EP2196625A1 - Aube de turbine dotée d'un passage agencé dans une paroi de séparation et noyau de coulage associé - Google Patents

Aube de turbine dotée d'un passage agencé dans une paroi de séparation et noyau de coulage associé Download PDF

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Publication number
EP2196625A1
EP2196625A1 EP08021447A EP08021447A EP2196625A1 EP 2196625 A1 EP2196625 A1 EP 2196625A1 EP 08021447 A EP08021447 A EP 08021447A EP 08021447 A EP08021447 A EP 08021447A EP 2196625 A1 EP2196625 A1 EP 2196625A1
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EP
European Patent Office
Prior art keywords
turbine blade
passage
coolant
sections
regions
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Withdrawn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
EP08021447A
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German (de)
English (en)
Inventor
Fathi Ahmad
Winfried Dr. Esser
Christian Lerner
Christoph Schiefer
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 EP08021447A priority Critical patent/EP2196625A1/fr
Publication of EP2196625A1 publication Critical patent/EP2196625A1/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Withdrawn legal-status Critical Current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

Definitions

  • the invention relates to a turbine blade with an aerodynamically curved airfoil, in which at least two areas of at least one cavity through which a coolant flows are largely separated from one another via a partition, such that at least one passage in the form of a cross-connection is provided in the partition at least in terms of a fluidic connection. Furthermore, the invention relates to a casting core for use in a casting device for producing a cast turbine blade according to the preamble of claim 1, by means of which, after its removal from the cast turbine blade, at least one cavity through which a coolant can flow remains in the turbine blade.
  • An initially mentioned turbine blade and a casting core for producing such a turbine blade for example, from WO 2003/042503 A1 known.
  • the turbine blade shown therein has four separate cooling systems for cooling the airfoil.
  • the cooling systems are formed by cavities provided inside the airfoil.
  • the cavities of the turbine blade produced in the casting process are produced using four lost casting cores.
  • the Indian WO 2003/042503 A1 denoted by the reference numeral 58 casting core is meandering formed with a total of two deflection, so that the remaining through him, the pressure side arranged cooling channel of the turbine blade has a total of three substantially parallel, mutually spaced cooling duct sections.
  • the third section of the respective cooling channel which can be flowed through by a cooling fluid as the last, serves to cool the leading edge and ends at the blade tip side.
  • the casting core required for the production of the cooling channel has a pin-shaped bridge between the first and the second Second cooling channel section to ensure sufficient stability of the casting core.
  • cross-connection is also known as the so-called "cross-over-hole”.
  • the casting cores are made of a ceramic material, wherein the pin-like bridge is usually designed as a pin made of quartz, which is attached to the respective sections of the casting core or adjacent casting cores.
  • the pins serve primarily to stabilize the casting core for the manufacturing process and for the exact storage of the casting core in the casting device during the filling process.
  • the object of the invention is to provide a turbine blade and to provide a casting core for producing the initially mentioned turbine blade, in which the amount of cooling air flowing through the cross connection is minimized while providing a further accurately positionable and stable casting core.
  • the task directed to the turbine blade is achieved with a turbine blade according to the features of claim 1.
  • the task directed to the casting core is achieved by a casting core according to the features of claim 8.
  • the invention proposes that the passage has a longitudinal sectional contour or orientation which makes it more difficult for it to flow through with coolant or the inflow of coolant, compared to a rectilinear passage extending perpendicularly through the dividing wall. It follows for a casting core that its support element, which in the turbine blade leaves the passage, is formed so that it forms the passage according to the invention.
  • the invention is based on the finding that a smaller amount of coolant flows through the passage when it has a longitudinal section contour or an orientation that makes it difficult to flow through the passage or the inflow. This ensures that the coolant flows as intended in those areas of a cavity of the turbine blade, which must be cooled by this. In this respect, the coolant is prevented from shortening its flow path via the passage in the dividing wall.
  • the coolant in one of the two regions, has a flow direction and the passage - based on the relevant area and on the coolant - on this side inlet and an other side outlet, wherein - based on the flow direction of the coolant in the area concerned - In the partition of the inlet is arranged downstream of the outlet.
  • This refinement is advantageous in particular in the case of regions which can be flowed through sequentially by one and the same coolant, which coolant can flow through in the opposite direction, for the most part. Since the inlet is provided downstream of the outlet in the partition wall and the pressure in the coolant decreases steadily with increasing channel length, the pressure difference between the inlet and the outlet is smaller than with a passage extending perpendicularly through the partition wall.
  • the passage may have an orientation that is partially opposite to the coolant flow direction in the inlet-side region.
  • a flow direction deflection is for the inflow of coolant required by more than 90 °, which makes the inflow of coolant difficult. This, too, reduces the flow rate of coolant through the passageway, resulting in an increase in the amount of coolant that flows properly through each of the respective areas of the cavity.
  • the passage can have differently sized cross sections along its extension. It is particularly advantageous if at least one transition between two different sized cross sections is formed in stages.
  • the abrupt change of the cross-section along the extension of the passage in accordance with a discontinuous pipe constriction or pipe extension leads to a loss of shock in the coolant flow as a result of the speed decrease or increase in speed.
  • the shock loss leads to an increased resistance coefficient for the flow in the passage, which in turn makes it difficult to flow through the passage.
  • the amount of coolant flowing through the passage can be reduced, which leads to an increase in the amount of coolant flowing through the regions.
  • the coolant in one of the regions, has a flow direction and the passage - with respect to the relevant region and to the coolant - on this side inlet and an other side outlet, wherein the passage has a cross-section centrally between inlet and outlet, which is substantially larger or smaller than the cross section of the inlet and / or the outlet.
  • the two areas are part of a cavity or are formed by two different cavities or whether the two areas are flowed through sequentially or in parallel by coolant.
  • the areas are connected to a coolant supply in such a way that the coolant that can flow in them flows in the opposite direction. This is given, for example, when the regions can be flowed through by coolant in a sequential manner and the cavity as a whole is meander-shaped, the regions being connected to one another via a deflection region.
  • the turbine blade may be formed both as a guide blade and as a blade.
  • both the areas and the cavities are part of a cooling system, which are arranged in the turbine blade.
  • the passages relating to the invention do not serve for the use of impingement cooling and are therefore not designed as impingement cooling openings.
  • the passages relating to the invention are solely due to the fact that for a secure and reliable positioning of the casting core and parts of its supporting elements are required, which support adjacent portions of the casting core or adjacent casting cores with each other.
  • the invention also relates to a casting core for use in a casting apparatus, by means of which a cast turbine blade according to the preamble of claim 1 can be produced, wherein after removal of the casting core from the cast turbine blade a permeable by a coolant cavity remains in the turbine blade.
  • the casting core comprises at least two sections, which are mainly separated from one another via a free space, and at least one support element arranged in the region of the free space and connecting the sections to one another.
  • the casting core is suitable after removal from a cast Turbine blade to connect the two hollow portions of the turbine blade left by the sections by means of a passage left by the support member in a partition wall, wherein the support member of the casting core is formed such that it leaves in the turbine blade a passage with a longitudinal sectional contour or orientation, which flows through the Passage or inflow of coolant difficult compared with a straight through the partition wall of the turbine blade extending rectilinear passage.
  • FIG. 1 shows the cross section through an airfoil 10 of a turbine blade 12.
  • the airfoil 10 is aerodynamically curved in a known manner and thus has an anströmbare of a working medium leading edge 14.
  • the air flowing around the blade 10 working fluid leaves the Turbine blade 12 at the trailing edge 16.
  • a suction-side airfoil wall 18 and a pressure-side airfoil wall 20 extend between the leading edge 14 and the trailing edge 16.
  • the airfoil 10 is hollow in the interior and has, by way of example, two cavities 21, 22 which are divided into regions 23, 24, 25, 26 are logically subdivided.
  • the front edge side region 23 is separated from the adjacent region 24 by a first partition wall 28, wherein the partition wall 28 extends from the suction side airfoil wall 18 to the pressure side airfoil wall 20.
  • the region 24 of the area 24 of the area inside the turbine blade 12 is immediately followed by the region 26.
  • the longitudinal section through the turbine blade 12 shows FIG. 2 ,
  • the turbine blade 12 according to FIG. 2 is designed as a guide vane for an axially effetströmbare gas turbine.
  • the turbine blade 12 is also equipped both with an outer platform 27 and an inner platform 30, whose respective, the airfoil 10 facing wall surface respectively forms parts of the radially outer and radially inner boundary surface of the annular cross-sectional flow channel of the working medium.
  • the turbine blade 12 has, for example, a feed opening 32 for a coolant 34, preferably cooling air, present on the inner platform 30.
  • the cavities 21, 22 are in fluid communication with the supply port 32 so that the coolant 34 flowing through the supply port 32 may enter both the cavity 21 and the cavity 22.
  • the coolant 34 entering the cavity 21 can exit through openings 36 arranged in the front edge 14 and mix with the working medium flowing around the airfoil 10.
  • the openings 36 are arranged in the form of a grid, known as a shower head.
  • the coolant 34 entering the region 24 of the cavity 22 flows to the radially outer end of the turbine blade 12, where it passes largely through the outer platform 27 in its direction is deflected by 180 °.
  • the coolant 34 flows through the region 25 and, after a further deflection, the region 26. From there, the coolant can exit from the turbine blade 12 via openings 38 arranged in the trailing edge 16.
  • the areas 24, 25 are separated by a separating wall 29 extending from the suction side airfoil wall 18 to the pressure side airfoil wall 20.
  • the areas 25, 26 are separated by a comparable partition 31 from each other.
  • two passages 40, 42 are arranged by way of example, which can be flowed through by coolant.
  • the coolant 34 flowing in the region 25 can partially pass into the region 26 without this coolant portion having to reach the platform-side deflection region 43 of the cavity 22, which is fluidically provided between the two regions 25, 26. Due to the presence of the passages 40, 42, the two regions 25, 26 are no longer completely separated from each other locally (in the region of the passage), but only to a large extent.
  • Both passages 40, 42 each have an inlet 44 and an outlet 46 for the coolant 34.
  • the orientation of the passage 40 is selected such that the coolant flowing through the region 25 for flowing into the passage 40 must be deflected by more than 90 ° from its previous flow direction, which makes it difficult for coolant 34 to flow into the passage 40. As a result, the amount of coolant 34 flowing through to the passage 40 can be reduced.
  • the inlet 44 with respect to the flow direction of the coolant 34, is arranged downstream of the outlet 46 in that region 25 in which the inlet 44 is arranged.
  • the longitudinal section contour of the passage 42 serves the same purpose.
  • the cross-sectional area is of the inlet 44 and the outlet 46 is smaller than the cross-sectional area which the passage 42 has centrally between the inlet 44 and the outlet 46.
  • the transition between the differently sized cross-sections of the passage 42 takes place in stages, so that in particular a pressure loss and circulations can be generated, which further complicate the passage of the passage 42 with coolant 34.
  • a diffuser-shaped contour (conical shape) could be selected.
  • the position of the passage according to the invention in the turbine blade 12 is not limited to the embodiment.
  • a further passage may be provided in each case.
  • a casting core 50 required for producing a turbine blade 12 according to the invention is shown FIG. 3 in a side view.
  • the casting core 50 is commonly used in a casting apparatus required to make a cast turbine blade.
  • the casting core 50 accommodates in the casting apparatus that space which is free of casting material in the cast turbine blade 12 after its completion and accordingly represents the cavity 21, 22 of the turbine blade 12.
  • the for the production of in 1 and 2 Casting core 50 thus required comprises a plurality of sections 52, 54, 56 and 58.
  • Each section 52, 54, 56 and 58 is of pin-shaped, oblong shape, which are each connected at one end to an adjacent section.
  • the sections 52, 54 are interconnected in area A.
  • the sections 54, 56 are in the area B and the sections 56, 58 are connected to each other in the area C.
  • the areas B, C leave in the cast turbine blade 12, the deflection areas 43. Due to the pin-like shape of the sections 52, 54, 56, 58 and only one side, end provided connection to adjacent sections on the areas B, C is from the Prior art known casting core only relatively poorly handled, so that it tends to break during its attachment in the casting apparatus.
  • supporting elements 60, 62 are provided between two immediately adjacent sections 56, 58, which connect the two relevant sections 56, 58 with each other and thus increase the rigidity of the casting core 50 as a whole.
  • the support elements 60, 62 which are also referred to as pins or quartz pins, according to the invention have a longitudinal sectional contour or orientation, which impedes the passage of the generated through them in the cast turbine blade 12 passage 40, 42, compared with a perpendicular through the partition wall of the turbine blade 12 extending rectilinear, d. H. cylindrical passage with constant diameter.
  • One of the two support elements 60 does not extend perpendicular to the longitudinal extent of the sections 56, 58.
  • the other support element 62 extends perpendicular to the surface of the section 56, 58, but this is not formed cylindrically with a constant diameter. It has two diameters of different sizes, with the diameter provided centrally between the section 56 and the section 58 being substantially larger than the diameter immediately adjacent to the sections 56, 58.
  • the support elements 60, 62 which are also removed like the casting core 50 after the casting of the turbine blade 12 from its interior, lead to arranged in the partition wall 31 passages 40, 42 (FIGS. FIG. 2 ).
  • the invention thus relates to a turbine blade 12 having an aerodynamically curved airfoil 10, in which at least two areas 23-26 of at least one cavity 21, 22 which can be flowed through by a coolant 34 are locally separated from each other by a rib-shaped partition wall 31, apart from a locally arranged passage 40 or passage 42. Due to the presence of at least a passage 40, 42 in the partition wall 28 - 31, the two areas 23 - 26 are only largely separated from each other.
  • the passage 40, 42 has a longitudinal sectional contour or orientation, which the through or Inflowing of the passage 40, 42 with coolant 34 difficult compared to a straight through the rib-shaped partition wall 31 extending rectilinear passage.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP08021447A 2008-12-10 2008-12-10 Aube de turbine dotée d'un passage agencé dans une paroi de séparation et noyau de coulage associé Withdrawn EP2196625A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP08021447A EP2196625A1 (fr) 2008-12-10 2008-12-10 Aube de turbine dotée d'un passage agencé dans une paroi de séparation et noyau de coulage associé

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP08021447A EP2196625A1 (fr) 2008-12-10 2008-12-10 Aube de turbine dotée d'un passage agencé dans une paroi de séparation et noyau de coulage associé

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EP2196625A1 true EP2196625A1 (fr) 2010-06-16

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EP08021447A Withdrawn EP2196625A1 (fr) 2008-12-10 2008-12-10 Aube de turbine dotée d'un passage agencé dans une paroi de séparation et noyau de coulage associé

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3147456A1 (fr) * 2015-09-28 2017-03-29 Siemens Aktiengesellschaft Aube de turbine dote d'encoche dans le sol de couronne

Citations (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO1994012766A1 (fr) * 1992-11-30 1994-06-09 United Technologies Corporation Structure d'aube pouvant etre refroidie
EP0641917A1 (fr) * 1993-09-08 1995-03-08 United Technologies Corporation Refroidissement du bord d'attaque d'une aube
EP1219780A2 (fr) * 2000-12-22 2002-07-03 ALSTOM Power N.V. Refroidissement d'un élément d'une turbomachine par jet
WO2003042503A1 (fr) 2001-11-14 2003-05-22 Honeywell International Inc. Aube ou pale de turbine a gaz refroidi interne
WO2003062607A1 (fr) * 2002-01-25 2003-07-31 Alstom (Switzerland) Ltd Élément refroidi pour turbine à gaz
EP1496203A1 (fr) * 2003-07-11 2005-01-12 Rolls-Royce Deutschland Ltd & Co KG Aube de turbine à gaz avec refroidissement par impact
US20050265837A1 (en) * 2003-03-12 2005-12-01 George Liang Vortex cooling of turbine blades
US20050281671A1 (en) * 2004-06-17 2005-12-22 Siemens Westinghouse Power Corporation Gas turbine airfoil trailing edge corner
US20060002795A1 (en) * 2004-07-02 2006-01-05 Siemens Westinghouse Power Corporation Impingement cooling system for a turbine blade
US20060083614A1 (en) * 2004-10-18 2006-04-20 United Technologies Corporation Airfoil with large fillet and micro-circuit cooling
GB2420156A (en) * 2004-11-16 2006-05-17 Rolls Royce Plc Heat transfer arrangement
EP1728970A2 (fr) * 2005-05-31 2006-12-06 United Technologies Corporation Système de refroidissement d'une aube de turbine

Patent Citations (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO1994012766A1 (fr) * 1992-11-30 1994-06-09 United Technologies Corporation Structure d'aube pouvant etre refroidie
EP0641917A1 (fr) * 1993-09-08 1995-03-08 United Technologies Corporation Refroidissement du bord d'attaque d'une aube
EP1219780A2 (fr) * 2000-12-22 2002-07-03 ALSTOM Power N.V. Refroidissement d'un élément d'une turbomachine par jet
WO2003042503A1 (fr) 2001-11-14 2003-05-22 Honeywell International Inc. Aube ou pale de turbine a gaz refroidi interne
WO2003062607A1 (fr) * 2002-01-25 2003-07-31 Alstom (Switzerland) Ltd Élément refroidi pour turbine à gaz
US20050265837A1 (en) * 2003-03-12 2005-12-01 George Liang Vortex cooling of turbine blades
EP1496203A1 (fr) * 2003-07-11 2005-01-12 Rolls-Royce Deutschland Ltd & Co KG Aube de turbine à gaz avec refroidissement par impact
US20050281671A1 (en) * 2004-06-17 2005-12-22 Siemens Westinghouse Power Corporation Gas turbine airfoil trailing edge corner
US20060002795A1 (en) * 2004-07-02 2006-01-05 Siemens Westinghouse Power Corporation Impingement cooling system for a turbine blade
US20060083614A1 (en) * 2004-10-18 2006-04-20 United Technologies Corporation Airfoil with large fillet and micro-circuit cooling
GB2420156A (en) * 2004-11-16 2006-05-17 Rolls Royce Plc Heat transfer arrangement
EP1728970A2 (fr) * 2005-05-31 2006-12-06 United Technologies Corporation Système de refroidissement d'une aube de turbine

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3147456A1 (fr) * 2015-09-28 2017-03-29 Siemens Aktiengesellschaft Aube de turbine dote d'encoche dans le sol de couronne
WO2017054996A1 (fr) * 2015-09-28 2017-04-06 Siemens Aktiengesellschaft Aube de turbine comprenant une rainure dans la base de couronne

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