EP2584148A1 - Aube de turbine refroidie par film pour une turbomachine - Google Patents

Aube de turbine refroidie par film pour une turbomachine Download PDF

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Publication number
EP2584148A1
EP2584148A1 EP11186222.3A EP11186222A EP2584148A1 EP 2584148 A1 EP2584148 A1 EP 2584148A1 EP 11186222 A EP11186222 A EP 11186222A EP 2584148 A1 EP2584148 A1 EP 2584148A1
Authority
EP
European Patent Office
Prior art keywords
turbine blade
inlet
edge portion
edge
wall
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Withdrawn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
EP11186222.3A
Other languages
German (de)
English (en)
Inventor
Tim-Adrian Albring
Volker Amedick
Tilman Auf Dem Kampe
Thomas Biesinger
Stefan Braun
Tobias Buchal
Jan Münzer
Markus Schmidt
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 EP11186222.3A priority Critical patent/EP2584148A1/fr
Publication of EP2584148A1 publication Critical patent/EP2584148A1/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Withdrawn legal-status Critical Current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
    • F01D5/186—Film cooling

Definitions

  • the invention relates to a turbine blade for a turbomachine, wherein the turbine blade is film-cooled.
  • a turbomachine in particular a gas turbine, has a turbine, in which hot gas, which was previously compressed in a compressor and heated to a combustion chamber, is expanded to obtain work.
  • the turbine is designed in axial construction, wherein the turbine is formed by a plurality of successively in the flow direction blade rings.
  • the blade rings have circumferentially disposed blades and vanes, the blades being mounted on a rotor of the gas turbine and the vanes secured to the housing of the gas turbine.
  • thermodynamic efficiency of the gas turbine is the higher, the higher the inlet temperature of the hot gas into the turbine.
  • thermal load capacity of the turbine blades there are limits with regard to the thermal load capacity of the turbine blades.
  • corresponding materials and material combinations are available for the turbine blades, which, however, according to the current state of the art, only allow an insufficient expansion of the potential for increasing the thermal efficiency of the gas turbine.
  • it is known to cool the turbine blades during operation of the gas turbine whereby the turbine blade itself is exposed to a lower thermal load, as without the cooling due to the thermal load of the hot gas would be the case.
  • the blades are cooled, for example, by means of film cooling.
  • the blades are provided on their surface with a plurality of cooling air holes, via which the blade inner cooling air is transported to the surface of the turbine blades. After leaving the cooling air holes, the cooling air flows in the form of a film along the surface of the turbine blade, thus isolating the surface of the turbine blade from the hot gas. Furthermore, the film cools the turbine blade by convection and additionally acts as a barrier, so that transport of the hot gas to the surface of the turbine blade is prevented.
  • the object of the invention is to provide a turbine blade for a turbomachine, which can be cooled with an effective film cooling.
  • the turbine blade for a turbomachine has an outer wall defining an inner cavity of the turbine blade in which cooling fluid for cooling the turbine blade is provided, wherein in the outer wall at least one passageway is provided through which the cooling fluid from the inner cavity to the outside of the turbine blade for training a cooling film on the outside of the outer wall is inclined and which is inclined towards the trailing edge of the turbine blade and has a shape such that at predetermined operating flow conditions for the turbine blade as it flows through the cooling fluid through the passage so asymmetrical over the cross section of the passage channel velocity distribution of the cooling fluid flow itself shows that in the passageway a pair of counter-rotating vortices is formed, after its exit from the passageway the velocity vectors show the cooling fluid flow between the vortex centers towards the outer wall.
  • the two whirl arms of the horseshoe vortex have the same sense of rotation as their each immediately adjacent arranged vortex of the pair of opposite vortex of the cooling fluid.
  • both the horseshoe vortex and the pair of opposing vortices are attenuated.
  • the pair of counter-rotating vortices are pressed with their counter-rotating turning inside towards the outer wall, whereby the pair of opposite vortices advantageously applies to the outside. Because the flow of cooling fluid between the vortex centers outside the passageway is directed towards the outer wall, there is an impact cooling effect on the outer wall in this region which is particularly effective for cooling the turbine blade.
  • the amount of cooling fluid necessary for cooling the turbine bucket is reduced compared to a cooling fluid amount necessary to cool a conventional turbine bucket.
  • the distance of the through-channels to one another in the turbine blade can be designed to be comparatively large, as a result of which fewer through-channels are required for the turbine blade according to the invention and the structural weakening of the turbine blade according to the invention through its through-channels is smaller.
  • the edge portion of the inlet of the passage channel at its upstream side is formed so sharp-edged relative to the other edge portions of the inlet, that in the passage on its upstream side a separation region of the cooling fluid flow can be formed, resulting in the asymmetric velocity distribution of the cooling fluid flow.
  • Preferred dimensions is at the edge portion of the inlet of the passage channel at the upstream side a detaching nose formed with a detaching edge such that this edge portion is formed by the detachment edge sharp-edged relative to the opposite edge portion of the inlet.
  • the detachment nose preferably has an acute angle at its separation edge, as a result of which the detachment and therefore also the pair of opposing vertebrae can advantageously be formed particularly strongly.
  • the detachment nose preferably projects with its detachment edge from the inside of the outer wall into the inner cavity. It is alternatively preferred that a recess is provided immediately adjacent or at a distance from the edge portion of the inlet of the passageway at its upstream side in the inside of the outer wall, whereby the detaching nose is formed as a projection of the outer wall. In this case, the detaching edge of the detaching nose is preferably aligned with the inside of the outer wall.
  • the side facing away from the passage channel of the detaching nose is concave and extends edge-free from the detachment edge to the inside.
  • the separation edge can be advantageously formed particularly sharp.
  • a bead is formed such that this edge portion is formed blunt relative to the opposite edge portion of the inlet.
  • the bead is preferably shaped convexly so that the cooling fluid flow can be flowed free of charge to the inlet of the through-channel on its downstream side.
  • the edge portion of the inlet of the through-channel is formed on its downstream side in such a round shape that this edge portion is formed blunt relative to the opposite edge portion of the inlet.
  • a possibly occurring separation region on the downstream side of the through-channel is advantageously avoidable.
  • the shape of the turbine blade according to the invention is suitable for casting, which makes it possible for the turbine blade to be produced by a casting method without any further adjustments.
  • the passageways are preferably by drilling, in particular laser drilling to produce, wherein the inventive shape of the inlet is preferably carried out by drilling from within the turbine blade.
  • the drilling inside the turbine blade is particularly advantageous if the turbine blade is provided on the outside with a coating, in particular a ceramic coating.
  • FIG. 1 shows a portion of an outer wall 2 of a turbine blade 1 of a turbomachine.
  • the outer wall 2 defines an inner cavity 3 and has an outer side 8 and an inner side 9.
  • a hot gas flow 13 occurs on the outside 8 with a hot gas main flow direction 14 which is substantially parallel to the outside 8 and which is directed from the leading edge to the trailing edge (not shown in the figure) of the turbine blade 1.
  • a plurality of through channels 4 is introduced, which are inclined in the direction of flow directed from the inside out to the rear edge of the turbine blade 1 and with the outside 8 include an acute angle of inclination 10.
  • the passage 4 in FIG. 1 has an inlet 11 facing the internal cavity 9 and an outlet 12 on the outside.
  • the passage 4 is cylindrical and has an axis of symmetry 5.
  • the through-channel 4 has an upstream side 16 relative to the hot-gas main flow direction 14 and a side 17 downstream of the hot-gas main flow direction 14.
  • the inlet 11 of the through-channel 4 has an upstream edge section 18 on the upstream side 16 and a downstream edge section 19 on the downstream side 17.
  • a cooling fluid 6, which flows into the through-channel 4 via the inlet 11 and flows out of the through-channel 4 via the outlet 12, is located in the inner hollow space 3. After leaving the through-channel 4, the cooling fluid 6 forms a cooling film 7 on the outside 8.
  • FIG. 1 it can be seen, on the inside 9 of the outer wall 2, formed on the upstream edge portion 18 of the inlet 11, a detaching nose 25 with a detaching edge.
  • the detachment nose 25 has a downstream detachment nose backside 27, which defines the passage 4 in the region of the inlet 11, and an upstream Ablösasenvorderseite 26, which extends to the inner side 9 of the outer wall 2.
  • the detachment edge is formed by the detachment nose front 26 and the detachment nose back 27.
  • a detachment nose 25 is provided for each of the through-channels 4. It is also conceivable that a plurality of passageways 4 is arranged in a row and the detachment nose is integrally formed and extends along the row.
  • a bead 28 is arranged on the inside 9 of the outer wall 2 at the downstream edge portion 19 of the inlet 11.
  • the bead inside 30 extending into the inner cavity 3 has a convex-shaped portion. It is conceivable that the convex-shaped region extends over the entire upstream bead front side 29 as far as the downstream edge section 19 and / or into the through-channel 4.
  • the bead 28 may also have a rectangular cross-section with rounded on the inside of the bead 30 corners. It is conceivable that for each of the through-channels 4 each a bump-like bead 28 is provided.
  • a plurality of passageways 4 is arranged in a row and the bead 28 extends along the row. It is also conceivable that the upstream edge portion 18 is rounded so that a detachment of this edge portion 18 is prevented.
  • the downstream edge portion 19 is blunt such that the cooling fluid flow 15 abuts against this edge portion 19.
  • the separating edge of the detaching nose (25) arranged at the upstream edge portion 18 is sharp enough so that the cooling fluid flow 15 can not follow this edge portion 18 so that a separation region 20 of the cooling fluid flow 15 forms in the passage 4 on the upstream side 16.
  • a centric transverse flow 21 is induced in the passage 4, which is directed from the upstream side 16 to the downstream side 17.
  • an opposite vortex pair 22 with two vortex centers 24 is produced in the passage 4, the velocity vectors between the two vortex centers 24 pointing to the downstream side 17 of the passage 4.
  • the velocity vectors of the cooling fluid flow 15 of the counter-rotating vortical pair 22 are directed between the vortex centers 24 to the outer wall 2 toward the exit of the counter-rotating vortex pair 22 from the passage 4.
  • the hot gas flow 13 flows around the counter-rotating vortex pair 22, whereby a horseshoe vortex 23 forms from the hot gas.
  • the horseshoe vortex 23 has two swirl arms which are arranged on opposite sides of the opposite swirl pair 22. Each of the vortex arms is formed by a vortex, wherein the velocity vectors of the hot gas flow 13 are directed between the vortex centers 24 of the vortex arms on the outer wall 2.
  • the swirl arms have the same sense of rotation as their each immediately adjacent arranged vortex of the opposite vortex pair 22, so that the horseshoe vortex 23 is attenuated and the transport of the hot gas is reduced to the outside 8 of the outer wall 2.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP11186222.3A 2011-10-21 2011-10-21 Aube de turbine refroidie par film pour une turbomachine Withdrawn EP2584148A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP11186222.3A EP2584148A1 (fr) 2011-10-21 2011-10-21 Aube de turbine refroidie par film pour une turbomachine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP11186222.3A EP2584148A1 (fr) 2011-10-21 2011-10-21 Aube de turbine refroidie par film pour une turbomachine

Publications (1)

Publication Number Publication Date
EP2584148A1 true EP2584148A1 (fr) 2013-04-24

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ID=45033725

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EP11186222.3A Withdrawn EP2584148A1 (fr) 2011-10-21 2011-10-21 Aube de turbine refroidie par film pour une turbomachine

Country Status (1)

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EP (1) EP2584148A1 (fr)

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2990606A1 (fr) * 2014-08-26 2016-03-02 Siemens Aktiengesellschaft Aube de turbine
EP2993304A1 (fr) * 2014-09-08 2016-03-09 United Technologies Corporation Composant de moteur à turbine à gaz avec trou formant un film de refroidissement
EP3012407A1 (fr) * 2014-10-20 2016-04-27 United Technologies Corporation Trou de refroidissement à film avec accumulateur d'écoulement en saillie
EP3239462A1 (fr) * 2016-04-26 2017-11-01 General Electric Company Profil aérodynamique pour un moteur à turbine
EP3375978A1 (fr) * 2017-03-15 2018-09-19 Mitsubishi Hitachi Power Systems, Ltd. Aube de turbine refroidie par film
EP3543477A1 (fr) * 2018-03-19 2019-09-25 United Technologies Corporation Entrée masquée vers des trous d'effusion
FR3111661A1 (fr) * 2020-06-22 2021-12-24 Safran Aircraft Engines Aube de turbine avec système de refroidissement

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2385900A1 (fr) * 1978-03-20 1978-10-27 Rolls Royce Aube mobile refroidie pour moteur a turbine a gaz
GB2262314A (en) * 1991-12-10 1993-06-16 Rolls Royce Plc Air cooled gas turbine engine aerofoil.
US5419039A (en) * 1990-07-09 1995-05-30 United Technologies Corporation Method of making an air cooled vane with film cooling pocket construction
US5624231A (en) * 1993-12-28 1997-04-29 Kabushiki Kaisha Toshiba Cooled turbine blade for a gas turbine
WO1998037310A1 (fr) * 1997-02-20 1998-08-27 Siemens Aktiengesellschaft Aube de turbine et son utilisation dans un systeme de turbine a gaz
DE10236676A1 (de) * 2002-08-09 2004-02-19 Rolls-Royce Deutschland Ltd & Co Kg Turbinenschaufel für eine Gasturbine mit zumindest einer Kühlungsausnehmung

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2385900A1 (fr) * 1978-03-20 1978-10-27 Rolls Royce Aube mobile refroidie pour moteur a turbine a gaz
US5419039A (en) * 1990-07-09 1995-05-30 United Technologies Corporation Method of making an air cooled vane with film cooling pocket construction
GB2262314A (en) * 1991-12-10 1993-06-16 Rolls Royce Plc Air cooled gas turbine engine aerofoil.
US5624231A (en) * 1993-12-28 1997-04-29 Kabushiki Kaisha Toshiba Cooled turbine blade for a gas turbine
WO1998037310A1 (fr) * 1997-02-20 1998-08-27 Siemens Aktiengesellschaft Aube de turbine et son utilisation dans un systeme de turbine a gaz
DE10236676A1 (de) * 2002-08-09 2004-02-19 Rolls-Royce Deutschland Ltd & Co Kg Turbinenschaufel für eine Gasturbine mit zumindest einer Kühlungsausnehmung

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2990606A1 (fr) * 2014-08-26 2016-03-02 Siemens Aktiengesellschaft Aube de turbine
EP2993304A1 (fr) * 2014-09-08 2016-03-09 United Technologies Corporation Composant de moteur à turbine à gaz avec trou formant un film de refroidissement
US10982552B2 (en) 2014-09-08 2021-04-20 Raytheon Technologies Corporation Gas turbine engine component with film cooling hole
EP3012407A1 (fr) * 2014-10-20 2016-04-27 United Technologies Corporation Trou de refroidissement à film avec accumulateur d'écoulement en saillie
US9957810B2 (en) 2014-10-20 2018-05-01 United Technologies Corporation Film hole with protruding flow accumulator
EP3239462A1 (fr) * 2016-04-26 2017-11-01 General Electric Company Profil aérodynamique pour un moteur à turbine
EP3375978A1 (fr) * 2017-03-15 2018-09-19 Mitsubishi Hitachi Power Systems, Ltd. Aube de turbine refroidie par film
CN108625905A (zh) * 2017-03-15 2018-10-09 三菱日立电力系统株式会社 涡轮叶片以及具备该涡轮叶片的燃气轮机
US10415398B2 (en) 2017-03-15 2019-09-17 Mitsubishi Hitachi Power Systems, Ltd. Turbine blades and gas turbine having the same
EP3543477A1 (fr) * 2018-03-19 2019-09-25 United Technologies Corporation Entrée masquée vers des trous d'effusion
US10823414B2 (en) 2018-03-19 2020-11-03 Raytheon Technologies Corporation Hooded entrance to effusion holes
FR3111661A1 (fr) * 2020-06-22 2021-12-24 Safran Aircraft Engines Aube de turbine avec système de refroidissement

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