EP1614859A1 - Aube de turbine refroidie par couche d'air - Google Patents

Aube de turbine refroidie par couche d'air Download PDF

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Publication number
EP1614859A1
EP1614859A1 EP04015805A EP04015805A EP1614859A1 EP 1614859 A1 EP1614859 A1 EP 1614859A1 EP 04015805 A EP04015805 A EP 04015805A EP 04015805 A EP04015805 A EP 04015805A EP 1614859 A1 EP1614859 A1 EP 1614859A1
Authority
EP
European Patent Office
Prior art keywords
turbine blade
blade
coolant
outlet
rows
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
EP04015805A
Other languages
German (de)
English (en)
Other versions
EP1614859B1 (fr
Inventor
Hans-Thomas Dr. Bolms
Ralf Müsgen
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 ES04015805T priority Critical patent/ES2282763T3/es
Priority to EP04015805A priority patent/EP1614859B1/fr
Priority to DE502004003477T priority patent/DE502004003477D1/de
Priority to US11/174,275 priority patent/US7500823B2/en
Priority to CNB2005100820514A priority patent/CN100350132C/zh
Publication of EP1614859A1 publication Critical patent/EP1614859A1/fr
Application granted granted Critical
Publication of EP1614859B1 publication Critical patent/EP1614859B1/fr
Active 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
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • 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
    • F05D2250/30—Arrangement of components
    • F05D2250/31—Arrangement of components according to the direction of their main axis or their axis of rotation
    • F05D2250/314—Arrangement of components according to the direction of their main axis or their axis of rotation the axes being inclined in relation to each other
    • 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
    • F05D2250/30—Arrangement of components
    • F05D2250/34—Arrangement of components translated
    • 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
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • the invention relates to a turbine blade for use in a gas turbine with an airfoil, which is provided with a number of coolant channels through which a coolant can pass, wherein a substantially longitudinally extending in the longitudinal direction of the turbine blade, the leading edge spaced coolant channel in the leading edge region of the airfoil in outlet openings Branch outgoing exit channels.
  • Gas turbines are used in many areas to drive generators or work machines.
  • the energy content of a fuel is used to generate a rotational movement of a turbine shaft.
  • the fuel is burned in a combustion chamber, compressed air being supplied by an air compressor.
  • the working medium produced in the combustion chamber by the combustion of the fuel, under high pressure and at high temperature, is guided via a turbine unit arranged downstream of the combustion chamber, where it relaxes to perform work.
  • a number of rotor blades which are usually combined into blade groups or rows of blades, are arranged thereon and drive the turbine shaft via a momentum transfer from the flow medium.
  • To guide the flow medium in the turbine unit also commonly associated guide blade rows are arranged between adjacent blade rows with the turbine housing.
  • the turbine blades in particular the guide vanes, usually have an airfoil extending along a blade axis for suitable guidance of the working medium, at the end for attachment of the turbine blade to the respective support body extending transversely to the blade axis Platform can be formed. But also on the other, free end, a platform or a platform-like shape may be appropriate.
  • cooling of the affected components in particular of rotor blades and / or guide vanes of the turbine unit, is usually provided.
  • the turbine blades are usually formed coolable, in particular, an effective and reliable cooling of the particular thermally loaded front edge of the respective turbine blade should be ensured.
  • Coolant is usually used as coolant. This is typically supplied to the respective turbine blade in the manner of open cooling via a number of coolant channels integrated into the airfoil or blade profile. Starting from these, the cooling air flows through the respectively provided regions of the turbine blade into outlet channels branching off from it, as a result of which convective cooling of the blade interior and of the blade wall is achieved. On the exit side, these channels are left open, so that the cooling air after flowing through the turbine blade from the exit openings, also referred to as film cooling holes, and forms a cooling air film on the surface of the airfoil. Due to this cooling air film, the material on the surface is largely protected against direct and overly intensive contact with the hot working medium flowing past at high speed.
  • the outlet openings there are usually arranged uniformly along at least two rows aligned parallel to the front edge.
  • the exit channels are also generally oriented obliquely to the longitudinal direction of the turbine blade, which supports the formation of the protective, flowing on the surface cooling air film. Since the outlet channels are usually introduced in the manufacture of the turbine blade for cost reasons only at the end from the outside, z. B. by laser drilling or other drilling methods, and in particular in the leading edge region of the airfoil access of the drilling instruments through the end molded platforms or platform-like formations may be hindered, it comes.
  • the exit channels often at an approximately mid-foot section and tip section of the respective Blade lying transition point for a change of orientation. This takes place in such a way that the coolant flowing out in a foot-side section of each row has a speed component pointing to the tip section in the region of the outlet openings, whereas the cooling medium flowing out in an adjoining tip-side section of each row has a speed component facing the foot section.
  • the outlet channels are inclined in the extension direction of the turbine blade, whereas they are inclined in the tip-side section opposite to the extension direction.
  • the invention is therefore an object of the invention to provide a turbine blade of the type mentioned above, for the simple means a particularly reliable and uniform cooling of the leading edge region while maintaining a particularly low demand for cooling air can be achieved.
  • transition points in which the orientation of the exit channels changes, are arranged offset from one another for each two adjacent rows in the longitudinal direction.
  • the invention is based on the consideration that the cooling medium emerging from the outlet openings in the leading edge region of the airfoil in order to form an effective cooling film has the largest possible velocity component in parallel should have to the surface. For this reason, the proven, oblique to the longitudinal direction alignment of the outlet channels should be maintained. With regard to the restrictions imposed on the manufacture of the airfoil and the restrictions on the access and orientation of the production tools, a change of orientation of the type described will continue to be desirable for the exit channels opening in the exit openings along each of the rows in which the exit openings are located. On the other hand, areas with a comparatively greatly reduced frequency density of the outlet channels in the blade wall should be avoided. For this purpose, it is to be ruled out that the gaps or intermediate spaces belonging to adjacent rows come to lie next to one another in the otherwise comparatively regular distribution pattern of the exit channels.
  • the associated transition points for each two adjacent rows are arranged offset from one another in the longitudinal direction.
  • the offset causes just a local entanglement of the outlet channels belonging to two adjacent rows and thus with respect to the totality of all rows a comparatively homogeneous distribution of the outlet channels over the entire leading edge region of the blade. Therefore, a relatively good and effective convective cooling of the blade interior is ensured in this area, so that a local overuse of the material is avoided by overheating.
  • the need for cooling medium can be kept comparatively low, which has a performance-enhancing effect for a equipped with such turbine blades gas turbine.
  • a flow behavior of the exiting cooling medium in the vicinity of the leading edge which is particularly favorable for effective film cooling, combined with good convective cooling of the adjacent blade wall, can be achieved by using the Outlets in the entire leading edge region are approximately evenly distributed in an advantageous embodiment of the invention, such that they lie on the vertices of an imaginary, curved to the leading edge of the airfoil, regular grid network. This causes a particularly homogeneous wetting of the blade surface with coolant.
  • angles of incidence of the outlet channels with respect to the longitudinal direction are preferably approximately equal for the foot-side and tip-side sections of all rows of outlet openings.
  • an optimized for the effect of the film cooling value which is known from experiments or calculations, can be adjusted.
  • the transition point belonging to the middle row in this case is displaced by three outlet openings in relation to the two outer rows.
  • the mutual offset is still low enough so that irritate the air currents flowing in the opposite direction in Verschränkungs Scheme only slightly irritate each other.
  • the turbine blade 2 is designed as a guide blade for a gas turbine not shown here. It comprises a foot section 4 and a tip section 6 with associated platforms 8, 10 and an intermediate airfoil 12 extending in the longitudinal direction L.
  • the profiled airfoil 12 has a leading edge 14 also extending substantially in the longitudinal direction L and a trailing edge 16 with side walls therebetween 18 on.
  • the turbine blade 2 is fixed via the foot section 4 on the inner casing of the turbine, wherein the associated platform 8 forms a wall element bounding the flow path of the working medium in the gas turbine.
  • the turbine shaft opposite the tip-side platform 10 forms another limit to the flowing working fluid.
  • the turbine blade 2 could also be designed as a moving blade, which is fastened in an analogous manner to the turbine shaft via a foot-side platform 8, also referred to as a blade root.
  • a coolant K is introduced into the blade interior via a number of inlet openings 20 arranged at the lower end of the foot section 4.
  • the coolant K is cooling air.
  • the coolant K After the coolant K has flowed through one or more coolant channels 22 adjoining the inlet openings 20 in the interior of the turbine blade 2, it exits from a number of outlet openings 24, also referred to as film cooling holes, corresponding to the coolant channels 22 in the area of the blade 12.
  • Different areas of the airfoil 12 provide in view of the various thermal and mechanical stress and the respective space conditions In Schaufelinneren to the arrangement and the design of the film cooling holes very different requirements.
  • the comparatively strongly curved leading edge region 28 immediately adjoining the leading edge 14 of the blade 12 requires effective cooling due to a relatively high load.
  • FIG. 2 shows the front region of the profiled airfoil 12 with the relatively strongly curved front edge region 28, which includes the leading edge 14 and adjoins the pressure side 30 and suction side 32.
  • coolant channel 22 From a substantially in the longitudinal direction L of the turbine blade 2 extending, spaced from the front edge 14 coolant channel 22 branch off outlet channels 34 of smaller cross section, which penetrate the blade wall 36 and open in the leading edge region 28 in outlet openings 24 or film cooling holes.
  • coolant K cooling of the adjacent areas of the blade wall 36 is achieved.
  • the effect of film cooling on the surface of the blade 12 caused by the cooling air flowing out of the outlet openings 24 occurs.
  • an air cushion or a protective film which prevents direct contact of the blade surface with the working medium having a high flow velocity, thus effectively forms on the surface through the cooling air flowing along it at a relatively low velocity.
  • the outlet openings 24 are arranged in the embodiment along three parallel to the leading edge 14 aligned rows, such that they form a regular grid pattern.
  • the outlet channels 34 are inclined relative to the longitudinal direction L of the turbine blade 2, so that in the region of their Outlet openings 24 for the outflowing coolant K a flat exit angle with respect.
  • the blade surface results. This also has a favorable effect on the formation of a protective cooling air film.
  • the inclination of the outlet channels 34 exists with regard to two different sections.
  • a foot-side section 38 of the illustrated row they are inclined so that the effluent from the outlet openings 24 coolant K has a pointing to the tip portion 6 of the turbine blade 2 speed component.
  • the orientation of the outlet channels 34 changes so that the coolant K flowing out of the tip-side section 42 of the row has a velocity component directed toward the foot section 4.
  • This change in orientation is due to the limited access of the drilling tools in the manufacture of the turbine blade 2 due to the platforms 8, 10 and the presence of a comparatively large gap 44 in the blade wall 36, which is otherwise uniformly traversed by outlet channels 34.
  • the turbine blade 2 is designed specifically for a particularly reliable cooling of the leading edge region 28 at the same time kept particularly low demand for coolant K.
  • the aforementioned transition points 40 are positioned offset from each other in the manner of a sectionally entangled arrangement of adjacent film cooling rows.
  • the partially cutaway perspective view of the leading edge 14 in FIG. 4 shows that the transition point 40 belonging to the middle row, in which the orientation of the exit channels 34 changes, is displaced in the longitudinal direction L with respect to the two outer rows.
  • the shift here in the exemplary embodiment three grid points.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP04015805A 2004-07-05 2004-07-05 Aube de turbine refroidie par couche d'air Active EP1614859B1 (fr)

Priority Applications (5)

Application Number Priority Date Filing Date Title
ES04015805T ES2282763T3 (es) 2004-07-05 2004-07-05 Alabe de turbina refrigerrada por pelicula.
EP04015805A EP1614859B1 (fr) 2004-07-05 2004-07-05 Aube de turbine refroidie par couche d'air
DE502004003477T DE502004003477D1 (de) 2004-07-05 2004-07-05 Filmgekühlte Turbinenschaufel
US11/174,275 US7500823B2 (en) 2004-07-05 2005-07-01 Turbine blade
CNB2005100820514A CN100350132C (zh) 2004-07-05 2005-07-05 透平叶片

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP04015805A EP1614859B1 (fr) 2004-07-05 2004-07-05 Aube de turbine refroidie par couche d'air

Publications (2)

Publication Number Publication Date
EP1614859A1 true EP1614859A1 (fr) 2006-01-11
EP1614859B1 EP1614859B1 (fr) 2007-04-11

Family

ID=34925626

Family Applications (1)

Application Number Title Priority Date Filing Date
EP04015805A Active EP1614859B1 (fr) 2004-07-05 2004-07-05 Aube de turbine refroidie par couche d'air

Country Status (5)

Country Link
US (1) US7500823B2 (fr)
EP (1) EP1614859B1 (fr)
CN (1) CN100350132C (fr)
DE (1) DE502004003477D1 (fr)
ES (1) ES2282763T3 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1927726A1 (fr) 2006-11-30 2008-06-04 Rolls-Royce plc Composant refroidi à l'air
EP3056675A1 (fr) * 2015-02-16 2016-08-17 United Technologies Corporation Systèmes et procédés de refroidissement d'aube

Families Citing this family (26)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7597540B1 (en) * 2006-10-06 2009-10-06 Florida Turbine Technologies, Inc. Turbine blade with showerhead film cooling holes
US7878761B1 (en) * 2007-09-07 2011-02-01 Florida Turbine Technologies, Inc. Turbine blade with a showerhead film cooling hole arrangement
US8317473B1 (en) * 2009-09-23 2012-11-27 Florida Turbine Technologies, Inc. Turbine blade with leading edge edge cooling
US8568085B2 (en) 2010-07-19 2013-10-29 Pratt & Whitney Canada Corp High pressure turbine vane cooling hole distrubution
US8545180B1 (en) * 2011-02-23 2013-10-01 Florida Turbine Technologies, Inc. Turbine blade with showerhead film cooling holes
JP5923936B2 (ja) * 2011-11-09 2016-05-25 株式会社Ihi フィルム冷却構造及びタービン翼
US20130156602A1 (en) * 2011-12-16 2013-06-20 United Technologies Corporation Film cooled turbine component
US8944750B2 (en) 2011-12-22 2015-02-03 Pratt & Whitney Canada Corp. High pressure turbine vane cooling hole distribution
US9322279B2 (en) * 2012-07-02 2016-04-26 United Technologies Corporation Airfoil cooling arrangement
US9109453B2 (en) * 2012-07-02 2015-08-18 United Technologies Corporation Airfoil cooling arrangement
US9121289B2 (en) 2012-09-28 2015-09-01 Pratt & Whitney Canada Corp. High pressure turbine blade cooling hole distribution
US9062556B2 (en) 2012-09-28 2015-06-23 Pratt & Whitney Canada Corp. High pressure turbine blade cooling hole distribution
US9228440B2 (en) 2012-12-03 2016-01-05 Honeywell International Inc. Turbine blade airfoils including showerhead film cooling systems, and methods for forming an improved showerhead film cooled airfoil of a turbine blade
US9562437B2 (en) 2013-04-26 2017-02-07 Honeywell International Inc. Turbine blade airfoils including film cooling systems, and methods for forming an improved film cooled airfoil of a turbine blade
US9464528B2 (en) * 2013-06-14 2016-10-11 Solar Turbines Incorporated Cooled turbine blade with double compound angled holes and slots
US9708915B2 (en) * 2014-01-30 2017-07-18 General Electric Company Hot gas components with compound angled cooling features and methods of manufacture
US10041356B2 (en) * 2014-08-15 2018-08-07 United Technologies Corporation Showerhead hole scheme apparatus and system
US9581029B2 (en) 2014-09-24 2017-02-28 Pratt & Whitney Canada Corp. High pressure turbine blade cooling hole distribution
US20160298464A1 (en) * 2015-04-13 2016-10-13 United Technologies Corporation Cooling hole patterned airfoil
CN104832218A (zh) * 2015-04-20 2015-08-12 西北工业大学 一种用于涡轮叶片前缘气膜冷却的错位对冲气膜孔排结构
CN106555617B (zh) * 2017-01-05 2018-07-10 西北工业大学 一种有斜下吹式气膜冷却孔的涡轮叶片
CN109030012B (zh) * 2018-08-24 2024-01-23 哈尔滨电气股份有限公司 一种带有冷却通道的透平叶根疲劳试验模拟件及试验方法
CN109110125A (zh) * 2018-09-03 2019-01-01 南京航空航天大学 一种旋翼桨叶结构设计方法
CN109736898A (zh) * 2019-01-11 2019-05-10 哈尔滨工程大学 一种交错复合角的叶片前缘气膜冷却孔结构
CN113404546A (zh) * 2021-07-09 2021-09-17 中国联合重型燃气轮机技术有限公司 叶片、透平和燃气轮机
US11959396B2 (en) * 2021-10-22 2024-04-16 Rtx Corporation Gas turbine engine article with cooling holes for mitigating recession

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JPS5851202A (ja) * 1981-09-24 1983-03-25 Hitachi Ltd ガスタ−ビンのタ−ビン翼前縁部の冷却装置
US5496151A (en) * 1994-02-03 1996-03-05 Societe Nationale D'etude Et De Construction De Moteures D'aviation "Snecma" Cooled turbine blade
GB2310896A (en) * 1996-03-05 1997-09-10 Rolls Royce Plc Air cooled wall
US5779437A (en) * 1996-10-31 1998-07-14 Pratt & Whitney Canada Inc. Cooling passages for airfoil leading edge
EP0894946A1 (fr) * 1997-02-04 1999-02-03 Mitsubishi Heavy Industries, Ltd. Pale fixe de refroidissement pour turbine a gaz
US6176676B1 (en) * 1996-05-28 2001-01-23 Kabushiki Kaisha Toshiba Cooling system for a main body used in a gas stream

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US6270317B1 (en) * 1999-12-18 2001-08-07 General Electric Company Turbine nozzle with sloped film cooling
US6331098B1 (en) * 1999-12-18 2001-12-18 General Electric Company Coriolis turbulator blade
CN1497128A (zh) * 2002-10-08 2004-05-19 通用电气公司 在翼型叶片上形成冷却孔的方法

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Publication number Priority date Publication date Assignee Title
JPS5851202A (ja) * 1981-09-24 1983-03-25 Hitachi Ltd ガスタ−ビンのタ−ビン翼前縁部の冷却装置
US5496151A (en) * 1994-02-03 1996-03-05 Societe Nationale D'etude Et De Construction De Moteures D'aviation "Snecma" Cooled turbine blade
GB2310896A (en) * 1996-03-05 1997-09-10 Rolls Royce Plc Air cooled wall
US6176676B1 (en) * 1996-05-28 2001-01-23 Kabushiki Kaisha Toshiba Cooling system for a main body used in a gas stream
US5779437A (en) * 1996-10-31 1998-07-14 Pratt & Whitney Canada Inc. Cooling passages for airfoil leading edge
EP0894946A1 (fr) * 1997-02-04 1999-02-03 Mitsubishi Heavy Industries, Ltd. Pale fixe de refroidissement pour turbine a gaz

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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1927726A1 (fr) 2006-11-30 2008-06-04 Rolls-Royce plc Composant refroidi à l'air
US8011890B2 (en) 2006-11-30 2011-09-06 Rolls-Royce Plc Air-cooled component
EP3056675A1 (fr) * 2015-02-16 2016-08-17 United Technologies Corporation Systèmes et procédés de refroidissement d'aube

Also Published As

Publication number Publication date
ES2282763T3 (es) 2007-10-16
EP1614859B1 (fr) 2007-04-11
CN100350132C (zh) 2007-11-21
US7500823B2 (en) 2009-03-10
US20060002796A1 (en) 2006-01-05
DE502004003477D1 (de) 2007-05-24
CN1724849A (zh) 2006-01-25

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