Abstract:
A process for producing a layer system is provided wherein the layer system has at least a substrate, a ceramic layer, which is applied to a surface structured in a targeted manner, in which process the intermediate layer, in particular the metallic layer, is applied in such a way that the recesses form during the coating. By introducing recesses into a surface, the stresses in the ceramic layer on the metallic substrate are reduced in such a manner that a longer lifespan for the ceramic layer is achieved.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is the U.S. National Stage of International Application No. PCT/EP2012/068048 filed Sep. 14, 2012, and claims the benefit thereof. The International Application claims the benefit of European Application No. EP11188032 filed Nov. 7, 2011. All of the applications are incorporated by reference herein in their entirety. 
     FIELD OF INVENTION 
     The invention relates to a process for producing a layer system. 
     BACKGROUND OF INVENTION 
     High-temperature components such as gas turbine components are often provided with ceramic thermal barrier layers, but these can also spall under the most extreme operating conditions. 
     This is caused by the occurrence of stresses, which lead to instances of spalling of the ceramic thermal barrier layer. 
     A solution to date was to provide the thermal barrier layer retrospectively with recesses. 
     SUMMARY OF INVENTION 
     It is therefore an object of the invention to further improve the solution to the aforementioned problem. 
     The object is achieved by a production process as claimed in the independent claims. 
     The dependent claims list further advantageous measures which can be combined with one another, as desired, in order to achieve further advantages. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIGS. 1-5  show exemplary embodiments of the invention, 
         FIG. 6  shows a turbine blade or vane, 
         FIG. 7  shows a combustion chamber, 
         FIG. 8  shows a gas turbine, and 
         FIG. 9  shows a list of superalloys. 
     
    
    
     The description and the figures represent exemplary embodiments of the invention. 
     DETAILED DESCRIPTION OF INVENTION 
       FIG. 5  shows a layer system  1 ,  120 ,  130 ,  155 . 
     The layer system  1 ,  120 ,  130 ,  155  comprises a substrate  4 , which in particular comprises a nickel-based or cobalt-based superalloy, in particular consists thereof, very particularly as per an alloy shown in  FIG. 9 . 
     An intermediate layer  10 , in particular a metallic bonding layer  10 , is optionally present on the surface  7  of the substrate  4 , and a ceramic thermal barrier layer  16  is present in turn on the surface  13  of said intermediate layer. 
     There are also combinations of substrates  4  with an aluminized surface region, in which case the ceramic thermal barrier layer can be applied directly to the substrate. 
     The metallic bonding layer  10  preferably comprises an MCrAlX alloy. 
     According to the invention, recesses  19 ′,  19 ″, . . . are present in or are introduced into the surface  7  of the substrate  4  or in the surface  13  of the layer  10  ( FIG. 1 ). 
     The recesses  19 ′,  19 ″, . . . have a certain depth b and a certain width a. 
     The width a of the recesses  19 ′,  19 ″, . . . is at least 10 μm, preferably 10 μm to 30 μm. 
     The depth b is at least 10%, preferably 10% to 30%, of the thickness of the underlying layer  10 , very particularly 10 μm to 30 μm. 
     The distance d between the recesses  19 ′,  19 ″, . . . lying opposite one another is at least 100 μm, preferably between 100 μm and 300 μm ( FIG. 2 ). 
     The parameters a, b, d can be varied depending on the operating conditions or locally (on the main blade or vane part  406  but not on the blade or vane platform  403 ) on the surface  7 ,  13 . 
     Similarly, the recesses  19 ′,  19 ″ can be present on the surface  7 ,  13  of the component  1 ,  120 ,  130  only in a locally limited manner. 
     The recesses  19 ′,  19 ″, . . . can preferably have a round configuration at the base  20  ( FIG. 1 ). 
     The recesses  19 ′,  19 ″, . . . can have a honeycomb structure ( FIG. 3 ) or a mesh structure ( FIG. 4 ). 
       FIG. 1  shows a cross section through such a surface structured in a targeted manner. 
     Depending on the size of the recesses  19 ′,  19 ″, . . . , the recess  19 ′,  19 ″ also continues into recesses  23 ′,  23 ″ at the surface  22  of the ceramic thermal barrier layer  16 . 
     Stresses are reduced and the metallic bonding layer  10  and ceramic thermal barrier layer  16  (or layer  16  and substrate  4 ) are mechanically braced. It is much easier to machine the metallic surface of the layer  10  or of the substrate  4  than a ceramic surface. 
     Similarly, the coating  16  can be configured in such a way that the outermost surface  22  is smooth, i.e. the underlying recesses  23 ′,  23 ″ would not be identifiable on the surface  22 . 
     The layers  10  are often applied by the application of material (e.g. powder) from a nozzle, in particular in a linear manner. By omitting a lane of coating when coating, or by targeted non-coating, no material is applied at that point and a recess  19 ′,  19 ″ is formed. 
     This is possible in particular in coating processes such as APS, VPS, LPPS, HVOF and cold gas spraying, in which powder is applied in tracks. 
     The structured surface  7 ,  13  is an integral part of a layer  10 . It therefore does not constitute a honeycomb structure filled with a ceramic material. 
       FIG. 6  shows, by way of example, a partial longitudinal section through a gas turbine  100 . 
     In the interior, 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. 
     An intake housing  104 , a compressor  105 , a, for example, toroidal combustion chamber  110 , in particular an annular combustion chamber, with a plurality of coaxially arranged burners  107 , a turbine  108  and the exhaust-gas housing  109  follow one another along the rotor  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 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 a turbine disk  133 . 
     A generator (not shown) is coupled to the rotor  103 . 
     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. 
     While the gas turbine  100  is operating, the components which are exposed to the hot working medium  113  are subject to thermal stresses. The guide vanes  130  and rotor blades  120  of the first turbine stage  112 , as seen in the direction of flow of the working medium  113 , together with the heat shield elements which line the annular 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 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  120 ,  130  and components of the 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  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 layer, 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 layer 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 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 . 
       FIG. 7  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 . For this purpose, the combustion chamber  110  overall is of annular configuration positioned around the axis of rotation  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, 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 . 
     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 layer, 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 layer 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 layer 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. 8  shows, by way of example, a partial longitudinal section through a gas turbine  100 . 
     In the interior, 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. 
     An intake housing  104 , a compressor  105 , a, for example, toroidal combustion chamber  110 , in particular an annular combustion chamber, with a plurality of coaxially arranged burners  107 , a turbine  108  and the exhaust-gas housing  109  follow one another along the rotor  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 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 a turbine disk  133 . 
     A generator (not shown) is coupled to the rotor  103 . 
     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. 
     While the gas turbine  100  is operating, the components which are exposed to the hot working medium  113  are subject to thermal stresses. The guide vanes  130  and rotor blades  120  of the first turbine stage  112 , as seen in the direction of flow of the working medium  113 , together with the heat shield elements which line the annular 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 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  120 ,  130  and components of the 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  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 layer, 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 layer 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 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 .