Abstract:
Known protective layers with a high Cr content form brittle phases which become even more brittle during use under the influence of carbon. The protective layer according to the invention has the composition 26% to 28% cobalt, 20% to 22% chromium, 7% to 8% aluminium, 0.5% to 0.7% yttrium and/or at least one equivalent metal from the group comprising scandium and the rare-earth elements, optionally silicon and/or rhenium and the rest made up of nickel.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is the US National Stage of International Application No. PCT/EP2006/067802, filed Oct. 26, 2006 and claims the benefit thereof. The International Application claims the benefits of European application No. 05024112.4 filed Nov. 4, 2005, both of the applications are incorporated by reference herein in their entirety. 
     
    
     FIELD OF INVENTION 
       [0002]    The invention relates to an alloy as claimed in the claims, to a protective layer for protecting a component against corrosion and/or oxidation at high temperatures as claimed in the claims and to a component as claimed in the claims. 
         [0003]    The invention relates in particular to a protective layer for a component which consists of a nickel- or cobalt-based superalloy. 
       BACKGROUND OF THE INVENTION 
       [0004]    Large numbers of protective layers for metal components, which are intended to increase their corrosion resistance and/or oxidation resistance, are known in the prior art. Most of these protective layers are known by the generic name MCrAlY, where M stands for at least one of the elements from the group comprising iron, cobalt and nickel and the other essential constituents are chromium, aluminum and yttrium. 
         [0005]    Typical coatings of this type are known from U.S. Pat. Nos. 4,005,989 and 4,034,142. 
         [0006]    U.S. Pat. No. 6,280,857 B1 discloses a protective layer which discloses the elements cobalt, chromium and aluminum based on nickel and mandatory additions of yttrium, rhenium and silicon. 
         [0007]    The endeavor to increase the intake temperatures both in static gas turbines and in aircraft engines is of great importance in the specialist field of gas turbines, since the intake temperatures are important determining quantities for the thermodynamic efficiencies achievable with gas turbines. Intake temperatures significantly higher than 1000° C. are possible when using specially developed alloys as base materials for components to be strongly heated, such as guide vanes and rotor blades, in particular by using monocrystalline superalloys. To date, the prior art permits intake temperatures of 950° C. or more for static gas turbines and 1100° C. or more in gas turbines of aircraft engines. 
         [0008]    Examples of the structure of a turbine blade with a monocrystalline substrate, which in turn may be complexly constructed, are disclosed by WO 91/01433 A1. 
         [0009]    While the physical loading capacity of the base materials so far developed for the components to be heavily loaded is substantially unproblematic in respect of possible further increases in the intake temperatures, it is necessary to resort to protective layers in order to achieve sufficient resistance against oxidation and corrosion. Besides sufficient chemical stability of a protective layer under the aggressions which are to be expected from exhaust gases at temperatures of the order of 1000° C., a protective layer must also have sufficiently good mechanical properties, not least in respect of the mechanical interaction between the protective layer and the base material. In particular, the protective layer must be ductile enough to be able to accommodate possible deformations of the base material and not crack, since points of attack would thereby be provided for oxidation and corrosion. The problem then typically arises that increasing the proportions of elements such as aluminum and chromium, which can improve the resistance of a protective layer against oxidation and corrosion, leads to a deterioration of the ductility of the protective layer so that mechanical failure is possible, in particular the formation of cracks, under a mechanical load conventionally occurring in a gas turbine. 
       SUMMARY OF INVENTION 
       [0010]    It is therefore an object of the invention to provide an alloy and a protective layer which has good high-temperature resistance to corrosion and oxidation, has good long-term stability and which is furthermore adapted particularly well to a mechanical load which is to be expected particularly in a gas turbine at a high temperature. 
         [0011]    The object is achieved by an alloy as claimed in the claims and a protective layer as claimed in the claims. 
         [0012]    It is another object of the invention to provide a component which has increased protection against corrosion and oxidation. 
         [0013]    The object is likewise achieved by a component as claimed in the claims, in particular a component of a gas turbine or steam turbine, which comprises a protective layer of the type described above for protection against corrosion and oxidation at high temperatures. 
         [0014]    Further advantageous measures are listed in the dependent claims. 
         [0015]    The measures listed in the dependent claims may advantageously be combined with one another in any desired way. 
         [0016]    The invention is based inter alia on the discovery that the protective layer exhibits brittle chromium-rhenium precipitates in the protective layer itself and in the transition region between the protective layer and the base material. These brittle phases, which are formed increasingly over time and with the temperature during use, lead during operation to very pronounced longitudinal cracks in the protective layer as well as in the layer-base material interface, with subsequent shedding of the protective layer. The brittleness of the precipitates is further increased by the interaction with carbon, which can diffuse into the protective layer from the base material or diffuses into the protective layer through the surface during a heat treatment in the furnace. The impetus to cracking is further enhanced by oxidation of these phases. 
         [0017]    The effect of cobalt, which determines the thermal and mechanical properties, is also important in this case. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS  
         [0018]    The invention will be explained in more detail below. 
           [0019]      FIG. 1  shows a layer system with a protective layer, 
           [0020]      FIG. 2  shows compositions of superalloys, 
           [0021]      FIG. 3  shows a gas turbine, 
           [0022]      FIG. 4  shows a perspective view of a combustion chamber and 
           [0023]      FIG. 5  shows a perspective view of a turbine blade. 
       
    
    
     DETAILED DESCRIPTION OF INVENTION 
       [0024]    According to the invention, a protective layer  7  ( FIG. 1 ) for protecting a component against corrosion and oxidation at a high temperature contains the following elements (in wt %): 
         [0025]    from 26% to 28% cobalt (Co) 
         [0026]    from 20% to 22% chromium (Cr) 
         [0027]    from 7% to 9% aluminum (Al) 
         [0028]    from 0.5% to 0.7% yttrium (Y) and/or at least one equivalent metal from the group comprising scandium (Sc) and the rare earth elements, remainder nickel (NiCoCrAlY). 
         [0029]    The alloy optionally contains up to 2 wt % silicon. 
         [0030]    The alloy may furthermore comprise up 11 wt % rhenium. 
         [0031]    The advantageous effect of the element rhenium can thereby be utilized while preventing the brittle phase formation. 
         [0032]    The alloy may optionally also comprise ruthenium. Ruthenium with a maximum proportion of 11 wt % may partially or fully replace the rhenium. 
         [0033]    It is preferable to use only rhenium. 
         [0034]    It is to be noted that the proportions of the individual elements are specially adapted with a view to their effects. If the proportions are dimensioned so that no chromium precipitates are formed, then advantageously no brittle phases are created during use of the protective layer so that the operating time performance is improved and extended. 
         [0035]    This is achieved not only by a low chromium content but also, taking into account the effect of aluminum on the phase formation, by accurately dimensioning the aluminum content. 
         [0036]    The choice of from 26 wt % to 28 wt % cobalt surprisingly improves the thermal and mechanical properties of the protective layer  7  significantly and superproportionally. 
         [0037]    With good corrosion resistance, the protective layer  7  has particularly good resistance against oxidation and is also distinguished by particularly good ductility properties, so that it is particularly qualified for use in a gas turbine with a further increase in the intake temperature. During operation, embrittlement scarcely takes place since the layer comprises hardly any chromium precipitates, in particular no chromium-rhenium precipitates, which become embrittled in the course of use. 
         [0038]    It is advantageous to set the proportion of aluminum at 8 wt % and to form of Al 2 O 3  during coating with the alloy. The proportion of aluminum can therefore be kept low. It is likewise advantageous to set the proportion of yttrium or the at least one equivalent element from the group comprising scandium and the rare earth elements at 0.6 wt %. Certain variations are encountered owing to industrial mass production, so that yttrium contents of from 0.4% to 0.5% or from 0.7% to 0.8% are also used and likewise exhibit good properties. 
         [0039]    It is particularly favorable to set the chromium content at about 21 wt %, the aluminum content at about 8 wt % and the cobalt content at about 27 wt %. 
         [0040]    The alloy preferably contains no other elements besides the elements nickel, chromium, cobalt, aluminum, yttrium (Sc, rare earths). 
         [0041]    Particularly good exemplary embodiments are:
       1) Ni-27Co-21Cr-8Al-0.6Y   2) Ni-27Co-21Cr-8Al-0.6Y-1.5Si   3) Ni-27Co-21Cr-8Al-0.6Y-1.5Si—Re       
 
         [0045]    The trace elements in the powder to be sprayed, which form precipitates and therefore represent embrittlements, play a likewise important role. 
         [0046]    The powders are for example applied by plasma spraying (APS, LPPS, VPS, . . . ). Other methods may likewise be envisaged (PVD, CVD, cold gas spraying, . . . ). 
         [0047]    The thickness of the protective layer  7  on the component  1  is preferably dimensioned at a value of between 100 μm and 300  82  m. 
         [0048]    In this component, the protective layer  7  is advantageously applied onto a substrate  4  made of a nickel-based or cobalt-based superalloy. 
         [0049]    The following composition in particular may be suitable as a substrate  4  (data in wt %): 
         [0050]    From 0.1% to 0.15% carbon 
         [0051]    from 18% to 22% chromium 
         [0052]    from 18% to 19% cobalt 
         [0053]    from 0% to 2% tungsten 
         [0054]    from 0% to 4% molybdenum 
         [0055]    from 0% to 1.5% tantalum 
         [0056]    from 0% to 1% niobium 
         [0057]    from 1% to 3% aluminum 
         [0058]    from  2 % to  4 % titanium 
         [0059]    from 0% to 0.75% hafnium 
         [0060]    optionally small proportions of boron and/or zirconium, remainder nickel. 
         [0061]    Compositions of this type are known as casting alloys under the references GDT222, IN939, IN6203 and Udimet 500. 
         [0062]    Other advantageous alternatives for the substrate  4  of the component are listed in  FIG. 2 . 
         [0063]    The protective layer  7  is particularly suitable for protecting a component against corrosion and oxidation while the component is being exposed to an exhaust gas at a material temperature of about 950° C., or even about 1100° C. in aircraft turbines. 
         [0064]    The protective layer  7  according to the invention is therefore particularly qualified for protecting a component of a gas turbine  100 , in particular a guide vane  130 , rotor blade  120  or other components, which are exposed to hot gas before or in the turbine of the gas turbine. 
         [0065]    The protective layer  7  may be used as an overlay (the protective layer is the outer layer) or as a bondcoat (the protective layer is an interlayer). 
         [0066]      FIG. 1  shows a layer system  1  as a component. 
         [0067]    The layer system  1  consists of a substrate  4 . 
         [0068]    The substrate  4  may be metallic and/or ceramic. Particularly in the case of turbine components, for example turbine rotor blades  120  ( FIG. 1 ) or guide vanes  130  ( FIGS. 3 ,  5 ), combustion chamber linings  155  ( FIG. 4 ) and other housing parts of a steam or gas turbine  100  ( FIG. 3 ), the substrate  4  consists of a nickel-, cobalt- or iron-based superalloy. 
         [0069]    Cobalt-based superalloys are preferably used. 
         [0070]    The protective layer  7  according to the invention is placed on the substrate  4 . 
         [0071]    This protective layer  7  is preferably applied by LPPS (low pressure plasma spraying). 
         [0072]    It may be used as an outer layer (not shown) or interlayer ( FIG. 1 ). 
         [0073]    In the latter case, there is a ceramic thermal insulation layer  10  on the protective layer  7 . 
         [0074]    The protective layer  7  may be applied onto newly produced components and refurbished components. 
         [0075]    Refurbishment means that components  1  are separated if need be from layers (thermal insulation layer) after their use and corrosion and oxidation products are removed, for example by an acid treatment (acid stripping). It may sometimes also be necessary to repair cracks. Such a component may subsequently be recoated, since the substrate  4  is very expensive. 
         [0076]      FIG. 3  shows a gas turbine  100  by way of example in a partial longitudinal section. 
         [0077]    The gas turbine  100  internally comprises a rotor  103 , which will also be referred to as the turbine rotor, mounted so as to rotate about a rotation axis  102 . 
         [0078]    Successively along the rotor  103 , there are an intake manifold  104 , a compressor  105 , an e.g. toroidal combustion chamber  110 , in particular a ring combustion chamber  106 , having a plurality of burners  107  arranged coaxially, a turbine  108  and the exhaust manifold  109 . 
         [0079]    The ring combustion chamber  106  communicates with an e.g. annular hot gas channel  111 . There, for example, four successively connected turbine stages  112  form the turbine  108 . 
         [0080]    Each turbine stage  112  is formed for example by two blade rings. As seen in the flow direction of a working medium  113 , a guide vane row  115  is followed in the hot gas channel  111  by a row  125  formed by rotor blades  120 . 
         [0081]    The guide vanes  130  are fastened on an inner housing  138  of a stator  143  while the rotor blades  120  of a row  125  are fitted on the rotor  103 , for example by means of a turbine disk  133 . Coupled to the rotor  103 , there is a generator or a work engine (not shown). 
         [0082]    During operation of the gas turbine  100 , air  135  is taken in and compressed by the compressor  105  through the intake manifold  104 . The compressed air provided at the turbine-side end of the compressor  105  is delivered to the burners  107  and mixed there with a fuel. The mixture is then burnt to form the working medium  113  in the combustion chamber  110 . 
         [0083]    From there, the working medium  113  flows along the hot gas channel  111  past the guide vanes  130  and the rotor blades  120 . At the rotor blades  120 , the working medium  113  expands by imparting momentum, so that the rotor blades  120  drive the rotor  103  and the work engine coupled to it. 
         [0084]    During operation of the gas turbine  100 , the components exposed to the hot working medium  113  experience thermal loads. Apart from the heat shield elements lining the ring combustion chamber  106 , the guide vanes  130  and rotor blades  120  of the first turbine stage  112 , as seen in the flow direction of the working medium  113 , are heated the most. 
         [0085]    In order to withstand the temperatures prevailing there, they may be cooled by means of a coolant. 
         [0086]    The substrates may likewise comprise a directional structure, i.e. they are monocrystalline (SX structure) or comprise only longitudinally directed grains (DS structure). 
         [0087]    Iron-, nickel- or cobalt-based superalloys used as the material. 
         [0088]    For example, superalloys such as are known from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949 are used. With respect to the chemical composition of the superalloys and their advantages, these documents are part of the disclosure. 
         [0089]    The blades  120 ,  130  comprise protective layers  7  according to the invention against corrosion and/or a thermal insulation layer. The thermal insulation layer consists for example of ZrO 2 , Y 2 O 3 —ZrO 2 , i.e. it is not stabilized or is partially or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide. 
         [0090]    Rod-shaped grains are produced in the thermal insulation layer by suitable coating methods, for example electron beam deposition (EB-PVD). 
         [0091]    The guide vanes  130  comprise a guide vane root (not shown here) facing the inner housing  138  of the turbine  108 , and a guide vane head lying opposite the guide vane root. The guide vane head faces the rotor  103  and is fixed on a fastening ring  140  of the stator  143 . 
         [0092]      FIG. 4  shows a combustion chamber  110  of a gas turbine  100 , which may comprise a layer system  1 . 
         [0093]    The combustion chamber  110  is designed for example as a so-called ring combustion chamber in which a multiplicity of burners  107 , arranged in the circumferential direction around the turbine shaft  103 , open into a common combustion chamber space. To this end, the combustion chamber  110  as a whole is designed as an annular structure which is positioned around the turbine shaft  103 . 
         [0094]    In order to achieve a comparatively high efficiency, the combustion chamber  110  is designed for a relatively high temperature of the working medium M, i.e. about 1000° C. to 1600° C. In order to permit a comparatively long operating time even under these operating parameters which are unfavorable for the materials, the combustion chamber wall  153  is provided with an inner lining formed by heat shield elements  155  on its side facing the working medium M. Each heat shield element  155  is equipped with a particularly heat-resistant protective layer on the working medium side or is made of refractory material and comprises the protective layer  7  according to  FIG. 1 . 
         [0095]    Owing to the high temperatures inside the combustion chamber  110 , a cooling system is also provided for the heat shield elements  155  or for their retaining elements. 
         [0096]    The materials of the combustion chamber wall and its coatings may be similar to the turbine blades  120 ,  130 . 
         [0097]    The combustion chamber  110  is in particular designed in order to detect losses of the heat shield elements  155 . To this end, a number of temperature sensors  158  are positioned between the combustion chamber wall  153  and the heat shield elements  155 . 
         [0098]      FIG. 5  shows a perspective view of a blade  120 ,  130  which comprises a layer system I with the protective layer  7  according to the invention. 
         [0099]    The blade  120 ,  130  extends along a longitudinal axis  121 . 
         [0100]    The blade  120 ,  130  comprises, successively along the longitudinal axis  121 , a fastening zone  400 , a blade platform  403  adjacent thereto as well as a blade surface zone  406 . The protective layer  7  or a layer system  1  according to  FIG. 1  is formed particularly in the blade surface zone  406 . 
         [0101]    A blade root  183  which is used to fasten the rotor blades  120 ,  130  on the shaft, is formed in the fastening zone  400 . The blade root  183  is configured as a hammerhead. Other configurations are possible, for example as a firtree or dovetail root. In conventional blades  120 ,  130 , for example solid metallic materials are used in all regions  400 ,  403 ,  406  of the rotor blade  120 ,  130 . 
         [0102]    The rotor blade  120 ,  130  may in this case be manufactured by a casting method, by a forging method, by a machining method or combinations thereof.