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EP2853689A1 - Agencement de canaux de refroidissement dans une aube de turbine - Google Patents

Agencement de canaux de refroidissement dans une aube de turbine Download PDF

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Publication number
EP2853689A1
EP2853689A1 EP13185944.9A EP13185944A EP2853689A1 EP 2853689 A1 EP2853689 A1 EP 2853689A1 EP 13185944 A EP13185944 A EP 13185944A EP 2853689 A1 EP2853689 A1 EP 2853689A1
Authority
EP
European Patent Office
Prior art keywords
cooling
blade
arrangement
cooling channels
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.)
Withdrawn
Application number
EP13185944.9A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
Thomas Burzych
Eugen Hummel
Gordon Emanuel Kunze
Frank PREUTEN
Thomas Alexis Schneider
Hannes Teuber
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 EP13185944.9A priority Critical patent/EP2853689A1/fr
Priority to US15/023,392 priority patent/US20160208622A1/en
Priority to JP2016516886A priority patent/JP2016533446A/ja
Priority to EP14772098.1A priority patent/EP3022397A1/fr
Priority to CN201480052859.5A priority patent/CN105593471A/zh
Priority to PCT/EP2014/069747 priority patent/WO2015044007A1/fr
Publication of EP2853689A1 publication Critical patent/EP2853689A1/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Withdrawn legal-status Critical Current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
    • 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
    • F01D5/12—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
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • 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
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • 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 an arrangement of cooling channels in a turbine blade.
  • Turbine blades in particular blades of gas turbines, are highly stressed components. The rotation takes place during operation with a high number of revolutions. Therefore, a high mechanical load capacity is required. In addition, high temperatures occur especially in gas turbine blades during operation. It generally applies that higher temperatures of the turbine blades driving gas mixture have a favorable effect on the efficiency of the gas turbine. In order to prevent too high temperatures of the turbine blades, the turbine blades are cooled. For this purpose, cooling channels are often arranged inside the turbine blades.
  • the object of the invention is to mitigate this disadvantage.
  • An arrangement of a plurality of cooling channels, that is to say at least two cooling channels, within a turbine blade for conveying cooling fluid is proposed.
  • the cooling fluid is usually air.
  • the cooling channels lead through the turbine blade to one or more cooling fluid outlets.
  • the turbine blade regularly has a blade root, an airfoil tip, an inlet edge and a trailing edge.
  • the cooling channels are connected to each other at selected locations and are separated from each other in other areas, that when the turbine blade is damaged in the region of a cooling channel, the cooling by the other cooling channels remains largely unimpaired.
  • a cooling passage generally runs from the blade root to the blade tip along the leading edge.
  • a leak caused by damage in this cooling channel causes the cooling fluid to escape there. This is problematic because cooling occurs downstream of the leak.
  • cooling fluid can pass from one cooling channel into another cooling channel. Should a leak have occurred in the other cooling channel upstream of the connection, cooling would be lost without the connection downstream. Through the connection, the cooling can be largely maintained downstream of the connection. But it is also necessary to separate the cooling channels in other areas from each other. Without the separation cooling fluid could pass unhindered in the event of a leak to the leak, so that the cooling would in turn be more affected. Above all, it is also normal, So in the absence of leakage, a channel structure, that is, a separation of the cooling channels, required to actually direct the cooling fluid through the entire turbine blade.
  • cooling fluid would flow from a cooling fluid inlet a short distance to a cooling fluid outlet. It is therefore always a reasonable compromise between connections of the cooling channels and separate areas to create.
  • those skilled in the art can provide a variety of different arrangements.
  • the cooling channels are connected to one another in such a way that, as the arrangement flows through, cooling fluid regularly flows from one cooling channel into another cooling channel. It would also be conceivable to provide this only in the event of a leak. For the purpose of an efficient flow, it has been found useful to provide this in normal operation.
  • the cooling channels are separated from an inner wall of the turbine blade by a perforated plate or a device in the manner of a perforated plate, so that the cooling fluid can pass largely perpendicular to the inner wall of the turbine blade.
  • This achieves so-called impingement cooling.
  • This is efficient because the cooling fluid is swirled on the inner wall and flows out again after the heating. If the cooling fluid only flow past the inner wall of the turbine blade, a film lying directly against the wall could form, in which the flow is comparatively weak. In addition, in one area just heated cooling fluid would be used to cool other areas.
  • At least one cooling passage begins at the blade root in a region near the leading edge of the turbine blade.
  • the inlet for the cooling fluid is, even with the arrangements known in the prior art, for structural reasons regularly at the blade root. Since at the leading edge, the turbine blade driving gas mixture is hottest, the thermal load of the turbine blade is highest there. Therefore, it makes sense that a cooling duct begins in the area of the leading edge.
  • At least one cooling channel begins in a region near the leading edge and near the blade root and leads as a diagonal channel through the turbine blade into a region near the trailing edge and near the blade tip. It should be made clear that the diagonal canal does not have to start at the blade root and not at the entrance but only in this area. A beginning at the blade root and at the leading edge but should not be excluded. The same applies to the end of the diagonal channel near the trailing edge and near the blade tip. The diagonal canal Allows the cooling fluid to flow well into different areas of the turbine blade and provide efficient cooling anywhere.
  • two cooling channels begin at the blade root in a region near the leading edge, which end in a region near the blade root and are connected to one another and to the diagonal channel.
  • This cooling fluid can pass fromméfluideinlässen the blade root to Diagonalkanal. If cooling fluid escapes from one of the aforementioned cooling channels due to a leak, the diagonal channel can continue to be supplied with cooling fluid through the other cooling channel.
  • cooling channels branch off from the diagonal channel, wherein, in particular, cooling channels branch off in the direction of the outlet edge and / or branches off cooling channels in the direction of the blade blade tip. In this way, the distribution of the cooling fluid in the entire region of the turbine blade can be further optimized.
  • a cooling channel runs parallel to the blade tip, into which opening the above-mentioned cooling channels extending in the direction of the blade tip.
  • the cooling channel running parallel to the blade tip can open into the same region as the diagonal channel.
  • the cooling channels branching off in the direction of the outlet edge run largely perpendicular to the outlet edge.
  • the cooling channels extending in the direction of the blade tip extend largely parallel to the outlet edge. This also serves to further optimize the distribution of the cooling fluid. It is always important to keep in mind that a leak at one point should affect the cooling of the turbine blade as little as possible.
  • cooling fluid outlets are provided in the region of the outlet edge through which cooling fluid can pass from the region inside the turbine blade into an area outside the turbine blade. This can be achieved in the region of the trailing edge on an outer wall, a further cooling.
  • the leaked cooling fluid can optionally be used to drive a further turbine stage.
  • At least one cooling fluid outlet is provided on the blade root in the region of the outlet edge.
  • the cooling fluid can flow from the cooling fluid inlet, which is normally located on the blade root in the region of the leading edge, through the turbine blade and flow back to the blade root in the area of the outlet edge.
  • the exiting cooling fluid can be reused to cool additional turbine blades.
  • a blade root 2 with which the turbine blade is attached to a rotor.
  • On the left is an entrance edge 3 can be seen.
  • the leading edge 3 is the area to which a gas mixture driving the turbine blade first impinges.
  • Above a blade tip 4 can be seen.
  • a trailing edge 5 is arranged.
  • the turbine blade is not flat, but curved. In this case, the leading edge 3 and the trailing edge 5 may be straight, but also curved.
  • the blade root 2 and the blade tip run as well as the rest of the blade area curved in any case. The curvature is due to an aerodynamic shape of the turbine blade.
  • the turbine blade has a front wall (not shown) extending from the leading edge to the trailing edge and a rear wall extending at a distance therefrom which again leads from the trailing edge to the leading edge.
  • a front wall (not shown) extending from the leading edge to the trailing edge and a rear wall extending at a distance therefrom which again leads from the trailing edge to the leading edge.
  • the distance between the front wall and the rear wall in the region of the leading edge 3 and the trailing edge 5 is very low and increases toward the blade center.
  • a first cooling channel 6 begins at the blade root 2 and runs directly along the leading edge 3.
  • a further cooling channel 7 extends away from the blade root 2 and is separated from the cooling channel 6.
  • the cooling channels 6 and 7 open into a region 8, which is near the leading edge 3 and near the blade root 2.
  • the cooling channels 6 and 7 are interconnected.
  • a diagonal channel 9 which leads into a region 10 near the trailing edge 5 and near the blade tip 4, also begins.
  • a cooling channel 11 extends parallel to the blade root 2.
  • the cooling channel 11 opens into a parallel to the trailing edge 5 extending cooling channel 12. If you follow the diagonal channel 9 from the area 8 near the leading edge 3 to the area 10 near the trailing edge 5 branches two cooling channels 13 and 14, which run parallel to the cooling channel 11 and open into the cooling channel 12.
  • cooling channels 15 and 16 extending parallel to the leading edge 3 branch off from the diagonal channel 8. These lead into a cooling channel 17, which runs parallel to the blade tip 4 in the vicinity of the blade tip 4 and opens into the region 10 and is connected there to the diagonal channel 9.
  • the region 10 is also connected to the cooling channel 12 running along the trailing edge 5.
  • the cooling channel 12 opens in the blade root 2 in a cooling fluid outlet 18.
  • cooling fluid outlets 19a to 19g are provided at the trailing edge 5.
  • the arrangement 1 of the cooling channels 6, 7, 9, 11, 12, 13, 14, 15, 16, 17 can also be referred to as "fir tree design".

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP13185944.9A 2013-09-25 2013-09-25 Agencement de canaux de refroidissement dans une aube de turbine Withdrawn EP2853689A1 (fr)

Priority Applications (6)

Application Number Priority Date Filing Date Title
EP13185944.9A EP2853689A1 (fr) 2013-09-25 2013-09-25 Agencement de canaux de refroidissement dans une aube de turbine
US15/023,392 US20160208622A1 (en) 2013-09-25 2014-09-17 Arrangement of cooling channels in a turbine blade
JP2016516886A JP2016533446A (ja) 2013-09-25 2014-09-17 タービンブレード内の冷却チャネルの配列
EP14772098.1A EP3022397A1 (fr) 2013-09-25 2014-09-17 Agencement de canaux de refroidissement dans une aube de turbine
CN201480052859.5A CN105593471A (zh) 2013-09-25 2014-09-17 涡轮机叶片内冷却通道的布置
PCT/EP2014/069747 WO2015044007A1 (fr) 2013-09-25 2014-09-17 Agencement de canaux de refroidissement dans une aube de turbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP13185944.9A EP2853689A1 (fr) 2013-09-25 2013-09-25 Agencement de canaux de refroidissement dans une aube de turbine

Publications (1)

Publication Number Publication Date
EP2853689A1 true EP2853689A1 (fr) 2015-04-01

Family

ID=49303737

Family Applications (2)

Application Number Title Priority Date Filing Date
EP13185944.9A Withdrawn EP2853689A1 (fr) 2013-09-25 2013-09-25 Agencement de canaux de refroidissement dans une aube de turbine
EP14772098.1A Withdrawn EP3022397A1 (fr) 2013-09-25 2014-09-17 Agencement de canaux de refroidissement dans une aube de turbine

Family Applications After (1)

Application Number Title Priority Date Filing Date
EP14772098.1A Withdrawn EP3022397A1 (fr) 2013-09-25 2014-09-17 Agencement de canaux de refroidissement dans une aube de turbine

Country Status (5)

Country Link
US (1) US20160208622A1 (fr)
EP (2) EP2853689A1 (fr)
JP (1) JP2016533446A (fr)
CN (1) CN105593471A (fr)
WO (1) WO2015044007A1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20180291743A1 (en) * 2017-04-07 2018-10-11 General Electric Company Turbine engine airfoil having a cooling circuit

Families Citing this family (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3037830B1 (fr) * 2015-06-29 2024-02-16 Snecma Ensemble de moulage d'une aube de turbomachine, comprenant une portion en relief de grande section
US10544684B2 (en) * 2016-06-29 2020-01-28 General Electric Company Interior cooling configurations for turbine rotor blades
FR3057906B1 (fr) * 2016-10-20 2019-03-15 Safran Aircraft Engines Aube de turbomachine a refroidissement optimise
US10422229B2 (en) * 2017-03-21 2019-09-24 United Technologies Corporation Airfoil cooling
US11644046B2 (en) * 2018-01-05 2023-05-09 Aurora Flight Sciences Corporation Composite fan blades with integral attachment mechanism
EP3832069A1 (fr) * 2019-12-06 2021-06-09 Siemens Aktiengesellschaft Aube de turbine pour turbine à gaz fixe

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GB827289A (en) * 1955-10-26 1960-02-03 Wiggin & Co Ltd Henry Improvements relating to hollow turbine or compressor blades
FR1209752A (fr) * 1958-09-10 1960-03-03 Wiggin & Co Ltd Henry Perfectionnements à la fabrication d'aubes de turbines et de machines analogues
US3014693A (en) * 1957-06-07 1961-12-26 Int Nickel Co Turbine and compressor blades
JPS59231103A (ja) * 1983-06-14 1984-12-25 Toshiba Corp ガスタ−ビン冷却翼
US5536143A (en) * 1995-03-31 1996-07-16 General Electric Co. Closed circuit steam cooled bucket
EP0939196A2 (fr) * 1998-02-26 1999-09-01 Kabushiki Kaisha Toshiba Aube de turbine à gaz
US6382914B1 (en) * 2001-02-23 2002-05-07 General Electric Company Cooling medium transfer passageways in radial cooled turbine blades
EP1471210A1 (fr) * 2003-04-24 2004-10-27 Siemens Aktiengesellschaft Composant de turbine avec une plaque d'impact de refroidissement
US20050084370A1 (en) * 2003-07-29 2005-04-21 Heinz-Jurgen Gross Cooled turbine blade

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US2641439A (en) * 1947-10-01 1953-06-09 Chrysler Corp Cooled turbine or compressor blade
US2687278A (en) * 1948-05-26 1954-08-24 Chrysler Corp Article with passages
DE1097212B (de) * 1956-10-22 1961-01-12 Her Majesty The Queen In The R Mit Kuehlkanaelen versehene Schaufel, insbesondere fuer Gasturbinen
US3017159A (en) * 1956-11-23 1962-01-16 Curtiss Wright Corp Hollow blade construction
US3171631A (en) * 1962-12-05 1965-03-02 Gen Motors Corp Turbine blade
US3554663A (en) * 1968-09-25 1971-01-12 Gen Motors Corp Cooled blade
JPS6285102A (ja) * 1985-10-11 1987-04-18 Hitachi Ltd ガスタ−ビン冷却翼
JP2851575B2 (ja) * 1996-01-29 1999-01-27 三菱重工業株式会社 蒸気冷却翼
EP2378073A1 (fr) * 2010-04-14 2011-10-19 Siemens Aktiengesellschaft Aube de rotor ou de stator pour turbomachine
CN201991570U (zh) * 2011-03-11 2011-09-28 北京华清燃气轮机与煤气化联合循环工程技术有限公司 燃气轮机的涡轮转子叶片
US9528379B2 (en) * 2013-10-23 2016-12-27 General Electric Company Turbine bucket having serpentine core

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB827289A (en) * 1955-10-26 1960-02-03 Wiggin & Co Ltd Henry Improvements relating to hollow turbine or compressor blades
US3014693A (en) * 1957-06-07 1961-12-26 Int Nickel Co Turbine and compressor blades
FR1209752A (fr) * 1958-09-10 1960-03-03 Wiggin & Co Ltd Henry Perfectionnements à la fabrication d'aubes de turbines et de machines analogues
JPS59231103A (ja) * 1983-06-14 1984-12-25 Toshiba Corp ガスタ−ビン冷却翼
US5536143A (en) * 1995-03-31 1996-07-16 General Electric Co. Closed circuit steam cooled bucket
EP0939196A2 (fr) * 1998-02-26 1999-09-01 Kabushiki Kaisha Toshiba Aube de turbine à gaz
US6382914B1 (en) * 2001-02-23 2002-05-07 General Electric Company Cooling medium transfer passageways in radial cooled turbine blades
EP1471210A1 (fr) * 2003-04-24 2004-10-27 Siemens Aktiengesellschaft Composant de turbine avec une plaque d'impact de refroidissement
US20050084370A1 (en) * 2003-07-29 2005-04-21 Heinz-Jurgen Gross Cooled turbine blade

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20180291743A1 (en) * 2017-04-07 2018-10-11 General Electric Company Turbine engine airfoil having a cooling circuit
US10697301B2 (en) * 2017-04-07 2020-06-30 General Electric Company Turbine engine airfoil having a cooling circuit

Also Published As

Publication number Publication date
WO2015044007A1 (fr) 2015-04-02
EP3022397A1 (fr) 2016-05-25
CN105593471A (zh) 2016-05-18
US20160208622A1 (en) 2016-07-21
JP2016533446A (ja) 2016-10-27

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