EP3112593A1 - Aube de turbine a refroidissement interieur - Google Patents

Aube de turbine a refroidissement interieur Download PDF

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Publication number
EP3112593A1
EP3112593A1 EP15175300.1A EP15175300A EP3112593A1 EP 3112593 A1 EP3112593 A1 EP 3112593A1 EP 15175300 A EP15175300 A EP 15175300A EP 3112593 A1 EP3112593 A1 EP 3112593A1
Authority
EP
European Patent Office
Prior art keywords
turbine blade
airfoil
cavity
blade
edge
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Withdrawn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
EP15175300.1A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
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 EP15175300.1A priority Critical patent/EP3112593A1/fr
Publication of EP3112593A1 publication Critical patent/EP3112593A1/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Withdrawn legal-status Critical Current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
    • 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, which successively comprises a foot region for fastening the turbine blade to a blade carrier, a platform region and an aerodynamically curved, internally cooled airfoil along a longitudinal axis, the airfoil having a suction-side airfoil wall and a pressure-side airfoil wall extending from a leading edge of the airfoil Blade blade extend to a trailing edge of the airfoil, wherein in the interior of the airfoil, a front edge adjacent cavity is provided, which is bounded by at least one of the two airfoil walls associated inner surface partially and which inner surface is prallkühlbar arranged by means disposed in the turbine blade means.
  • Such turbine blades are well known in the art and include in their leading edge mostly a variety of so-called spray head holes, which are also referred to in English as a showerhead.
  • spray head holes are also referred to in English as a showerhead.
  • the introduction of these spray holes is relatively expensive, which is considered disadvantageous.
  • the spray head holes cause notch stresses in the material surrounding the holes, which under operating stress can cause premature failure of the turbine blade end of life.
  • safe operation of a gas turbine equipped with this turbine blade is no longer reliably possible, which is why it is then set quiet and the turbine blades are replaced. This reduces the availability of the gas turbine.
  • the object of the invention is therefore to provide a turbine blade with an improved life, which is also inexpensive to produce.
  • the front edge is free of spray head holes and that for discharging the cooling medium from the cavity at least one connected to the cavity, extending through the platform area and opening at an edge of the platform area outflow channel is provided.
  • the invention is based on the finding that the use of film cooling that protects the leading edge is not always absolutely necessary. It has been found that the impingement cooling of the leading edge by means disposed in the interior of the turbine blade can be sufficient to adequately cool the material of the blade in the region of its leading edge while achieving the desired service life. After the impingement cooling, however, the cooling medium is discharged from the front room. Contrary to the usual procedure that dissipate heated by the impingement cooling medium through spray head holes, it is now proposed according to the invention that at least one, but preferably each platform region more outflow channels are provided.
  • outflow channels are arranged such that they essentially extend through the platform region and thus fluidly connect the cavity to a space adjoining the edge of the platform region so that the warmed cooling medium can leave the turbine blade at the edge of the platform region.
  • the material forming the platform region of the turbine blade can thereby also cool down.
  • the emerging at the edge of the platform area cooling medium can then be advantageously used for blocking of columns, which by the juxtaposition of platforms adjacent turbine blades are inherently inherent within the gas turbine structure.
  • one and the same coolant is used for several consecutive functions: first, the cooling medium cools the material of the leading edge of the airfoil by impingement cooling. Subsequently, the cooling medium flows through the discharge channels, which are located in the platform area and meanwhile cools the platform of the turbine blade. Finally, the now heated baffle coolant serves to block turbine gaps and thus avoids the local hot gas intake.
  • Another advantage of the turbine blade according to the invention is that the expenses for the introduction of Sprühkopf holes omitted, so that the turbine blade is cheaper to manufacture than that known from the prior art. Likewise, their cooling medium requirement is reduced, compared to the known variants.
  • the outflow channel opens into a groove sunk in the edge of the platform region.
  • the recessed groove in the edge of the platform area serves to receive plate-shaped sealing elements.
  • the plate-shaped sealing elements can be better protected from the effects of hot gas.
  • the means comprises a rib connecting the pressure-side blade wall to the suction-side blade wall, in which the impact cooling openings are provided.
  • Such ribs are produced simultaneously during casting of the turbine blade comprising at least the platform and the blade, resulting in a simple and economically producible turbine blade.
  • the means may also be designed as an impingement cooling insert, which is produced in a casting process Turbine blade is used later.
  • the cavity referred to above is then that space in the interior of the turbine blade, which is arranged between the inner surface of the cast blade walls and the impact cooling insert in the region of the leading edge.
  • the turbine blade may be configured either as a turbine blade or as a turbine vane.
  • FIG. 1 shows a longitudinal section through a turbine blade 10 according to the invention, which is designed according to the embodiment shown as a guide vane.
  • the turbine blade 10 comprises, along a longitudinal axis 12, a hook-shaped foot region 14, a first platform region 16, and an aerodynamically curved, internally cooled airfoil blade 18 in succession
  • FIG. 1 is the inner structure of the airfoil 18 is shown.
  • the airfoil 18 includes a suction side airfoil wall (not shown) and a pressure side airfoil wall extending from a leading edge 20 of the airfoil to a trailing edge 22 of the airfoil.
  • a further platform region 24 may be arranged.
  • a total of four juxtaposed cavities 26a to 26d are arranged in the interior of the airfoil 18.
  • the cavities 26 extend from the first platform region 16 to the further platform region 24.
  • the cavities 26b and 26c are interconnected via a first deflection region 28a, just as the cavities 26c and 26d are fluidically connected to each other via a second deflection region 28b.
  • a plurality of so-called sockets 30 are arranged in the cavity 26d. These sockets 30 are also known as pin-fins. Via openings 32 separated from each other by webs 32, the cavity 26 is in fluid communication with the surroundings of the turbine blade 10.
  • the cavities 26 are partially separated from each other by ribs 27a, 27b and 27c oriented parallel to the longitudinal axis 12.
  • a plurality of impingement cooling holes 36 are provided as a means enabling impingement cooling of the leading edge, whereby the cavities 26a and 26b can communicate with each other.
  • outflow channels 38 are provided in the platform regions 16, 24, whereby the cavity 26a is in fluid communication with the surroundings of the turbine blade 10.
  • the cavity 26b also has an opening 29, via which the airfoil 18 of the turbine blade 10, a cooling medium 42 can be fed.
  • the blade 18 is flowed around by a hot gas, which flows first to the front edge 20, then both the suction side as well as the pressure-side blade wall and flows around at the trailing edge 22.
  • a hot gas which flows first to the front edge 20, then both the suction side as well as the pressure-side blade wall and flows around at the trailing edge 22.
  • the annular flow channel of the gas turbine which is annular in cross-section, is bounded radially by the platforms 16, 24. So that the material of the turbine blade 10 reaches the predetermined stability, this is cooled by supplying the cooling medium 42 from the inside. For this purpose, the cooling medium 42 is guided into the interior of the turbine blade 10 via the feed opening 29. It then flows into the cavity 26b from where it is distributed in different partial streams.
  • a significant partial flow passes to the end 29 of the cavity 26b opposite the opening 29, where it flows into the first deflection region 28a, after which it subsequently flows through the cavity 26c. Thereafter, the cooling medium 42 flows through the second deflection region 28b and then flows into the fourth cavity 26d past the bases 30, whereafter it leaves the turbine blade 10 by flowing through the outlet openings 34.
  • cooling medium 42 flows through the individual impingement cooling openings 36, whereby individual impingement cooling jets 44 arise, which impinges on that inner surface of the blade walls which are formed in the region of the front edge 20 of the suction-side airfoil wall and / or the pressure-side airfoil wall. Subsequently, the cooling medium 42 leaves the cavity 26a through the outflow channels 38.
  • a plurality of outflow channels 38 may be provided, wherein the front edge 20 has no spray head holes.
  • the outflow channels 38 can open into a groove 45 arranged on the edge 43 of the platform.
  • FIG. 2 shows a second embodiment of a turbine blade 10 according to the invention.
  • first embodiment shows a second embodiment of a turbine blade 10 according to the invention.
  • an impingement cooling insert 50 is provided as a means enabling impingement cooling of the leading edge.
  • the impact cooling insert 50 is tubular and has on its inner surface opposite the tube wall 52, the impact cooling openings 36.
  • the turbine blade 10 has, in addition to a first feed opening 29 for cooling medium, a further feed opening 31.
  • the feed opening 31 is part of the impingement cooling insert 50, as a result of which the cooling medium can be fed thereto.
  • the cooling medium which can be fed to the turbine blade 10 via the feed opening 29 passes to the outlet openings 34 via the cavities 26b, 26c and 26d.
  • the impact cooling insert 50 has, at its end 54 opposite the feed opening 31, a passage 58 through which a partial stream 56 of the cooling medium supplied to the impact cooling insert 50 can exit to be used there for further purposes.
  • the invention relates to a turbine blade 10 which comprises, along a longitudinal axis 12, a foot region 14 for fastening the turbine blade to a blade carrier, a platform region 16 and an aerodynamically curved internally coolable airfoil 18, wherein the airfoil 18 has a suction-side airfoil wall and a pressure-side airfoil wall which extends each extend from a front edge 20 to a trailing edge 22, wherein in the interior of the blade 18, a front edge 20 adjacent to the cavity 26 a is provided, which is partially bounded by an at least one of the two airfoil walls associated inner surface and which inner surface by in the turbine blade 10 Means 27a, 36, 50 is prallkühlbar.
  • the Leading edge 20 is free of spray head holes, and that for discharging the impingement cooling medium 42 from the cavity 26a at least one connected to the cavity 26a extending through the platform portion 16, 24 and opening out at an edge of the platform region outflow channel 38 is provided.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP15175300.1A 2015-07-03 2015-07-03 Aube de turbine a refroidissement interieur Withdrawn EP3112593A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP15175300.1A EP3112593A1 (fr) 2015-07-03 2015-07-03 Aube de turbine a refroidissement interieur

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP15175300.1A EP3112593A1 (fr) 2015-07-03 2015-07-03 Aube de turbine a refroidissement interieur

Publications (1)

Publication Number Publication Date
EP3112593A1 true EP3112593A1 (fr) 2017-01-04

Family

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Family Applications (1)

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EP15175300.1A Withdrawn EP3112593A1 (fr) 2015-07-03 2015-07-03 Aube de turbine a refroidissement interieur

Country Status (1)

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EP (1) EP3112593A1 (fr)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3594449A1 (fr) * 2018-07-13 2020-01-15 Honeywell International Inc. Aube de turbine avec système de refroidissement à tolérance à la poussière
CN111485957A (zh) * 2020-04-29 2020-08-04 中国航发湖南动力机械研究所 涡轮导向冷却叶片
CN111927563A (zh) * 2020-07-31 2020-11-13 中国航发贵阳发动机设计研究所 一种适用于高温环境的涡轮叶片
US11230929B2 (en) 2019-11-05 2022-01-25 Honeywell International Inc. Turbine component with dust tolerant cooling system

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4293275A (en) * 1978-09-14 1981-10-06 Hitachi, Ltd. Gas turbine blade cooling structure
GB2210415A (en) * 1987-09-25 1989-06-07 Toshiba Kk Turbine vane with cooling features
US20120315150A1 (en) * 2011-06-09 2012-12-13 Mitsubishi Heavy Industries, Ltd. Turbine rotor blade
US20130280091A1 (en) * 2012-04-24 2013-10-24 Mark F. Zelesky Gas turbine engine airfoil impingement cooling

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4293275A (en) * 1978-09-14 1981-10-06 Hitachi, Ltd. Gas turbine blade cooling structure
GB2210415A (en) * 1987-09-25 1989-06-07 Toshiba Kk Turbine vane with cooling features
US20120315150A1 (en) * 2011-06-09 2012-12-13 Mitsubishi Heavy Industries, Ltd. Turbine rotor blade
US20130280091A1 (en) * 2012-04-24 2013-10-24 Mark F. Zelesky Gas turbine engine airfoil impingement cooling

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3594449A1 (fr) * 2018-07-13 2020-01-15 Honeywell International Inc. Aube de turbine avec système de refroidissement à tolérance à la poussière
US10989067B2 (en) 2018-07-13 2021-04-27 Honeywell International Inc. Turbine vane with dust tolerant cooling system
US11448093B2 (en) 2018-07-13 2022-09-20 Honeywell International Inc. Turbine vane with dust tolerant cooling system
US11713693B2 (en) 2018-07-13 2023-08-01 Honeywell International Inc. Turbine vane with dust tolerant cooling system
US11230929B2 (en) 2019-11-05 2022-01-25 Honeywell International Inc. Turbine component with dust tolerant cooling system
CN111485957A (zh) * 2020-04-29 2020-08-04 中国航发湖南动力机械研究所 涡轮导向冷却叶片
CN111485957B (zh) * 2020-04-29 2022-11-25 中国航发湖南动力机械研究所 涡轮导向冷却叶片
CN111927563A (zh) * 2020-07-31 2020-11-13 中国航发贵阳发动机设计研究所 一种适用于高温环境的涡轮叶片

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