EP2260180B1 - Leitschaufel für eine gasturbine - Google Patents

Leitschaufel für eine gasturbine Download PDF

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Publication number
EP2260180B1
EP2260180B1 EP09726037.6A EP09726037A EP2260180B1 EP 2260180 B1 EP2260180 B1 EP 2260180B1 EP 09726037 A EP09726037 A EP 09726037A EP 2260180 B1 EP2260180 B1 EP 2260180B1
Authority
EP
European Patent Office
Prior art keywords
blade
guide vane
gas turbine
cross
cooling
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.)
Active
Application number
EP09726037.6A
Other languages
German (de)
English (en)
French (fr)
Other versions
EP2260180A2 (de
Inventor
Willy Heinz Hofmann
Roland DÜCKERSHOFF
Brian Kenneth Wardle
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.)
Ansaldo Energia IP UK Ltd
Original Assignee
Ansaldo Energia IP UK Ltd
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Publication date
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Publication of EP2260180A2 publication Critical patent/EP2260180A2/de
Application granted granted Critical
Publication of EP2260180B1 publication Critical patent/EP2260180B1/de
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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
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F05D2220/321—Application in turbines in gas turbines for a special turbine stage
    • F05D2220/3215—Application in turbines in gas turbines for a special turbine stage the last stage of the turbine
    • 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
    • F05D2230/21—Manufacture essentially without removing material by casting
    • 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
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/301—Cross-sectional characteristics
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/185—Two-dimensional patterned serpentine-like
    • 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
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Definitions

  • the present invention relates to the field of gas turbine technology. It relates to a guide vane for a gas turbine according to the preamble of claim 1. It also relates to a gas turbine equipped with such a vane.
  • Such a gas turbine which has become known in the art as GT24 / 26, for example, from an article of Joos, F. et al., "Field Experience of the Sequential Combustion System for the ABB GT24 / GT26 Gas Turbine Family", IGTI / ASME 98-GT-220, 1998 Sweden ,
  • the local Fig. 1 shows the basic structure of such a gas turbine, where the local Fig. 1 in the present application as Fig. 1 is reproduced. Furthermore, such a gas turbine goes out EP-B1-0 620 362 out.
  • US 4,930,980A discloses a cooled turbine nozzle according to the preamble of claim 1.
  • Fig. 1 shows a gas turbine 10 with sequential combustion, in which along an axis 19, a compressor 11, a first combustion chamber 14, a high pressure turbine (HDT) 15, a second combustion chamber 17 and a low pressure turbine (NDT) 18 are arranged.
  • the compressor 11 and the two turbines 15, 18 are part of a rotor which rotates about the axis 19.
  • the compressor 11 sucks in air and compresses it.
  • the compressed air flows into a plenum, and from there into premix burners, where this air is mixed with at least one fuel, fuel supplied at least via the fuel feed 12.
  • premix burners go out in principle EP-A1-0 321 809 or EP-A2-0 704 657 out.
  • the compressed air flows into the premix burners, where the mixing, as stated above, takes place with at least one fuel.
  • This fuel / air mixture then flows into the first combustion chamber 14, into which this mixture passes to form a stable flame front for combustion.
  • the hot gas thus provided is partially expanded in the subsequent high-pressure turbine 15 under working performance and then flows into the second combustion chamber 17, where a further fuel supply 16 takes place. Due to the high temperatures, which still has the hot gas partially released in the high-pressure turbine 15, combustion takes place in the second combustion chamber 17, which combustion is based on autoignition.
  • the hot gas reheated in the second combustion chamber 17 is then expanded in a multistage low-pressure turbine 18.
  • the low-pressure turbine 18 comprises a plurality of rows of blades and vanes arranged alternately in the flow direction, which are arranged alternately.
  • the vanes of the third directional blade row in the flow direction are in Fig. 1 provided with the reference numeral 20 '.
  • a gaseous cooling medium eg compressed air from the compressor of the gas turbine is shown or supplied with steam.
  • the cooling medium is sent through cooling channels formed in the blade (often in serpentines) and / or at different points of the blade through holes (holes, Slits) to form a cooling film, in particular on the outside of the blade (film cooling)
  • An example of such a cooled blade is in the document US-A-5,813,835 described and illustrated.
  • the guide vanes 20 'in the known gas turbine Fig. 1 are formed as cooled blades, which have inside extending in the radial direction of the cooling channels, as for example from the document WO-A1-2006029983 have become known.
  • Such vanes are based on a manufactured high-tech casting process, wherein the casting material from both sides (blade head and cover plate) is fed to the mold. Because of the comparatively thin walls of the airfoil and the channels and openings for the cooling air produced by the casting process, the service life, the cooling air consumption and the cooling effect achieved depend strongly on the precision achievable in the casting process. This is especially the case when such blades still have a pronounced curvature in space.
  • the invention aims to remedy this situation.
  • the invention wants to suggest an improvement here. It is an object of the invention to provide a guide vane, which is able to maximize the service life and cooling, taking into account the casting conditions.
  • the airfoil has in the radial direction a cross-sectional area of the blade material that varies over the height of the airfoil.
  • the cross-sectional area of the blade material is understood to be the difference between the total cross-sectional area of the blade leaf and the cross-sectional area of the cooling passages.
  • the cross-sectional area of the blade material is a minimum, depending on the height of the blade.
  • the minimum cross-sectional area of the blade material is in the range between 20% and 40% of the total height of the airfoil.
  • Another embodiment of the guide vane according to the invention is characterized in that it has a curved shape in space, that arranged in the interior of the airfoil a number extending in the radial direction of cooling channels in the direction of the hot gas flow one behind the other and by respectively at the ends of the airfoil.
  • the cooling channels arranged deflection are connected to each other, that the cooling medium flows through the cooling channels successively in alternating directions, and that the cooling channels in the radial direction of the curvature of the airfoil follow in space.
  • a gas turbine is equipped with such a guide vane according to the invention, wherein the vane is arranged in a turbine of the gas turbine.
  • the gas turbine is a sequential combustion gas turbine having a first combustion chamber with a high pressure turbine downstream and a second combustion chamber with a downstream low pressure turbine with the nozzle disposed in the low pressure turbine. (See the above-mentioned Fig. 1 ).
  • the low-pressure turbine preferably has a plurality of rows of guide vanes one behind the other in the flow direction, wherein the guide vane according to the invention is arranged in a middle row of guide vanes.
  • FIG. 2 is a side view of a vane in the low-pressure turbine of a gas turbine with sequential combustion after Fig. 1 illustrated according to a preferred embodiment of the invention.
  • the guide vane 20 comprises a strongly curved airfoil 22 which extends in the longitudinal direction (in the radial direction of the gas turbine) between a vane head 23 and a cover plate 21 and extends in the direction of the hot gas stream 29 from a front edge 27 to a trailing edge 28. Between the two edges 27 and 28, the blade 22 is outwardly through a pressure side (in Fig. 2 on the side facing away from the viewer) and a suction side 26 limited.
  • the vane 20 is secured by means of the formed on the top of the cover plate 21 hook-shaped fastening elements 24 and 25 on the turbine housing, while it rests sealingly with the blade head 23 on the rotor.
  • the inner structure of the vane 20 is in Fig. 3
  • the airfoil is traversed in the longitudinal direction by three cooling channels 30, 31 and 32, which follow the curvature of the airfoil in space and are arranged one behind the other in the direction of the hot gas stream 29 and are connected to one another by deflection regions arranged at the ends of the airfoil, that the cooling medium flows through the cooling channels 30, 31, 32 successively in alternating directions.
  • the airfoil 22 with its inner cooling channels 30, 31, 32 is bounded outwardly by walls 33, 36, while the cooling channels 30, 31, 32 are delimited from each other by walls 34 and 35.
  • the total cross-sectional area of the walls 33,..., 36 in the radial direction, ie in the direction of the height h of the airfoil 22, is the difference between the airfoil cross-section and the cross-section of the cooling channels 30, 31, 32. This area difference is the integral cross-sectional area of the airfoil blade material.
  • Cross-sectional area of the blade material over the height h varies, in particular, this cross-sectional area passes through a minimum.
  • the minimum of the cross-sectional area is in the range between 20% and 40% of the height h of the airfoil 22 or in the range of 0.2h to 0.4h, as in Fig. 3 indicated by the dashed borders.
  • This design influences the shape of the blade profile in terms of cross-sectional area, wall thickness, chord length and cooling channel cross-section. With an appropriate distribution of these parameters over the blade height, the underlying requirements with respect to the life of the blade, the achievable cooling and the cooling air consumption are achieved.
  • the guide vanes according to the invention can be used in gas turbines with sequential combustion, in particular in the middle rows of guide vanes of the low-pressure turbine, which is connected downstream of the second combustion chamber.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP09726037.6A 2008-03-28 2009-03-05 Leitschaufel für eine gasturbine Active EP2260180B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
CH4682008 2008-03-28
PCT/EP2009/052570 WO2009118235A2 (de) 2008-03-28 2009-03-05 Leitschaufel für eine gasturbine

Publications (2)

Publication Number Publication Date
EP2260180A2 EP2260180A2 (de) 2010-12-15
EP2260180B1 true EP2260180B1 (de) 2017-10-04

Family

ID=40001498

Family Applications (1)

Application Number Title Priority Date Filing Date
EP09726037.6A Active EP2260180B1 (de) 2008-03-28 2009-03-05 Leitschaufel für eine gasturbine

Country Status (5)

Country Link
US (1) US8459934B2 (zh)
EP (1) EP2260180B1 (zh)
JP (1) JP5490091B2 (zh)
CN (1) CN102016234B (zh)
WO (1) WO2009118235A2 (zh)

Families Citing this family (14)

* Cited by examiner, † Cited by third party
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US8757961B1 (en) * 2011-05-21 2014-06-24 Florida Turbine Technologies, Inc. Industrial turbine stator vane
US8720526B1 (en) * 2012-11-13 2014-05-13 Siemens Energy, Inc. Process for forming a long gas turbine engine blade having a main wall with a thin portion near a tip
ITCO20120059A1 (it) * 2012-12-13 2014-06-14 Nuovo Pignone Srl Metodi per produrre pale cave sagomate in 3d di turbomacchine mediante produzione additiva, pale cave di turbomacchina e turbomacchine
EP3034798B1 (en) * 2014-12-18 2018-03-07 Ansaldo Energia Switzerland AG Gas turbine vane
US12048611B2 (en) 2015-01-08 2024-07-30 Operart Llc Dental implant prosthesis
EP3081751B1 (en) * 2015-04-14 2020-10-21 Ansaldo Energia Switzerland AG Cooled airfoil and method for manufacturing said airfoil
EP3112589A1 (de) 2015-07-03 2017-01-04 Siemens Aktiengesellschaft Turbinenschaufel
US10174622B2 (en) * 2016-04-12 2019-01-08 Solar Turbines Incorporated Wrapped serpentine passages for turbine blade cooling
US10641174B2 (en) 2017-01-18 2020-05-05 General Electric Company Rotor shaft cooling
US11274569B2 (en) 2017-12-13 2022-03-15 Pratt & Whitney Canada Corp. Turbine shroud cooling
US10533454B2 (en) 2017-12-13 2020-01-14 Pratt & Whitney Canada Corp. Turbine shroud cooling
US10570773B2 (en) * 2017-12-13 2020-02-25 Pratt & Whitney Canada Corp. Turbine shroud cooling
US10502093B2 (en) * 2017-12-13 2019-12-10 Pratt & Whitney Canada Corp. Turbine shroud cooling
US11365645B2 (en) 2020-10-07 2022-06-21 Pratt & Whitney Canada Corp. Turbine shroud cooling

Citations (2)

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EP1013882A2 (en) * 1998-12-24 2000-06-28 Rolls-Royce Plc Gas turbine engine internal air system
EP1626162A1 (en) * 2004-08-11 2006-02-15 United Technologies Corporation Temperature tolerant vane assembly

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EP1626162A1 (en) * 2004-08-11 2006-02-15 United Technologies Corporation Temperature tolerant vane assembly

Also Published As

Publication number Publication date
WO2009118235A2 (de) 2009-10-01
US8459934B2 (en) 2013-06-11
WO2009118235A3 (de) 2010-11-25
EP2260180A2 (de) 2010-12-15
JP2011517480A (ja) 2011-06-09
JP5490091B2 (ja) 2014-05-14
US20110076155A1 (en) 2011-03-31
CN102016234B (zh) 2015-05-20
CN102016234A (zh) 2011-04-13

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