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WO2010112360A1 - Elément refroidi pour une turbine à gaz - Google Patents

Elément refroidi pour une turbine à gaz Download PDF

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Publication number
WO2010112360A1
WO2010112360A1 PCT/EP2010/053691 EP2010053691W WO2010112360A1 WO 2010112360 A1 WO2010112360 A1 WO 2010112360A1 EP 2010053691 W EP2010053691 W EP 2010053691W WO 2010112360 A1 WO2010112360 A1 WO 2010112360A1
Authority
WO
WIPO (PCT)
Prior art keywords
cooling
impingement
impingement cooling
wall
blade
Prior art date
Application number
PCT/EP2010/053691
Other languages
German (de)
English (en)
Inventor
Jörg KRÜCKELS
Tanguy Arzel
Jose Anguisola Mcfeat
Martin Schnieder
Original Assignee
Alstom Technology Ltd
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
Family has litigation
First worldwide family litigation filed litigation Critical https://patents.darts-ip.com/?family=40627674&utm_source=google_patent&utm_medium=platform_link&utm_campaign=public_patent_search&patent=WO2010112360(A1) "Global patent litigation dataset” by Darts-ip is licensed under a Creative Commons Attribution 4.0 International License.
Application filed by Alstom Technology Ltd filed Critical Alstom Technology Ltd
Priority to EP10709734.7A priority Critical patent/EP2414639B8/fr
Publication of WO2010112360A1 publication Critical patent/WO2010112360A1/fr
Priority to US13/247,429 priority patent/US20120063891A1/en

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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • 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
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving blades
    • F05D2240/81—Cooled platforms
    • 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
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • 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
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • 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
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/205—Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes

Definitions

  • the present invention relates to the field of gas turbines. It relates to a cooled component for a gas turbine according to the preamble of claim 1.
  • Blades, vanes, heat shields or other exposed to the hot gas flow of a gas turbine components must be intensively cooled in order to cope with the thermal and mechanical stresses occurring in the machine during operation.
  • cooling is film cooling, in which through openings in the wall of the loaded component, a cooling medium, usually compressed air from the compressor part of the gas turbine, exits into the hot gas channel and forms a cooling film on the hot gas facing surface of the component.
  • impingement cooling in which a pressurized cooling medium flows through an impeller cooling plate provided with distributed openings, and the resulting radiation on the spaced inner wall of the loaded component meet, and wherein the cooling medium at this impact receives intense heat from the wall and dissipates.
  • the effectiveness of the impingement cooling depends in particular on the type and distribution of the openings provided for this purpose in the impingement cooling plate, on the distance of the impingement cooling plate from the wall to be cooled, on the selected one Fortströmungstechnik the cooling medium after impact, so the outflow of the cooling medium after performing cooling work, and in general from the difference of prevailing before and behind the impingement cooling plate pressures of the cooling medium. If, however, the cooling medium flows out of the chamber formed between the impingement cooling plate and the wall to be cooled through channels in the wall into the hot gas flow, the pressure difference is decisively influenced by the static pressure prevailing in the hot gas duct at this point.
  • Fig. 1 shows a section of a gas turbine component 20 with a body 21 which is thermally highly loaded on the outside (the lower side in Fig. 1).
  • a baffle cooling chamber 23 is formed which is closed by a baffle cooling plate 22 spaced from the inner wall.
  • the impingement cooling plate 22 has a plurality of distributed impingement cooling holes 25 through which a cooling medium radially enters the impingement cooling chamber 23 and meets the inner wall of the body 21 (arrows in FIG. 1). After the cooling medium has absorbed heat from there when it hits the wall, it flows outward into the hot-gas channel via one (or more) cooling hole (s).
  • impingement cooling chambers have no connection to the hot gas duct.
  • a cooled gas turbine blade is known in which the inner platform is cooled by impingement cooling.
  • the impingement cooling area is divided by a rib into two zones to reduce cross-flows affecting the cooling and localized To reduce heat transfer coefficients.
  • film cooling holes are arranged through which the cooling air exits both zones into the hot gas channel.
  • the invention aims to remedy this situation. It is therefore an object of the invention to provide a cooled component for a gas turbine, which avoids the disadvantages of known solutions and is particularly characterized in that a local thermal loads and static pressure conditions optimally adapted, uniform cooling of the loaded surface is achieved.
  • Film cooling holes are in communication with the hot gas channel, and that the density and / or the distribution and / or the diameter of the impingement cooling holes resp. the flow cross section of the openings in the impingement cooling plates of the individual impingement cooling chambers is adapted to the respective thermal load and / or the respective static pressure prevailing in operation on the outside of the wall.
  • An embodiment of the invention is characterized in that the component is a blade provided with a platform of the gas turbine, in particular a guide vane, and the impingement-cooled wall is a wall of the platform.
  • Another embodiment of the invention is characterized in that the platform is involved in a sequential cooling of the blade, in which the platform and the blade of the blade are flowed through by the same cooling medium.
  • a further embodiment is characterized in that the blade comprises an airfoil having a leading edge and a trailing edge, and that an impingement cooling chamber disposed downstream of the trailing edge is configured with its impingement cooling plate for increased cooling due to the increased thermal stress imposed by the trailing edge training wake vortex is tuned.
  • the impingement cooling chamber located downstream of the trailing edge may also be in a heat shield.
  • a further embodiment of the invention is characterized in that the component is a heat shield, where the impingement cooling is arranged in different chambers.
  • FIG. 1 shows a section through a conventionally cooled gas turbine component with a continuous impingement cooling chamber, which is covered by a uniform impingement cooling plate.
  • Fig. 2 is a perspective view of a blade platform with several separate, parallel-working impingement cooling chambers according to an embodiment of the invention and
  • FIG. 3 shows in a simplified representation the section in the plane III-III through the blade provided with impingement cooling blades according to FIG. 2.
  • FIGS. 2 and 3 show an exemplary embodiment of a baffled component in the form of a blade 10 of a gas turbine.
  • the blade 10 has an extending in the blade longitudinal direction blade 15, to which at one end a transverse platform 11 connects.
  • FIG. 2 shows the view from the rear onto the platform 11, in FIG. 3 the section through the platform 11 along the line III-III in FIG. 2.
  • the platform 11 has a wall 28 oriented essentially perpendicular to the blade longitudinal direction , with which it adjoins the hot gas channel 27 of the gas turbine. In the hot gas channel 27, the hot gas acts on the rotor and guide blades arranged there, as indicated in FIG. 3 by the horizontal arrows.
  • the airfoil 15 of the blade 10 is bounded upstream by a leading edge 16 and downstream by a trailing edge 17. It has a convex curved suction side and a concave curved pressure side, which are not directly visible in the figures.
  • a cooling medium in particular cooling air
  • the thermal loading of the wall 28 of the platform 11 on the outer side facing the hot gas channel 27 is also locally different. If the local temperature is high, the thermal load is usually high, and vice versa. In addition, the local flow conditions play a role, because the heat transfer between the hot gas and the wall depends on whether the local flow is laminar or turbulent, or whether there is even a rest zone at the local place.
  • the present solution therefore proposes correspondingly different cooling of thermally differently loaded areas of the wall 28 in order to achieve the most uniform possible cooling or temperature distribution of the platform 1 1 and on the other hand to consume as little cooling air as the cooling air consumption in the efficiency of the machine received.
  • the impingement cooling on the rear side of the wall 28 is specifically subdivided into different regions, each having its own impingement cooling chamber 13 or 13a, 13b, the cooling-technical configuration of these impingement cooling chambers reflecting the respective load ratio.
  • Each of the partition walls 14 and webs separated from each other, working in parallel impingement cooling chambers 13a, 13b is associated with its own baffle cooling plate 19a, 19b, in the baffle openings 18a, 18b of different numbers and / or distribution and / or different diameters, respectively.
  • Flow cross-section are provided.
  • the impingement cooling plates 19a, 19b may be individual separate sheets; but they can also be designed so that different areas are detected by a large common baffle cooling plate, so that this baffle cooling plate then extends over all of the impact cooling chambers 13a, 13b.
  • the exit of the heated cooling air into the hot gas duct 27 takes place separately for each impingement cooling chamber 13a, 13b via its own film cooling bores 26a, 26b.
  • the heated cooling air reaches the outside of the wall 28 and forms there a film that cools and protects against the hot gas in the hot gas channel 27.
  • the respective impact cooling chamber 13a, 13b simultaneously couples to the one in there
  • Hot gas channel 27 prevailing static pressure.
  • the static pressure p a in the hot gas duct 27 is relevant for the impingement cooling chamber 13 a, while the static pressure p b in the hot gas duct 27 is of importance for the impingement cooling chamber 13 b .
  • the pressure on the other side of the wall 28 can be lowered to limit the cooling air mass flow through this chamber. This means at the same time that the pressure drop across the associated baffle cooling plate is greater and thus a higher heat transfer coefficient can be achieved in the impingement cooling.
  • the local static pressure in the hot gas passage 27 for an impingement cooling chamber is high, the pressure on the other side of the wall 28 must be made higher to prevent the hot gas from entering the chamber.
  • the platform 1 1 can be integrated with advantage in a sequential cooling of the blade 10.
  • the cooling air is through the interior 12 of the Leaflet 15 and then flows through the impingement cooling plates 19a, 19b through into the impingement cooling chambers 13a, 13b, to exit after the impingement cooling of the wall through the film cooling holes 26a, 26b in the hot gas channel 27.
  • the locally adapted and optimized impingement cooling is advantageous in the case of a blade, in which an increased thermal load occurs due to a vortice train forming at the trailing edge 17 of the blade 15.
  • An impingement cooling chamber 13b arranged downstream of the trailing edge 17 is then designed with its impingement cooling plate 19b for increased cooling (increased hole density in FIG. 3) in order to absorb or compensate for the increased thermal load.
  • the final purpose of the invention is to be seen to make the cooling of the platform so that on the different pressure or thermal conditions, is taken into account by the prevailing flow conditions below this platform, whereby a uniform cooling of the entire platform can be achieved .
  • the different pressure conditions below the platform are due to the fact that the impinged blade contour, especially immediately below the platform, is exposed to different pressures, which is why it is also provided in a proposal according to the invention to provide the cooling of the platform sectorally by individual impingement cooling chambers, as can be seen from the figures , As a result of the chamber, this leads to separated air outlets, which is why it is important that the static pressure prevailing in the cooling system is taken into account.
  • the blow-out is designed open.
  • the solution according to the invention is based on the idea of carrying out the individual design of the impingement cooling chambers as a function of the thermal conditions which form effectively below the platform.
  • the invention provides the following advantages:
  • the cooling can be customized to specific thermal loads in specific areas of the bucket.
  • Cooling can be combined with sequential cooling arrangements that first cool an outer platform and then parallel the airfoil and an inner platform.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne un élément refroidi (10, 11) pour une turbine à gaz, délimitant le canal de gaz chauds (27) de la turbine à gaz avec la face extérieure d'une paroi (28) et présentant, sur la face intérieure, un dispositif de refroidissement par impact qui comprend une pluralité de chambres de refroidissement par impact (13a, 13b) agencées les unes à côté des autres, de préférence contiguës les unes aux autres et agissant parallèlement. Ces chambres sont recouvertes par des tôles de refroidissement par impact (19a, 19b) présentant des trous de refroidissement par impact (18a, 18b) et sont exposées à de l'air de refroidissement. Un refroidissement optimal dudit élément est obtenu du fait que les chambres de refroidissement par impact (13a, 13b) communiquent respectivement avec le canal de gaz chauds (27) par leurs propres trous de refroidissement par film (26a, 26b), et en ce que l'épaisseur, la répartition et/ou le diamètre des trous de refroidissement par impact (18a, 18b) dans les tôles de refroidissement par impact (19a, 19b) des différentes chambres de refroidissement par impact (13a, 13b) sont adaptés à la contrainte thermique respective et/ou à la pression statique respective existant en fonctionnement sur la face extérieure de la paroi (28).
PCT/EP2010/053691 2009-03-30 2010-03-22 Elément refroidi pour une turbine à gaz WO2010112360A1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP10709734.7A EP2414639B8 (fr) 2009-03-30 2010-03-22 Elément refroidi pour une turbine à gaz
US13/247,429 US20120063891A1 (en) 2009-03-30 2011-09-28 Cooled component for a gas turbine

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
CH00503/09 2009-03-30
CH00503/09A CH700687A1 (de) 2009-03-30 2009-03-30 Gekühltes bauteil für eine gasturbine.

Related Child Applications (1)

Application Number Title Priority Date Filing Date
US13/247,429 Continuation US20120063891A1 (en) 2009-03-30 2011-09-28 Cooled component for a gas turbine

Publications (1)

Publication Number Publication Date
WO2010112360A1 true WO2010112360A1 (fr) 2010-10-07

Family

ID=40627674

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2010/053691 WO2010112360A1 (fr) 2009-03-30 2010-03-22 Elément refroidi pour une turbine à gaz

Country Status (4)

Country Link
US (1) US20120063891A1 (fr)
EP (1) EP2414639B8 (fr)
CH (1) CH700687A1 (fr)
WO (1) WO2010112360A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2505787A1 (fr) * 2011-03-28 2012-10-03 Rolls-Royce plc Composant de moteur à turbine à gaz et moteur à turbine à gaz associé
EP2469034A3 (fr) * 2010-12-22 2014-01-01 United Technologies Corporation Aube statorique de turbine ayant une plateforme avec circuit de refroidissement et procédé de manufacture associé

Families Citing this family (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2863011A1 (fr) * 2013-10-16 2015-04-22 Siemens Aktiengesellschaft Aube de turbine, stator, turbine et centrale électrique associés
US10760432B2 (en) 2017-10-03 2020-09-01 Raytheon Technologies Corporation Airfoil having fluidly connected hybrid cavities
US10697309B2 (en) 2018-04-25 2020-06-30 Raytheon Technologies Corporation Platform cover plates for gas turbine engine components
JP6508499B1 (ja) * 2018-10-18 2019-05-08 三菱日立パワーシステムズ株式会社 ガスタービン静翼、これを備えているガスタービン、及びガスタービン静翼の製造方法
JP6799702B1 (ja) * 2020-03-19 2020-12-16 三菱パワー株式会社 静翼及びガスタービン

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GB2210415A (en) * 1987-09-25 1989-06-07 Toshiba Kk Turbine vane with cooling features
EP0698723A2 (fr) 1994-08-23 1996-02-28 General Electric Company Circuit de refroidissement fermé pour aube distributeur de turbine
EP0937863A2 (fr) * 1998-02-23 1999-08-25 Mitsubishi Heavy Industries, Ltd. Plateforme pour une aube mobile d'une turbine à gaz
WO2002050402A1 (fr) 2000-12-19 2002-06-27 General Electric Company Systeme de refroidissement par impact de jet pour plate-forme d'aube de turbine
EP1808575A2 (fr) * 2006-01-12 2007-07-18 Siemens Power Generation, Inc. Aube de turbine inclinée avec refroidissement par impact
WO2008052846A1 (fr) * 2006-10-30 2008-05-08 Siemens Aktiengesellschaft Aube de turbine

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DE19733148C1 (de) 1997-07-31 1998-11-12 Siemens Ag Kühlluftverteilung in einer Turbinenstufe einer Gasturbine
US6471480B1 (en) * 2001-04-16 2002-10-29 United Technologies Corporation Thin walled cooled hollow tip shroud
GB0117110D0 (en) 2001-07-13 2001-09-05 Siemens Ag Coolable segment for a turbomachinery and combustion turbine
US6779597B2 (en) 2002-01-16 2004-08-24 General Electric Company Multiple impingement cooled structure
DE10217388A1 (de) * 2002-04-18 2003-10-30 Siemens Ag Luft- und dampfgekühlte Plattform einer Turbinenschaufel
US6805533B2 (en) 2002-09-27 2004-10-19 Siemens Westinghouse Power Corporation Tolerant internally-cooled fluid guide component
US20060056968A1 (en) 2004-09-15 2006-03-16 General Electric Company Apparatus and methods for cooling turbine bucket platforms
US7568882B2 (en) 2007-01-12 2009-08-04 General Electric Company Impingement cooled bucket shroud, turbine rotor incorporating the same, and cooling method
US8123466B2 (en) * 2007-03-01 2012-02-28 United Technologies Corporation Blade outer air seal
US8439629B2 (en) 2007-03-01 2013-05-14 United Technologies Corporation Blade outer air seal

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2210415A (en) * 1987-09-25 1989-06-07 Toshiba Kk Turbine vane with cooling features
EP0698723A2 (fr) 1994-08-23 1996-02-28 General Electric Company Circuit de refroidissement fermé pour aube distributeur de turbine
EP0937863A2 (fr) * 1998-02-23 1999-08-25 Mitsubishi Heavy Industries, Ltd. Plateforme pour une aube mobile d'une turbine à gaz
WO2002050402A1 (fr) 2000-12-19 2002-06-27 General Electric Company Systeme de refroidissement par impact de jet pour plate-forme d'aube de turbine
EP1808575A2 (fr) * 2006-01-12 2007-07-18 Siemens Power Generation, Inc. Aube de turbine inclinée avec refroidissement par impact
WO2008052846A1 (fr) * 2006-10-30 2008-05-08 Siemens Aktiengesellschaft Aube de turbine

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2469034A3 (fr) * 2010-12-22 2014-01-01 United Technologies Corporation Aube statorique de turbine ayant une plateforme avec circuit de refroidissement et procédé de manufacture associé
US8714909B2 (en) 2010-12-22 2014-05-06 United Technologies Corporation Platform with cooling circuit
EP2505787A1 (fr) * 2011-03-28 2012-10-03 Rolls-Royce plc Composant de moteur à turbine à gaz et moteur à turbine à gaz associé

Also Published As

Publication number Publication date
EP2414639B1 (fr) 2016-12-28
EP2414639A1 (fr) 2012-02-08
CH700687A1 (de) 2010-09-30
EP2414639B8 (fr) 2017-03-15
US20120063891A1 (en) 2012-03-15

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