EP1201879A2 - Composant refroidi, noyau de coulage et procédé pour la fabrication dudit composant - Google Patents

Composant refroidi, noyau de coulage et procédé pour la fabrication dudit composant Download PDF

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Publication number
EP1201879A2
EP1201879A2 EP01123193A EP01123193A EP1201879A2 EP 1201879 A2 EP1201879 A2 EP 1201879A2 EP 01123193 A EP01123193 A EP 01123193A EP 01123193 A EP01123193 A EP 01123193A EP 1201879 A2 EP1201879 A2 EP 1201879A2
Authority
EP
European Patent Office
Prior art keywords
channel
cooling
component
cooling channel
diameter
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.)
Granted
Application number
EP01123193A
Other languages
German (de)
English (en)
Other versions
EP1201879B1 (fr
EP1201879A3 (fr
Inventor
Hartmut Haehnle
Ibrahim Dr. El-Nashar
Rudolf Dr. Kellerer
Beat Von Arx
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.)
GE Vernova GmbH
Original Assignee
Alstom Technology AG
Alstom Schweiz 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 Alstom Technology AG, Alstom Schweiz AG filed Critical Alstom Technology AG
Publication of EP1201879A2 publication Critical patent/EP1201879A2/fr
Publication of EP1201879A3 publication Critical patent/EP1201879A3/fr
Application granted granted Critical
Publication of EP1201879B1 publication Critical patent/EP1201879B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime 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
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22C—FOUNDRY MOULDING
    • B22C9/00—Moulds or cores; Moulding processes
    • B22C9/10—Cores; Manufacture or installation of cores
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22C—FOUNDRY MOULDING
    • B22C9/00—Moulds or cores; Moulding processes
    • B22C9/22—Moulds for peculiarly-shaped castings
    • B22C9/24—Moulds for peculiarly-shaped castings for hollow articles
    • 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
    • Y10T29/49341—Hollow blade with cooling passage

Definitions

  • the present invention relates to the field of gas turbine technology. It relates to a cooled component for gas turbines according to the preamble of Claim 1.
  • Such a component is in the form of a turbine blade, e.g. from the publication GB-A-2 202 907.
  • the invention further relates to a cast core for the production of such Component and a method for producing such a component.
  • the efficiency of gas turbines which is closely related to the level of the inlet temperature for the hot combustion gases and for reasons of efficient Fuel efficiency and economy should be as high as possible material-related reasons to a particular extent depending on an efficient Use of the cooling air, which as a coolant is usually the compressor stage is removed.
  • the operational safety and service life of the gas turbine are essential adequate cooling of the thermally highly stressed turbine components or components, in particular the inlet-side guide vanes and blades of the first turbine stages.
  • the cooling can be effected in different ways, e.g. by means of internal cooling (cooling the component by cooling air circulating inside) and / or by means of Film cooling (generating a cooling air film through suitably arranged outlet openings on the loaded outside of the component).
  • a known method for efficient internal cooling is a so-called “cyclone” (or “vortex chamber” in GB-A-2 202 907).
  • a "cyclone” will an elongated cooling duct with a mostly circular or elliptical cross section through a series of tangentially opening feed bores with cooling air applied.
  • the incoming cooling air forms a vortex in the cooling channel the longitudinal axis of the channel rotates and due to the high speed and Turbulence in the edge area a particularly effective cooling of the channel wall and so that the cooled component causes.
  • Fig. 1 is a turbine blade in a simplified perspective view 10 reproduced with such a known cyclone cooling.
  • the turbine blade 10 is shown "transparently" so that the inner ones Cavities and channels are recognizable as solid lines.
  • the Turbine blade 10 has a leading edge 13 and a trailing edge ("trailing edge") 14, each between the longitudinal direction of the blade extend the blade root 11 and the blade tip 12.
  • the special one Formation of the blade root 11 for fastening the blade to the rotor and Supply of the blade with cooling air, as described, for example, in US-A-4,293,275 or US-A-5,002,460 is disclosed in Figure 1 for the sake of Simplification not reproduced.
  • Coolant channel 15 For internal cooling of the turbine blade 10 is from the blade root 11 through a connecting channel, not shown, cooling air in a longitudinal direction extending coolant channel 15 fed (vertical arrows in Fig. 1). Parallel to the coolant channel 15 and parallel to the one to be cooled, particularly thermally loaded front edge 13 of the turbine blade 10 runs a cylindrical cooling channel 16, which forms the cyclone. From the coolant channel 15 a number of transverse feed bores 17 to the cooling channel 16 and opens out there approximately tangential one. The tangential through the feed holes 17 in the cooling channel 16 Incoming cooling air (horizontal arrows in Fig. 1) forms one over the Channel-extending vortex, the heat from the surrounding channel wall receives.
  • the heated cooling air either emerges from the cooling channel 16 at the end from, or - as shown in GB-A-2 202 907 - through tangential outlets in Form of holes or slots.
  • Other internal cooling facilities, the simultaneously serve for film cooling and / or with the rear edge 14 in connection are omitted in Fig. 1 for the sake of simplicity.
  • cyclone cooling depends to a large extent on the feed (Boundary conditions, position and cross sections of the feed bores, etc.).
  • feed bores 17 with a bore diameter that is smaller than half the hydraulic diameter of the cooling channel (cyclone) 16. Since a turbine blade 10 of the type shown in FIG. 1 is usually by a Metal casting process is required for the formation of the coolant channel 15, the cooling channel 16 and the two connecting feed bores 17 a corresponding multi-connected cast core can be used. Weaknesses of such a core are those due to the above Diameter condition comparatively thin connecting webs, which the the form later feed bores. At this point it can easily become one Core breakage come, which questions the success of the casting.
  • the object is achieved by the entirety of the features of claim 1.
  • the essence of the invention is through a suitable formation of the whole the feed bores increase the rigidity of the associated cast core improve without adhering to the specified diameter conditions for having to give up the feed holes. This happens because the Feed bores predominantly have a bore diameter that is smaller than half the hydraulic diameter of the cooling channel, and that for Improvement of the output rate when casting the component selected feed bores have a bore diameter that is larger than that half hydraulic diameter of the cooling channel.
  • the selected feed holes at the ends of the cooling channel arranged, in particular the bottom and the top feed hole are used as the selected feed hole. This can the desired swirl of cooling air over the entire interior of the cooling duct train practically unhindered and develop its maximum cooling effect.
  • the component e.g. a turbine blade, particularly long, can, however, in the Be advantageous in terms of core stability if according to another Embodiment additionally selected feed bores in the central area of the cooling channel are provided.
  • the cast core according to the invention for the production of such a component which Cast core a first channel part to form the coolant channel and one includes second channel part to form the cooling channel, and a plurality of Connecting webs, which run transversely between the two channel parts and serve to form the feed holes, is characterized in that the Connecting webs predominantly have an outer diameter that is smaller is selected as half the hydraulic diameter of the cooling channel, and that Connecting webs have an outer diameter that is larger than half the hydraulic diameter of the cooling channel.
  • the selected connecting webs are preferably at the ends of the arranged second channel part, in particular the bottom and the top Connection bridge are used as the selected connection bridge.
  • the method according to the invention for producing a component according to the invention by means of a metal casting process is characterized in that a cast core according to the invention is used.
  • FIG. 3 is an exemplary embodiment of an internally cooled gas turbine component
  • a turbine blade 10 'comparable to FIG. 1 is reproduced.
  • the same parts of the turbine blade 10 ' are given the same reference numerals provided, as with the turbine blade 10 from FIG. 1.
  • the majority of the feed holes, namely the feed holes 17, meet the criterion characteristic of a cyclone in diameter, that their bore diameter is smaller than half the hydraulic one Diameter of the cooling channel 16.
  • Only a few selected feed holes, namely the feed bores 25, 26 and 27 have a bore diameter which is larger than half the hydraulic diameter of the cooling duct 16. Through these selected feed bores 25, .., 27 can - as will be explained below - the production output rate the blades are increased significantly.
  • the cast core 18 comprises a first channel part 19, which is required to form the coolant channel 15 and a second channel part 20, which is used for the formation of the cooling channel 16 responsible is.
  • Both channel parts 19 and 20 are one above the other by a series arranged connecting webs 21 and 22, .., 24 connected, each one have a round cross-section.
  • the majority of the connecting webs, namely the "Thin” connecting webs 21 serve to form the feed bores the above "Cyclone criterion" with regard to the diameter are sufficient.
  • Only a few selected connecting webs, namely connecting webs 22, 23 and 24, are “thicker” and thus strengthen the connection between the core parts 19 and 20 and thus the mechanical rigidity of the cast core 18 as a whole.
  • cooling channel 16 or the second channel part 20 is not very long, it is sufficient off, the two outer connecting webs 22 and 24 as selected Form connecting webs with an enlarged cross section. In this way can be practically on the entire length of the cooling channel 16 of the cooling air vortex train undisturbed because the "cyclone criterion" is met there. With longer cooling channels 16 or channel parts 20, however, it can be expedient and advantageous to also provide individual selected connecting webs 26 in the middle area, to make the casting core 18 stiffer there.
  • the diameter of the selected feed holes 25, .., 27 or the selected one Connecting webs 22, .., 24 is chosen larger than that in any case half hydraulic diameter. How big the diameter actually is depends largely on the geometry of the casting core and the casting process and must be determined in individual cases.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Molds, Cores, And Manufacturing Methods Thereof (AREA)
EP01123193A 2000-10-27 2001-09-28 Composant refroidi, noyau de coulage et procédé pour la fabrication dudit composant Expired - Lifetime EP1201879B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE10053356 2000-10-27
DE10053356A DE10053356A1 (de) 2000-10-27 2000-10-27 Gekühltes Bauteil, Gusskern für die Herstellung eines solchen Bauteils, sowie Verfahren zum Herstellen eines solchen Bauteils

Publications (3)

Publication Number Publication Date
EP1201879A2 true EP1201879A2 (fr) 2002-05-02
EP1201879A3 EP1201879A3 (fr) 2003-07-16
EP1201879B1 EP1201879B1 (fr) 2004-11-10

Family

ID=7661311

Family Applications (1)

Application Number Title Priority Date Filing Date
EP01123193A Expired - Lifetime EP1201879B1 (fr) 2000-10-27 2001-09-28 Composant refroidi, noyau de coulage et procédé pour la fabrication dudit composant

Country Status (3)

Country Link
US (1) US6547525B2 (fr)
EP (1) EP1201879B1 (fr)
DE (2) DE10053356A1 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2434093A3 (fr) * 2010-09-23 2013-08-07 Rolls-Royce Deutschland Ltd & Co KG Aubes de turbine refroidies pour une turbine à gaz
EP3091183A1 (fr) * 2015-05-08 2016-11-09 United Technologies Corporation Canaux de régulation thermique pour composants de turbomachine
US10328485B2 (en) 2014-09-04 2019-06-25 Safran Aircraft Engines Method for producing a ceramic core

Families Citing this family (24)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE50311059D1 (de) 2003-10-29 2009-02-26 Siemens Ag Gussform
US7128533B2 (en) * 2004-09-10 2006-10-31 Siemens Power Generation, Inc. Vortex cooling system for a turbine blade
US7690894B1 (en) 2006-09-25 2010-04-06 Florida Turbine Technologies, Inc. Ceramic core assembly for serpentine flow circuit in a turbine blade
WO2011019672A2 (fr) * 2009-08-09 2011-02-17 Rolls-Royce Corporation Support pour un article cuit
GB0921818D0 (en) * 2009-12-15 2010-01-27 Rolls Royce Plc Casting of internal features within a product (
DE102012017491A1 (de) 2012-09-04 2014-03-06 Rolls-Royce Deutschland Ltd & Co Kg Turbinenschaufel einer Gasturbine mit Drallerzeugungselement
WO2014137470A1 (fr) 2013-03-05 2014-09-12 Vandervaart Peter L Agencement de composant pour moteur à turbine à gaz
WO2014163698A1 (fr) 2013-03-07 2014-10-09 Vandervaart Peter L Pièce refroidie de turbine à gaz
US10012090B2 (en) * 2014-07-25 2018-07-03 United Technologies Corporation Airfoil cooling apparatus
US10273810B2 (en) 2016-10-26 2019-04-30 General Electric Company Partially wrapped trailing edge cooling circuit with pressure side serpentine cavities
US10465521B2 (en) 2016-10-26 2019-11-05 General Electric Company Turbine airfoil coolant passage created in cover
US10450875B2 (en) 2016-10-26 2019-10-22 General Electric Company Varying geometries for cooling circuits of turbine blades
US10309227B2 (en) 2016-10-26 2019-06-04 General Electric Company Multi-turn cooling circuits for turbine blades
US10598028B2 (en) 2016-10-26 2020-03-24 General Electric Company Edge coupon including cooling circuit for airfoil
US10352176B2 (en) 2016-10-26 2019-07-16 General Electric Company Cooling circuits for a multi-wall blade
US10450950B2 (en) 2016-10-26 2019-10-22 General Electric Company Turbomachine blade with trailing edge cooling circuit
US10240465B2 (en) * 2016-10-26 2019-03-26 General Electric Company Cooling circuits for a multi-wall blade
US10301946B2 (en) 2016-10-26 2019-05-28 General Electric Company Partially wrapped trailing edge cooling circuits with pressure side impingements
EP3832069A1 (fr) 2019-12-06 2021-06-09 Siemens Aktiengesellschaft Aube de turbine pour turbine à gaz fixe
US11814965B2 (en) 2021-11-10 2023-11-14 General Electric Company Turbomachine blade trailing edge cooling circuit with turn passage having set of obstructions
CN114215607A (zh) * 2021-11-29 2022-03-22 西安交通大学 一种涡轮叶片前缘旋流冷却结构
CN114109518A (zh) * 2021-11-29 2022-03-01 西安交通大学 一种涡轮叶片前缘带肋旋流-气膜复合冷却结构
CN114412577B (zh) * 2022-01-24 2024-03-15 杭州汽轮动力集团股份有限公司 涡轮动叶长叶片
US11998974B2 (en) * 2022-08-30 2024-06-04 General Electric Company Casting core for a cast engine component

Citations (3)

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Publication number Priority date Publication date Assignee Title
US4293275A (en) 1978-09-14 1981-10-06 Hitachi, Ltd. Gas turbine blade cooling structure
GB2202907A (en) 1987-03-26 1988-10-05 Secr Defence Cooled aerofoil components
US5002460A (en) 1989-10-02 1991-03-26 General Electric Company Internally cooled airfoil blade

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GB893706A (en) * 1960-01-05 1962-04-11 Rolls Royce Blades for fluid flow machines
US3542486A (en) * 1968-09-27 1970-11-24 Gen Electric Film cooling of structural members in gas turbine engines
FR2516165B1 (fr) * 1981-11-10 1986-07-04 Snecma Aube de turbine a gaz a chambre de refroidissement par circulation de fluide et son procede de realisation
US4669957A (en) * 1985-12-23 1987-06-02 United Technologies Corporation Film coolant passage with swirl diffuser
US5603606A (en) * 1994-11-14 1997-02-18 Solar Turbines Incorporated Turbine cooling system
US5498133A (en) * 1995-06-06 1996-03-12 General Electric Company Pressure regulated film cooling
EP0892151A1 (fr) * 1997-07-15 1999-01-20 Asea Brown Boveri AG Système de refroidissement pour le bord d'attaque d'une aube de turbine à gas
DE19738065A1 (de) * 1997-09-01 1999-03-04 Asea Brown Boveri Turbinenschaufel einer Gasturbine

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4293275A (en) 1978-09-14 1981-10-06 Hitachi, Ltd. Gas turbine blade cooling structure
GB2202907A (en) 1987-03-26 1988-10-05 Secr Defence Cooled aerofoil components
US5002460A (en) 1989-10-02 1991-03-26 General Electric Company Internally cooled airfoil blade

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2434093A3 (fr) * 2010-09-23 2013-08-07 Rolls-Royce Deutschland Ltd & Co KG Aubes de turbine refroidies pour une turbine à gaz
US9051841B2 (en) 2010-09-23 2015-06-09 Rolls-Royce Deutschland Ltd & Co Kg Cooled turbine blades for a gas-turbine engine
US10328485B2 (en) 2014-09-04 2019-06-25 Safran Aircraft Engines Method for producing a ceramic core
EP3091183A1 (fr) * 2015-05-08 2016-11-09 United Technologies Corporation Canaux de régulation thermique pour composants de turbomachine
US9988912B2 (en) 2015-05-08 2018-06-05 United Technologies Corporation Thermal regulation channels for turbomachine components

Also Published As

Publication number Publication date
EP1201879B1 (fr) 2004-11-10
DE10053356A1 (de) 2002-05-08
US20020051706A1 (en) 2002-05-02
DE50104476D1 (de) 2004-12-16
EP1201879A3 (fr) 2003-07-16
US6547525B2 (en) 2003-04-15

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