EP2990598A1 - Aube de turbine et turbine - Google Patents

Aube de turbine et turbine Download PDF

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Publication number
EP2990598A1
EP2990598A1 EP14182462.3A EP14182462A EP2990598A1 EP 2990598 A1 EP2990598 A1 EP 2990598A1 EP 14182462 A EP14182462 A EP 14182462A EP 2990598 A1 EP2990598 A1 EP 2990598A1
Authority
EP
European Patent Office
Prior art keywords
turbine blade
turbine
material recess
blade
wall
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Withdrawn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
EP14182462.3A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
Nihal Kurt
Radan RADULOVIC
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 EP14182462.3A priority Critical patent/EP2990598A1/fr
Priority to US15/504,361 priority patent/US20170234136A1/en
Priority to CN201580045956.6A priority patent/CN106795770B/zh
Priority to PCT/EP2015/069615 priority patent/WO2016030449A1/fr
Priority to EP15756631.6A priority patent/EP3158168B1/fr
Publication of EP2990598A1 publication Critical patent/EP2990598A1/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Withdrawn legal-status Critical Current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
    • 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
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • 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
    • F05D2260/94—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF]
    • 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
    • F05D2260/94—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF]
    • F05D2260/941—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF] particularly aimed at mechanical or thermal stress reduction

Definitions

  • the invention relates to a turbine blade with an internally cooled turbine blade in which a cavity is divided by rib elements in at least one coolant-carrying cooling channel.
  • the invention further relates to a turbine, in particular a gas turbine, with at least one turbine stage comprising a plurality of turbine blades.
  • a turbine blade of this type is equipped with an internally cooled turbine blade in order to withstand even high temperatures prevailing in the turbine, especially in a hot gas turbine, thermally and mechanically.
  • the turbine blades are often thermally and mechanically loaded higher, in which case it hardly plays a role, whether it is the turbine blade to a vane or a blade of the turbine.
  • an internally cooled turbine blade has a cavity through which a cooling medium can be passed.
  • a fin element or a plurality of rib elements is often additionally arranged to form in the cavity at least one cooling channel with an often meandering cooling channel profile.
  • both a front side wall and a corresponding rear wall of the turbine airfoil may be thermo-mechanically highly stressed in the region of a fin member stiffening the turbine airfoil be.
  • partially critical states of stress can be set on the turbine blade, whereby the turbine blade is exposed to particularly disadvantageous loading conditions in some areas, which there can lead to faster material fatigue over time.
  • the transition regions between the rib element and the front or the rear side wall of the turbine blade leaf are to be mentioned in particular.
  • the object of the invention is achieved by a turbine blade with an internally cooled turbine blade in which a cavity is divided by rib elements in at least one coolant-carrying cooling channel, wherein a arranged next to at least one of the rib members on a turbine blade wall material recess is configured such that within the turbine blade leaf voltages can be reduced in a surrounding region of the at least one rib element.
  • the turbine blade on the front side or the back side itself, but also in the rib element, can be used especially in transition areas between the rib element and the outer walls, ie the front or rear side walls per se, especially thermo-mechanically induced stresses are significantly reduced, whereby a material fatigue in this critical areas can be correspondingly well delayed.
  • thermo-mechanical stresses caused by temperature differences between the suction side and the pressure side of the turbine airfoil can be significantly reduced in critical areas of the turbine airfoil.
  • the present material recess is designed such that it enables an improved stress distribution within the rib element, in transition areas between the actual rib element and the front side wall of the turbine blade leaf and / or the rear wall of the turbine blade leaf, but also in the actual outer walls of the turbine blade leaf.
  • a stress reduction of at least 10% or preferably of more than 20% or 25% can be achieved, in particular in critical surrounding areas or regions around the fin element end.
  • the term “material fatigue” particularly includes fatigue cracking, which is caused especially by thermo-mechanical fatigue of the airfoil material.
  • LCF fatigue low cycle fatigue
  • low-load cycle fatigue ie short-term or low-load cycle fatigue
  • the possible number of load cycles can be increased by more than twice, in particular by more than three times, the previous usual number of load cycles.
  • the number of load changes that can be achieved can be considerably increased, and thus, in particular, the risk of premature LCF fatigue can be significantly reduced, if according to the invention a corresponding material recess is provided in the surrounding area of the rib element. It has been shown that by the material recess according to the invention a related LCF life expectancy of a turbine blade can be significantly increased.
  • a preferred embodiment provides that the material recess is arranged in an environmental region of a rib element end that ends free in the cooling channel. Specifically, in an environmental region of a fin element end that terminates exposed in a turbine blade airfoil, increased and / or increased thermo-mechanical stresses may occur and cause faster material fatigue there.
  • a particularly preferred embodiment provides that the material recess is arranged axially in front of a head side of a free end in the cooling channel fin element end on a turbine blade leaf wall. Especially in a region axially in front of the fin element end, which also formulates an inner curve boundary of the cooling channel, higher critical stress states can occur, which then promote early material fatigue there.
  • thermo-mechanical stresses occurring there, in particular in the turbine blade can be reduced more favorably.
  • the material recess may be located at different distances away from the fin member, particularly from the fin member end, especially with respect to different designs of different turbine blades.
  • the material recess is arranged less than 30 mm or less than 20 mm, preferably less than 10 mm, spaced from the fin element on a turbine blade leaf wall.
  • the material recess can hereby reach as far as the rib element or even be worked into the rib element.
  • the rib element may have the material recess at least partially.
  • the material recess is preferably arranged more than 1 mm or 5 mm away from the rib element.
  • the material recess is configured as at least a partial wall thickness reduction of a turbine blade leaf wall.
  • the material recess is designed shell-shaped.
  • the present material recess is arranged, for example, as a concave dent in the turbine blade outer wall.
  • the material recess can be designed differently. It is particularly advantageous if the material recess is designed as at least one concave cavity on a turbine blade leaf wall. A concavely shaped cavity has very little, if any, aerodynamic effect on the turbine bucket blade.
  • the material recess is configured on the inside of a turbine blade leaf wall. Specifically, by means of a concave material recess thermo-mechanical stresses can be advantageously redirected in the surrounding region of the rib member, in particular within the turbine blade outer wall. In addition, at least one material recess which is provided on the inner side of the turbine blade outer wall facing the cavity or the cooling channel becomes fluidically noticeable.
  • the material recess is arranged on the rear wall of the turbine blade. It is understood that the present material recess can be produced with different geometrical base shapes.
  • the material recess has a circular or oval base surface shape.
  • differently configured base shapes may be advantageous.
  • the material recess has a straight elongated or a curved elongated base surface shape.
  • the material recess may also have a base surface shape combined therewith or an entirely different base surface shape.
  • the material recess is characterized by a trough-shaped or cup-shaped recess made on the inside of the turbine blade outer wall.
  • the turning region of the coolant channel in this case corresponds to a curve of the meandering cooling channel profile of the coolant channel.
  • the object of the invention is also achieved by a turbine, in particular a gas turbine, with at least one turbine stage comprising a plurality of turbine blades, wherein the at least one turbine stage with a plurality of turbine blades and / or turbine vanes according to a turbine blade according to one of the features described herein.
  • a turbine whose turbine blades are less affected or endangered by material fatigue can not only operate more reliable and low-maintenance, but also has a total of a longer life, and thus can be operated more economically.
  • each at least partially shown turbine blade 1 is a blade 2 of a hot gas turbine, not shown here.
  • the turbine blade 1 has an internally cooled turbine blade 3, in which case the inside 4 of the front side wall 5 of the turbine blade 3 is shown (FIG. 1).
  • a leading edge portion 6 of the turbine blade 3 is located on the right hand a leading edge portion 6 of the turbine blade 3.
  • Left hand is accordingly a trailing edge region 7 of the turbine blade 3, at which a plurality of cooling air outlet holes 8 (only exemplified) are present.
  • the trailing edge region 7 is particularly in the FIG. 2 only partially shown.
  • the turbine blade 3 has a cavity 10, wherein in the present case, this cavity 10 is only partially illustrated by the inside 4 with respect to FIG.
  • rib elements 11 and 12 there are two rib elements 11 and 12 in the cavity 10, by means of which a multiply wound cooling channel 13 with a meandering cooling channel course inside the cavity 10 is configured.
  • cooling air can be passed as coolant through the turbine blade 3 in order to cool it from the inside.
  • cooling channel 13 flows from a foot area and thus from the direction of an opening 14 (see only FIG. 2 ) of a turbine blade root 15 coming cooling air substantially directly to the leading edge region 6 facing first cooling channel portion 16 and the rear edge region 7 facing another cooling channel portion 17th
  • the meandering cooling channel course of the coiled cooling channel 13 is configured, at least in the region of the partial view shown, in particular by the two rib elements 11 and 12, whereby the first rib element 11 spatially separates the two cooling channel sections 16 and 17.
  • the first rib element 11 terminates freely in the cooling channel 13 with its rib element end 24 defined by its head side 23, specifically in the turning region 19.
  • thermo-mechanical stress states in particular in the transition areas between the first rib element 11 and the front side wall 5 of the turbine blade 3 and / or the rear side wall of the turbine blade 3, which may cause increased material fatigue there.
  • a material recess 29 is formed on the inner side 4 in order to achieve an advantageous stress reduction in this surrounding region 28 of the rib element end 24.
  • the material recess 29 is arranged at a distance of less than 10 mm from the head side 23 axially in front of the rib element end 24.
  • the material recess 29 is designed as a concave, trough-shaped cavity 30 with a substantially oval base surface (not explicitly numbered) on the inside 4 of the front side wall 5 of the turbine blade 3.
  • the material recess 29 also embodies a partial wall thickness reduction on the front side wall 5 of the turbine blade 3.
  • reinforcing and guide rib elements 37 (numbered only by way of example) and reinforcing web elements 38 (numbered only by way of example), which additionally stabilize the turbine blade 3 in the thinner trailing edge region 7.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP14182462.3A 2014-08-27 2014-08-27 Aube de turbine et turbine Withdrawn EP2990598A1 (fr)

Priority Applications (5)

Application Number Priority Date Filing Date Title
EP14182462.3A EP2990598A1 (fr) 2014-08-27 2014-08-27 Aube de turbine et turbine
US15/504,361 US20170234136A1 (en) 2014-08-27 2015-08-27 Turbine blade and turbine
CN201580045956.6A CN106795770B (zh) 2014-08-27 2015-08-27 涡轮叶片和涡轮机
PCT/EP2015/069615 WO2016030449A1 (fr) 2014-08-27 2015-08-27 Aube de turbine et turbine
EP15756631.6A EP3158168B1 (fr) 2014-08-27 2015-08-27 Aube de turbine et turbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP14182462.3A EP2990598A1 (fr) 2014-08-27 2014-08-27 Aube de turbine et turbine

Publications (1)

Publication Number Publication Date
EP2990598A1 true EP2990598A1 (fr) 2016-03-02

Family

ID=51398570

Family Applications (2)

Application Number Title Priority Date Filing Date
EP14182462.3A Withdrawn EP2990598A1 (fr) 2014-08-27 2014-08-27 Aube de turbine et turbine
EP15756631.6A Active EP3158168B1 (fr) 2014-08-27 2015-08-27 Aube de turbine et turbine

Family Applications After (1)

Application Number Title Priority Date Filing Date
EP15756631.6A Active EP3158168B1 (fr) 2014-08-27 2015-08-27 Aube de turbine et turbine

Country Status (4)

Country Link
US (1) US20170234136A1 (fr)
EP (2) EP2990598A1 (fr)
CN (1) CN106795770B (fr)
WO (1) WO2016030449A1 (fr)

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6390775B1 (en) * 2000-12-27 2002-05-21 General Electric Company Gas turbine blade with platform undercut
EP1757773A1 (fr) * 2005-08-26 2007-02-28 Siemens Aktiengesellschaft Aube creuse de turbine
FR2924156A1 (fr) * 2007-11-26 2009-05-29 Snecma Sa Aube de turbomachine
US20100329888A1 (en) * 2006-05-18 2010-12-30 Nadvit Gregory M Turbomachinery blade having a platform relief hole, platform cooling holes, and trailing edge cutback
EP2497903A2 (fr) * 2011-03-07 2012-09-12 Alstom Technology Ltd Composant pour turbomachines

Family Cites Families (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5296308A (en) * 1992-08-10 1994-03-22 Howmet Corporation Investment casting using core with integral wall thickness control means
US5431537A (en) * 1994-04-19 1995-07-11 United Technologies Corporation Cooled gas turbine blade
US6142734A (en) * 1999-04-06 2000-11-07 General Electric Company Internally grooved turbine wall
US6607355B2 (en) * 2001-10-09 2003-08-19 United Technologies Corporation Turbine airfoil with enhanced heat transfer
DE10217390A1 (de) * 2002-04-18 2003-10-30 Siemens Ag Turbinenschaufel
US20040094287A1 (en) * 2002-11-15 2004-05-20 General Electric Company Elliptical core support and plug for a turbine bucket
EP1512489B1 (fr) * 2003-09-05 2006-12-20 Siemens Aktiengesellschaft Aube pour turbine

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6390775B1 (en) * 2000-12-27 2002-05-21 General Electric Company Gas turbine blade with platform undercut
EP1757773A1 (fr) * 2005-08-26 2007-02-28 Siemens Aktiengesellschaft Aube creuse de turbine
US20100329888A1 (en) * 2006-05-18 2010-12-30 Nadvit Gregory M Turbomachinery blade having a platform relief hole, platform cooling holes, and trailing edge cutback
FR2924156A1 (fr) * 2007-11-26 2009-05-29 Snecma Sa Aube de turbomachine
EP2497903A2 (fr) * 2011-03-07 2012-09-12 Alstom Technology Ltd Composant pour turbomachines

Also Published As

Publication number Publication date
CN106795770B (zh) 2018-12-11
WO2016030449A1 (fr) 2016-03-03
US20170234136A1 (en) 2017-08-17
EP3158168B1 (fr) 2023-05-17
EP3158168A1 (fr) 2017-04-26
CN106795770A (zh) 2017-05-31

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