US20100028128A1 - Component with diagonally extending recesses in the surface and process for operating a turbine - Google Patents

Component with diagonally extending recesses in the surface and process for operating a turbine Download PDF

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Publication number
US20100028128A1
US20100028128A1 US12/521,912 US52191207A US2010028128A1 US 20100028128 A1 US20100028128 A1 US 20100028128A1 US 52191207 A US52191207 A US 52191207A US 2010028128 A1 US2010028128 A1 US 2010028128A1
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United States
Prior art keywords
component
recess
turbine
layer
angle
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Abandoned
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US12/521,912
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English (en)
Inventor
Marcus Fischer
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Siemens AG
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Siemens AG
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Assigned to SIEMENS AKTIENGESELLSCHAFT reassignment SIEMENS AKTIENGESELLSCHAFT ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: FISCHER, MARCUS
Publication of US20100028128A1 publication Critical patent/US20100028128A1/en
Abandoned 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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/288—Protective coatings for blades
    • 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
    • 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
    • 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
    • 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
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002—Wall structures
    • 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
    • F05D2230/10—Manufacture by removing material
    • 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
    • F05D2230/10—Manufacture by removing material
    • F05D2230/13—Manufacture by removing material using lasers
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/19—Two-dimensional machined; miscellaneous
    • F05D2250/192—Two-dimensional machined; miscellaneous bevelled
    • 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
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23M—CASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M2900/00—Special features of, or arrangements for combustion chambers
    • F23M2900/05003—Details of manufacturing specially adapted for combustion chambers
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23M—CASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M2900/00—Special features of, or arrangements for combustion chambers
    • F23M2900/05004—Special materials for walls or lining
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00018—Manufacturing combustion chamber liners or subparts
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the invention relates to a component with diagonally extending recesses on the surface, and to a process for operating a turbine.
  • turbomachines for example gas turbines
  • U.S. Pat. No. 6,703,137 B2 discloses recesses which extend perpendicularly with respect to the surface in a turbine blade or vane and have an outer thermal barrier coating on a bonding layer.
  • the object is achieved by means of a component having recesses extending diagonally with respect to the direction of flow and by means of a process for operating a turbine comprising such components.
  • the recesses preferably extend only in one layer, i.e. are preferably present within one layer.
  • the recesses are then present only in the outermost layer.
  • the outermost layer is a ceramic layer in which the recesses are present.
  • FIGS. 1 , 2 , 3 , 4 , 5 , 6 show exemplary embodiments
  • FIG. 7 shows a gas turbine
  • FIG. 8 shows a perspective view of a turbine blade or vane
  • FIG. 9 shows a perspective view of a combustion chamber.
  • FIG. 1 shows a component 1 in cross section.
  • the component 1 is a turbine rotor blade or guide vane 120 , 130 ( FIGS. 7 , 8 ) or a combustion chamber element 155 ( FIG. 9 ).
  • the invention is explained merely by way of example with reference to turbine blades or vanes 120 , 130 of gas turbines 100 , but may be used for any desired component which has a medium flowing over or around it, that is to say also in gas turbines for aircraft or in steam turbines or compressors.
  • the component 1 , 120 , 130 , 155 comprises a substrate 4 which, particularly in the case of high-temperature applications, such as in gas turbines, consists of a nickel-base or cobalt-base alloy. Iron-base superalloys are also used in the case of components of steam turbines.
  • a bonding layer 7 which preferably consists of an alloy of the MCrAlX type and to which an outer ceramic thermal barrier coating 10 has been applied is preferably present on the substrate 4 .
  • the recesses 19 start from a surface 16 of the component 120 , 130 , 155 and may be present in a solid component 120 , 130 , 155 (component comprising only a substrate 4 ) or in layers 7 , 10 ( FIGS. 1 , 2 , 3 ).
  • the recesses 19 may also extend through one or more layers 7 , 10 (not shown).
  • the component 1 has a medium flowing over or around it in the direction of flow 13 .
  • the recesses 19 preferably extend diagonally in the direction of flow 13 ( FIGS. 1 , 2 , 3 ). Equally, however, they may also extend diagonally counter to the direction of flow 13 .
  • the recess 19 represents a blind hole or always has a base 28 . It is therefore not used as a film-cooling hole.
  • the recesses 19 have a longitudinal direction 22 which extends within the recess 19 from the base 28 of the recess 19 as far as the surface 16 of the component and which extends at an angle ⁇ diagonally with respect to the direction of flow 13 or with respect to the surface 16 ( FIG. 2 ).
  • the penetration depth d of a recess 19 extends perpendicularly with respect to the surface 16 of the component 120 , 130 and may be dimensioned in values relative to the layer thickness s of the individual layers 7 , 10 and to the overall layer thickness.
  • a penetration depth d of the recess into one layer 10 or into the layers 7 , 10 is preferably defined in values relative to the layer thickness s of the outermost layer.
  • the penetration depth extends perpendicularly with respect to the outer surface 16 . It is preferably 10%-120% of the layer thickness s, i.e. in the case of 120%, it extends into the substrate 4 or an underlying and/or underlying layer 7 and into the substrate 4 via the outer layer 10 .
  • the penetration depth d is preferably between 10% and 90% of the layer thickness s of the outermost layer 10 , i.e. it is arranged only within the outermost layer 10 . Particular preference is given to using penetration depths of 50%-80% of the layer thickness of the outermost layer 10 ( FIG. 3 ).
  • the outermost layer 10 preferably has a thickness of from 1-2 mm and, for the recess 19 , has a penetration depth d of 1 mm.
  • the recesses 19 preferably have the same penetration depth d ( FIG. 2 ) from the surface 16 of the component.
  • a penetration depth d is preferably from 10% to 120% of the layer thickness s.
  • the angle ⁇ is not 90° ( ⁇ 90°, i.e. ⁇ >90° or ⁇ 90°).
  • the difference from 90° is selected such that it is outside a tolerance range given for the production of perpendicularly extending recesses, as is known from U.S. Pat. No. 6,703,137 B2.
  • the angle ⁇ is preferably ⁇ 80° or >100°.
  • the angle ⁇ is preferably between 20° and 80°.
  • the recess 19 is preferably of elongated form in the plane of the surface 16 of the component 1 , 120 , 130 , 155 , i.e. the extent 1 in the plane of the surface 16 is preferably at least ten times greater than the penetration depth d ( FIGS. 3 , 4 ).
  • the recess 19 may also be bent ( FIG. 4 ).
  • the recess 19 may also surround a component 120 , 130 , 155 , i.e. may surround the main blade or vane part 406 in the case of a turbine blade or vane 120 , 130 .
  • the angle ⁇ is defined by the direction of flow 13 and a lateral direction 25 , which represents an edge of the recess 19 level with the surface 16 .
  • FIG. 2 shows a further exemplary embodiment.
  • film-cooling holes 418 , 419 are present in the substrate 4 and/or also in the layers 7 , 10 .
  • the film-cooling holes 418 extend from a cavity of the component preferably until they are level with the penetration depth d of the recesses 19 .
  • the film-cooling holes 418 may also extend as far as the surface 16 (not shown), where recesses 19 are located or else where no recesses 19 are located.
  • concealed film-cooling holes 419 are present, and these are present underneath the thermal barrier coating 10 and underneath the bonding layer 7 .
  • the film-cooling hole 418 may be as wide as the recess at the level of the plane 20 , and may be thinner or else wider than the extent of the recess 19 in the direction of flow 13 .
  • the recess may have any desired cross section.
  • the recesses are in the form of a parallelogram. In cross section parallel to the surface 16 , the edges of the recess 19 have edges extending in parallel in cross section.
  • the recess 19 may also be wider in the region of the surface 16 than in the region of the base 28 of the recess 19 ( FIG. 4 ).
  • the width of the recess 19 ′′ on the surface 16 may also be smaller than on the base 28 level with the penetration depth d.
  • the longitudinal direction 22 is always formed by a line which extends in the plane of a side wall 23 , 26 and has the smallest distance between the base 28 of the recess 19 and the surface 16 of the recess 19 .
  • the recesses 19 may be introduced in different ways. In the case of metallic layers 7 or metallic substrates 4 , this can be done using a known mechanical method. In the case of ceramics and under ceramic layers 10 , this is preferably done by means of a laser, as is also explained in U.S. Pat. No. 6,703,137 B2, or by means of electron irradiation.
  • the recesses 19 have the effect that the air molecules do not move and thus form a type of open porosity, in which case the air remains in the recesses or slots 19 as a result of the diagonal position in the direction of flow 13 .
  • FIG. 6 shows by way of example a partial longitudinal section through a gas turbine 100 .
  • the gas turbine 100 has a rotor 103 which is mounted such that it can rotate about an axis of rotation 102 , has a shaft 101 , and is also referred to as the turbine rotor.
  • the annular combustion chamber 110 is in communication with a for example annular hot gas duct 111 .
  • a for example annular hot gas duct 111 There, by way of example, four successive turbine stages 112 form the turbine 108 .
  • Each turbine stage 112 is formed for example from two blade rings. As seen in the direction of flow of a working medium 113 , a guide vane row 115 is followed in the hot gas duct 111 by a row 125 formed from rotor blades 120 .
  • the guide vanes 130 are secured to an inner casing 138 of a stator 143 , whereas the rotor blades 120 belonging to a row 125 are arranged on the rotor 103 , for example by means of a turbine disk 133 .
  • a generator (not shown) is coupled to the rotor 103 .
  • air 135 is drawn in through the intake casing 104 and compressed by the compressor 105 .
  • the compressed air provided at the turbine end of the compressor 105 is passed to the burners 107 , where it is mixed with a fuel.
  • the mixture is then burnt in the combustion chamber 110 , forming the working medium 113 .
  • the working medium 113 flows along the hot gas duct 111 past the guide vanes 130 and the rotor blades 120 .
  • the working medium 113 is expanded at the rotor blades 120 , transferring its momentum, so that the rotor blades 120 drive the rotor 103 and the latter in turn drives the generator coupled to it.
  • Substrates of the components may likewise have a directional structure, i.e. they are in single-crystal form (SX structure) or have only longitudinally oriented grains (DS structure).
  • SX structure single-crystal form
  • DS structure longitudinally oriented grains
  • iron-base, nickel-base or cobalt-base superalloys are used as material for the components, in particular for the turbine blade or vane 120 , 130 and components of the combustion chamber 110 .
  • the guide vane 130 has a guide vane root (not shown here) facing the inner casing 138 of the turbine 108 and a guide vane head at the opposite end from the guide vane root.
  • the guide vane head faces the rotor 103 and is fixed to a securing ring 140 of the stator 143 .
  • FIG. 7 shows a perspective view of a rotor blade 120 or guide vane 130 of a turbomachine, which extends along a longitudinal axis 121 .
  • the turbomachine may be a gas turbine of an aircraft or of a power plant for generating electricity, a steam turbine or a compressor.
  • the blade or vane 120 , 130 has, in succession along the longitudinal axis 121 , a securing region 400 , an adjoining blade or vane platform 403 , a main blade or vane part 406 and a blade or vane tip 415 .
  • the vane 130 may have a further platform (not shown) at its vane tip 415 .
  • a blade or vane root 183 which is used to secure the rotor blades 120 , 130 to a shaft or a disk (not shown), is formed in the securing region 400 .
  • the blade or vane root 183 is designed, for example, in hammerhead form. Other configurations, such as a fir-tree or dovetail root, are possible.
  • the blade or vane 120 , 130 has a leading edge 409 and a trailing edge 412 for a medium which flows past the main blade or vane part 406 .
  • the blade or vane 120 , 130 may in this case be produced by a casting process, also by means of directional solidification, by a forging process, by a milling process or combinations thereof.
  • Workpieces with a single-crystal structure or structures are used as components for machines which, in operation, are exposed to high mechanical, thermal and/or chemical stresses.
  • Single-crystal workpieces of this type are produced, for example, by directional solidification from the melt. This involves casting processes in which the liquid metallic alloy solidifies to form the single-crystal structure, i.e. the single-crystal workpiece, or solidifies directionally.
  • dendritic crystals are oriented along the direction of heat flow and form either a columnar crystalline grain structure (i.e. grains which run over the entire length of the workpiece and are referred to here, in accordance with the language customarily used, as directionally solidified) or a single-crystal structure, i.e. the entire workpiece consists of one single crystal.
  • a transition to globular (polycrystalline) solidification needs to be avoided, since non-directional growth inevitably forms transverse and longitudinal grain boundaries, which negate the favorable properties of the directionally solidified or single-crystal component.
  • directionally solidified microstructures refers in general terms to directionally solidified microstructures, this is to be understood as meaning both single crystals, which do not have any grain boundaries or at most have small-angle grain boundaries, and columnar crystal structures, which do have grain boundaries running in the longitudinal direction but do not have any transverse grain boundaries.
  • This second form of crystalline structures is also described as directionally solidified microstructures (directionally solidified structures).
  • the blades or vanes 120 , 130 may likewise have coatings protecting against corrosion or oxidation, e.g. (MCrAlX; M is at least one element selected from the group consisting of iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon and/or at least one of the rare earth elements, or hafnium (Hf)). Alloys of this type are known from EP 0 486 489 B1, EP0 786 017 B1, EP0 412 397 B1 or EP 1 306 454 A1, which are intended to form part of this disclosure with regard to the chemical composition of the alloy.
  • MrAlX M is at least one element selected from the group consisting of iron (Fe), cobalt (Co), nickel (Ni)
  • X is an active element and stands for yttrium (Y) and/or silicon and/or at least one of the rare earth elements, or hafnium (Hf)). Allo
  • the density is preferably 95% of the theoretical density.
  • thermal barrier coating which is preferably the outermost layer and consists for example of ZrO 2 , Y 2 O 3 —ZrO 2 , i.e. unstabilized, partially stabilized or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide, to be present on the MCrAlX.
  • the thermal barrier coating covers the entire MCrAlX layer.
  • Columnar grains are produced in the thermal barrier coating by means of suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).
  • suitable coating processes such as for example electron beam physical vapor deposition (EB-PVD).
  • the thermal barrier coating may have grains that are porous and/or include micro-cracks or macro-cracks in order to improve the resistance to thermal shocks. Therefore, the thermal barrier coating is preferably more porous than the MCrAlX layer.
  • the blade or vane 120 , 130 may be hollow or solid in form. If the blade or vane 120 , 130 is to be cooled, it is hollow and may also have film-cooling holes 418 (indicated by dashed lines).
  • FIG. 8 shows a combustion chamber 110 of the gas turbine 100 .
  • the combustion chamber 110 is configured, for example, as what is known as an annular combustion chamber, in which a multiplicity of burners 107 , which generate flames 156 and are arranged circumferentially around an axis of rotation 102 , open out into a common combustion chamber space 154 .
  • the combustion chamber 110 overall is of annular configuration positioned around the axis of rotation 102 .
  • the combustion chamber 110 is designed for a relatively high temperature of the working medium M of approximately 1000° C. to 1600° C.
  • the combustion chamber wall 153 is provided, on its side which faces the working medium M, with an inner lining formed from heat shield elements 155 .
  • a cooling system may also be provided for the heat shield elements 155 and/or their holding elements, on account of the high temperatures in the interior of the combustion chamber 110 .
  • the heat shield elements 155 are then for example hollow and may also have cooling holes (not shown) which open out into the combustion chamber space 154 .
  • each heat shield element 155 made from an alloy is equipped with a particularly heat-resistant protective layer (MCrAlX layer and/or ceramic coating) or is made from material that is able to withstand high temperatures (solid ceramic bricks).
  • M is at least one element selected from the group consisting of iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon and/or at least one of the rare earth elements, or hafnium (Hf). Alloys of this type are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1, which are intended to form part of this disclosure with regard to the chemical composition of the alloy.
  • a for example ceramic thermal barrier coating consisting for example of ZrO 2 , Y 2 O 3 —ZrO 2 , i.e. unstabilized, partially stabilized or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide, may also be present on the MCrAlX.
  • Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).
  • EB-PVD electron beam physical vapor deposition
  • the thermal barrier coating may have grains that are porous and/or include micro-cracks or macro-cracks in order to improve the resistance to thermal shocks.
  • Refurbishment means that after they have been used, protective layers may have to be removed from turbine blades or vanes 120 , 130 , heat shield elements 155 (e.g. by sand-blasting). Then, the corrosion and/or oxidation layers and products are removed. If appropriate, cracks in the turbine blade or vane 120 , 130 or the heat shield element 155 are also repaired. This is followed by recoating of the turbine blades or vanes 120 , 130 , heat shield elements 155 , after which the turbine blades or vanes 120 , 130 or the heat shield elements 155 can be reused.
  • protective layers may have to be removed from turbine blades or vanes 120 , 130 , heat shield elements 155 (e.g. by sand-blasting). Then, the corrosion and/or oxidation layers and products are removed. If appropriate, cracks in the turbine blade or vane 120 , 130 or the heat shield element 155 are also repaired. This is followed by recoating of the turbine blades or vanes 120 , 130 , heat
US12/521,912 2007-01-05 2007-10-29 Component with diagonally extending recesses in the surface and process for operating a turbine Abandoned US20100028128A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
EP07000189.6 2007-01-05
EP07000189A EP1942250A1 (de) 2007-01-05 2007-01-05 Bauteil mit schräg verlaufenden Vertiefungen in der Oberfläche und Verfahren zum Betreiben einer Turbine
PCT/EP2007/061609 WO2008080655A1 (de) 2007-01-05 2007-10-29 Bauteil mit schräg verlaufenden vertiefungen in der oberfläche und verfahren zum betreiben einer turbine

Publications (1)

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US20100028128A1 true US20100028128A1 (en) 2010-02-04

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US12/521,912 Abandoned US20100028128A1 (en) 2007-01-05 2007-10-29 Component with diagonally extending recesses in the surface and process for operating a turbine

Country Status (8)

Country Link
US (1) US20100028128A1 (de)
EP (2) EP1942250A1 (de)
JP (1) JP2010514984A (de)
KR (1) KR20090107520A (de)
CN (1) CN101573510B (de)
AT (1) ATE506524T1 (de)
DE (1) DE502007007024D1 (de)
WO (1) WO2008080655A1 (de)

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20120167573A1 (en) * 2010-12-30 2012-07-05 United Technologies Corporation Thermal Barrier Coatings and Methods of Application
US20140349065A1 (en) * 2011-11-24 2014-11-27 Siemens Aktiengesellschaft Modified interface around a hole
US20160201912A1 (en) * 2013-09-11 2016-07-14 Siemens Aktiengesellschaft Wedge-shaped ceramic heat shield of a gas turbine combustion chamber
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DE502007007024D1 (de) 2011-06-01
ATE506524T1 (de) 2011-05-15
EP2097616A1 (de) 2009-09-09
CN101573510B (zh) 2012-11-14
EP1942250A1 (de) 2008-07-09
KR20090107520A (ko) 2009-10-13
EP2097616B1 (de) 2011-04-20
CN101573510A (zh) 2009-11-04
WO2008080655A1 (de) 2008-07-10

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