012-12-19 2016-04-20 United Technologies Corp Diffuseur pour système d'alimentation
EP2770260A3 (fr) * 2013-02-26 2015-09-30 Rolls-Royce Deutschland Ltd & Co KG Bardeau à refroidissement par impact effusion d'une chambre de combustion de turbine à gaz comprenant des perçages d'effusion prolongés
US9518738B2 (en) 2013-02-26 2016-12-13 Rolls-Royce Deutschland Ltd & Co Kg Impingement-effusion cooled tile of a gas-turbine combustion chamber with elongated effusion holes
DE102013003444A1 (de) * 2013-02-26 2014-09-11 Rolls-Royce Deutschland Ltd & Co Kg Prall-effusionsgekühlte Schindel einer Gasturbinenbrennkammer mit verlängerten Effusionsbohrungen
EP3008392A4 (fr) * 2013-06-14 2016-07-20 United Technologies Corp Panneau de chemisage de chambre de combustion à géométrie ondulée de moteur à turbine à gaz
WO2014201249A1 (fr) 2013-06-14 2014-12-18 United Technologies Corporation Panneau de chemisage de chambre de combustion à géométrie ondulée de moteur à turbine à gaz
WO2015065579A1 (fr) 2013-11-04 2015-05-07 United Technologies Corporation Ensemble paroi pour moteur à turbine à gaz comprenant un rail décalé
EP3066390A4 (fr) * 2013-11-04 2016-11-23 United Technologies Corp Ensemble paroi pour moteur à turbine à gaz comprenant un rail décalé
EP3009744A1 (fr) * 2014-10-13 2016-04-20 Rolls-Royce plc Élément de garniture pour une chambre de combustion et procédé associé
US10451277B2 (en) 2014-10-13 2019-10-22 Rolls-Royce Plc Liner element for a combustor, and a related method
US10408452B2 (en) 2015-10-16 2019-09-10 Rolls-Royce Plc Array of effusion holes in a dual wall combustor

Also Published As

Publication number Publication date
EP1983265A3 (fr) 2011-04-27
US8099961B2 (en) 2012-01-24
US20080264065A1 (en) 2008-10-30
DE102007018061A1 (de) 2008-10-23

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