EP2860360A1 - Système de refroidissement pour le refroidissement d'une aube de turbine - Google Patents

Système de refroidissement pour le refroidissement d'une aube de turbine Download PDF

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Publication number
EP2860360A1
EP2860360A1 EP20130187901 EP13187901A EP2860360A1 EP 2860360 A1 EP2860360 A1 EP 2860360A1 EP 20130187901 EP20130187901 EP 20130187901 EP 13187901 A EP13187901 A EP 13187901A EP 2860360 A1 EP2860360 A1 EP 2860360A1
Authority
EP
European Patent Office
Prior art keywords
cooling
cooling air
turbine
air holes
cooling system
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Withdrawn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
EP20130187901
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German (de)
English (en)
Inventor
Gregoire Etienne Witz
Hans-Peter Bossmann
Matthias Hoebel
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.)
General Electric Technology GmbH
Original Assignee
Alstom Technology 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 Alstom Technology AG filed Critical Alstom Technology AG
Priority to EP20130187901 priority Critical patent/EP2860360A1/fr
Publication of EP2860360A1 publication Critical patent/EP2860360A1/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Withdrawn legal-status Critical Current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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/122—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with erodable or abradable material
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/246—Fastening of diaphragms or stator-rings
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/90—Coating; Surface treatment
    • 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
    • F05D2240/307—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 tip 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 present invention relates to a cooling system for cooling a turbine blade of a gas turbine comprising a rotor with multiple turbine blades and a stator surrounding the rotor.
  • the invention relates in particular to such a cooling system of high-performance gas turbines, in which due to the increased temperatures a stator heat shield is provided, including at least one abradable thermal barrier coating.
  • some cooling of the turbine blades is achieved through convection by providing passages for a flow of cooling air from the compressor internally within the blades, so that heat may be discharged from the metal structure of the blade by the cooling air.
  • a cooling arrangement for a turbine blade shroud in which the turbine blade is provided with several internal cooling passages and cooling air openings such that a cooling air film is directed to the forward face, the rearward face and/or the tip of the turbine blade.
  • Such a cooling arrangement is rather complex with respect to its design and requires a time-consuming and costly manufacturing method for its realization.
  • a very critical part in view of the cooling is the tip region of turbine blades, as it is the hottest location of the turbine blade and because it is difficult to provide sufficient cooling at this specific location of the turbine.
  • This problem comes from the rather complex cooling geometries required to cool the tip of the turbine blade and the heat pick-up of any cooling air that is brought through the interior of the turbine blade to the area of the turbine tip.
  • the possibilities of an internal cooling by means of cooling air passages are rather limited, and it is not simple to create a sufficient film of cooling air on the blade tip in a homogeneous way. Nevertheless, such high-performance gas turbines require a cooling system, which provides a sufficient cooling, in particular of the tip portion of the turbine blades, in order to have an efficient operation performance of the turbine.
  • a cooling system for cooling a turbine blade of a gas turbine comprising a rotor with multiple turbine blades and a stator surrounding the rotor and provided with a heat shield
  • the stator heat shield comprises a base element covered with at least one abradable thermal barrier coating facing to a tip region of the turbine blade
  • the cooling system comprises means for generating a cooling air flow, characterized in that said stator heat shield is provided with a plurality of cooling air holes terminating in the area of said tip region of said turbine blades such that in operation of the turbine a cooling air flow is directed to the tip of respective turbine blades.
  • the cooling system for generating a cooling air flow is integrated within the stator heat shield, and the cooling air flow is directly supplied to the area of the adjacent blade tip of turbine blades.
  • a more efficient cooling in particular of the critical tip region of turbine blades is hereby possible. It is not necessary to provide complex and additional cooling means at the turbine blade itself such that the all-in-all design of the cooling system is comparatively simple.
  • the plurality of cooling air holes in the stator heat shield which may be realized, for example, by drilling, is arranged such that the cooling air flow efficiently cools the tip region of the turbine blades during the operation of the gas turbine.
  • the plurality of cooling air holes is arranged in a specific pattern and at respective areas of the stator heat shield adjacent to the tip regions of the turbine blades.
  • the cooling air holes in the stator heat shield extend through said base element as well as said at least one thermal barrier coating of said stator heat shield. That means, the cooling air holes pass completely through the elements of the stator heat shield.
  • the cooling air holes of this preferred form of realization are therefore always open, and, when connected to the means for generating a cooling air flow, the cooling air flow surrounding the blade tip of the turbine blade is always generated.
  • the cooling air holes are inclined with respect to a tip front line of the turbine blades in an acute angle of in particular approximately 45°. That means, the cooling air holes are inclined to an outer surface of the heat shield and therefore to the front of the turbine blade. When cooling air flows out of the cooling air holes, the cooling air will therefore pass laterally along the complete width of the tip region of the turbine blade, and a more efficient cooling is achieved.
  • the cooling holes have an exit opening at a lateral position with regard to said turbine blades.
  • a cooling air flow is hereby generated from a side portion of the tip region of the turbine blade.
  • the cooling holes are initially closed by at least a part of said thermal barrier coating before an operation of the turbine.
  • the cooling holes in the stator heat shield are, for example, initially closed by means of the abradable thermal barrier coating or a portion of this coating.
  • Only the cooling air holes in the stator heat shield are used for the purpose of cooling the turbine blades, which are in the respective pertinent area of the turbine blades.
  • the initial closure of the cooling air holes can be achieved, for example, by only drilling cooling air holes in the base element of the stator heat shield and by covering the base element with the abradable thermal barrier coating afterwards.
  • the turbine blade will then abrade this abradable thermal barrier coating only in this area of the turbine blades such that an efficient cooling of the tip region of the turbine blades is provided with a reduced need of cooling air.
  • the cooling air holes have a constant diameter and a straight form through said heat shield.
  • the cooling air holes may therefore be realized by a simple drilling processing.
  • a maximum pressure of the cooling air flow at the exit opening close to the blade tip is achieved.
  • the cooling air holes have a larger diameter in the base element as compared to the abradable thermal barrier coating.
  • the thermal barrier coating is formed of a first inner layer and of a second outer layer.
  • the two separate layers of the thermal barrier coating may be realized with the same material or with a different material.
  • the second outer layer of the thermal barrier coating can specifically be used for the initial closing of the cooling air holes, whereas the cooling air holes are only drilled in the first inner layer of the thermal barrier coating and/or a base element of the stator heat shield.
  • the cooling of the tip region of the turbine blade will therefore only come into effect in the respective pertinent areas, in which the turbine blade is running adjacent to the stator heat shields.
  • FIG. 1 A first example of realization of the cooling system for turbine blades of the present invention is shown in respective schematic cross-section views in Figs. 1 and 2 .
  • the cooling system is integrated within a stator heat shield 2 having in this embodiment a base element 3 and an abradable thermal barrier coating 4, in which a plurality of cooling air holes 5 is formed.
  • the cooling air holes 5 in this embodiment of the invention are always open, with an exit opening 6 on the upper side facing to the critical tip region of a turbine blade (not shown in Fig. 1 ).
  • the cooling system is furthermore provided with means for generating a cooling air flow, which is arranged outside of the stator heat shield 2 shown in Figs. 1 and 2 .
  • the cooling air holes 5 in this first embodiment are inclined at approximately 45° with respect to the upper surface of the stator heat shield 2 and therefore also with respect to the front line of a turbine blade 1 (cf. Fig. 2 ).
  • An exit opening 6 of the cooling air holes 5 is arranged in a lateral position of the tip region of the turbine blade 1 when the turbine is operated.
  • the cooling air flow according to the arrow in Fig. 2 will therefore pass from one lateral side of the tip region of the turbine blade 1 to the other side and will provide an efficient cooling of in particular the tip region of this turbine blade 1. This cooling effect is also supported by the inclined form of the plurality of cooling air holes 5.
  • the cooling air holes 5 go completely through the material of both elements of the stator heat shield 2, namely the base element 3 and the abradable thermal barrier coating 4, which is in this field also denominated as a TBC.
  • the plurality of cooling air holes 5 is, for example, realized by means of a drilling processing and with appropriate diameters for providing a sufficient cooling air flow to the tip region of the turbine blade 1.
  • the base element 3 of the stator heat shield 2 is realized through an appropriate metal or metal alloy, whereas the TBC or thermal barrier coating 4 consists preferably of a ceramic material, such as an Yttria-stabilized zirconia (YSZ), which provides the desired low heat-conductivity and the protection of the outer stator elements of the gas turbine.
  • YSZ Yttria-stabilized zirconia
  • a bond coat 7 is provided for connecting the two elements to one another.
  • the cooling air hole 5 being provided at a lateral position of the tip region of the turbine blade 1 remains open, with its exit opening 6 in such a form that a homogeneous air flow of cooling air is applied to the tip region of the turbine blade 1.
  • the cooling system may easily be integrated within a given turbine design. Independent of the amount of removal of material in the abradable thermal barrier coating 4, still a sufficient cooling air flow is provided through the exit openings 6 at any circumstance by means of the cooling system of the invention.
  • the multiple cooling air holes 5 in the stator heat shield 2 may preferably be arranged in a specified pattern and with respect to the position of the turbine blades 1 running within the gas turbine in operation. For example, the multiple cooling air holes 5 are arranged in a concentrated form in the areas where the turbine blades 1 will be mounted, facing to the respective stator heat shields 2.
  • FIG. 3 and 4 of the attached drawings show a second embodiment of the cooling system for cooling a turbine blade of the invention.
  • the cooling air holes 5 are initially closed by means of an outer layer 42 of the thermal barrier coating 4.
  • the thermal barrier coating 4 is here realized in a two-layer form with a first inner layer 41 and a second outer layer 42, which can be made of the same ceramic material or of different types of ceramic materials.
  • the multiple cooling air holes 5 are closed by means of the outer layer 42 of the thermal barrier coating 4.
  • the cooling blade will be run as close as possible to the heat shields 2 in order to achieve a sufficient cooling effect.
  • the cooling air holes 5 of this form of realization are designed and manufactured such that they are closed by default and will only be opened if the gap between the stator heat shields 2 and the turbine blade 1 is small enough.
  • this is achieved by means of the second outer layer 42 of the thermal barrier coating 4, which initially closes the cooling air holes 5 formed in the base element 3, the bond coat 7 and the inner layer 41 of the thermal barrier coating 4, as can be seen when comparing Fig. 3 and Fig. 4 of the attached drawings.
  • the outer layer 42 will be abraded and consequently the cooling air holes 5 will be opened, so that the efficient cooling by means of the cooling air flow supplied to the tip region of the turbine blade 1 is achieved.
  • This form of realization allows using only the respective cooling air holes 5 of the plurality of cooling air holes 5, which will bring an effective cooling of the blade tip, and the requirement of cooling air pressure and the amount of cooling air are reduced.
  • the cooling air holes 5 are arranged in a lateral position with regard to the turbine blade 1 and are inclined at approximately 45°with respect to the front line of the tip of the turbine blade.
  • the cooling air holes 5 may be not inclined or inclined at another degree and may be arranged at a different position or may be arranged over the whole extension of the respective stator heat shields 2 such that a kind of regular pattern of cooling air holes 5 is provided all around the outer turbine parts.
  • the thermal barrier coating 4 is here also realized on the basis of a ceramic material, which can be the same for the outer layer 42 and the inner layer, but which can also be different in any of these two layers.
  • the bond coat 7 provides a secure fitting of the abradable thermal barrier coating 4 to the base element 3, which is preferably made from a metal alloy.
  • FIG. 5 of the attached drawings shows a third example of realization of a cooling system for cooling a turbine blade integrated within a stator heat shield 2 in a schematic cross-section.
  • This third embodiment is similar to the above-described second embodiment of the invention and comprises a thermal barrier coating 4 made of a first inner layer 41 and a second outer layer 42, which initially closes the plurality of cooling air holes 5.
  • the outer layer 42 of the thermal barrier coating 4 will be abraded and the exit opening 6 will afterwards be open, so that a cooling air flow is supplied through the cooling air holes 5, which are in a position adjacent to the tip region of the turbine blade.
  • the cooling system is furthermore provided with an emergency cooling element in the form of an emergency cooling hole 8 in the base element 3 and the bond coat 7, having a larger diameter as compared to the cooling diameter of the cooling air holes 5 in the thermal barrier coating 4.
  • an emergency cooling is achieved for the case that the thermal barrier coating falls off.
  • the larger diameter emergency cooling hole 8 will automatically be opened, and the amount of cooling air flow will thereby be increased. Any deterioration of the turbine blades 1 is hereby avoided also in cases in which the complete thermal barrier coating 4 spalls off and until the stator heat shield 2 is exchanged during the upcoming maintenance of the gas turbine.
  • the diameter of the emergency cooling hole 8 is approximately twice the diameter of the cooling air hole 5. Also in this example of realization, the cooling air holes 5 extend in an inclined angle at approximately 45° compared to the upper surface of the stator heat shield 2, and therefore to a front line of the turbine blade 1.
  • the present invention is not limited to this exemplary form of realization, and the cooling air hole 5 may have a different extension and a different position than shown in Fig. 5 of the drawings. It is to be noted that in the drawings only a single cooling air hole 5 is shown, but that the cooling system of the invention has a plurality of cooling air holes 5 in any of the stator heat shields 2.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP20130187901 2013-10-09 2013-10-09 Système de refroidissement pour le refroidissement d'une aube de turbine Withdrawn EP2860360A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP20130187901 EP2860360A1 (fr) 2013-10-09 2013-10-09 Système de refroidissement pour le refroidissement d'une aube de turbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP20130187901 EP2860360A1 (fr) 2013-10-09 2013-10-09 Système de refroidissement pour le refroidissement d'une aube de turbine

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EP2860360A1 true EP2860360A1 (fr) 2015-04-15

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114687810A (zh) * 2022-03-30 2022-07-01 沈阳航空航天大学 一种带有非均匀预扩张气模孔的涡轮叶片

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5375973A (en) * 1992-12-23 1994-12-27 United Technologies Corporation Turbine blade outer air seal with optimized cooling
EP1669545A1 (fr) * 2004-12-08 2006-06-14 Siemens Aktiengesellschaft Système de couches, utilisation et procédé pour la fabrication d'un système multicouche
GB2434842A (en) 2006-02-02 2007-08-08 Rolls Royce Plc Cooling arrangement for a turbine blade shroud
EP2613015A1 (fr) * 2012-01-04 2013-07-10 United Technologies Corporation Joint d'air externe de lames hybrides pour moteur à turbine à gaz

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5375973A (en) * 1992-12-23 1994-12-27 United Technologies Corporation Turbine blade outer air seal with optimized cooling
EP1669545A1 (fr) * 2004-12-08 2006-06-14 Siemens Aktiengesellschaft Système de couches, utilisation et procédé pour la fabrication d'un système multicouche
GB2434842A (en) 2006-02-02 2007-08-08 Rolls Royce Plc Cooling arrangement for a turbine blade shroud
EP2613015A1 (fr) * 2012-01-04 2013-07-10 United Technologies Corporation Joint d'air externe de lames hybrides pour moteur à turbine à gaz

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114687810A (zh) * 2022-03-30 2022-07-01 沈阳航空航天大学 一种带有非均匀预扩张气模孔的涡轮叶片
CN114687810B (zh) * 2022-03-30 2023-08-18 沈阳航空航天大学 一种带有非均匀预扩张气模孔的涡轮叶片

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