US20140255652A1 - Surface having specially formed recesses and component - Google Patents
Surface having specially formed recesses and component Download PDFInfo
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- US20140255652A1 US20140255652A1 US14/352,233 US201214352233A US2014255652A1 US 20140255652 A1 US20140255652 A1 US 20140255652A1 US 201214352233 A US201214352233 A US 201214352233A US 2014255652 A1 US2014255652 A1 US 2014255652A1
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- component
- recess
- longitudinal direction
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- recesses
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- 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
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
- F01D5/284—Selection of ceramic materials
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- 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
- F23M5/00—Casings; Linings; Walls
- F23M5/02—Casings; Linings; Walls characterised by the shape of the bricks or blocks used
-
- 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
- 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
- F23R3/007—Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
-
- 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
- F05D2250/184—Two-dimensional patterned sinusoidal
-
- 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
- F05D2250/185—Two-dimensional patterned serpentine-like
-
- 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
- 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
- F05D2300/00—Materials; Properties thereof
- F05D2300/20—Oxide or non-oxide ceramics
-
- 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
- F05D2300/00—Materials; Properties thereof
- F05D2300/20—Oxide or non-oxide ceramics
- F05D2300/21—Oxide ceramics
-
- 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
- F05D2300/00—Materials; Properties thereof
- F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
- F05D2300/611—Coating
-
- 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
- F23R2900/00019—Repairing or maintaining combustion chamber liners or subparts
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- 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
-
- 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
- Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
- Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
- Y10T428/00—Stock material or miscellaneous articles
- Y10T428/24—Structurally defined web or sheet [e.g., overall dimension, etc.]
- Y10T428/24479—Structurally defined web or sheet [e.g., overall dimension, etc.] including variation in thickness
Definitions
- the invention relates to the special configuration of elongate recesses within a surface and to a component.
- ceramic materials Compared to metals, ceramic materials have a relatively low ductility, and cracks can arise as a result of stresses. Particularly in the case of components coated with ceramic, such as gas turbine components, instances of spalling can occur in the ceramic layer. This takes place primarily in regions of cooling air bore outlets in the form of what are known as shaped holes at the surface of the ceramic layer. High thermal stresses, which lead to the premature onset of cracking in the ceramic layer and then to spalling, arise here during the operation of a hollow cast and cooled turbine blade or vane. Therefore, recesses are often introduced, as in WO 2009/126194 A1.
- the object is achieved by a surface and by a component as claimed.
- FIG. 1 shows a component having a recess
- FIGS. 2 , 3 show a special form of the recess
- FIG. 4 shows a turbine blade or vane
- FIG. 5 shows a combustion chamber
- FIG. 6 shows a gas turbine
- FIG. 7 shows a list of superalloys.
- the invention relates to surfaces of solid components, layers, in particular ceramic surfaces, but also metals, which can have a certain brittleness, such as NiCoCrAlY alloys in a certain temperature range.
- FIG. 1 shows a merely exemplary high-temperature component 120 , 130 , 155 ( FIGS. 4 , 5 ) having a surface 19 around which medium flows in a direction of overflow 10 .
- Elongate recesses 4 are not arranged parallel with respect to a direction of overflow 10 over the surface 19 , but rather at an angle which differs considerably from 0°, preferably at an angle of 90°+/ ⁇ 20°, with respect to the direction of overflow 10 .
- the recesses 4 have a longitudinal direction 11 .
- the recesses 4 are made in particular where the greatest thermomechanical loads are to be expected.
- this is the region around the leading edge 409 ( FIG. 4 ) and the region around the cooling air bores.
- FIG. 2 shows a cross section through a recess 4 , which has a front edge 25 and a rear edge 28 at the surface 19 .
- the widened portion 26 is preferably formed at the rear edge 28 , i.e. the flow-side end.
- the recess 4 in a substrate or layer, in particular a ceramic layer 13 widens proceeding from the base 16 as far as the surface 19 , preferably constantly ( FIG. 2 ) or only above a certain height 22 ( FIG. 3 ), i.e. up to the height 22 , the cross section transverse to the longitudinal direction 11 of the recess 4 is constant.
- the recess 4 preferably does not extend over the entire thickness of the layer 13 (not shown).
- the front edge 25 preferably runs perpendicular to the surface of a substrate of the layer 13 (or to the surface 19 ).
- the turbine blade or vane 120 , 130 preferably has a substrate made of a nickel-based or cobalt-based superalloy, in particular a material as shown in FIG. 7 .
- FIG. 4 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 rare earth element, 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.
- the density is preferably 95% of the theoretical density.
- the layer preferably has a composition Co—30Ni—28Cr—8Al—0.6Y—0.7Si or Co—28Ni—24Cr—10Al—0.6Y.
- nickel-based protective layers such as Ni—10Cr—12Al—0.6Y—3Re or Ni—12Co—21Cr—11Al—0.4Y—2Re or Ni—25Co—17Cr—10Al—0.4Y—1.5Re.
- 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 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 include grains that are porous or have micro-cracks or macro-cracks, in order to improve the resistance to thermal shocks.
- the thermal barrier coating is therefore preferably more porous than the MCrAlX layer.
- Refurbishment means that after they have been used, protective layers may have to be removed from components 120 , 130 (e.g. by sand-blasting). Then, the corrosion and/or oxidation layers and products are removed. If appropriate, cracks in the component 120 , 130 are also repaired. This is followed by recoating of the component 120 , 130 , after which the component 120 , 130 can be reused.
- 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. 5 shows a combustion chamber 110 of a gas turbine.
- 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 .
- 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 rare earth element 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.
- 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 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 heat shield element 155 are also repaired. This is followed by recoating of the heat shield elements 155 , after which the heat shield elements 155 can be reused.
- 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 .
- FIG. 6 shows, by way of example, a partial longitudinal section through a gas turbine 100 .
- the gas turbine 100 has a rotor 103 with a shaft 101 which is mounted such that it can rotate about an axis of rotation 102 and is also referred to as the turbine rotor.
- the annular combustion chamber 110 is in communication with a, for example, annular hot-gas passage 111 , where, by way of example, four successive turbine stages 112 form the turbine 108 .
- Each turbine stage 112 is formed, for example, from two blade or vane rings. As seen in the direction of flow of a working medium 113 , in the hot-gas passage 111 a row of guide vanes 115 is followed by a row 125 formed from rotor blades 120 .
- the guide vanes 130 are secured to an inner housing 138 of a stator 143 , whereas the rotor blades 120 of a row 125 are fitted to the rotor 103 for example by means of a turbine disk 133 .
- a generator (not shown) is coupled to the rotor 103 .
- the compressor 105 While the gas turbine 100 is operating, the compressor 105 sucks in air 135 through the intake housing 104 and compresses it. The compressed air provided at the turbine-side end of the compressor 105 is passed to the burners 107 , where it is mixed with a fuel. The mix is then burnt in the combustion chamber 110 , forming the working medium 113 . From there, the working medium 113 flows along the hot-gas passage 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-based, nickel-based or cobalt-based 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 blades or vanes 120 , 130 may likewise have coatings protecting against corrosion (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, scandium (Sc) and/or at least one rare earth element, or hafnium). 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.
- a 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 guide vane 130 has a guide vane root (not shown here), which faces the inner housing 138 of the turbine 108 , and a guide vane head which is 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 .
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Combustion & Propulsion (AREA)
- Materials Engineering (AREA)
- Ceramic Engineering (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A surface having specially formed recesses and component is provided. The component has a layer with a surface, having elongate recesses, which have a longitudinal direction, wherein the recesses are arranged at an angle which differs considerably from 0°, in particular 90°+/−20°, with respect to a direction of overflow over the surface, and are at least partially widened transversely to the longitudinal direction thereof in the region of the surface with respect to the base of the recess, in which the recess is widened proceeding from the base only above a certain height within the recess, and in particular the recess has a constant width before it, in which the recess does not extend through the entire thickness of the layer.
Description
- This application is the US National Stage of International Application No. PCT/EP2012/066062 filed Aug. 17, 2012, and claims the benefit thereof. The International Application claims the benefit of European Application No. EP11186464 filed Oct. 25, 2011. All of the applications are incorporated by reference herein in their entirety.
- The invention relates to the special configuration of elongate recesses within a surface and to a component.
- Compared to metals, ceramic materials have a relatively low ductility, and cracks can arise as a result of stresses. Particularly in the case of components coated with ceramic, such as gas turbine components, instances of spalling can occur in the ceramic layer. This takes place primarily in regions of cooling air bore outlets in the form of what are known as shaped holes at the surface of the ceramic layer. High thermal stresses, which lead to the premature onset of cracking in the ceramic layer and then to spalling, arise here during the operation of a hollow cast and cooled turbine blade or vane. Therefore, recesses are often introduced, as in WO 2009/126194 A1.
- It is therefore an object of the invention to solve this problem. The object is achieved by a surface and by a component as claimed.
- The dependent claims list further advantageous measures which can be combined with one another, as desired, in order to achieve further advantages.
-
FIG. 1 shows a component having a recess, -
FIGS. 2 , 3 show a special form of the recess, -
FIG. 4 shows a turbine blade or vane, -
FIG. 5 shows a combustion chamber, -
FIG. 6 shows a gas turbine, and -
FIG. 7 shows a list of superalloys. - The description and the figures represent merely exemplary embodiments of the invention.
- In general terms, the invention relates to surfaces of solid components, layers, in particular ceramic surfaces, but also metals, which can have a certain brittleness, such as NiCoCrAlY alloys in a certain temperature range.
-
FIG. 1 shows a merely exemplary high-temperature component FIGS. 4 , 5) having asurface 19 around which medium flows in a direction ofoverflow 10. -
Elongate recesses 4 are not arranged parallel with respect to a direction ofoverflow 10 over thesurface 19, but rather at an angle which differs considerably from 0°, preferably at an angle of 90°+/−20°, with respect to the direction ofoverflow 10. - The
recesses 4 have alongitudinal direction 11. - The
recesses 4 are made in particular where the greatest thermomechanical loads are to be expected. For the gas turbine blade orvane FIG. 4 ) and the region around the cooling air bores. -
FIG. 2 shows a cross section through arecess 4, which has afront edge 25 and arear edge 28 at thesurface 19. - The widened
portion 26 is preferably formed at therear edge 28, i.e. the flow-side end. - The
recess 4 in a substrate or layer, in particular aceramic layer 13, widens proceeding from thebase 16 as far as thesurface 19, preferably constantly (FIG. 2 ) or only above a certain height 22 (FIG. 3 ), i.e. up to theheight 22, the cross section transverse to thelongitudinal direction 11 of therecess 4 is constant. - The
recess 4 preferably does not extend over the entire thickness of the layer 13 (not shown). - The
front edge 25 preferably runs perpendicular to the surface of a substrate of the layer 13 (or to the surface 19). - If the
front edge 25 is inclined, this inclination is in the direction of overflow. - This improves the spalling behavior, i.e. the growth of cracks in the ceramic protective layer is stopped. Similarly, the aerodynamics of a circulating cooling air, which flows out of the cooling air holes, are improved compared to conventional recesses by the form of the
recess 4. The surface temperature too is reduced here compared to conventional recesses. The geometries of the recesses which are proposed here can be subsequently introduced into theprotective layer 13 by a laser, for example. Similarly, it is possible to already produce therecesses 4 during the coating operation. - The turbine blade or
vane FIG. 7 . -
FIG. 4 shows a perspective view of arotor blade 120 orguide vane 130 of a turbomachine, which extends along alongitudinal 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 longitudinal axis 121, asecuring region 400, an adjoining blade orvane platform 403, a main blade orvane part 406 and a blade orvane tip 415. - As a
guide vane 130, thevane 130 may have a further platform (not shown) at itsvane tip 415. - A blade or
vane root 183, which is used to secure therotor blades 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 edge 409 and atrailing edge 412 for a medium which flows past the main blade orvane part 406. - In the case of conventional blades or
vanes regions vane - Superalloys of this type are known, for example, from
EP 1 204 776 B1,EP 1 306 454,EP 1 319 729 A1, WO 99/67435 or WO 00/44949. - The blade or
vane - 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.
- In this case, 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. In these processes, 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.
- Where the text 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).
- Processes of this type are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1.
- The blades or
vanes EP 1 306 454 A1. - The density is preferably 95% of the theoretical density.
- A protective aluminum oxide layer (TGO=thermally grown oxide layer) is formed on the MCrAlX layer (as an intermediate layer or as the outermost layer).
- The layer preferably has a composition Co—30Ni—28Cr—8Al—0.6Y—0.7Si or Co—28Ni—24Cr—10Al—0.6Y. In addition to these cobalt-based protective coatings, it is also preferable to use nickel-based protective layers, such as Ni—10Cr—12Al—0.6Y—3Re or Ni—12Co—21Cr—11Al—0.4Y—2Re or Ni—25Co—17Cr—10Al—0.4Y—1.5Re.
- It is also possible for a thermal barrier coating, which is preferably the outermost layer and consists for example of ZrO2, Y2O3—ZrO2, 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 suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).
- Other coating processes are possible, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may include grains that are porous or have micro-cracks or macro-cracks, in order to improve the resistance to thermal shocks. The thermal barrier coating is therefore preferably more porous than the MCrAlX layer.
- Refurbishment means that after they have been used, protective layers may have to be removed from
components 120, 130 (e.g. by sand-blasting). Then, the corrosion and/or oxidation layers and products are removed. If appropriate, cracks in thecomponent component component - The blade or
vane vane -
FIG. 5 shows acombustion chamber 110 of a gas turbine. - The
combustion chamber 110 is configured, for example, as what is known as an annular combustion chamber, in which a multiplicity ofburners 107, which generate flames 156 and are arranged circumferentially around an axis ofrotation 102, open out into a common combustion chamber space 154. For this purpose, thecombustion chamber 110 overall is of annular configuration positioned around the axis ofrotation 102. - To achieve a relatively high efficiency, the
combustion chamber 110 is designed for a relatively high temperature of the working medium M of approximately 1000° C. to 1600° C. To allow a relatively long service life even with these operating parameters, which are unfavorable for the materials, thecombustion chamber wall 153 is provided, on its side which faces the working medium M, with an inner lining formed fromheat shield elements 155. - On the working medium side, 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). - These protective layers may be similar to the turbine blades or vanes, i.e. for example 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 rare earth element 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. - A for example ceramic thermal barrier coating, consisting for example of ZrO2, Y2O3—ZrO2, 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).
- Other coating processes are conceivable, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. 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 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
heat shield element 155 are also repaired. This is followed by recoating of theheat shield elements 155, after which theheat shield elements 155 can be reused. - 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 thecombustion chamber 110. Theheat 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. -
FIG. 6 shows, by way of example, a partial longitudinal section through agas turbine 100. - In the interior, the
gas turbine 100 has arotor 103 with a shaft 101 which is mounted such that it can rotate about an axis ofrotation 102 and is also referred to as the turbine rotor. - An
intake housing 104, acompressor 105, a, for example,toroidal combustion chamber 110, in particular an annular combustion chamber, with a plurality of coaxially arrangedburners 107, aturbine 108 and the exhaust-gas housing 109 follow one another along therotor 103. - The
annular combustion chamber 110 is in communication with a, for example, annular hot-gas passage 111, where, by way of example, four successive turbine stages 112 form theturbine 108. - Each
turbine stage 112 is formed, for example, from two blade or vane rings. As seen in the direction of flow of a workingmedium 113, in the hot-gas passage 111 a row of guide vanes 115 is followed by a row 125 formed fromrotor blades 120. - The guide vanes 130 are secured to an
inner housing 138 of astator 143, whereas therotor blades 120 of a row 125 are fitted to therotor 103 for example by means of aturbine disk 133. - A generator (not shown) is coupled to the
rotor 103. - While the
gas turbine 100 is operating, thecompressor 105 sucks inair 135 through theintake housing 104 and compresses it. The compressed air provided at the turbine-side end of thecompressor 105 is passed to theburners 107, where it is mixed with a fuel. The mix is then burnt in thecombustion chamber 110, forming the workingmedium 113. From there, the workingmedium 113 flows along the hot-gas passage 111 past theguide vanes 130 and therotor blades 120. The workingmedium 113 is expanded at therotor blades 120, transferring its momentum, so that therotor blades 120 drive therotor 103 and the latter in turn drives the generator coupled to it. - While the
gas turbine 100 is operating, the components which are exposed to the hot workingmedium 113 are subject to thermal stresses. The guide vanes 130 androtor blades 120 of thefirst turbine stage 112, as seen in the direction of flow of the workingmedium 113, together with the heat shield elements which line theannular combustion chamber 110, are subject to the highest thermal stresses. - To be able to withstand the temperatures which prevail there, they may be cooled by means of a coolant.
- 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).
- By way of example, iron-based, nickel-based or cobalt-based superalloys are used as material for the components, in particular for the turbine blade or
vane combustion chamber 110. - Superalloys of this type are known, for example, from
EP 1 204 776 B1,EP 1 306 454,EP 1 319 729 A1, WO 99/67435 or WO 00/44949. - The blades or
vanes EP 1 306 454 A1. - A thermal barrier coating, consisting for example of ZrO2, Y2O3—ZrO2, 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).
- The
guide vane 130 has a guide vane root (not shown here), which faces theinner housing 138 of theturbine 108, and a guide vane head which is at the opposite end from the guide vane root. The guide vane head faces therotor 103 and is fixed to a securingring 140 of thestator 143.
Claims (16)
1.-11. (canceled)
12. A component having a layer with a surface, comprising
elongate recesses, which have a longitudinal direction, wherein the recesses are arranged at an angle which differs considerably from 0°, with respect to a direction of overflow over the surface, and are at least partially widened transversely to the longitudinal direction thereof in the region of the surface with respect to the base of the recess,
wherein the recess is widened proceeding from the base only above a certain height within the recess, and
wherein the recess does not extend through the entire thickness of the layer.
13. The component as claimed in claim 12 ,
wherein the recess has a front edge and a rear edge, and
wherein the widened portion is formed at the rear edge.
14. The component as claimed in claim 12 ,
wherein the recess has a rectangular cross section at least in certain points transversely to the longitudinal direction as far as the widened portion.
15. The component as claimed in claim 12 ,
wherein the recesses are arranged at an angle of 90°+/−20° with respect to a direction of overflow over the surface.
16. The component as claimed in claim 12 , wherein the recess has a corrugated form along the longitudinal direction.
17. The component as claimed in claim 12 ,
wherein the widened portion has a cross section which is at least 10% larger than that of the base.
18. The component as claimed in claim 12 ,
wherein the component is in the form of a surface of a solid component.
19. The component as claimed in claim 12 ,
wherein the component is in the form of a ceramic layer.
20. The component as claimed in claim 12 ,
wherein the front edge extends virtually perpendicular to the surface.
21. The component as claimed in claim 12 , wherein the surface comprises a ceramic surface.
22. The component as claimed in claim 12 , wherein the angle is 90°+/−20°.
23. The component as claimed in claim 12 , wherein the recess has a constant width before the certain height.
24. The component as claimed in claim 13 , wherein the widened portion is formed at the outflow-side end.
25. The component as claimed in claim 14 , wherein the recess has a completely rectangular cross section transverse to the longitudinal direction as far as the widened portion.
26. The component as claimed in claim 16 , wherein the recess has an s shape along the longitudinal direction.
Applications Claiming Priority (3)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP11186464.1 | 2011-10-25 | ||
EP11186464.1A EP2586985A1 (en) | 2011-10-25 | 2011-10-25 | Surface with specially formed depressions and component |
PCT/EP2012/066062 WO2013060499A1 (en) | 2011-10-25 | 2012-08-17 | Surface having specially formed recesses and component |
Publications (1)
Publication Number | Publication Date |
---|---|
US20140255652A1 true US20140255652A1 (en) | 2014-09-11 |
Family
ID=46704634
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US14/352,233 Abandoned US20140255652A1 (en) | 2011-10-25 | 2012-08-17 | Surface having specially formed recesses and component |
Country Status (3)
Country | Link |
---|---|
US (1) | US20140255652A1 (en) |
EP (2) | EP2586985A1 (en) |
WO (1) | WO2013060499A1 (en) |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20160281511A1 (en) * | 2012-11-16 | 2016-09-29 | Siemens Aktiengesellschaft | Modified surface around a hole |
Families Citing this family (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE102016222401A1 (en) * | 2016-11-15 | 2018-05-17 | Siemens Aktiengesellschaft | Creation of a forming hole in a wall |
Citations (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5558922A (en) * | 1994-12-28 | 1996-09-24 | General Electric Company | Thick thermal barrier coating having grooves for enhanced strain tolerance |
US20090097970A1 (en) * | 2007-10-16 | 2009-04-16 | United Technologies Corp. | Systems and Methods Involving Abradable Air Seals |
Family Cites Families (15)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE58908611D1 (en) | 1989-08-10 | 1994-12-08 | Siemens Ag | HIGH-TEMPERATURE-RESISTANT CORROSION PROTECTION COATING, IN PARTICULAR FOR GAS TURBINE COMPONENTS. |
DE3926479A1 (en) | 1989-08-10 | 1991-02-14 | Siemens Ag | RHENIUM-PROTECTIVE COATING, WITH GREAT CORROSION AND / OR OXIDATION RESISTANCE |
JP3370676B2 (en) | 1994-10-14 | 2003-01-27 | シーメンス アクチエンゲゼルシヤフト | Protective layer for protecting members against corrosion, oxidation and thermal overload, and method of manufacturing the same |
EP0861927A1 (en) | 1997-02-24 | 1998-09-02 | Sulzer Innotec Ag | Method for manufacturing single crystal structures |
EP0892090B1 (en) | 1997-02-24 | 2008-04-23 | Sulzer Innotec Ag | Method for manufacturing single crystal structures |
EP1306454B1 (en) | 2001-10-24 | 2004-10-06 | Siemens Aktiengesellschaft | Rhenium containing protective coating protecting a product against corrosion and oxidation at high temperatures |
WO1999067435A1 (en) | 1998-06-23 | 1999-12-29 | Siemens Aktiengesellschaft | Directionally solidified casting with improved transverse stress rupture strength |
US6231692B1 (en) | 1999-01-28 | 2001-05-15 | Howmet Research Corporation | Nickel base superalloy with improved machinability and method of making thereof |
DE50006694D1 (en) | 1999-07-29 | 2004-07-08 | Siemens Ag | HIGH-TEMPERATURE-RESISTANT COMPONENT AND METHOD FOR PRODUCING THE HIGH-TEMPERATURE-RESISTANT COMPONENT |
US8357454B2 (en) | 2001-08-02 | 2013-01-22 | Siemens Energy, Inc. | Segmented thermal barrier coating |
EP1319729B1 (en) | 2001-12-13 | 2007-04-11 | Siemens Aktiengesellschaft | High temperature resistant part, made of single-crystal or polycrystalline nickel-base superalloy |
EP1712739A1 (en) * | 2005-04-12 | 2006-10-18 | Siemens Aktiengesellschaft | Component with film cooling hole |
EP1942250A1 (en) * | 2007-01-05 | 2008-07-09 | Siemens Aktiengesellschaft | Component with bevelled grooves in the surface and method for operating a turbine |
US8079806B2 (en) * | 2007-11-28 | 2011-12-20 | United Technologies Corporation | Segmented ceramic layer for member of gas turbine engine |
EP2385155B1 (en) * | 2008-05-26 | 2015-06-24 | Siemens Aktiengesellschaft | Ceramic thermal barrier coating system with two ceramic layers |
-
2011
- 2011-10-25 EP EP11186464.1A patent/EP2586985A1/en not_active Withdrawn
-
2012
- 2012-08-17 WO PCT/EP2012/066062 patent/WO2013060499A1/en active Application Filing
- 2012-08-17 EP EP12748448.3A patent/EP2771546A1/en not_active Withdrawn
- 2012-08-17 US US14/352,233 patent/US20140255652A1/en not_active Abandoned
Patent Citations (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5558922A (en) * | 1994-12-28 | 1996-09-24 | General Electric Company | Thick thermal barrier coating having grooves for enhanced strain tolerance |
US20090097970A1 (en) * | 2007-10-16 | 2009-04-16 | United Technologies Corp. | Systems and Methods Involving Abradable Air Seals |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20160281511A1 (en) * | 2012-11-16 | 2016-09-29 | Siemens Aktiengesellschaft | Modified surface around a hole |
Also Published As
Publication number | Publication date |
---|---|
WO2013060499A1 (en) | 2013-05-02 |
EP2771546A1 (en) | 2014-09-03 |
EP2586985A1 (en) | 2013-05-01 |
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Owner name: SIEMENS AKTIENGESELLSCHAFT, GERMANY Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:AHMAD, FATHI;MENKE, CHRISTIAN;PAUL, UWE;REEL/FRAME:032689/0170 Effective date: 20140318 |
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STCB | Information on status: application discontinuation |
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