EP1709298A1 - Aube refroidie pour une turbine a gaz - Google Patents

Aube refroidie pour une turbine a gaz

Info

Publication number
EP1709298A1
EP1709298A1 EP05701516A EP05701516A EP1709298A1 EP 1709298 A1 EP1709298 A1 EP 1709298A1 EP 05701516 A EP05701516 A EP 05701516A EP 05701516 A EP05701516 A EP 05701516A EP 1709298 A1 EP1709298 A1 EP 1709298A1
Authority
EP
European Patent Office
Prior art keywords
blade
cooling
cooling channel
inlet
flow
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
EP05701516A
Other languages
German (de)
English (en)
Other versions
EP1709298B1 (fr
Inventor
Shailendra Naik
Sacha Parneix
Ulrich Rathmann
Helene Saxer-Felici
Stefan Schlechtriem
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.)
General Electric Technology GmbH
Original Assignee
Alstom Technology 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 filed Critical Alstom Technology AG
Publication of EP1709298A1 publication Critical patent/EP1709298A1/fr
Application granted granted Critical
Publication of EP1709298B1 publication Critical patent/EP1709298B1/fr
Not-in-force legal-status Critical Current
Anticipated expiration legal-status Critical

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
    • F01D5/187—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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the 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
    • F05D2250/00—Geometry
    • F05D2250/50—Inlet or outlet
    • 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
    • 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

Definitions

  • the present invention relates to the field of gas turbine technology. It relates to a cooled blade for a gas turbine according to the preamble of claim 1.
  • Such a blade is e.g. known from US-A-4,278,400.
  • blades with a cover band are used which are exposed to hot gases with temperatures of more than 1200 K and pressures of more than 6 bar during operation.
  • the blade 10 comprises an airfoil 11 which merges downwards into a blade root 12 via a blade shaft 25. At the upper end, that is to say at the blade tip or blade tip, the blade blade 11 merges into a shroud section 21 which, together with the shroud sections of the other blades, forms a closed, annular shroud with a complete blade ring.
  • the airfoil has a span with which it extends from the airfoil shaft to the airfoil tip. If the blade is installed in a turbine, the span of the blade is oriented in the radial direction of the turbine cross section, which is why the direction of the span is also referred to below as the radial direction.
  • the airfoil 11 has a leading edge 19, which is flown by the hot gas, and a trailing edge 20. Inside the airfoil 11, a plurality of radial cooling channels 13, 14 and 15 are arranged, which are connected to one another in terms of flow by means of deflection regions 17, 18 and one serpentine with several Form turns (see the flow arrows in the cooling channels 13, 14, 15 of FIG. 1).
  • the cooling medium Due to the one-time passage of the cooling medium through the serpentine cooling channels 13, 14, 15, the cooling medium flows through the cooling channels with increasing temperature and reaches the highest temperature in the last cooling channel 15 of the rear edge 20.
  • the rear edge 20 of the blade 10 can therefore be below certain operating conditions reach excessively high temperatures of the cooling medium and the blade material or metal.
  • the resulting mismatch of the metal temperature over the axial length of the blade can lead to high temperature creep and consequently to the deformation of the trailing edge 20.
  • the secondary effect of the rear edge deformation is a tilting of the shroud segments 21 in the axial, radial and circumferential directions.
  • the tilting of the shroud segments 21 can lead to the gaps between individual shroud segments opening and the entry of high-temperature hot gas into the shroud cavity.
  • the temperatures of the shroud metal can increase significantly and quickly cause the shroud to creep and ultimately lead to the high-temperature failure of the shroud.
  • This known type of multiple supply with cooling medium has various disadvantages: the injector changes the pressure conditions and flow conditions in the cooling ducts massively compared to the configuration with single supply through the entrance of the cooling duct at the front edge. In particular, a balance must be found between the cooling medium flowing out at the front edge for film cooling and the cooling medium drawn in by the injector. This requires a completely new design of the blade cooling, which is difficult to adapt to changing requirements.
  • the injector principle and the associated vacuum generation are not suitable for blades without film cooling of the leading edge and blades with a cooled shroud.
  • an additional flow of cooling medium is branched off directly from the main cooling inlet and via a between the Main cooling inlet and the second deflection area opening, in an exemplary embodiment, a bore or an opening made during casting, is fed into the cooling channel running along the rear edge. Since the flow of the cooling medium is branched off from the main cooling flow through the bypass hole and added again later, the cooling medium flow remains unchanged overall.
  • An advantageous embodiment of the invention is characterized in that the bore is designed and arranged such that the cooling medium flowing through the bore flows directly into the second cooling duct through the second deflection region. This results in a particularly efficient temperature reduction through the bypass flow in the cooling channel of the rear edge.
  • FIG. 1 shows in longitudinal section the configuration of a cooled gas turbine blade with multiple supply of the cooling medium and cooled cover band according to a preferred embodiment of the invention
  • Fig. 2 shows the foot area of the blade from Fig. 1 in an enlarged
  • FIG. 3 shows a top view of the shroud section of the blade from FIGS. 1, 2; and Fig. 4-6 different sections through the shroud area of the blade from Fig. 1, 2 along the parallel section planes AA, BB and CC shown in Fig. 5.
  • FIGS. 1 and 2 A preferred embodiment of a cooled gas turbine blade with multiple supply of the cooling medium according to the invention is shown in FIGS. 1 and 2.
  • the main flow of the cooling medium in the area of the blade shaft 25 enters the cooling channel 13 from below through a main cooling inlet 16 and partly passes through openings in the shroud section 21 (bores 27,..., 29 in FIGS. 3 to 6) and partly along the rear edge 20 again (see the arrows drawn in FIG. 1 on the shroud section 21 and on the rear edge 20).
  • a portion of the cooling medium flowing into the main cooling inlet 16 is branched off through a bore 23 and fed via the second deflection region 18 to the cooling channel 15 at the rear edge.
  • the bore 23 is preferably designed and arranged in this case (i.e., in the present case leading obliquely upwards) in such a way that the cooling medium flow flowing through it is directed directly into the cooling duct 15.
  • the purpose of the bypass bore 23 is to introduce cooler cooling medium directly into the rear edge area of the blade 10.
  • cooling bores 27, 28, 29 are provided in the shroud section 21 of the blade 10 (FIGS. 3 to 6).
  • the cooling medium emerging through the bores 27, 28, 29 serves to actively cool the shroud section 21.
  • the cooling bores 27, 28, 29 in the shroud section 21 preferably have an inner diameter in the range between 0.6 mm and 4 mm. All three bores 27, 28, 29 are positioned and dimensioned on the shroud section 21 such that an uniform beam penetration into the main stream of the shroud cavity takes place.

Landscapes

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

Abstract

L'invention concerne une aube (10) refroidie, destinée à une turbine à gaz et présentant une pale (11) qui s'étend à partir d'une emplanture (12) et d'un arbre (25) et qui comporte un bord d'attaque (19) et un bord de fuite (20) ainsi que, à l'intérieur de la pale (11), une pluralité de canaux de refroidissement (13,14, 15) s'étendant dans le sens radial et situés les uns derrière les autres dans le sens d'écoulement. Un premier canal de refroidissement (13) situé le long du bord d'attaque (19) et un deuxième canal de refroidissement (15) situé le long du bord de fuite (20), s'étendant de l'emplanture (12) jusqu'à l'extrémité de la pale (11), sont parcourus par un flux principal d'un fluide de refroidissement. L'entrée du premier canal de refroidissement (13) communique avec l'entrée (16) du flux principal de fluide de refroidissement et la sortie du premier canal de refroidissement (13) communique avec l'entrée du deuxième canal de refroidissement (15) par l'intermédiaire d'une première zone de déviation (17), d'un troisième canal de refroidissement (14) situé entre le premier (13) et le deuxième canal de refroidissement (15), et d'une deuxième zone de déviation (18). Des premiers moyens (23) permettent d'ajouter, de l'extérieur, au flux principal échauffé, provenant du troisième canal de refroidissement (14) et pénétrant dans le deuxième canal de refroidissement (15), un flux supplémentaire de fluide de refroidissement à une température inférieure. L'objectif de l'invention est de réaliser cette aube de manière particulièrement simple. A cet effet, les premiers moyens comprennent un orifice (23) s'étendant de l'entrée (16) du flux principal de fluide de refroidissement jusqu'à la deuxième zone de déviation (18).
EP05701516.6A 2004-01-16 2005-01-14 Aube refroidie pour une turbine a gaz Not-in-force EP1709298B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102004002327A DE102004002327A1 (de) 2004-01-16 2004-01-16 Gekühlte Schaufel für eine Gasturbine
PCT/EP2005/050137 WO2005068783A1 (fr) 2004-01-16 2005-01-14 Aube refroidie pour une turbine a gaz

Publications (2)

Publication Number Publication Date
EP1709298A1 true EP1709298A1 (fr) 2006-10-11
EP1709298B1 EP1709298B1 (fr) 2015-11-11

Family

ID=34716622

Family Applications (1)

Application Number Title Priority Date Filing Date
EP05701516.6A Not-in-force EP1709298B1 (fr) 2004-01-16 2005-01-14 Aube refroidie pour une turbine a gaz

Country Status (6)

Country Link
US (1) US7520724B2 (fr)
EP (1) EP1709298B1 (fr)
CN (1) CN100408812C (fr)
DE (1) DE102004002327A1 (fr)
TW (1) TWI356870B (fr)
WO (1) WO2005068783A1 (fr)

Families Citing this family (16)

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Publication number Priority date Publication date Assignee Title
FR2898384B1 (fr) * 2006-03-08 2011-09-16 Snecma Aube mobile de turbomachine a cavite commune d'alimentation en air de refroidissement
US7731483B2 (en) * 2007-08-01 2010-06-08 General Electric Company Airfoil shape for a turbine bucket and turbine incorporating same
US7988420B2 (en) * 2007-08-02 2011-08-02 General Electric Company Airfoil shape for a turbine bucket and turbine incorporating same
ES2398303T3 (es) 2008-10-27 2013-03-15 Alstom Technology Ltd Álabe refrigerado para una turbina de gas y turbina de gas que comprende un tal álabe
EP2230383A1 (fr) 2009-03-18 2010-09-22 Alstom Technology Ltd Aube de turbine avec refroidissement de l'extrémité
WO2013167513A1 (fr) 2012-05-07 2013-11-14 Alstom Technology Ltd Procédé de fabrication d'éléments en superalliages monocristallins (sx) ou solidifiés de manière directionnelle (ds)
US10145269B2 (en) 2015-03-04 2018-12-04 General Electric Company System and method for cooling discharge flow
GB201506728D0 (en) * 2015-04-21 2015-06-03 Rolls Royce Plc Thermal shielding in a gas turbine
GB201512810D0 (en) * 2015-07-21 2015-09-02 Rolls Royce Plc Thermal shielding in a gas turbine
JP5905631B1 (ja) * 2015-09-15 2016-04-20 三菱日立パワーシステムズ株式会社 動翼、これを備えているガスタービン、及び動翼の製造方法
US10683763B2 (en) * 2016-10-04 2020-06-16 Honeywell International Inc. Turbine blade with integral flow meter
US10378363B2 (en) 2017-04-10 2019-08-13 United Technologies Corporation Resupply hole of cooling air into gas turbine blade serpentine passage
US10961854B2 (en) * 2018-09-12 2021-03-30 Raytheon Technologies Corporation Dirt funnel squealer purges
US11021961B2 (en) 2018-12-05 2021-06-01 General Electric Company Rotor assembly thermal attenuation structure and system
US11118462B2 (en) * 2019-01-24 2021-09-14 Pratt & Whitney Canada Corp. Blade tip pocket rib
US11371359B2 (en) 2020-11-26 2022-06-28 Pratt & Whitney Canada Corp. Turbine blade for a gas turbine engine

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US4278400A (en) * 1978-09-05 1981-07-14 United Technologies Corporation Coolable rotor blade
FR2468727A1 (fr) * 1979-10-26 1981-05-08 Snecma Perfectionnement aux aubes de turbine refroidies
US4775296A (en) * 1981-12-28 1988-10-04 United Technologies Corporation Coolable airfoil for a rotary machine
US4474532A (en) * 1981-12-28 1984-10-02 United Technologies Corporation Coolable airfoil for a rotary machine
EP0340149B1 (fr) * 1988-04-25 1993-05-19 United Technologies Corporation Moyens de dépoussiérage pour une aube refroidie par de l'air
WO1995014848A1 (fr) * 1993-11-24 1995-06-01 United Technologies Corporation Profil de turbine a refroidissement ameliore
US5915923A (en) * 1997-05-22 1999-06-29 Mitsubishi Heavy Industries, Ltd. Gas turbine moving blade
JPH11241602A (ja) * 1998-02-26 1999-09-07 Toshiba Corp ガスタービン翼
US6832889B1 (en) * 2003-07-09 2004-12-21 General Electric Company Integrated bridge turbine blade
US6966756B2 (en) * 2004-01-09 2005-11-22 General Electric Company Turbine bucket cooling passages and internal core for producing the passages
US7137780B2 (en) * 2004-06-17 2006-11-21 Siemens Power Generation, Inc. Internal cooling system for a turbine blade

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Also Published As

Publication number Publication date
CN1910343A (zh) 2007-02-07
CN100408812C (zh) 2008-08-06
TWI356870B (en) 2012-01-21
WO2005068783A1 (fr) 2005-07-28
US7520724B2 (en) 2009-04-21
EP1709298B1 (fr) 2015-11-11
TW200532096A (en) 2005-10-01
DE102004002327A1 (de) 2005-08-04
US20060292006A1 (en) 2006-12-28

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