Abstract:
At least two ply layers of ceramic on a substrate which is applied to protect a surface in a heated hot environment. Each of the outer and bottom layers including zirconium oxide and stabilizers of yttrium oxide in different respective proportions of the yttrium oxide; the outer layer has fully stabilized zirconium oxide and the bottom layer has partially stabilized zirconium oxide.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present application is a 35 U.S.C. §§371 national phase conversion of PCT/EP2012/065458, filed Aug. 7, 2012, which claims priority of European Patent Application No. 11185023.6, filed Oct. 13, 2011, the contents of which are incorporated by reference herein. The PCT International Application was published in the German language. 
     
    
     FIELD OF THE INVENTION 
       [0002]    The invention relates to a two-ply ceramic layer system based on zirconium oxide. 
       TECHNICAL BACKGROUND 
       [0003]    The use of zirconium oxide as a single layer on turbine blades or vanes is known. 
         [0004]    It is similarly known to use a plurality of ceramic layers having an inner zirconium oxide layer and an outer oxide layer based on pyrochlore. 
       SUMMARY OF THE INVENTION 
       [0005]    It is an object of the invention to further simplify and to improve existing ceramic coatings. 
         [0006]    The two ply ceramic layer system of the invention includes at least two ply layers of ceramic on a substrate for protecting a surface in a heated hot environment. Each of the outer and bottom layers includes zirconium oxide and stabilizers of yttrium oxide in different respective proportions of the yttrium oxide. The outer layer has fully stabilized zirconium oxide and the bottom layer has partially stabilized zirconium oxide. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0007]      FIG. 1  shows a layer system, 
           [0008]      FIG. 2  shows a list of superalloys, 
           [0009]      FIG. 3  shows a turbine blade or vane, 
           [0010]      FIG. 4  shows a combustion chamber, and 
           [0011]      FIG. 5  shows a gas turbine. 
       
    
    
     DESCRIPTION OF EMBODIMENTS 
       [0012]    The description and the Figures represent only exemplary embodiments of the invention. 
         [0013]      FIG. 1  shows an example of a layer system  1  according to the invention. 
         [0014]    The layer system  1  comprises a substrate  4 . 
         [0015]    This is preferably metallic and comprises a nickel-based or cobalt-based superalloy. In particular, use is made of superalloys as per  FIG. 2 . 
         [0016]    A metallic bonding and corrosion-resistant layer  7  is optionally present on the substrate  4 . These can be aluminide layers in various variations or preferably MCrAlX layers (X optionally=Y, Re, Si, Ta, Fe, . . . ). 
         [0017]    An aluminum oxide layer which contributes to oxidation protection (not shown) is formed on the MCrAlX layer  7  or on the substrate  4  during operation (TGO) or preferably beforehand. 
         [0018]    The preferably two-ply ceramic layer  15  in each case based on zirconium oxide, i.e. the proportion of zirconium oxide is at least 50% by weight, in particular at least 60% by weight, is present on the metallic bonding layer  7  or the substrate  4 . 
         [0019]    The ceramic layers  10 ,  13  have different properties. 
         [0020]    The bottom ceramic layer  10  based on zirconium oxide is preferably partially stabilized. 
         [0021]    The partial stabilization can preferably be effected by yttrium oxide with known proportions, preferably between 4% by weight and 12% by weight. 
         [0022]    Other known stabilizers—alone or in addition—corresponding to the known additions for partial stabilization are likewise possible, for example magnesium oxide and/or other oxides. 
         [0023]    The outer ceramic layer  13  is preferably a fully stabilized zirconium oxide layer. 
         [0024]    The full stabilization is preferably likewise achieved by yttrium oxide in corresponding higher proportions, preferably between 20% by weight and 50% by weight, in particular between 30% and 40%. Other known stabilizers—alone or in addition—corresponding to the known additions for partial stabilization are likewise possible, for example magnesium oxide or other oxides. 
         [0025]    This ceramic double layer system has the advantage that very high operating temperatures are achievable through the fully stabilized system of the outer layer  13 , since the C phase in this system is stable throughout the temperature range. Therefore, no phase transitions arise in the top layer  13 . 
         [0026]    A good bond between the ceramic layers  10 ,  13  is provided as a result of the lower temperature at the inner ceramic layer  10  and the similar chemical systems of the layers  10 ,  13 . Similarly, the high proportions of the stabilizer, in particular yttrium, make it possible to contain the losses thereof and to keep the system in the C phase region stable. 
         [0027]    The two layers  10 ,  13  preferably have a porosity of &gt;16%, in particular of 16% to 24%. 
         [0028]    The outer layer  13  is at least 20%, in particular at least 30%, thicker than the bottom layer  10 . The layer thickness ratio (10/13) is preferably 1/3 to 2/3. 
         [0029]    The bottom ceramic layer  10  preferably has a thickness of up to 500 micrometers. 
         [0030]    The second ceramic layer  13  preferably has a thickness of up to 1000 micrometers. 
         [0031]    A layer system of this type can be used for components  120 ,  130 ,  155  ( FIGS. 3 and 4 ) that are employed at high temperatures. These are in particular combustion chamber blocks  155 , turbine blades or vanes  120 ,  130  for aircraft, gas turbines  100  ( FIG. 5 ) and/or steam turbines. 
         [0032]      FIG. 3  shows a perspective view of a rotor blade  120  or guide vane  130  of a turbomachine, which extends along a longitudinal axis  121 . 
         [0033]    The turbomachine may be a gas turbine of an aircraft or of a power plant for generating electricity, a steam turbine or a compressor. 
         [0034]    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 . 
         [0035]    As a guide vane  130 , the vane  130  may have a further platform (not shown) at its vane tip  415 . 
         [0036]    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 . 
         [0037]    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. 
         [0038]    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 . 
         [0039]    In the case of conventional blades or vanes  120 ,  130 , by way of example solid metallic materials, in particular superalloys, are used in all regions  400 ,  403 ,  406  of the blade or vane  120 ,  130 . 
         [0040]    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. 
         [0041]    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. 
         [0042]    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. 
         [0043]    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. 
         [0044]    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). 
         [0045]    Processes of this type are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1. 
         [0046]    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. 
         [0047]    The density is preferably 95% of the theoretical density. 
         [0048]    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). 
         [0049]    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. 
         [0050]    It is also possible for a 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. 
         [0051]    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). 
         [0052]    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. 
         [0053]    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. 
         [0054]    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). 
         [0055]      FIG. 4  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 . 
         [0056]    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 . 
         [0057]    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). 
         [0058]    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. 
         [0059]    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. 
         [0060]    Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD). 
         [0061]    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. 
         [0062]    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. 
         [0063]    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 . 
         [0064]      FIG. 5  shows, by way of example, a partial longitudinal section through a gas turbine  100 . 
         [0065]    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. 
         [0066]    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 . 
         [0067]    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 . 
         [0068]    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 . 
         [0069]    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 . 
         [0070]    A generator (not shown) is coupled to the rotor  103 . 
         [0071]    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. 
         [0072]    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 means of a coolant. 
         [0073]    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. 
         [0074]    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. 
         [0075]    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. 
         [0076]    Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD). 
         [0077]    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 .