Abstract:
Thermal barrier coating layer systems, in addition to good thermal barrier properties, also have to have a long service life of the thermal barrier coating. The layer system according to the invention comprises a specially adapted layer sequence of metallic bonding layer inner ceramic layer and outer ceramic layer.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS  
       [0001]     This application claims the benefits of European Patent application No. 05007225.5 filed Apr. 1, 2005 and is incorporated by reference herein in its entirety.  
       FIELD OF THE INVENTION  
       [0002]     The invention relates to a layer system as described in the claims.  
       BACKGROUND OF THE INVENTION  
       [0003]     A layer system of this type has a substrate comprising a metal alloy based on nickel, cobalt or iron. Products of this type are used in particular as components of a gas turbine, in particular as gas turbine blades or vanes or heat shields. The components are exposed to a hot-gas stream of aggressive combustion gases, and consequently they have to be able to withstand high thermal stresses. Furthermore, it is necessary for these components to be resistant to oxidation and corrosion. Furthermore, mechanical demands are imposed in particular on moving components, for example gas turbine blades or vanes, but also on static parts. The power and efficiency of a gas turbine in which components that can be exposed to hot gases are used rise with an increasing operating temperature. To achieve a high efficiency and a high power, components of the gas turbines which are subject to particularly high stresses from the high temperatures are coated with a ceramic material. This ceramic material acts as a thermal barrier coating between the hot-gas stream and the metallic substrate.  
         [0004]     The metallic base body is protected from the aggressive hot-gas stream by coatings. Modem components generally have a plurality of coatings which each perform specific tasks. Therefore, a multilayer system is present.  
         [0005]     Since power and efficiency of gas turbines rise with increasing operating temperature, constant attempts have been made to achieve a higher gas turbine performance by improving the coating system.  
         [0006]     EP 0 944 746 B1 discloses the use of pyrochlores as thermal barrier coating.  
         [0007]     However, to be used as material for a thermal barrier coating, it is necessary for materials not only to have good thermal barrier properties but also good bonding to the substrate.  
         [0008]     EP 0 992 603 A1 discloses a thermal barrier coating system comprising gadolinium oxide and zirconium oxide, which is not supposed to have a pyrochlore structure.  
       SUMMARY OF THE INVENTION  
       [0009]     Therefore, it is an object of the invention to provide a layer system which has good thermal barrier properties and good bonding to the substrate and therefore provides a long service life of the overall layer system.  
         [0010]     The invention is based on the discovery that the entire system has to be considered as a single unit, rather than regarding and optimizing individual layers or combinations of individual layers in isolation, with a view to achieving a long service life.  
         [0011]     The object is achieved by the layer system as claimed in the claims.  
         [0012]     The subclaims list further advantageous measures which can be combined in any desired, advantageous way. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0013]     In the drawing:  
         [0014]      FIG. 1  shows a layer system according to the invention,  
         [0015]      FIG. 2  shows a turbine blade or vane,  
         [0016]      FIG. 3  shows a gas turbine. 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0017]      FIG. 1  shows a layer system  1  according to the invention.  
         [0018]     The layer system  1  comprises a metallic substrate  4 , which in particular for components used at high temperatures consists of a nickel-base or cobalt-base superalloy.  
         [0019]     Directly on the substrate  4  there is a metallic bonding layer  7 , which consists either of 11-13 wt % cobalt, 20-22 wt % chromium, 10.5-11.5 wt % aluminum, 0.3-0.5 wt % yttrium, 1.5-2.5 wt % rhenium, remainder nickel, or 24-26 wt % cobalt, 16-18 wt % chromium, 9.5-11 wt % aluminum, 0.3-0.5 wt % yttrium, 0.5-2 wt % rhenium, remainder nickel.  
         [0020]     Even before the application of further ceramic layers, an aluminum oxide layer has formed on this metallic bonding layer  7 , or an aluminum oxide layer of this type is formed during operation.  
         [0021]     A fully or partially stabilized zirconium oxide layer is present as inner ceramic layer  10  on the metallic bonding layer  7  or on the aluminum oxide layer (not shown). It is preferable to use yttrium-stabilized zirconium oxide. It is also possible to use calcium oxide, cerium oxide or hafnium oxide to stabilize zirconium oxide.  
         [0022]     The zirconium oxide is preferably applied as a plasma-spray layer, but also may be applied as a columnar structure by means of electron beam physical vapor deposition.  
         [0023]     An outer ceramic layer  13 , which mostly comprises a pyrochlore phase, i.e. is made up to an extent of at least 80 wt % of the pyrochlore phase and comprises Gd 2 Hf 2 O 7  or Gd 2 Zr 2 O 7 , has been applied to the stabilized zirconium oxide layer  10 . It is preferable for the outer layer  13  to consist of 100 wt % of one of the two pyrochlore phases.  
         [0024]     Amorphous phases or pure GdO 2  or pure ZrO 2  or pure HfO 2  have been disregarded. Mixed phases of GdO 2  and ZrO 2  and/or HfO 2  which do not comprise the pyrochlore phase are undesirable and should be minimized.  
         [0025]     The crucial factor in the invention is the discovery that not only does the interaction between the outer ceramic layer  13  and an inner ceramic layer  10  need to be optimized, but also the metallic bonding layer  7  has a significant influence on the service life and function of the outer ceramic layer  13  of this two-layer ceramic structure.  
         [0026]      FIG. 2  shows a perspective view of a rotor blade  120  or guide vane  130  of a turbomachine, which extends along a longitudinal axis  121 .  
         [0027]     The turbomachine may be a gas turbine of an aircraft or of a power plant for generating electricity, a steam turbine or a compressor.  
         [0028]     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  and a main blade or vane part  406 .  
         [0029]     As a guide vane  130 , the vane  130  may have a further platform (not shown) at its vane tip  415 .  
         [0030]     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 .  
         [0031]     The blade or vane root  183  is designed, for example, in hammerhead form. Other configurations as a fir-tree root or dovetail root are possible.  
         [0032]     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 .  
         [0033]     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 .  
         [0034]     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; these documents form part of the disclosure with regard to the chemical composition of the alloy.  
         [0035]     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.  
         [0036]     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.  
         [0037]     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.  
         [0038]     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.  
         [0039]     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).  
         [0040]     Processes of this type are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1; these documents form part of the present disclosure.  
         [0041]     The blades or vanes  120 ,  130  may also 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, which are intended to form part of the present disclosure with regard to the chemical composition of the alloy.  
         [0042]     There may also be a thermal barrier coating consisting, for example, of ZrO 2 , Y 2 O 4 —ZrO 3 , i.e. unstabilized, partially stabilized or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide, 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).  
         [0043]     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.  
         [0044]     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  (illustrated in dashed lines).  
         [0045]      FIG. 3  shows, by way of example, a partial longitudinal section through a gas turbine  100 .  
         [0046]     In the interior, the gas turbine  100  has a rotor  103  which is mounted such that it can rotate about an axis of rotation  102  and has a shaft  101  and is also referred to as the turbine rotor.  
         [0047]     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 .  
         [0048]     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 .  
         [0049]     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 .  
         [0050]     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 .  
         [0051]     A generator (not shown) is coupled to the rotor  103 .  
         [0052]     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.  
         [0053]     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.  
         [0054]     To be able to withstand the temperatures which prevail there, they may be cooled by means of a coolant.  
         [0055]     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).  
         [0056]     By way of example, 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 .  
         [0057]     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; these documents form part of the disclosure with regard to the chemical composition of the alloys.  
         [0058]     The blades or vanes  120 ,  130  may also have coatings which protect 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 represents yttrium (Y) and/or silicon and/or at least one rare earth element and/or hafnium).  
         [0059]     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 the present disclosure with regard to the chemical composition.  
         [0060]     A thermal barrier coating may also be present on the MCrAlX, 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.  
         [0061]     Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).  
         [0062]     The guide vane  130  has a guide vane root (not shown here) facing the inner housing  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 .  
         [0063]     List of Designations  
         [0064]      1  Layer system 
    Substrate     Bonding layer     Inner ceramic layer     Outer ceramic layer     Gas turbine     Axis of rotation     Rotor     Intake housing     Compressor     Annular combustion chamber     Burner     Turbine     Exhaust-gas housing     Combustion chamber     Hot-gas passage     Turbine stage     Working medium     Row of guide vanes     Rotor blade     Longitudinal axis     Row     Guide vane     Turbine disk     Air     Inner housing     Securing ring     Stator     Combustion chamber wall     Heat shield element     Blade or vane root     Securing region     Blade or vane platform     Main blade or vane part     Leading edge     Trailing edge     Blade or vane tip     Film-cooling holes