EP3333366A1 - Aube de turbine avec refroidissement du bord d'attaque - Google Patents

Aube de turbine avec refroidissement du bord d'attaque Download PDF

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Publication number
EP3333366A1
EP3333366A1 EP16202826.0A EP16202826A EP3333366A1 EP 3333366 A1 EP3333366 A1 EP 3333366A1 EP 16202826 A EP16202826 A EP 16202826A EP 3333366 A1 EP3333366 A1 EP 3333366A1
Authority
EP
European Patent Office
Prior art keywords
leading edge
edge portion
turbine
holes
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Withdrawn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
EP16202826.0A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
Daniela Koch
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 EP16202826.0A priority Critical patent/EP3333366A1/fr
Publication of EP3333366A1 publication Critical patent/EP3333366A1/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Withdrawn legal-status Critical Current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
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • F05D2220/321—Application in turbines in gas turbines for a special turbine stage
    • 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
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/303—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the leading edge of a rotor blade
    • 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 a turbine blade for a turbomachine, comprising a blade platform and an airfoil projecting from the blade platform with a peripheral wall having a leading edge, a trailing edge, a pressure-side wall portion and a suction-side wall portion and defining a cavity extending at least along the Front edge extends, wherein in the region of the leading edge for cooling a plurality of the cavity with the outside of the airfoil connecting and spaced apart cooling bores is provided which form a plurality of rows of bores, each extending in a circumferential direction of the airfoil and between the blade platform and the opposite end of the airfoil are stacked in a longitudinal direction of the leading edge. Furthermore, the invention relates to a gas turbine with such turbine blades.
  • Such turbine blades are known in the prior art in different designs and are used in turbomachines to convert the thermal and kinetic energy of a working fluid, in particular a hot gas, into rotational energy.
  • a turbine blade includes a blade platform and an airfoil that protrudes from the blade platform.
  • the airfoil has a peripheral wall that includes an upstream side leading edge, a downstream side trailing edge, a pressure side wall portion, and a suction side wall portion, and defining a cavity therein that extends along the leading edge.
  • Gas turbine engines include a housing through which an annular flow passage extends in an axial direction.
  • a plurality of turbine stages are arranged one behind the other in the axial direction and spaced from each other.
  • Each turbine stage includes a plurality of turbine blades that form a stator vane (stator) connected to the housing and a rotor blade (rotor) connected to a rotor centrally supported and passing through the housing in the axial direction.
  • the flow channel is flowed through by an expanding hot gas.
  • the hot gas flowing through the flow channel is deflected by the guide vanes in such a way that it optimally flows against the rotor blades arranged behind it.
  • the torque generated by the hot gas by means of the blades puts the rotor in rotation.
  • the rotational energy of the rotor can then be converted into electrical energy, for example by means of a generator.
  • thermodynamic efficiency of a gas turbine is the higher, the higher the inlet temperature of the hot gas is in the gas turbine.
  • the height of the inlet temperature is limited, inter alia, by the thermal load capacity of the turbine blades. Accordingly, it is an object to provide turbine blades, which have sufficient mechanical resistance for the operation of the gas turbine even at very high inlet temperatures of the hot gas.
  • turbine blades are usually provided with elaborate coating systems.
  • a further increase in the allowable inlet temperature of the hot gas during operation of the gas turbine can be achieved by cooling the turbine blades.
  • cavities are provided in the blades, through which a cooling fluid is passed.
  • the cooling fluid flows through the cavity of the airfoil, for example, by impingement cooling, in which the cooling fluid is guided so that it bounces from the inside to the peripheral wall of the airfoil, and / or cooled by film cooling, in the Cooling fluid flows through the provided in the peripheral wall of the airfoil cooling holes from the cavity to the outside of the airfoil and forms an outer cooling film there.
  • a plurality of cooling holes are provided in the region of the leading edge of the turbine blade for cooling thereof, which connect the cavity in the interior of the airfoil with its outer side and are arranged at a distance from each other.
  • the cooling holes thereby form a plurality of rows of holes, each extending in a circumferential direction of the airfoil and are arranged one above the other between the blade platform and the opposite end of the airfoil in a longitudinal direction of the leading edge.
  • each row of holes may have three cooling holes, as exemplified in FIG FIG. 5 is shown.
  • the object is achieved by a turbine blade of the aforementioned type, in which in a longitudinal direction of the middle leading edge portion, the rows of holes have a greater number of cooling holes than in a lower leading edge region and / or in an upper leading edge portion.
  • the invention is based on the idea of variably selecting the number of cooling holes per row of holes - viewed in the longitudinal direction of the leading edge - in accordance with the actual thermal load by the stagnation point migrating during operation of the turbomachine. For this purpose, in a central, thermally more heavily loaded by the stagnation point leading edge portion, the rows of holes on a larger number of cooling holes than in an outer lower and / or upper leading edge portion. In this way, the total number of cooling holes in the region of the leading edge is reduced, which is accompanied on the one hand by a low coolant consumption.
  • At least one thermally little loaded outer leading edge portion is less cooled than the middle leading edge portion, whereby with respect to the longitudinal direction of the leading edge of the turbine blade the temperature gradient reduced accordingly.
  • a small temperature gradient can reduce the material stresses in the region of the front edge, in particular stresses between a coating system provided on the outside of the peripheral wall and the peripheral wall of the airfoil and thus increase the service life of the turbine blade. Apart from that, a small number of cooling bores result in correspondingly low production costs of the turbine blade.
  • the rows of holes in the central leading edge portion each have three cooling holes. This corresponds to a conventional shower head design in the middle leading edge portion.
  • the rows of holes in the lower leading edge portion and / or in the upper leading edge portion each have two cooling holes.
  • fewer cooling holes are thus provided in the latter, whereby the cooling fluid consumption can be reduced and the temperature gradient with respect to the longitudinal direction of the leading edge can be reduced.
  • the length of the lower leading edge portion and / or the upper leading edge portion is each in the range of 20% to 40% of the total length of the leading edge and is preferably about 30%.
  • the length of the middle leading edge portion is in the range of 20% to 60% of the total length of the leading edge and is preferably about 40%, as viewed in the longitudinal direction. In other words, fewer cooling holes per row of holes are provided over a length of between 40% and 80% of the total length of the leading edge.
  • the number of rows of holes in the central leading edge portion may be greater than that Number of rows of holes in the lower leading edge portion and / or in the upper leading edge portion. This number ratio is established at the above aspect ratio between at least one outer leading edge portion and the middle leading edge portion when the rows of holes are formed equidistantly in the longitudinal direction of the leading edge over the entire length thereof.
  • the distance between adjacently arranged rows of holes can be in the range of 1 mm to 5 mm and preferably 3 mm. In this way, a sufficient cooling film can be produced at the leading edge of the turbine blade.
  • the distance between adjacently arranged rows of holes increases - as viewed in the longitudinal direction - towards the outside. This is accompanied by a further reduction of the rows of holes and correspondingly reduces the cooling fluid consumption more.
  • the distance between adjacently arranged cooling bores preferably lies in the range from 1 mm to 5 mm and is preferably 3 mm. Such distances between the cooling holes of a row of holes are conducive to producing a sufficient cooling film at the leading edge of the airfoil.
  • the turbine blade may be formed as a guide vane.
  • the turbine blade may be formed as a blade.
  • the inventively improved cooling bore arrangement at the leading edge of the airfoil is equally suitable for guide vanes and rotor blades, in particular also vanes with two blade platforms, which are arranged at the opposite outer ends of the airfoil.
  • turbomachine comprising at least two turbine stages arranged one behind the other in a flow direction, each turbine stage comprising a vane ring formed by a plurality of vanes and a blade ring formed therefrom of a plurality of blades, the vanes and / or the vanes Blades are turbine blades according to the invention.
  • Turbomachines with turbine blades according to the invention have a relatively low cooling fluid consumption and allow longer maintenance intervals, which is associated with lower downtime and thus higher efficiency.
  • the length of the lower leading edge portions and / or the upper leading edge portions is in the range of 25% to 35% of the total length of the leading edge and is preferably 30%.
  • the length of the lower leading edge portion and / or the upper leading edge portion may range from 30% to 40% of the total length of the leading edge, and preferably 35%. This is accompanied by a reduction in the length of the central leading edge portion of the turbine blades of the second turbine stage by up to 10 percentage points from the first turbine stage the total consumption of cooling fluid of the turbomachine is further reduced.
  • FIGS. 1 to 3 show three preferred embodiments of a turbine blade for a turbomachine according to the present invention, which is designed here as a guide vane for a gas turbine.
  • the FIG. 1 shows a turbine blade 1 according to a first embodiment of the present invention.
  • the turbine blade 1 comprises a blade platform 2, an airfoil 3, which protrudes from the blade platform 2 and a Blade foot 4, which projects from the blade platform 2 opposite to the blade 3.
  • the airfoil 3 has a peripheral wall 5 with a front edge 6 upstream of a flow direction S of a turbomachine, a downstream trailing edge 7, a pressure-side wall section 8 and a suction-side wall section 9. Furthermore, the circumferential wall 5 defines in the interior of the airfoil 3 a cavity, not visible in the figures, which extends from the blade platform 2 along the front edge 6.
  • a plurality of cooling bores 10 is provided for the cooling thereof, which connect the cavity to the outside of the airfoil 3 and are arranged at a distance from one another.
  • the cooling holes 10 form a plurality of rows of holes 11, each extending in a circumferential direction U of the blade 3.
  • the rows of holes 11 are arranged one above the other between the blade platform 2 and the opposite end of the blade 3 in a longitudinal direction L of the front edge 6.
  • the turbine blade 1 can have further cooling bores which are provided, for example, at the trailing edge 7 or the wall sections 8, 9 and are not shown in the figures for the sake of clarity.
  • the rows of holes 11 have a greater number of cooling holes 10, as in an upper leading edge portion 14 and a lower leading edge portion 13.
  • the rows of holes 11 in the middle leading edge portion 12 each have three cooling holes 10 while the rows of holes 11 in the lower leading edge portion 13 and in the upper leading edge portion 14 each have two cooling holes 10.
  • the length L 1 of the lower leading edge portion 13 and the length L 3 of the upper leading edge portion 14 are - the longitudinal direction L of view - 30% each of the total length L V of the leading edge 6. Accordingly, the length L 2 of the central leading edge portion 12 is - in the longitudinal direction L of view - 40% of the total length L V of the front edge. 6
  • the distances d L between adjacently arranged rows of holes 11 are identical over the total length L V of the front edge and amount to 3 mm. Accordingly, the number of rows of holes 11 in the central leading edge portion 12 is greater than the number of rows of holes 11 in the lower leading edge portion 13 and in the upper leading edge portion 14.
  • the distance d U between adjacently located cooling holes 10 within a row of holes 11 is in the range of 1 mm up to 5 mm.
  • cooling holes 10 can also be provided in the region of the leading edge of a guide vane, in particular a guide vane with two opposing vane platforms, between which an airfoil is arranged.
  • the FIG. 2 shows a turbine blade 1 according to another embodiment of the present invention.
  • the turbine blade 1 has the same basic structure as the turbine blade according to the first embodiment of the present invention described above.
  • the arrangement of the cooling holes 10 at the leading edge 6 of the turbine blade 6 differs in that the length L 1 of the lower leading edge portion 13 and the length L 3 of the upper leading edge portion 14 - viewed in the longitudinal direction L - each 35% of the total length L V of the leading edge 6 amount.
  • the length L 2 of the central leading edge portion 12 is - in the longitudinal direction L of view - 30% of the total length L V of the front edge. 6
  • the FIG. 3 shows a turbine blade 1 according to a third embodiment of the present invention.
  • the turbine blade 1 has the same basic structure as the above-described turbine blade according to the first embodiment of the present invention.
  • the lengths L 1 , L 2 , L 3 of the lower leading edge portion 13, the middle leading edge portion 12 and the upper leading edge portion 14 are identical.
  • the rows of holes 11 in the lower leading edge portion 13 and in the upper leading edge portion 14 - viewed in the longitudinal direction L - are not arranged equidistantly one above the other, but increase towards the outside.
  • one advantage of the turbine blade 1 according to the invention is that its leading edge 6 can be cooled more efficiently.
  • Both the lower leading edge portion 13 and the upper leading edge portion 14 are subjected to a lower thermal load during operation of the turbomachine than the central leading edge portion 12, since only the latter is hit by the stagnation point of the hot gas flow. As a result, less cooling fluid is needed to cool the lower leading edge portion 13 and the upper leading edge portion 14 than cooling the middle leading edge portion 12. Accordingly, in the lower leading edge portion 13 and in the upper leading edge portion 14 are each provided a smaller number of cooling holes 10 than in the middle leading edge portion 12, which is associated with a small cooling fluid consumption of the turbine blade 1.
  • FIG. 4 shows a partial cross section of a turbomachine 15 according to an embodiment of the present invention, which is a gas turbine.
  • the turbomachine 15 comprises five turbine stages 16, which are arranged one behind the other in the flow direction S.
  • each turbine stage 16 comprises a vane ring 17, which is formed from guide vanes 18, and a blade ring 19 arranged behind it, which is formed from a plurality of rotor blades 20.
  • the blades 20a of the first turbine stage 16a are turbine blades according to the first embodiment of the present invention.
  • the blades 20b of the second turbine stage 16b disposed behind the first turbine stage 16a are turbine blades according to the second embodiment of the present invention.
  • the ratio of the length L 1 of the lower leading edge portion 13 and the length L 3 of the upper leading edge portion 14 to the length L 2 of the central lower edge portion 12 is greater than that of blades 20a of the first turbine stage 16a.
  • the guide vanes 18 may be provided as turbine blades according to the invention with corresponding, in particular in the direction of the rear turbine stages increasing aspect ratio of the leading edge portions.
  • Advantages of the flow machine according to the invention are that its cooling fluid consumption is low and the material stresses at the leading edges 6 of the blades 20, in particular between the peripheral wall and provided on the outside of the peripheral wall coating system due to temperature gradients are low. Therefore, 15 long maintenance intervals may be provided for the turbomachine, which is associated with correspondingly low downtime and good economic efficiency of the turbomachine 15.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP16202826.0A 2016-12-08 2016-12-08 Aube de turbine avec refroidissement du bord d'attaque Withdrawn EP3333366A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP16202826.0A EP3333366A1 (fr) 2016-12-08 2016-12-08 Aube de turbine avec refroidissement du bord d'attaque

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP16202826.0A EP3333366A1 (fr) 2016-12-08 2016-12-08 Aube de turbine avec refroidissement du bord d'attaque

Publications (1)

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EP3333366A1 true EP3333366A1 (fr) 2018-06-13

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EP16202826.0A Withdrawn EP3333366A1 (fr) 2016-12-08 2016-12-08 Aube de turbine avec refroidissement du bord d'attaque

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11939882B2 (en) 2019-01-17 2024-03-26 Mitsubishi Heavy Industries, Ltd. Turbine rotor blade and gas turbine

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0473991A2 (fr) * 1990-09-04 1992-03-11 Westinghouse Electric Corporation Turbine à gaz avec des aubes refroidies
US20050135932A1 (en) * 2003-05-23 2005-06-23 Dodd Alec G. Turbine blade
EP1760267A2 (fr) * 2005-08-31 2007-03-07 General Electric Company Aube refroidie de turbine
EP1798379A2 (fr) * 2005-12-19 2007-06-20 General Electric Company Aube statorique à refroidissement à contre-courant
US20120117975A1 (en) * 2007-09-06 2012-05-17 Sharma Om P Gas turbine engine systems and related methods involving vane-blade count ratios greater than unity
US8545180B1 (en) * 2011-02-23 2013-10-01 Florida Turbine Technologies, Inc. Turbine blade with showerhead film cooling holes
WO2014137686A1 (fr) * 2013-03-04 2014-09-12 United Technologies Corporation Refroidissement de turbine à gaz à profil aérodynamique haute portance dans une zone de stagnation
EP3091185A1 (fr) * 2015-05-08 2016-11-09 United Technologies Corporation Profil d'aube et procédé associé de fabrication

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0473991A2 (fr) * 1990-09-04 1992-03-11 Westinghouse Electric Corporation Turbine à gaz avec des aubes refroidies
US20050135932A1 (en) * 2003-05-23 2005-06-23 Dodd Alec G. Turbine blade
EP1760267A2 (fr) * 2005-08-31 2007-03-07 General Electric Company Aube refroidie de turbine
EP1798379A2 (fr) * 2005-12-19 2007-06-20 General Electric Company Aube statorique à refroidissement à contre-courant
US20120117975A1 (en) * 2007-09-06 2012-05-17 Sharma Om P Gas turbine engine systems and related methods involving vane-blade count ratios greater than unity
US8545180B1 (en) * 2011-02-23 2013-10-01 Florida Turbine Technologies, Inc. Turbine blade with showerhead film cooling holes
WO2014137686A1 (fr) * 2013-03-04 2014-09-12 United Technologies Corporation Refroidissement de turbine à gaz à profil aérodynamique haute portance dans une zone de stagnation
EP3091185A1 (fr) * 2015-05-08 2016-11-09 United Technologies Corporation Profil d'aube et procédé associé de fabrication

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11939882B2 (en) 2019-01-17 2024-03-26 Mitsubishi Heavy Industries, Ltd. Turbine rotor blade and gas turbine
DE112019004841B4 (de) 2019-01-17 2026-04-23 Mitsubishi Heavy Industries, Ltd. Turbinenlaufschaufel und Gasturbine

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