EP2918779A1 - Aube de turbine - Google Patents

Aube de turbine Download PDF

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Publication number
EP2918779A1
EP2918779A1 EP14158871.5A EP14158871A EP2918779A1 EP 2918779 A1 EP2918779 A1 EP 2918779A1 EP 14158871 A EP14158871 A EP 14158871A EP 2918779 A1 EP2918779 A1 EP 2918779A1
Authority
EP
European Patent Office
Prior art keywords
hot gas
hole
turbine blade
platform
airfoil
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Withdrawn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
EP14158871.5A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
Björn Buchholz
Stefan Dahlke
Daniela Koch
Nihal Kurt
Ralf Müsgen
Radan RADULOVIC
Marco Schüler
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
Siemens Corp
Original Assignee
Siemens AG
Siemens Corp
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, Siemens Corp filed Critical Siemens AG
Priority to EP14158871.5A priority Critical patent/EP2918779A1/fr
Publication of EP2918779A1 publication Critical patent/EP2918779A1/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Withdrawn legal-status Critical Current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
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • 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
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving blades
    • F05D2240/81—Cooled platforms
    • 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

Definitions

  • the invention relates to a turbine blade with an aerodynamically curved airfoil comprising a pressure side wall and a suction side wall, each extending from a leading edge of the airfoil to a trailing edge of the airfoil and in a transverse thereto Spannweitecardi of a blade foot end to a blade end, on the fuß devisem
  • a transversely projecting platform is provided at the end and / or at the head end of the airfoil, such that a hot gas exposable surface of the respective side wall of the airfoil merges via a groove into a surface of the platform exposed to the hot gas, wherein the platform forms one of the hot gas Surface opposite cold gas surface, from which at least one hole extends into the hot gas surface.
  • Such turbine blades are for example from the EP 1 669 544 A1 .
  • the cooling channels opening out in the transitions between the blade leaf and the platform are used to energize the near-wall flow and to manipulate secondary flows.
  • the object of the invention is to provide an alternative turbine blade whose life is further improved.
  • an impact cooling insert is arranged at a distance from the hole wall for cooling the wall surrounding the hole in question.
  • the invention is based on the recognition that the hollow-throat-like transitions between the blade and the platform enriched in their mass compared to the side walls can not be adequately cooled despite the passage through the holes with coolant. For this reason, the invention proposes that in the respective holes an impact cooling insert is arranged at a distance from the hole wall for cooling the wall surrounding the hole in question.
  • an impact cooling insert is arranged at a distance from the hole wall for cooling the wall surrounding the hole in question.
  • the holes are in their cross-section, which is oriented perpendicular to the direction of penetration, slit-like, so that they preferably have a mouth opening, which lies partially in the region of the groove and partially in the region of that hot gas surface, which is attributable to the platform.
  • slot-shaped holes which because of their shape could also be referred to as elongated holes, lead to a reduced rigidity of the platform, which makes the turbine blade more adaptable to different thermal expansions. This leads to the reduction of mechanical stresses.
  • the impact cooling insert arranged in the hole comprises a closed end on the hot gas side, whose surface is matched step-free to the surface of the environment of the hole.
  • the step-free adaptation avoids the occurrence of aerodynamic losses in the hot gas at flows of the airfoil and the platform along the groove.
  • the hot gas side closed end of the baffle insert unnecessarily high consumption of coolant, which moves the amount of coolant for cooling the hole walls within reasonable limits.
  • the coolant exiting through the impingement cooling apertures of the impingement baffle can flow to the hot gas side surface of the turbine blade after impingement cooling, from where the refrigerant faces the turbine blade upstream of the hole to form a film cooling Protect influences of the hot gas.
  • downstream refers to the flow direction of the hot gas, which imposes its direction on the exiting coolant.
  • the turbine blade according to the invention can also be a turbine blade that is already operationally required, which can be upgraded as part of a reprocessing by the subsequent introduction of the said holes or slots and the impact cooling inserts.
  • the cold gas surface of the platform is opposed to a baffle cooling element at a distance at which the baffle cooling insert - or better said, an inflow-side end of the baffle cooling insert - is attached.
  • a particularly simple manner of fastening the impact cooling insert can be provided. Spacers between the perforated wall and impact cooling insert are not required or only to a very limited extent.
  • the relevant holes scoop blade pressure side in one area between 20% and 80% of a chord whose normalized length extends from 0% at the leading edge to 100% at the trailing edge of the airfoil. Since in operation the pressure side wall of the turbine blade is exposed to a different temperature than the suction side wall of the turbine blade and thus both said side walls thermally stretch differently thermally, the turbine blade described above can better compensate for the different thermal strains, since the stiffnesses of the pressure side wall and the suction side wall in the region Attachment - ie in the region of the groove - by the aforementioned features are locally together.
  • the invention thus relates to a turbine blade having an aerodynamically curved airfoil comprising a pressure sidewall and a suction sidewall each extending from a leading edge of the airfoil to a trailing edge of the airfoil and in a transverse spanwise direction from a blade root end to a blade end at the foot end and / or at the head end of the airfoil a transversely projecting platform is provided, such that a hot gas exposable surface of the respective side wall via a groove in a hot gas exposable surface of the platform monolithic, wherein the platform is one of the hot gas Surface has opposite cold gas surface from which extends at least one hole to the hot gas surface.
  • FIG. 1 shows in perspective a part of a cast turbine blade 10 comprising an airfoil 12 which is aerodynamically curved.
  • the airfoil 12 includes in known manner a pressure side wall 14 and a suction side wall 16 opposite thereto, both walls extending from a common front edge 18 to a common rear edge 20.
  • the turbine blade 10 is configured as a so-called cut-back turbine blade, which immediately upstream of the trailing edge 20 includes a plurality of openings 22, from which a coolant, preferably cooling air, can escape from the interior of the turbine blade 10.
  • the openings 22 are separated from each other by webs.
  • Both the pressure side wall 14 and the suction side wall 16 extend from a foot end not shown to a head end 24 of the airfoil 12.
  • a platform 26 is monolithically arranged at the head end 24 of the airfoil 12, with respect to the Extension of the airfoil 12 extending transversely thereof.
  • the surface 28 of the side walls 14, 16 as well as the surface 30 of the platform 26 are each continuously via a groove 32 into each other.
  • Airfoil pressure side, four holes 34 are arranged in the transition region between the platform 26 and the airfoil 12.
  • the holes 34 have a slot shape that is rounded to avoid stress concentrations at their respective ends. Consequently, the holes 34 can also be referred to as elongated holes. In the embodiment shown a total of four slots are provided. It goes without saying that more or less holes can also be provided.
  • the openings of the holes 34 opening in the hot gas surface lie, on the one hand, in that hot gas surface attributable to the platform 26 and, on the other hand, in the hot gas surface attributable to the groove 32.
  • the blade 12 can be characterized in particular by a chord 37, which represents an imaginary line between the front edge 18 and the rear edge 20.
  • the chord has a normalized length of 100%, the beginning coincides with the leading edge 18 and the end coincides with the trailing edge 20 of the airfoil 12.
  • the holes 34 lie in a section of 20% to 70% of the chord 37.
  • FIG. 1 can show the holes 34, if they are slit-shaped, with the chord 37 form an angle ⁇ , which is of the order of about 60 °.
  • the impact cooling inserts according to the invention are in the holes 34 in FIG. 1 not shown.
  • FIG. 2 shows a section through the turbine blade 10 according to FIG. 1 along the section line II-II, but with arranged in the holes 34 impact cooling inserts.
  • Each baffle insert 36 includes a hot gas side closed end 38.
  • the baffles 36 are also made hollow inside so that impact cooling holes 40 are provided in the walls thereof. At the same time a small distance between the outer surface of the baffle insert 36 and the hole walls 42 is provided.
  • the impingement cooling inserts 36 are attached to an impingement cooling element 44 so that they have an inflow end 46 through which cooling air can flow into the interior of the impingement cooling insert 36 during operation.
  • the cooling air in the form of impingement cooling jets exits through the impingement cooling openings 40 and impinges on the perforated walls 42 in an impingement-cooling manner. From there, the cooling gap flows away from the hot gas side surface.
  • FIG. 3 shows in perspective view alone the impingement cooling element 44 with impingement cooling openings 40 arranged thereon and four impact cooling inserts 36 attached thereto. Impact cooling openings 40 are likewise provided in the impingement cooling inserts 36.
  • To form the turbine blade according to the invention is from the cold gas side of the platform 26 forth in FIG. 3 shown impingement cooling element 44 together with the attached impact cooling inserts 36 in the in FIG. 1 illustrated turbine blade 10 is used. The impingement cooling element 44 is then in a known manner at the in FIG. 1 attached turbine blade 10 secured, for example by soldering or welding.
  • the turbine blade 10 can be used as intended in a gas turbine, in which case from the back of the platform forth cooling air to the baffle cooling element 44 and in the baffles 36 can be fed or flowed to the holes 34 surrounding walls 42 in more efficient way to cool than previously form of impingement cooling.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP14158871.5A 2014-03-11 2014-03-11 Aube de turbine Withdrawn EP2918779A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP14158871.5A EP2918779A1 (fr) 2014-03-11 2014-03-11 Aube de turbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP14158871.5A EP2918779A1 (fr) 2014-03-11 2014-03-11 Aube de turbine

Publications (1)

Publication Number Publication Date
EP2918779A1 true EP2918779A1 (fr) 2015-09-16

Family

ID=50241200

Family Applications (1)

Application Number Title Priority Date Filing Date
EP14158871.5A Withdrawn EP2918779A1 (fr) 2014-03-11 2014-03-11 Aube de turbine

Country Status (1)

Country Link
EP (1) EP2918779A1 (fr)

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2253443A (en) 1991-03-05 1992-09-09 Rolls Royce Plc Gas turbine nozzle guide vane arrangement
EP1669544A1 (fr) 2004-12-13 2006-06-14 The General Electric Company Etage de turbine avec contour de raccordement refroidi par couche d'air
EP1688587A2 (fr) 2005-01-10 2006-08-09 General Electric Company Étage de turbine à raccord en forme d'entonnoir

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2253443A (en) 1991-03-05 1992-09-09 Rolls Royce Plc Gas turbine nozzle guide vane arrangement
EP1669544A1 (fr) 2004-12-13 2006-06-14 The General Electric Company Etage de turbine avec contour de raccordement refroidi par couche d'air
EP1688587A2 (fr) 2005-01-10 2006-08-09 General Electric Company Étage de turbine à raccord en forme d'entonnoir

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