EP1247939A1 - Aube de turbine et son procédé de production - Google Patents

Aube de turbine et son procédé de production Download PDF

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Publication number
EP1247939A1
EP1247939A1 EP01108759A EP01108759A EP1247939A1 EP 1247939 A1 EP1247939 A1 EP 1247939A1 EP 01108759 A EP01108759 A EP 01108759A EP 01108759 A EP01108759 A EP 01108759A EP 1247939 A1 EP1247939 A1 EP 1247939A1
Authority
EP
European Patent Office
Prior art keywords
turbine blade
feed
chamber
turbine
cooling medium
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Withdrawn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
EP01108759A
Other languages
German (de)
English (en)
Inventor
Peter Tiemann
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 EP01108759A priority Critical patent/EP1247939A1/fr
Priority to JP2002100928A priority patent/JP2002317601A/ja
Priority to US10/116,873 priority patent/US6619912B2/en
Priority to CN02119076A priority patent/CN1380486A/zh
Publication of EP1247939A1 publication Critical patent/EP1247939A1/fr
Withdrawn legal-status Critical Current

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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
    • F01D5/18—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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—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
    • F05D2230/00—Manufacture
    • F05D2230/10—Manufacture by removing material
    • 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
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49316—Impeller making
    • Y10T29/49336—Blade making
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49316—Impeller making
    • Y10T29/49336—Blade making
    • Y10T29/49339—Hollow blade

Definitions

  • the present invention relates to a method for manufacturing a turbine blade which has at least one chamber and at least one feeder for loading the chamber with has a cooling medium, at least one feed runs at an angle to a longitudinal axis of the turbine blade. It also relates to a turbine blade, in particular for a gas turbine that has at least one chamber and at least one a feed to apply a cooling medium to the chamber having.
  • Such a method and such a turbine blade are known from US 5,599,166.
  • the turbine blade has two separate, meandering chambers on, each with a feed for loading with are connected to a coolant.
  • the two feeders run essentially parallel to the longitudinal axis of the turbine blade.
  • US 5,413,458 describes another turbine blade which likewise at least one chamber for the application of a Has cooling medium.
  • the cooling medium is in one Direction fed, which is also essentially parallel to the longitudinal axis of the turbine blade.
  • a disadvantage of the known turbine blades and manufacturing processes is the imperative to determine the direction of the Feed.
  • the turbine blades generally have a Wing profile on that of a crossing the turbine Medium flows around.
  • To attach to a housing or a platform serves as a rotor.
  • the cooling medium must first the platform flow through before it enters the wing profile. This leads to the platform and the airfoil always must be cooled with the same cooling medium, in particular with a cooling medium that has the same pressure and temperature having. A targeted cooling of more demanding Part of the turbine blade is not possible.
  • the object of the present invention is therefore a method for the manufacture of a turbine blade and a turbine blade itself to provide a targeted exposure enable with a cooling medium.
  • this object is achieved in a method of type mentioned solved in that to form the A core is used with one approach and the feed Approach is spaced from a shape so that the Feeding the turbine blade after removal from the Form is closed, and that post-processing to expose the feed is made.
  • Turbine blade is provided that the feed is angled to a longitudinal axis of the turbine blade and essentially parallel to a direction of flow of a medium the turbine runs.
  • the one or more used to manufacture the turbine blade Cores are inserted and held in the mold as before.
  • the cores are supported in the form over the base not before.
  • the cores can therefore, as in the known Move the process during the pouring process. An interference the core position through a contact between the approach and the form is not there.
  • the invention deviates for the first time from the concept of the feeders essentially parallel to the longitudinal axis of the turbine blade from.
  • a feeder is provided that is angled is arranged to the longitudinal axis and substantially parallel to a direction of flow of the medium through the turbine. This feed is a targeted application highly stressed parts of the turbine blade with a cooling medium allows.
  • a second is preferred in the method according to the invention Feed substantially parallel to the longitudinal axis of the turbine blade intended.
  • the two feeders can then with different cooling media. This Difference can be particularly in pressure and / or temperature of the coolant supplied in each case. It results targeted, highly efficient cooling of individuals Parts of the turbine blade.
  • the feeders can be on a Leading edge, a trailing edge or both edges of the Turbine blade to be arranged. Through the targeted arrangement the turbine blade can be optimally cooled.
  • the is at an angle to Longitudinal axis tapered, especially conical. It then has a relatively large one Inlet cross-section.
  • the cooling medium can therefore comparatively low pressure to be fed and is compressed as it flows in.
  • the feeder is designed that flow losses are minimized.
  • the feed extending in the axial direction is advantageous between a platform and a wing profile of the turbine blade arranged.
  • the fed through this feed Cooling medium can thus directly in the wing profile enter.
  • the second one then serves to cool the platform, feed substantially parallel to the longitudinal axis.
  • the proposed distribution of the cooling medium according to the invention is particularly advantageous in the case of a turbine blade which has at least two chambers.
  • the first chamber is then there with the first feed and the second chamber with the second Feeder in connection.
  • the first chamber is here advantageous in the area of a front edge of the turbine blade arranged.
  • This arranged in the area of the front edge has in generally a higher cooling requirement than the second chamber. If the front edge is provided with openings through which the cooling medium can leak, must also be applied with a cooling medium of higher pressure. reason for this is that the cooling medium to flow out of the first Chamber the steel pressure of the medium flowing through the turbine must overcome. According to the first chamber can now the first supply with a cooling medium with higher Pressure as the second chamber. This first Chamber can thus be specifically cooled more. This Cooling effort is not required for the second chamber. It can therefore optimize the consumption of the cooling medium and thereby the overall efficiency can be increased. alternative or in addition, targeted cooling of the rear edge respectively.
  • Figure 1 shows a schematic longitudinal section through a Gas turbine 10 with a housing 11 and a rotor 12. On the housing 11 are rows of guide vanes 13 and on the rotor 12 Rows of blades 14 are provided.
  • the gas turbine 10 a hot gas flows through it in the direction of arrow 15, that rotates the rotor 12 about its axis of rotation 16 in Arrow direction 17 offset.
  • a cooling medium is used for cooling, which is supplied according to arrows 18, 19.
  • this feed is only shown for a guide vane 13.
  • the present invention is not based on a guide vane 13 is limited, but can also with a Blade 14 are used.
  • Figure 2 shows a longitudinal section and Figure 3 shows a cross section through a guide vane 13.
  • the guide vane 13 has a platform 38 for attachment to the housing 11 and a wing profile 39 on which the hot gas flows.
  • This Wing profile 39 is supported by a suction-side wall 20 and a pressure side wall 21 is formed.
  • Between the walls 20, 21 are a first chamber 22 and three others, together communicating chambers 23, 24, 25 are provided.
  • the individual chambers 22, 23, 24, 25 are separated from one another by walls 26 Cut.
  • a retrofitted one is used for covering Platform 38, for example in the form of a sheet or Perforated plate.
  • the first chamber 22 is here on a front edge 32 of the wing profile 39 of the guide vane 13 is arranged.
  • the chamber 23 servess to apply a cooling medium to the chamber 22 an approach 30 that forms a supply for the cooling medium.
  • the chamber 23 is supplied with a cooling medium via openings 31, that successively the first chamber 23 and then the Flows through chambers 24, 25.
  • the openings 34 also form a feeder.
  • the cooling medium is according to the chamber 22 Arrow direction 18 approximately perpendicular to a longitudinal axis 37 of the Guide vane 13 supplied. Acting on chamber 23 takes place in the direction of arrow 19 approximately parallel to the longitudinal axis.
  • the approach 30 allows feeding between the platform 38 and the wing profile 39.
  • the chamber 22 is filled with a cooling medium of a higher pressure the chamber 23 acts.
  • the reason for this is that this chamber 22 in the area of the highly stressed leading edge 32 of the Guide vane 13 is located.
  • the higher pressure level is special required when chamber 22 has a series of Openings 27, 28 is provided. Through these openings it can Coolant emerge and form a cooling film that runs along of the walls 20, 21 extends in the region of the front edge 32. Since the leading edge 32 flows directly from the hot gas not only the static pressure of the hot gas, but also to overcome its back pressure.
  • Gap 29 In the area of a rear edge 34 of the guide vane 13 there is a Gap 29 provided.
  • the chamber enters through this gap 23 supplied cooling medium. Since the gap 29 only with the static pressure of the hot gas is applied to the cooling of the chambers 23, 24, 25 a lower pressure of the Cooling medium sufficient.
  • the turbine blade 13, 14 according to the invention is thus the more heavily used chamber 22 with a higher cooling medium Pressure cooled than the other chambers 23, 24, 25.
  • a separate feed in the form of the approach 30 provided.
  • This feed 30 is angled to the longitudinal axis 37 of the turbine blade 13, 14 and is between the platform 38 and the wing profile 39 arranged. It is conical and has a streamlined Shape up.
  • the cooling medium is essentially parallel via this feed 31 supplied to the longitudinal axis 37.
  • FIG. 4 shows a further exemplary embodiment of a turbine blade 13 in a view similar to FIG. 2.
  • This turbine blade 13 has two lugs 30a, 30b, one of which arranged on the front edge 32 and one on the rear edge 33 is.
  • Both approaches 30a, 30b are conical and flow-friendly educated.
  • the supplied via the approaches 30a, 30b Cooling medium acts in each case on chambers 22, 25 which are in the Area of the front edge 32 or the rear edge 34 lie.
  • the middle area with the chambers 23, 24 is via a feed 31 applied substantially parallel to the longitudinal axis 37.
  • FIG. 5 shows a top view of the device used to manufacture the in Figure 2 shown turbine blade 13 used core 35a, 35b, 35c and FIG. 6 shows a section along the line VI-VI through this turbine blade 13.
  • the neck 33 of the core 35a, 35b, 35c tapers, so that the feeder also serving approach 30 of the turbine blade 13 tapers.
  • the inside of the approach 30 is smooth, so that the Flow resistance is minimized.
  • FIG. 7 schematically shows a multi-part core 35a, 35b, 35c in a form 40.
  • the individual parts 35a, 35b, 35c are over Connection pins 36 fixed relative to each other.
  • the core 35a, 35b, 35c protrudes beyond the shape 40 and is there held.
  • the resulting openings in the turbine blade 13, 14 are subsequently closed by the platform 38.
  • the lugs 33a, 33b are not in contact with the mold 40.
  • the core 35a, 35b, 35c can therefore be like the known ones Moving procedure when pouring.
  • the illustrated core 35a, 35b, 35c is introduced into the mold 40 and the mold 40 closed. After filling and As the material cools, the mold 40 is opened and the turbine blade 13, 14 taken together with the core 35a, 35b, 35c. The core 35a, 35b, 35c is then removed, for example, drained. The approach 30 of the turbine blade 13, 14 is then initially closed. He is through a suitable post-processing exposed. The completed one Turbine blades 13, 14 then provide a supply of the cooling medium both in the axial direction at an angle to the longitudinal axis 37 as well as parallel to the longitudinal axis 37.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP01108759A 2001-04-06 2001-04-06 Aube de turbine et son procédé de production Withdrawn EP1247939A1 (fr)

Priority Applications (4)

Application Number Priority Date Filing Date Title
EP01108759A EP1247939A1 (fr) 2001-04-06 2001-04-06 Aube de turbine et son procédé de production
JP2002100928A JP2002317601A (ja) 2001-04-06 2002-04-03 タービン翼の製造方法およびタービン翼
US10/116,873 US6619912B2 (en) 2001-04-06 2002-04-05 Turbine blade or vane
CN02119076A CN1380486A (zh) 2001-04-06 2002-04-06 制造涡轮叶片的方法及涡轮叶片

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP01108759A EP1247939A1 (fr) 2001-04-06 2001-04-06 Aube de turbine et son procédé de production

Publications (1)

Publication Number Publication Date
EP1247939A1 true EP1247939A1 (fr) 2002-10-09

Family

ID=8177083

Family Applications (1)

Application Number Title Priority Date Filing Date
EP01108759A Withdrawn EP1247939A1 (fr) 2001-04-06 2001-04-06 Aube de turbine et son procédé de production

Country Status (4)

Country Link
US (1) US6619912B2 (fr)
EP (1) EP1247939A1 (fr)
JP (1) JP2002317601A (fr)
CN (1) CN1380486A (fr)

Families Citing this family (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7216694B2 (en) * 2004-01-23 2007-05-15 United Technologies Corporation Apparatus and method for reducing operating stress in a turbine blade and the like
US7198467B2 (en) * 2004-07-30 2007-04-03 General Electric Company Method and apparatus for cooling gas turbine engine rotor blades
US7144215B2 (en) * 2004-07-30 2006-12-05 General Electric Company Method and apparatus for cooling gas turbine engine rotor blades
US7131817B2 (en) * 2004-07-30 2006-11-07 General Electric Company Method and apparatus for cooling gas turbine engine rotor blades
US20070122280A1 (en) * 2005-11-30 2007-05-31 General Electric Company Method and apparatus for reducing axial compressor blade tip flow
US20090074588A1 (en) * 2007-09-19 2009-03-19 Siemens Power Generation, Inc. Airfoil with cooling hole having a flared section
US8657574B2 (en) * 2010-11-04 2014-02-25 General Electric Company System and method for cooling a turbine bucket
US20130318996A1 (en) * 2012-06-01 2013-12-05 General Electric Company Cooling assembly for a bucket of a turbine system and method of cooling
US10669887B2 (en) 2018-02-15 2020-06-02 Raytheon Technologies Corporation Vane airfoil cooling air communication
US10808572B2 (en) * 2018-04-02 2020-10-20 General Electric Company Cooling structure for a turbomachinery component

Citations (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2883151A (en) * 1954-01-26 1959-04-21 Curtiss Wright Corp Turbine cooling system
US3623825A (en) * 1969-11-13 1971-11-30 Avco Corp Liquid-metal-filled rotor blade
US3799696A (en) * 1971-07-02 1974-03-26 Rolls Royce Cooled vane or blade for a gas turbine engine
US4127358A (en) * 1976-04-08 1978-11-28 Rolls-Royce Limited Blade or vane for a gas turbine engine
US4177010A (en) * 1977-01-04 1979-12-04 Rolls-Royce Limited Cooled rotor blade for a gas turbine engine
US4453888A (en) * 1981-04-01 1984-06-12 United Technologies Corporation Nozzle for a coolable rotor blade
US4529357A (en) * 1979-06-30 1985-07-16 Rolls-Royce Ltd Turbine blades
US4596281A (en) * 1982-09-02 1986-06-24 Trw Inc. Mold core and method of forming internal passages in an airfoil
US4672727A (en) * 1985-12-23 1987-06-16 United Technologies Corporation Method of fabricating film cooling slot in a hollow airfoil
US5291654A (en) * 1993-03-29 1994-03-08 United Technologies Corporation Method for producing hollow investment castings
US5413458A (en) 1994-03-29 1995-05-09 United Technologies Corporation Turbine vane with a platform cavity having a double feed for cooling fluid
US5599166A (en) 1994-11-01 1997-02-04 United Technologies Corporation Core for fabrication of gas turbine engine airfoils
US5669759A (en) * 1995-02-03 1997-09-23 United Technologies Corporation Turbine airfoil with enhanced cooling
DE19921644A1 (de) * 1999-05-10 2000-11-16 Abb Alstom Power Ch Ag Kühlbare Schaufel für eine Gasturbine

Family Cites Families (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5498126A (en) * 1994-04-28 1996-03-12 United Technologies Corporation Airfoil with dual source cooling
US5827043A (en) * 1997-06-27 1998-10-27 United Technologies Corporation Coolable airfoil

Patent Citations (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2883151A (en) * 1954-01-26 1959-04-21 Curtiss Wright Corp Turbine cooling system
US3623825A (en) * 1969-11-13 1971-11-30 Avco Corp Liquid-metal-filled rotor blade
US3799696A (en) * 1971-07-02 1974-03-26 Rolls Royce Cooled vane or blade for a gas turbine engine
US4127358A (en) * 1976-04-08 1978-11-28 Rolls-Royce Limited Blade or vane for a gas turbine engine
US4177010A (en) * 1977-01-04 1979-12-04 Rolls-Royce Limited Cooled rotor blade for a gas turbine engine
US4529357A (en) * 1979-06-30 1985-07-16 Rolls-Royce Ltd Turbine blades
US4453888A (en) * 1981-04-01 1984-06-12 United Technologies Corporation Nozzle for a coolable rotor blade
US4596281A (en) * 1982-09-02 1986-06-24 Trw Inc. Mold core and method of forming internal passages in an airfoil
US4672727A (en) * 1985-12-23 1987-06-16 United Technologies Corporation Method of fabricating film cooling slot in a hollow airfoil
US5291654A (en) * 1993-03-29 1994-03-08 United Technologies Corporation Method for producing hollow investment castings
US5413458A (en) 1994-03-29 1995-05-09 United Technologies Corporation Turbine vane with a platform cavity having a double feed for cooling fluid
US5599166A (en) 1994-11-01 1997-02-04 United Technologies Corporation Core for fabrication of gas turbine engine airfoils
US5669759A (en) * 1995-02-03 1997-09-23 United Technologies Corporation Turbine airfoil with enhanced cooling
DE19921644A1 (de) * 1999-05-10 2000-11-16 Abb Alstom Power Ch Ag Kühlbare Schaufel für eine Gasturbine

Also Published As

Publication number Publication date
JP2002317601A (ja) 2002-10-31
US20020155000A1 (en) 2002-10-24
US6619912B2 (en) 2003-09-16
CN1380486A (zh) 2002-11-20

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