EP1512489B1 - Aube pour turbine - Google Patents

Aube pour turbine Download PDF

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Publication number
EP1512489B1
EP1512489B1 EP03020211A EP03020211A EP1512489B1 EP 1512489 B1 EP1512489 B1 EP 1512489B1 EP 03020211 A EP03020211 A EP 03020211A EP 03020211 A EP03020211 A EP 03020211A EP 1512489 B1 EP1512489 B1 EP 1512489B1
Authority
EP
European Patent Office
Prior art keywords
blade
slot
trailing edge
length
turbine blade
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.)
Expired - Lifetime
Application number
EP03020211A
Other languages
German (de)
English (en)
Other versions
EP1512489A1 (fr
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 DE50306044T priority Critical patent/DE50306044D1/de
Priority to EP03020211A priority patent/EP1512489B1/fr
Priority to US10/931,089 priority patent/US7160084B2/en
Publication of EP1512489A1 publication Critical patent/EP1512489A1/fr
Application granted granted Critical
Publication of EP1512489B1 publication Critical patent/EP1512489B1/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
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/142—Shape, i.e. outer, aerodynamic form of the blades of successive rotor or stator blade-rows
    • F01D5/143—Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour
    • 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

Definitions

  • the invention relates to a, according to the preamble of claim 1, directed along a blade axis and formed from a base blade.
  • Such a blade is known from US-2003/0138322 A.
  • a flow medium is transported in a flow channel in order to gain energy from this.
  • turbine blades are arranged in the flow channel.
  • vanes vane rings and rotor vanes formed from rotor blades are arranged in the flow channel.
  • the vanes suitably direct the flow medium onto the blades, which are connected to a rotor and rotated so that kinetic energy of the flow medium is converted to rotational energy.
  • Such blades in turbomachines are often exposed to considerable mechanical stress. Especially at the same time high temperature and high speeds, such as in a gas turbine, it comes to a high material stress of the blade material. As a result, cracks may form in the blade material, which spread over time with continued stress. Eventually, the blade may fail, shattering the blade or breaking up debris. For subsequent blades in the flow direction, this can lead to considerable damage. Cracking and crack propagation should therefore be monitored. Depending on the speed of the processes, this can lead to a significant reduction in the availability of the turbine come as regular service intervals lead to downtime of the turbine.
  • JP 2000018001 a gas turbine blade is shown, are introduced in the relief slots in the direction of the blade axis to the edge of the head area. These relief slots are used to reduce thermal stresses in this area. The reduction of thermal stresses should reduce cracking.
  • the relief slots are limited to the head area.
  • JP 10299408 shows a gas turbine blade in which elliptical holes are introduced in areas of high thermal stresses, which are intended to reduce crack propagation.
  • the bores are arranged in the transition region of the blade and the platform, wherein in the airfoil region the ellipse main axis is directed perpendicular to the blade axis. A corresponding orientation of the holes can be found at the trailing edge.
  • the object of the invention is the specification of a turbine blade which is exposed to particularly low thermal stresses.
  • a turbine blade which is directed along a blade axis and formed from a base body, comprising a foot region A head region and an airfoil having a blade height reaching from the root region to the tip region and having a blade width reaching from a blade leading edge to a blade trailing edge, wherein a fillet is formed in a transition region between the blade trailing edge and the root region, wherein a relief slot is formed across the blade trailing edge ,
  • the invention is based on the recognition that the blade trailing edge of a turbine blade itself is exposed to particularly high mechanical stresses in a region above the rounded transition region between the blade trailing edge and the base region and in this rounded transition region itself. Furthermore, the invention is based on the finding that the blade trailing edge is not unduly mechanically destabilized by slots extending transversely to it with appropriate dimensioning. By introducing a relief slot transversely to the blade trailing edge through the blade trailing edge, a considerable relief from thermal stresses is now achieved in that thermal expansion can be compensated through the slot.
  • the relief slot is in the vicinity of the rounding. Especially in an area in the vicinity of the rounding the blade trailing edge is subject to particularly high thermal stresses.
  • the relief slot effectively dissipates these stresses in this particularly affected area.
  • the relief slot is located less than 20% of the blade height of the rounding. Particularly preferred is a distance of the relief slot from the rounding of less than 10% of the blade height.
  • the slot has a length of at least 2% of the blade width. In this extension, a particularly high effect in the discharge through the slot is achieved.
  • the slot has at most a length of 5% of the sheet width.
  • a greater extension of the slot length than this 5% of the blade width leads only to a comparatively small further relief of thermal stresses, while on the other hand would suffer from too great a slot length, the mechanical stability of the blade trailing edge.
  • At least two, more preferably at least three relief slots are provided. With more than two or three relief slots following each other along the blade axis, a larger portion of the blade trailing edge can be relieved of thermal stresses. In addition, higher overall thermal stresses can be counteracted. Preferably, all relief slots in an area smaller than 25% of the blade height are removed from the rounding.
  • Three relief slots are preferably provided, wherein a first slot lying closest to the rounding has a first length, a second slot following the first slot has a second length and a third slot has a third length following the second slot along the blade axis wherein the third length is greater than the second length and the second length is greater than the first length.
  • the turbine blade is a gas turbine blade.
  • Gas turbine blades are exposed to particularly high temperatures. Accordingly, it is precisely here to build particularly high thermal stresses.
  • the relief slot preferably has an approximately circular extension on its end opposite the blade trailing edge. Such a circular widening reduces the radii of curvature of the surfaces delimiting the slot at the end and thus reduces the stresses that occur particularly at such curvatures.
  • the circular extension is a circular bore from which the slot extends through the blade trailing edge.
  • the slot is preferably cut by means of a laser beam or it is milled.
  • the gas turbine 1 shows a gas turbine 1.
  • the gas turbine 1 is directed along a turbine axis 10 and has along the turbine axis 10 successively a compressor 3, a Combustion chamber 5 and a turbine part 7.
  • the compressor 3 and the turbine part 7 are arranged on a common turbine shaft 9.
  • a hot gas duct 12 is formed in the guide vanes 11 and blades 13, which are arranged on the turbine shaft 9, protrude into it.
  • the compressor air 15 is burned with fuel in the combustion chamber 5 to hot gas 17, which flows through the hot gas channel 12. It puts the turbine shaft 9 in motion via the effect on the rotor blades 13.
  • the rotational energy of the turbine shaft 9 may be e.g. be used for generating electrical energy.
  • FIG. 2 shows a gas turbine guide vane 31.
  • the gas turbine guide vane 31 has a foot region 33 with a platform 34.
  • an airfoil 35 connects.
  • the airfoil 35 ends in a head region 37, which in particular also has a platform, which is not shown here.
  • the platform 34 and also the illustrated platform of the head region 37 serve to delimit the hot gas channel 12.
  • the airfoil 35 has a blade height h .
  • the airfoil 35 has a blade width b .
  • the blade airfoil 35 extends from a blade leading edge 39 to a blade trailing edge 41.
  • the gas turbine guide blade 31 has a base body 32 which is hollow a blade outer wall 63 encloses the cavity. Stabilizing side ribs 65 are arranged in the cavity between the suction side 47 and the pressure side 45.
  • a rounding 71 is formed in the region of the blade leading edge 39, and a rounding 73 is formed in the region of the blade trailing edge 41.
  • These fillets 71, 73 or also called thickenings or notches, are exposed to particularly high mechanical stresses during operation.
  • discharge slots 51 are provided in the blade trailing edge.
  • FIG. 3 shows a detail of a longitudinal section through the gas turbine guide blade 31 in the region of the rounding 73 between the blade trailing edge 41 and the platform 34.
  • the relief slots 51 extend transversely to the blade trailing edge 41 through the blade trailing edge 41.
  • the blade trailing edge 41 may be formed, for example, solely by the suction side 47, while the pressure side 45 ends in a stepped manner and cooling air openings are provided in this stage, which cool the blade trailing edge 41. This would be an open blade trailing edge 41.
  • a closed blade trailing edge 41 may also be present in which the pressure side 45, rounded, merges into the suction side 47 and thereby forms the blade trailing edge 41.
  • the relief slots 51 in the suction side 47, the pressure side 45 or in both sides extend.
  • the relief slots 51 end with their blade trailing edge 41 opposite ends in circular extensions 53, in which comparatively little stress caused by a relatively low curvature.
  • the relief slot 51 closest to the rounding has a smaller amount than the second relief slot following in the blade axis direction.
  • the second relief slot is again shorter than the one in the direction of the blade axis following and the rounding 73 farthest third relief slot 51st
  • thermal stresses are reduced by a thermal expansion can be compensated in the discharge slots 51.
  • thermal stresses are minimized both in the region of the trailing edge 4 and in the rounding 73.
  • cooling air 67 is introduced for cooling.
  • This cooling air 67 exits the slot 51 from the hollow interior of the gas turbine blade 34.
  • the slot 51 is shaped so that the cooling air 67 forms a cooling film on the surface of the airfoil 35.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Claims (10)

  1. Aube (31) de turbine dirigée suivant un axe (30) d'aube et formée d'une pièce (32) de base, comprenant une partie (33) d'emplanture, une partie (37) de tête et une lame (35) d'aube ayant une hauteur h de lame allant de la partie (33) d'emplanture à la partie (37) de tête et ayant une largeur b de lame allant d'un bord (39) avant d'aube à un bord (41) arrière d'aube, un arrondi (73) étant formé dans la partie de transition entre le bord (41) arrière d'aube et la partie (33) d'emplanture,
    caractérisée en ce que
    une fente (51) de soulagement traverse le bord (41) arrière d'aube.
  2. Aube (31) de turbine suivant la revendication 1,
    dans laquelle la fente (51) de soulagement est à proximité de l'arrondi (73).
  3. Aube (31) de turbine suivant la revendication 2,
    dans laquelle la fente (51) de soulagement est éloignée de l'arrondi (73) de moins de 20 % de la hauteur h de la lame.
  4. Aube (31) de turbine suivant l'une des revendications précédentes,
    dans laquelle la fente (51) a une longueur représentant au moins 2 % de la largeur b de la lame.
  5. Aube (31) de turbine suivant l'une des revendications précédentes,
    dans laquelle la fente (51) a une largeur représentant 5 % au plus de la larguer b de la lame.
  6. Aube (31) de turbine suivant l'une des revendications précédentes,
    dans laquelle il est prévu au moins deux fentes (51) de soulagement.
  7. Aube (31) de turbine suivant l'une des revendications précédentes,
    dans laquelle il est prévu au moins trois fentes (51) de soulagement.
  8. Aube (31) de turbine suivant l'une des revendications précédentes, constituée sous la forme d'une aube de turbine à gaz.
  9. Aube (31) de turbine suivant l'une des revendications précédentes,
    dans laquelle la fente (51) de soulagement a du côté opposé au bord (41) arrière d'aube, un élargissement (53) à peu près circulaire.
  10. Aube (31) de turbine suivant la revendication 7,
    dans laquelle une première fente (51) la plus proche de l'arrondi (73) a une première longueur une deuxième fente (52) suivant la première fente (51), le long de l'axe (30) de l'aube, a une deuxième longueur et une troisième fente (51) suivant la deuxième fente, le long de l'axe (30) de l'aube, a une troisième longueur, la troisième longueur étant plus grande que la deuxième longueur et la deuxième longueur étant plus grande que la première longueur.
EP03020211A 2003-09-05 2003-09-05 Aube pour turbine Expired - Lifetime EP1512489B1 (fr)

Priority Applications (3)

Application Number Priority Date Filing Date Title
DE50306044T DE50306044D1 (de) 2003-09-05 2003-09-05 Schaufel einer Turbine
EP03020211A EP1512489B1 (fr) 2003-09-05 2003-09-05 Aube pour turbine
US10/931,089 US7160084B2 (en) 2003-09-05 2004-08-31 Blade of a turbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP03020211A EP1512489B1 (fr) 2003-09-05 2003-09-05 Aube pour turbine

Publications (2)

Publication Number Publication Date
EP1512489A1 EP1512489A1 (fr) 2005-03-09
EP1512489B1 true EP1512489B1 (fr) 2006-12-20

Family

ID=34130157

Family Applications (1)

Application Number Title Priority Date Filing Date
EP03020211A Expired - Lifetime EP1512489B1 (fr) 2003-09-05 2003-09-05 Aube pour turbine

Country Status (3)

Country Link
US (1) US7160084B2 (fr)
EP (1) EP1512489B1 (fr)
DE (1) DE50306044D1 (fr)

Families Citing this family (25)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
KR101070904B1 (ko) * 2004-08-20 2011-10-06 삼성테크윈 주식회사 레이디얼 터빈 휠
CA2503879A1 (fr) * 2005-04-07 2006-10-07 General Electric Canada Rainures de detente pour aubes de roue mobile de turbine francis
EP1757773B1 (fr) * 2005-08-26 2008-03-19 Siemens Aktiengesellschaft Aube creuse de turbine
US9840931B2 (en) * 2008-11-25 2017-12-12 Ansaldo Energia Ip Uk Limited Axial retention of a platform seal
EP2299056A1 (fr) * 2009-09-02 2011-03-23 Siemens Aktiengesellschaft Refroidissement d'un composant de turbine à gaz sous la forme d'un disque de rotor ou d'une aube
US8556584B2 (en) * 2011-02-03 2013-10-15 General Electric Company Rotating component of a turbine engine
CH705838A1 (de) * 2011-12-05 2013-06-14 Alstom Technology Ltd Abgasgehäuse für eine Gasturbine sowie Gasturbine mit einem Abgasgehäuse.
US20130177430A1 (en) * 2012-01-05 2013-07-11 General Electric Company System and method for reducing stress in a rotor
US9228448B2 (en) * 2013-09-20 2016-01-05 United Technologies Corporation Background radiation measurement system
EP2863010A1 (fr) * 2013-10-21 2015-04-22 Siemens Aktiengesellschaft Aube de turbine
US9506365B2 (en) 2014-04-21 2016-11-29 Honeywell International Inc. Gas turbine engine components having sealed stress relief slots and methods for the fabrication thereof
CA2954716C (fr) * 2014-07-14 2022-10-18 Lm Wp Patent Holding A/S Coin profile pour fixation d'une piece d'extension d'enveloppe aerodynamique
US20160047251A1 (en) * 2014-08-13 2016-02-18 United Technologies Corporation Cooling hole having unique meter portion
EP2990598A1 (fr) * 2014-08-27 2016-03-02 Siemens Aktiengesellschaft Aube de turbine et turbine
US11280214B2 (en) 2014-10-20 2022-03-22 Raytheon Technologies Corporation Gas turbine engine component
DE102015207760A1 (de) * 2015-04-28 2016-11-03 Siemens Aktiengesellschaft Heißgasführendes Gehäuse
CN108430691B (zh) * 2015-12-21 2021-03-16 通用电气公司 修理好的涡轮机部件和对应修理方法
US10415408B2 (en) * 2016-02-12 2019-09-17 General Electric Company Thermal stress relief of a component
FR3048998B1 (fr) * 2016-03-16 2019-12-13 Safran Aircraft Engines Rotor de turbine comprenant une entretoise de ventilation
FR3056630B1 (fr) * 2016-09-26 2018-12-07 Safran Aircraft Engines Disque aubage monobloc de soufflante pour turbomachine d'aeronef
DE102017208707A1 (de) * 2017-05-23 2018-11-29 Siemens Aktiengesellschaft Verfahren zur Herstellung einer Turbinenschaufel
KR102048874B1 (ko) * 2018-04-09 2019-11-26 두산중공업 주식회사 유연성이 향상된 터빈 베인
DE102019103640A1 (de) 2019-02-13 2020-08-13 Mitsubishi Hitachi Power Systems Europe Gmbh Brennstoffdüse mit Dehnungsschlitzen für einen Kohlenstaubbrenner
US20210115796A1 (en) * 2019-10-18 2021-04-22 United Technologies Corporation Airfoil component with trailing end margin and cutback
CN114227180A (zh) * 2021-12-29 2022-03-25 哈尔滨汽轮机厂有限责任公司 一种提升汽轮机叶片加工精度的方法

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CA2334071C (fr) * 2000-02-23 2005-05-24 Mitsubishi Heavy Industries, Ltd. Aube mobile de turbine a gaz
US6490791B1 (en) 2001-06-22 2002-12-10 United Technologies Corporation Method for repairing cracks in a turbine blade root trailing edge
FR2835015B1 (fr) * 2002-01-23 2005-02-18 Snecma Moteurs Aube mobile de turbine haute pression munie d'un bord de fuite au comportement thermique ameliore
US6929451B2 (en) * 2003-12-19 2005-08-16 United Technologies Corporation Cooled rotor blade with vibration damping device

Also Published As

Publication number Publication date
US20050106028A1 (en) 2005-05-19
US7160084B2 (en) 2007-01-09
EP1512489A1 (fr) 2005-03-09
DE50306044D1 (de) 2007-02-01

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