EP2131010A1 - Aube de turbine à gaz et procédé de fabrication d'une aube de turbine à gaz - Google Patents

Aube de turbine à gaz et procédé de fabrication d'une aube de turbine à gaz Download PDF

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Publication number
EP2131010A1
EP2131010A1 EP08010288A EP08010288A EP2131010A1 EP 2131010 A1 EP2131010 A1 EP 2131010A1 EP 08010288 A EP08010288 A EP 08010288A EP 08010288 A EP08010288 A EP 08010288A EP 2131010 A1 EP2131010 A1 EP 2131010A1
Authority
EP
European Patent Office
Prior art keywords
film cooling
gas turbine
turbine blade
blade
openings
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Withdrawn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
EP08010288A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
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 EP08010288A priority Critical patent/EP2131010A1/fr
Publication of EP2131010A1 publication Critical patent/EP2131010A1/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Withdrawn legal-status Critical Current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
    • F01D5/186—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
    • F01D5/005—Repairing methods or devices
    • 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
    • F05D2230/80—Repairing, retrofitting or upgrading methods

Definitions

  • the invention relates to a gas turbine blade with a surface which can be flowed around by a hot gas and in which a film cooling opening or a plurality of film cooling openings is or are arranged. Furthermore, the invention relates to a method for producing a gas turbine blade with film cooling openings arranged in a row.
  • Film cooled gas turbine blades formed as vanes or as blades, are well known in the art.
  • the gas turbine blades are known to serve to deflect the gas turbine flowing through the hot gas, wherein the hot gas umströmbare surface is usually formed by an outer wall which is part of a cast blade body.
  • the surface can be formed on the one hand by an aerodynamically curved blade of the gas turbine blade.
  • the surface which can be flowed around by the hot gas forms part of a platform which is arranged on one or both sides on the blade and which delimits the annular flow channel of the gas turbine radially inwards and / or radially outwards. So that the material of the blade body can withstand the hot temperatures for a particularly long time, a so-called. Film cooling is often used in addition to other measures.
  • Cooling air is guided through the film cooling openings to the hot gas-charged surface through the outer wall which is flowed around by the hot gas, wherein the film cooling openings are designed and arranged such that the exiting cooling air deposits a protective film over the surface, ie between the outer wall and the hot gas flow.
  • Film cooling holes arranged in rows transverse to the hot gas flow direction in the outer wall.
  • the object of the invention is therefore to provide a gas turbine blade of the type mentioned, which is particularly easy to manufacture or editable.
  • Another object of the invention is the disclosure of a method for producing a gas turbine blade according to the invention.
  • the object directed to the device is achieved with a gas turbine blade according to the features of claim 1, the object directed to the method is achieved with the features of claim 6.
  • the invention is based on the recognition that a gas turbine blade can also be made modular in the area of film cooling openings.
  • a few or all of the film cooling openings of a row are to be arranged in a separately produced film cooling element, which is at least partially provided by an outer wall formed hot gas surface of the gas turbine blade is used.
  • the film cooling holes are introduced into the cast outer wall of the generally produced by casting and thus largely integral gas turbine blade - in the invention, however, the film cooling holes are arranged in a separate film cooling element, which is used in the cast outer wall and embedded in the surface without offset.
  • the gas turbine blade preferably the blade body, is first produced in the casting process.
  • a recess provided for receiving the film cooling element can be produced simultaneously with the casting. It is also conceivable to introduce the recess later in the casting.
  • the film cooling element with the film cooling openings arranged therein is produced by an arbitrary manufacturing method. Subsequently, the prefabricated film cooling element is inserted and attached as a separate component in the gas turbine blade. As a result, the manufacturing costs can be reduced, although a separate component must be made.
  • film cooling openings are lined up in the film cooling element.
  • it can reduce the number of film cooling elements to be inserted and fixed in the gas turbine blade.
  • all arranged in a row film cooling openings are arranged in a single film cooling element.
  • not every row of film cooling openings arranged in the gas turbine blade has to be designed according to the invention.
  • a film cooling opening or a plurality of film cooling openings arranged in a row can be introduced directly in the outer wall of the preferably one-piece gas turbine blade.
  • the outer wall is at least partially formed as a cast blade base body with a recess in which the film cooling element is used.
  • the blade base body produced in a casting process and therefore essentially integral with one another comprises at least one fastening region - in the case of turbine vanes at least one hook disposed on the cold side, in turbine blades a dovetail-shaped, hammer-shaped or fir tree-shaped blade root, at least one platform radially limiting the flow duct and / or aerodynamically curved blade for deflecting the hot gas flow in the tangential direction of the annular channel.
  • the production costs can be reduced in an advantageous manner.
  • the recess may be slot-shaped. Due to the juxtaposition of film cooling openings, a recess shaped as a slot is suitable, in which the film cooling element corresponding to the slot can be inserted into the blade main body or into the outer wall of the gas turbine blade.
  • a not detected by the hot gas side wall of the slot may be formed stepped, in which case the opposite side wall of the film cooling element is formed corresponding thereto. The step of the side wall can then serve as a stop for the film cooling element to be used in order to allow a particularly simple positioning thereof.
  • the corresponding step arranged on the film cooling element is then selected such that the surface present on the film cooling element, which can be acted upon by the hot gas, lies in one plane with the immediately adjacent surface of the outer wall.
  • this can a stepless resp. wind-slippery surface of the gas turbine blade can be achieved.
  • the method according to the invention also has the particular advantage that film cooling openings can also be positioned at a position and also with an alignment on the otherwise almost one-piece gas turbine blade on which the film cooling opening with a predetermined orientation does not engage directly with the conventional production methods (laser boring, eroding) Blade body is introduced.
  • film cooling openings can also be attached to those positions of the cast blade base body or those blow-off directions that were previously inaccessible to the tools with which the film cooling openings have been introduced into the blade base body.
  • Another essential advantage of the invention lies in restoring a gas turbine blade which has already been used as intended and which has one or more film cooling openings arranged in a blade main body. If one of these openings is already damaged by cracks, which heretofore prohibits further use of the gas turbine blade in question due to insufficient service life, the gas turbine blade can nevertheless be restored by removing the relevant film cooling opening or film cooling openings from the blade body to form a gas turbine blade Recess or a slot to be removed, wherein then a film cooling element according to the invention is inserted and secured in the slot.
  • the attachment can be done with standard methods, such as soldering or welding, of course, where appropriate, the film cooling element may optionally be made of a different material than the blade body. Thus, the previously existing in the blade body cracks are removed, causing their Growth in the material of the blade body is no longer possible.
  • FIG. 1 A gas turbine blade 10 relating to the invention is shown in FIG FIG. 1 shown in perspective.
  • the gas turbine blade 10 is according to FIG. 1 designed as a blade.
  • the invention also relates to a guide vane, not shown, of a gas turbine.
  • the rotor blade 10 comprises a cross-sectionally fir-tree-shaped blade root 12 and a platform 14 arranged thereon.
  • the platform 14 is adjoined by an aerodynamically curved blade 16, which has a front edge 18 and a trailing edge 20.
  • At the front edge 18 are arranged as a so-called "shower head” arranged cooling holes, from which a flowing inside coolant, preferably cooling air, can escape.
  • the blade 16 includes a respect FIG. 1 rear suction side wall 22 and a front side pressure side wall 24.
  • a row 26 of film cooling openings 28 is arranged by way of example.
  • the series 26 is not strictly straight-line; the film cooling opening 28 lie on a rather slightly curved line.
  • FIG. 2 shows the section through the airfoil 16 of the gas turbine blade 10 according to Fig. 1 ,
  • the airfoil 16 is formed substantially hollow between the suction side wall 22 and the pressure side wall 24 and divided by a plurality of the two side walls 22, 24 mutually supporting ribs 30 in a plurality of cavities 32.
  • the cavities 32 are successively or also in parallel by a coolant, preferably cooling air, flowed through, so that inside flowing cooling air can escape through the in the pressure side wall 24 exemplarily arranged film cooling openings 28 to on the hot gas 33 exposable surface 34 a protective film of Cooling air to form, which protects the side wall 24 in front of the hot gas 33 flowing from the front edge 18 to the trailing edge 20.
  • a coolant preferably cooling air
  • the gas turbine blade 10 is substantially in one piece, ie the blade root 12, the platform 14 and / or the blade 16 is or have emerged as a contiguous blade body 38 from a casting process.
  • the blade main body 38 thus comprises at least the outer wall 22, 24, which may be formed both as a suction side wall 22 or as a pressure side wall 24.
  • a recess 42 formed as a slot 42 is arranged, in which the prefabricated film cooling element 40 is inserted flush with the outside.
  • a gas turbine blade 10 that is modular in the area of the film cooling openings 28 is provided.
  • a single film cooling opening 28, some film cooling openings 28 and / or all lying in a row 26 film cooling openings 28 may not be located directly in the blade body 38, but in a separately prepared film cooling element 40th
  • a film cooling element 40 according to the invention also to be embedded in that surface of the gas turbine blade 10 which is formed by the platform 14.
  • Another particular advantage of the invention lies in the (repro) treatment of a gas turbine blade 10 already used as intended, thereby extending its period of use.
  • the use of the invention when one of the film cooling openings 28 is located at a position of the gas turbine blade 10, at the film cooling opening with the otherwise conventional manufacturing methods (Laser drilling, erosion) can not be produced due to lack of accessibility or in which the desired orientation, ie blow-out direction of the film cooling opening for the same reason so far can not be produced. This is the case, for example, with film cooling openings, which are arranged particularly close to the transition from platform 14 to blade 16.
  • FIG. 4 and FIG. 5 show the section through an inventive outer wall of a gas turbine blade according to the section lines IV and V.
  • FIG. 3
  • both side walls of the recess 44 on a gradation both side walls of the recess 44 on a gradation.
  • the side surfaces of the film cooling element which are opposite these side surfaces are formed with contact surfaces corresponding to the step, whereby when the film cooling element 40 is inserted, it can be positioned exactly opposite the surface 34 and a too deep insertion of the film cooling element can be avoided.
  • the surfaces 34 of the blade main body 38 and the film cooling element 40 which can be exposed to the hot gas 33 are arranged and curved in such a way that the surfaces exposed to the hot gas flow are altogether offset and allow a wind-slippery flow around hot gas.
  • the invention proposes a gas turbine blade 10 with a surface 34 which can be flowed around by a hot gas 33 in which a film cooling opening 28 or a plurality of film cooling openings 28 are arranged, wherein the surface is reprocessed for a gas turbine blade 10 already used as intended or for producing a particularly inexpensive gas turbine blade 10 34 is at least partially formed by an outer wall 22, 24 of the gas turbine blade 10 and at least one film cooling element 40 inserted therein, wherein the film cooling opening 28 and / or at least one of the film cooling openings 28 is or are arranged in the film cooling element 40.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP08010288A 2008-06-05 2008-06-05 Aube de turbine à gaz et procédé de fabrication d'une aube de turbine à gaz Withdrawn EP2131010A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP08010288A EP2131010A1 (fr) 2008-06-05 2008-06-05 Aube de turbine à gaz et procédé de fabrication d'une aube de turbine à gaz

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP08010288A EP2131010A1 (fr) 2008-06-05 2008-06-05 Aube de turbine à gaz et procédé de fabrication d'une aube de turbine à gaz

Publications (1)

Publication Number Publication Date
EP2131010A1 true EP2131010A1 (fr) 2009-12-09

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Family Applications (1)

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EP08010288A Withdrawn EP2131010A1 (fr) 2008-06-05 2008-06-05 Aube de turbine à gaz et procédé de fabrication d'une aube de turbine à gaz

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EP (1) EP2131010A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3396108A1 (fr) * 2017-04-26 2018-10-31 General Electric Company Procédé permettant de fournir une structure de refroidissement pour un composant
DE102009044584B4 (de) * 2008-11-20 2021-05-20 General Electric Co. Schaufelanordnung mit Kühlöffnungen für ein Turbinentriebwerk

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3650635A (en) * 1970-03-09 1972-03-21 Chromalloy American Corp Turbine vanes
US20030082048A1 (en) * 2001-10-22 2003-05-01 Jackson Melvin Robert Airfoils with improved strength and manufacture and repair thereof
EP1847696A1 (fr) * 2006-04-21 2007-10-24 Siemens Aktiengesellschaft Composant pour un système de post-combustion dans une turbine à gaz et turbine à gaz associée.

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3650635A (en) * 1970-03-09 1972-03-21 Chromalloy American Corp Turbine vanes
US20030082048A1 (en) * 2001-10-22 2003-05-01 Jackson Melvin Robert Airfoils with improved strength and manufacture and repair thereof
EP1847696A1 (fr) * 2006-04-21 2007-10-24 Siemens Aktiengesellschaft Composant pour un système de post-combustion dans une turbine à gaz et turbine à gaz associée.

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102009044584B4 (de) * 2008-11-20 2021-05-20 General Electric Co. Schaufelanordnung mit Kühlöffnungen für ein Turbinentriebwerk
EP3396108A1 (fr) * 2017-04-26 2018-10-31 General Electric Company Procédé permettant de fournir une structure de refroidissement pour un composant
US10583489B2 (en) 2017-04-26 2020-03-10 General Electric Company Method of providing cooling structure for a component

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