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EP1757773A1 - Aube creuse de turbine - Google Patents

Aube creuse de turbine Download PDF

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Publication number
EP1757773A1
EP1757773A1 EP05018595A EP05018595A EP1757773A1 EP 1757773 A1 EP1757773 A1 EP 1757773A1 EP 05018595 A EP05018595 A EP 05018595A EP 05018595 A EP05018595 A EP 05018595A EP 1757773 A1 EP1757773 A1 EP 1757773A1
Authority
EP
European Patent Office
Prior art keywords
turbine blade
profile
profile wall
slot
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.)
Granted
Application number
EP05018595A
Other languages
German (de)
English (en)
Other versions
EP1757773B1 (fr
Inventor
Fathi Ahmad
Michael Dr. Dankert
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 ES05018595T priority Critical patent/ES2303163T3/es
Priority to DE502005003344T priority patent/DE502005003344D1/de
Priority to EP05018595A priority patent/EP1757773B1/fr
Priority to JP2006228883A priority patent/JP4689558B2/ja
Priority to CNA2006101463720A priority patent/CN1936273A/zh
Priority to US11/510,239 priority patent/US7845905B2/en
Publication of EP1757773A1 publication Critical patent/EP1757773A1/fr
Application granted granted Critical
Publication of EP1757773B1 publication Critical patent/EP1757773B1/fr
Not-in-force legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • F01D5/16—Form or construction for counteracting blade vibration
    • 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 hollow turbine blade with an airfoil profile formed by a suction-side profile wall and a pressure-side profile wall and having a profile height directed along a blade axis from a platform to a profile tip, with at least one support rib provided in the interior of the turbine blade which connects the pressure-side profile wall with the suction-side profile wall in each case in a connection region and with at least one slot provided in the profile wall on the hot gas side and extending along the blade axis. Furthermore, the invention relates to the use of a generic turbine blade.
  • the object of the invention is therefore to provide a turbine blade with extended life.
  • the object is achieved by a generic turbine blade in which the slot opposite hot gas side in the profile wall of the connecting rib formed by the support rib and the profile wall.
  • the invention is based on the recognition that the material of the airfoil profile heats up due to the outside flowing hot gas.
  • the extending between the pressure-side profile wall and the suction-side profile wall inside support rib is colder than the heated material of the profile walls.
  • the support rib merges integrally into the pressure-side or suction-side profile wall, heat energy is locally conducted from the respective profile wall into the support rib via the inner connection region and dissipated, so that a reduced material temperature along the overflow in the profile wall in the mouth region of the support rib Profile height extending connection area occurs.
  • the profile wall is hotter in many areas. As a result, thermal stresses occur in the material, which can cause crack generation and promote crack growth.
  • the invention proposes that the slot opposite the hot gas side in the profile wall of the connecting rib formed by the support rib and the profile wall.
  • the slot relieves the material by allowing locally larger thermal expansion of the profile wall.
  • the relief slot leads to a life-prolonging reduction of the thermally induced stresses in the profile wall.
  • the thermal stresses still occurring in the airfoil then occur in an order of magnitude that is harmless to the material. Cracks and / or crack growth are less likely to occur at this point, extending the life of the turbine blade.
  • the slot can also serve as a crack stopper or
  • Crack limiter serve, which in turn can extend the life of the turbine blade.
  • a gas turbine equipped with this long-life turbine blade has a longer operating time and a shorter service life, since the turbine blades are less likely to be examined for cracks with critical lengths and replaced if necessary.
  • the maintenance costs of gas turbines can be lowered and their cost-effectiveness can be further improved.
  • the slot extends along the blade axis and has at least a length of 10%, preferably at least 20% of the profile height.
  • this measure extends the service life of the gas turbine, since the support ribs provided in the interior of the turbine blade also extend along the blade axis and connect the pressure-side profile wall with the suction-side profile wall in each case in a connection region.
  • the slot or slits may also extend into the transition region.
  • the transition region can be protected from cracking.
  • crack propagation is delayed or limited in the transition region.
  • the slot provided in the hot gas-circumscribed outer surface can also extend beyond the transition region into the platform.
  • the locally occurring cold entry ie the heat extraction locally occurring due to the cooler support rib
  • the locally occurring cold entry can be lowered in a particularly effective manner, whereby the material of the profile wall, between the connection region and the outer surface opposite thereto, in comparison to The state of the art is warmer.
  • the thermal stresses are reduced, which leads to an extension of the service life of the turbine blade.
  • the slot is filled with a filler material to avoid possibly occurring aerodynamic losses in the hot gas due to edges.
  • the filling material is softer than the material of the profile wall. The occurring thermally induced strains of the profile wall can be compensated particularly well in this case of the soft filler.
  • FIG. 1 shows a gas turbine 1 in a longitudinal partial section. It has inside a rotatably mounted about a rotation axis 2 rotor 3, which is also referred to as a turbine runner. Along the rotor 3 successive an intake 4, a compressor 5, a toroidal annular combustion chamber 6 with a plurality of rotationally symmetrical to each other arranged burners 7, a turbine unit 8 and an exhaust housing 9.
  • the annular combustion chamber 6 forms a combustion chamber 17, which with a annular hot gas channel 18 communicates.
  • There four successive turbine stages 10 form the turbine unit 8. Each turbine stage 10 is formed of two blade rings.
  • the turbine blade 30 has a platform 32, on the surface 34 of which a hot air 11 umströmbares airfoil 36 is arranged.
  • the airfoil 36 extends from a leading edge 38 to a trailing edge 40. It also has an intake side profile wall 42 extending therebetween and a pressure side profile wall 44 also extending therebetween.
  • three cavities 46 are provided, which are separated from each other by two support ribs 48.
  • the support ribs 48 connect the suction-side profile wall 42 with the pressure-side profile wall 44 and serve to increase the rigidity of the airfoil profile 36.
  • the turbine blade 30 is typically manufactured by a casting process. In a casting apparatus three casting cores are used for this purpose, which are removed after the manufacture of the turbine blade 30 from this. At this point, the cavities remain 46, between which the support ribs 48 are arranged. Therefore, with cast turbine blade 30, the support ribs 48 integrally merge into and interconnect with the suction side and / or pressure side profile walls 42, 44 in a connection region 50, providing particularly good thermal coupling of the profile wall 42, 44 to the support rib.
  • the airfoil 36 surrounded by the hot gas 11 is completely heated.
  • a temperature profile has hitherto been established which has a local temperature minimum in the region of each support rib 48. This caused by the cooler support rib 48 uneven heating of the airfoil profile 36 caused in the near-surface portion of the profile walls 42, 44 so large, thermally induced stresses that there cracks could occur and crack growth occurred more. This limited the life of the known turbine blade.
  • the slot 56 provided in a profile wall 42, 44 is now arranged in a section of the profile wall 42, 44 which corresponds to the connection area 50 and thus also the support rib 48 is opposite.
  • the slot 56 raises the local temperature minimum occurring in its area since the thermal conductivity of the connection area 50 has been reduced due to the reduced cross-section. Accordingly, the temperature gradients along the tread walls 42, 44 are reduced from the leading edge 38 to the trailing edge 40, which reduces stress in the portion having the slot 50. The thermal stresses are then at a harmless level and the material of the airfoil 36 can thereby withstand the loads occurring longer.
  • the slots 56 have a minimum length L, which corresponds to at least 10%, preferably at least 20% of the height H of the airfoil 36.
  • the height H of the airfoil 36 is determined between the surface 34 of the platform 32 and the tip 58 of the airfoil 36.
  • the slot 56 can extend into a rounded transition region 60, which is arranged between the platform 32 and the airfoil profile 36.
  • This configuration of the slots 56 is represented by the contours 62 shown in dashed line style.
  • a particularly good protection against crack-like wear can be achieved if the slot 62 also extends into the platform 32.
  • FIG. 3 shows a section through the turbine blade 30 according to the invention along the section line III from FIG. 2.
  • the turbine blade 30 can be designed both as a moving blade and / or as a guide blade for a particularly stationary gas turbine 1.
  • the airfoil profile 36 shown in cross-section shows the leading edge 38, the trailing edge 40, the suction-side profile wall 42, the pressure-side profile wall 44 and two support ribs 48 separating the cavities 46 which merge into the profile walls 42, 44 in a connection region 50.
  • the slots 56 shown are filled with a filling material, as a result of which a particular aerodynamic surface contour of the airfoil profile 36 can be brought about. Transverse to the flow direction of the hot gas 11 extending projections and edges are thereby avoided in the profile walls 42, 44.
  • the slots 56 each protrude into the profile walls 42, 44 with a penetration depth E. This can be so large that the slots 56 protrude into the connection region 50 and possibly even beyond it into the support ribs 48. This ensures that the temperature difference along the airfoil profile 36 from the leading edge 38 to the trailing edge 48 is made particularly effective, so as to further increase the life of the turbine blade 30.
  • the invention is particularly effective when the hollow turbine blade 30 and the airfoil profile 36 are flowed through by a coolant, for example compressor air taken from the compressor 5 of the gas turbine 1.
  • a coolant for example compressor air taken from the compressor 5 of the gas turbine 1.
  • the profile walls 42, 44 are cooled from the inside, as well as the support ribs 48.
  • the unwanted local cold entry or the local heat removal from the profile wall 42, 44 via the connecting portion 50 and the support ribs 48 is due to the particularly good thermoelectric coupling particularly effective. Accordingly, the temperature differences along the profile walls 42, 44 and thus also the thermal stresses in an internally cooled turbine blade 30 are particularly large.
  • the life, in particular of internally cooled turbine blades 30 can be extended particularly effective with the invention.
  • the serving for thermal relief slots 56 may also be provided in only one profile wall, for example, the suction-side profile wall 42 or the pressure-side profile wall 44.
  • the slots 56, 62 serve as boundaries for cracks occurring in the adjacent blade material. For example, if in the region of the central cavity 46 in one of the two profile walls 42, 44, a crack is present and this lengthens in the flow direction of the hot gas 11, this forcibly increased to a maximum of one of the two slots 56. The extension of the crack on the slot 56 is not possible.
  • the invention provides a measure for equalizing the thermal stress in a wing profile 36 of a turbine blade 30 in order to increase the service life of the turbine blade 30 and, accordingly, the operating times of a gas turbine 1 equipped therewith.
  • the invention proposes that the hollow turbine blade 30 in the region of the support ribs 48, which has a suction-side profile wall 42 with a pressure-side profile wall 44 in each case connect a connecting portion 50 together, has hot gas side arranged slots 56 for relief.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP05018595A 2005-08-26 2005-08-26 Aube creuse de turbine Not-in-force EP1757773B1 (fr)

Priority Applications (6)

Application Number Priority Date Filing Date Title
ES05018595T ES2303163T3 (es) 2005-08-26 2005-08-26 Alabe de turbina hueco.
DE502005003344T DE502005003344D1 (de) 2005-08-26 2005-08-26 Hohle Turbinenschaufel
EP05018595A EP1757773B1 (fr) 2005-08-26 2005-08-26 Aube creuse de turbine
JP2006228883A JP4689558B2 (ja) 2005-08-26 2006-08-25 中空タービン翼
CNA2006101463720A CN1936273A (zh) 2005-08-26 2006-08-25 空心透平叶片
US11/510,239 US7845905B2 (en) 2005-08-26 2006-08-25 Hollow turbine blade

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP05018595A EP1757773B1 (fr) 2005-08-26 2005-08-26 Aube creuse de turbine

Publications (2)

Publication Number Publication Date
EP1757773A1 true EP1757773A1 (fr) 2007-02-28
EP1757773B1 EP1757773B1 (fr) 2008-03-19

Family

ID=36584564

Family Applications (1)

Application Number Title Priority Date Filing Date
EP05018595A Not-in-force EP1757773B1 (fr) 2005-08-26 2005-08-26 Aube creuse de turbine

Country Status (6)

Country Link
US (1) US7845905B2 (fr)
EP (1) EP1757773B1 (fr)
JP (1) JP4689558B2 (fr)
CN (1) CN1936273A (fr)
DE (1) DE502005003344D1 (fr)
ES (1) ES2303163T3 (fr)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2990598A1 (fr) * 2014-08-27 2016-03-02 Siemens Aktiengesellschaft Aube de turbine et turbine
EP2998507A1 (fr) * 2014-09-16 2016-03-23 Siemens Aktiengesellschaft Une aube de turbine refroidie comportant des entretoises internes entre les cavités de refroidissement qui comportent des points de rupture pour reduir les gradients thermiques
DE102017208707A1 (de) 2017-05-23 2018-11-29 Siemens Aktiengesellschaft Verfahren zur Herstellung einer Turbinenschaufel
EP3473431A1 (fr) * 2017-10-23 2019-04-24 MTU Aero Engines GmbH Pale et rotor pour une turbomachine ainsi que turbomachine

Families Citing this family (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE502006003548D1 (de) * 2006-08-23 2009-06-04 Siemens Ag Beschichtete Turbinenschaufel
US8500411B2 (en) * 2010-06-07 2013-08-06 Siemens Energy, Inc. Turbine airfoil with outer wall thickness indicators
EP2863010A1 (fr) * 2013-10-21 2015-04-22 Siemens Aktiengesellschaft Aube de turbine
CN104006000A (zh) * 2014-05-29 2014-08-27 安徽银龙泵阀股份有限公司 一种内置冷却管的泵芯
CN104747241A (zh) * 2015-01-23 2015-07-01 朱晓义 一种压力装置及涡轮发动机
US10731469B2 (en) 2016-05-16 2020-08-04 Raytheon Technologies Corporation Method and apparatus to enhance laminar flow for gas turbine engine components
US20170328206A1 (en) * 2016-05-16 2017-11-16 United Technologies Corporation Method and Apparatus to Enhance Laminar Flow for Gas Turbine Engine Components
US20180051566A1 (en) * 2016-08-16 2018-02-22 General Electric Company Airfoil for a turbine engine with a porous tip
CN106703897B (zh) * 2016-12-21 2018-01-12 中国南方航空工业(集团)有限公司 一种空心叶片内腔低熔点合金清理装置
CN106757044B (zh) * 2016-12-21 2018-12-14 中国南方航空工业(集团)有限公司 一种空心叶片内腔低熔点合金清理方法
CN106761953B (zh) * 2016-12-21 2018-05-08 中国南方航空工业(集团)有限公司 一种叶片内腔低熔点合金去除装置
CN106757045B (zh) * 2016-12-21 2018-12-14 中国南方航空工业(集团)有限公司 一种叶片内腔低熔点合金去除方法
US11118462B2 (en) 2019-01-24 2021-09-14 Pratt & Whitney Canada Corp. Blade tip pocket rib
JP6799702B1 (ja) * 2020-03-19 2020-12-16 三菱パワー株式会社 静翼及びガスタービン
US11371359B2 (en) 2020-11-26 2022-06-28 Pratt & Whitney Canada Corp. Turbine blade for a gas turbine engine

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JPH1061406A (ja) * 1996-08-27 1998-03-03 Mitsubishi Heavy Ind Ltd ガスタービン翼の冷却構造
JP2000018001A (ja) * 1998-06-30 2000-01-18 Mitsubishi Heavy Ind Ltd 動翼熱応力軽減装置
EP1508399A1 (fr) 2003-08-22 2005-02-23 Siemens Aktiengesellschaft Aube de turbomachine et procédé pour empêcher la propagation des fissures dans une aube de turbomachine
EP1512489A1 (fr) * 2003-09-05 2005-03-09 Siemens Aktiengesellschaft Aube pour turbine
EP1525942A1 (fr) * 2003-10-23 2005-04-27 Siemens Aktiengesellschaft Turbine à gas et aube mobile pour turbomachine

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Publication number Priority date Publication date Assignee Title
JPH1061406A (ja) * 1996-08-27 1998-03-03 Mitsubishi Heavy Ind Ltd ガスタービン翼の冷却構造
JP2000018001A (ja) * 1998-06-30 2000-01-18 Mitsubishi Heavy Ind Ltd 動翼熱応力軽減装置
EP1508399A1 (fr) 2003-08-22 2005-02-23 Siemens Aktiengesellschaft Aube de turbomachine et procédé pour empêcher la propagation des fissures dans une aube de turbomachine
EP1512489A1 (fr) * 2003-09-05 2005-03-09 Siemens Aktiengesellschaft Aube pour turbine
EP1525942A1 (fr) * 2003-10-23 2005-04-27 Siemens Aktiengesellschaft Turbine à gas et aube mobile pour turbomachine

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Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2990598A1 (fr) * 2014-08-27 2016-03-02 Siemens Aktiengesellschaft Aube de turbine et turbine
WO2016030449A1 (fr) * 2014-08-27 2016-03-03 Siemens Aktiengesellschaft Aube de turbine et turbine
EP2998507A1 (fr) * 2014-09-16 2016-03-23 Siemens Aktiengesellschaft Une aube de turbine refroidie comportant des entretoises internes entre les cavités de refroidissement qui comportent des points de rupture pour reduir les gradients thermiques
WO2016041761A1 (fr) * 2014-09-16 2016-03-24 Siemens Aktiengesellschaft Aube de turbine refroidie pourvue, entre les compartiments de refroidissement, de nervures de raccordement internes présentant des points de rupture destinés à la réduction de tensions thermiques
US10287892B2 (en) 2014-09-16 2019-05-14 Siemens Aktiengesellschaft Turbine blade and turbine
DE102017208707A1 (de) 2017-05-23 2018-11-29 Siemens Aktiengesellschaft Verfahren zur Herstellung einer Turbinenschaufel
EP3473431A1 (fr) * 2017-10-23 2019-04-24 MTU Aero Engines GmbH Pale et rotor pour une turbomachine ainsi que turbomachine
US10844726B2 (en) 2017-10-23 2020-11-24 MTU Aero Engines AG Blade and rotor for a turbomachine and turbomachine

Also Published As

Publication number Publication date
EP1757773B1 (fr) 2008-03-19
JP2007064219A (ja) 2007-03-15
US7845905B2 (en) 2010-12-07
JP4689558B2 (ja) 2011-05-25
CN1936273A (zh) 2007-03-28
US20070128035A1 (en) 2007-06-07
DE502005003344D1 (de) 2008-04-30
ES2303163T3 (es) 2008-08-01

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