Abstract:
The invention relates to a metal protective coating which protects against corrosion and oxidation, in particular for components in turbines, said components being used at high temperatures. The advantageous composition is as follows: 11.5 20.0 wt % chrome, 0.3 1.5 wt % silicone, 0.0 1.0 wt % aluminium, 0,0-0.7 wt % yttrium and/or at least one metal from the group comprising Sc and rare earth elements, and the remainder being iron and production-related impurities.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS  
       [0001]     This application is the US National Stage of International Application No. PCT/EP2004/013661, filed Dec. 1, 2004 and claims the benefit thereof. The International Application claims the benefits of European Patent application No. 03028577.9 filed Dec. 11, 2003. All of the applications are incorporated by reference herein in their entirety. 
     
    
     FIELD OF THE INVENTION  
       [0002]     The invention relates to a metallic protective layer as described in the claims and a layer system as described in the claims.  
       BACKGROUND OF THE INVENTION  
       [0003]     Metallic protective layers for protecting a component, in particular a component which consists of a superalloy based on iron, nickel or cobalt, against corrosion and oxidation in particular at high temperatures, with the component, in particular a component of a steam or gas turbine, being exposed to a flue gas or the like at a high temperature, are generally known.  
         [0004]     Most of these protective layers are known under the collective name MCrAlX, where M represents at least one of the elements selected from the group consisting of iron, cobalt and nickel and further essential constituents are chromium, aluminum and X=yttrium, although the latter may also be partially or entirely replaced by an equivalent element selected from the group consisting of scandium and the rare earth elements.  
         [0005]     Typical coatings of this type are known from U.S. Pat. Nos. 4,005,989 and 4,034,142. Moreover, it is known from the latter patent that an additional silicon content can further improve the properties of protective layers of the type described above.  
         [0006]     Furthermore, EP-A 0 194 392 has disclosed numerous special compositions for protective layers with admixtures of further elements for various applications. In this context, the element rhenium as an admixture forming up to 10% by weight, as well as many other elements that can optionally be added, is mentioned. In view of the lack of more specific further ranges for possible admixtures, however, none of the protective layers indicated is qualified for special conditions, such as for example on rotor blades and guide vanes of steam or gas turbines with high inlet temperatures which have to be operated for prolonged periods of time.  
         [0007]     Protective layers which contain rhenium are also known from U.S. Pat. No. 5,154,885, EP-A 0 412 397, DE 694 01 260 T2 and WO 91/02108 A1. The overall disclosure revealed by these documents is incorporated in its entirety in the present disclosure.  
         [0008]     EP 0 253 754 91 reveals embodiments for applying a protective layer to a gas turbine component that is to be exposed to high thermal stresses.  
         [0009]     Efforts to increase the inlet temperatures of both stationary steam and gas turbines and aircraft engines are of considerable significance in the specialist field of gas turbines, since the inlet temperatures are important variables in determining the thermodynamic efficiencies which can be achieved by gas turbines. The use of specially developed alloys as base materials for components which are to be exposed to high thermal stresses, such as guide vanes and rotor blades, in particular the use of single-crystal superalloys, allows inlet temperatures of well over 1000° C. Nowadays, the state of the art allows inlet temperatures of 950° C. and more in the case of stationary gas turbines and 1100° C. and more in gas turbines of aircraft engines.  
         [0010]     Examples of the structure of a turbine blade or vane with a single-crystal substrate, which for its part may be of complex construction, are disclosed by WO 91/01433 A1.  
         [0011]     Whereas the physical load-bearing capacity of the by now highly developed base materials for the highly stressed components is substantially free of problems with regard to possible future increases in the inlet temperatures, to achieve a sufficient resistance to oxidation and corrosion it is necessary to employ protective layers. In addition to a protective layer being sufficiently chemically stable under the attacks expected from flue gases at temperatures of the order of 1000° C., a protective layer must also have sufficiently good mechanical properties, not least with regard to the mechanical interaction between the protective layer and the base material. In particular, the protective layer must be sufficiently ductile to be able to follow any deformation of the base material and not to crack, since this can give rise to points of attack for oxidation and corrosion. This typically presents the problem that increasing the levels of elements such as aluminum and chromium, which are able to improve the resistance of a protective layer to oxidation and corrosion, leads to a deterioration in the ductility of the protective layer, with the result that mechanical failure, in particular the formation of cracks, is likely under the mechanical stresses which customarily occur in a gas turbine. Examples of the ductility of the protective layer being reduced by the elements chromium and aluminum are known from the prior art.  
       SUMMARY OF THE INVENTION  
       [0012]     Accordingly, the invention is based on the object of providing a protective layer and a layer system which have a good high-temperature stability in corrosion and oxidation, a good long-term stability and, moreover, are especially well matched to mechanical stresses which are expected in particular in a steam or gas turbine at high temperature.  
         [0013]     To achieve this object, the invention provides a protective layer and a layer system comprising this protective layer for protecting a component against corrosion and oxidation at a high temperature, which substantially comprises the following elements (details of amounts in percent by weight): 
    11.5 to 20.0 wt % chromium,     0.3 to 1.5 wt % silicon,     to 1.0 wt % aluminum,     to 0.7 wt % yttrium and/or at least one metal selected from the group consisting of scandium and the rare earth elements, 
 
 remainder iron and production-related impurities. 
   
 
         [0018]     In particular, the metallic protective layer consists of 
    12.5 to 14.0 wt % chromium,     0.5 to 1.0 wt % silicon,     0.0 to 0.5 wt % aluminum,     0.0 to 0.7 wt % yttrium and/or at least one metal selected from the group consisting of scandium and the rare earth elements, 
 
 remainder iron and production-related impurities.
   
 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0023]     The invention is explained in more detail in the figures, in which: 
     FIGS. 1, 2  show examples of arrangements of the protective layer,      FIG. 3  shows a gas turbine,      FIG. 4  shows a combustion chamber, and      FIG. 5  shows a steam turbine.   
 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0028]      FIG. 1  shows an example of an arrangement of a metallic protective layer  7  of a layer system  1 .  
         [0029]     The metallic protective layer  7  is arranged on a substrate  4  and in this case forms the outer layer of the layer system  1 .  
         [0030]     In  FIG. 2 , the metallic protective layer  7  constitutes an intermediate layer in the layer system  1 .  
         [0031]     The metallic protective layer  7  is likewise arranged on a substrate  4 , but a further, for example ceramic layer  10  is also present on the metallic protective layer  7 .  
         [0032]     The protective layer  7  described also acts, for example, as a bonding layer for improving the bonding of the layer  10  to the substrate  4 .  
         [0033]     Other or further metallic and/or ceramic layers may be present.  
         [0034]     In particular an aluminum oxide layer may be applied to or produced on this layer  7 .  
         [0035]     The ceramic layer  10  is in particular a thermal barrier coating based on zirconium oxide. This may be partially or fully stabilized zirconium oxide. Further ceramic materials for the ceramic thermal barrier coating  10  are conceivable.  
         [0036]     Likewise conceivable are all coating processes for applying the metallic protective layer  7  and/or the ceramic layer  10  to the substrate  4  or to the metallic protective layer  7 .  
         [0037]     As has already been explained above, layer systems  1  of this type can be used for components in a gas turbine  100  ( FIG. 3 ) and in a steam turbine  300 ,  303  ( FIG. 5 ) or aircraft turbine.  
         [0038]     The layer systems  1  can be used for newly produced components or refurbished components.  
         [0039]     Highly stressed components, in particular turbine blades or vanes  354 ,  357 ,  366  ( FIG. 5 ),  120 ,  130  ( FIG. 3 ) are in many cases refurbished after use by the outer layers  7 ,  10  as well as further corrosion or oxidation layers being removed. The component (substrate  4 ) is also inspected for cracks, which are repaired if appropriate.  
         [0040]     The component (substrate  4 ) can then again be provided with a metallic protective layer  7  in order to form a layer system  1 .  
         [0041]     The protective layer  7  combines a good resistance to corrosion with a particularly high stability with respect to oxidation and is also distinguished by particularly good ductility properties, making it especially well qualified for use in a steam turbine in particular in the event of a further increase in the inlet temperature.  
         [0042]     The composition of the protective layer  7  based on iron has particularly good properties; in particular, the protective layer  7  can be very successfully applied to ferritic substrates  4 .  
         [0043]     In this case, the coefficients of thermal expansion a of substrate  4  and protective layer  7  can be very well matched, i.e. differences of up to 10% are possible, or are identical, so that there is no thermally induced build-up of stresses between substrate  4  and protective layer  7  (thermal mismatch), which could cause the protective layer  7  to flake off.  
         [0044]     Identical coefficients of thermal expansion means that the differences are at most such that no thermally induced stresses occur at the temperatures of use.  
         [0045]     This is particularly important since in the case of ferritic materials being used for the substrate  4 , it is often the case that there is no heat treatment carried out for the diffusion bonding of the layer  7  to the substrate  4 , since the ferritic substrate  4  has undergone a final heat treatment and should not be exposed to any further heat treatment close to or above the temperature of the final heat treatment (tempering treatment).  
         [0046]     The protective layer  7  is particularly suitable for protecting a ferritic component against corrosion and oxidation at temperatures of up to 800° C., in particular up to 650° C.  
         [0047]     The protective layer  7  bonds to the substrate  4  mostly or exclusively through adhesion.  
         [0048]     The thickness of the protective layer  7  on the component  1  is preferably set to between approximately 100 μm and 300 μm.  
         [0049]     The protective layer  7  is also particularly suitable for protecting a component against corrosion and oxidation while the component is exposed to a flue gas with the material at a temperature of around 950° C., or in the case of aircraft turbines even around 1100° C.  
         [0050]     The protective layer  7  according to the invention is therefore particularly suitable for protecting a component of a steam turbine  300 ,  303  ( FIG. 5 ) or gas turbine  100  ( FIG. 3 ), in particular a guide vane  130 , rotor blade  120  or other component (housing parts) which is exposed to hot steam or gas upstream or in the turbine part of the steam or gas turbine.  
         [0051]     The substrate  4  may be metallic or ceramic.  
         [0052]     In particular, the substrate  4  is a ferritic base alloy in the case of a steam turbine, a nickel-base or cobalt-base superalloy in the case of a gas turbine or a steel, in particular a 1% CrMoV steel or a 10% to 12% chromium steel.  
         [0053]     Further advantageous ferritic substrates  4  for the layer system  1  may consist of: 
    1% to 2% Cr steel for shafts ( 309 ,  FIG. 4 ):     such as for example 30CrMoNiV5-11 or 23CrMoNiWV8-8,     1% to 2% Cr steel for housings (for example  333 ,  FIG. 4 ): G17CrMoV5-10 or G17CrMo9-10     10% Cr steel for shafts ( 309 ,  FIG. 4 ):     X12CrMoWVNbN10-1-1     10% Cr steel for housings (for example  333 ,  FIG. 4 ):     GX12CrMoWVNbN10-1-1 or GX12CrMoVNbN9-1.    
 
         [0061]     Furthermore, the following composition is suitable as substrate  4  (details in percent by weight): 
    0.03 to 0.05% carbon     18 to 19% chromium     12 to 15% cobalt     3 to 6% molybdenum     1 to 1.5% tungsten     2 to 2.5% aluminum     3 to 5% titanium 
 
 optionally small amounts of tantalum, niobium, boron and/or zirconium, remainder nickel. 
   
 
         [0069]     Materials of this type are known as forging alloys under the names Udimet 520 and Udimet 720.  
         [0070]     Alternatively, the following composition is suitable for the substrate  4  of the component  1  (details in percent by weight): 
    0 to 0.15% carbon     18 to 22% chromium     18 to 19% cobalt     0 to 2% tungsten     0 to 4% molybdenum     0 to 1.5% tantalum     0 to 1% niobium     1 to 3% aluminum     2 to 4% titanium     0 to 0.75% hafnium 
 
 optionally small amounts of boron and/or zirconium, remainder nickel. 
   
 
         [0081]     Compositions of this type are known as casting alloys under the names GTD222, IN939, IN6203 and Udimet 500.  
         [0082]     Another alternative for the substrate  4  of the component  1  is the following composition (details in percent by weight): 
    0.07 to 0.1% carbon     12 to 16% chromium     8 to 10% cobalt     1.5 to 2% molybdenum     2.5 to 4% tungsten     1.5 to 5% tantalum     0 to 1% niobium     3 to 4% aluminum     3.5 to 5% titanium     0 to 0.1% zirconium     0 to 1% hafnium 
 
 optionally a small amount of born, remainder nickel. 
   
 
         [0094]     Compositions of this type are known as casting alloys PWA1483SX, IN738LC, GTD111, IN792CC and IN792DS; the material IN738LC is considered particularly preferred.  
         [0095]     The following composition is considered a further alternative for the substrate  4  of the component  1  (details in percent by weight): 
    approximately 0.25% carbon     24 to 30% chromium     10 to 11% nickel     7 to 8% tungsten     0 to 4% tantalum     0 to 0.3% aluminum     0 to 0.3% titanium     0 to 0.6% zirconium    
 
         [0104]     optionally a small amount of boron, remainder cobalt.  
         [0105]     Compositions of this type are known as casting alloys under the names FSX414, X45, ECY768 and MAR-M-509.  
         [0106]      FIG. 3  shows, by way of example, a partial longitudinal section through a gas turbine  100 .  
         [0107]     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 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  106 , 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  106  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 .  
         [0108]     The guide vanes  130  are secured to the stator  143 , whereas the rotor blades  120  of a row  125  are fitted to the rotor  103  by means of a turbine disk  133 . A generator (not shown) is coupled to the rotor  103 .  
         [0109]     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.  
         [0110]     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 bricks which line the annular combustion chamber  106 , are subject to the highest thermal stresses. To be able to withstand the temperatures which prevail there, they have to be cooled by means of a coolant. The blades or vanes  120 ,  130  may also have above-described protective layers  7  protecting against corrosion (MCrAlX; M═Fe, Co, Ni, X═Y, rare earths) and heat (thermal barrier coating, for example ZrO 2 , Y 2 O 4 -ZrO 2 ).  
         [0111]     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 .  
         [0112]      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  102  arranged circumferentially around the turbine shaft  103  open out into a common combustion chamber space. For this purpose, the combustion chamber  110  overall is of annular configuration positioned around the turbine shaft  103 .  
         [0113]     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  is equipped with a particularly heat-resistant protective layer or is made from material that is able to withstand high temperatures. Moreover, on account of the high temperatures in the interior of the combustion chamber  110 , a cooling system is provided for the heat shield elements  155  and/or for their holding elements.  
         [0114]     The materials of the combustion chamber wall and their coatings may be similar to the turbine blades or vanes.  
         [0115]      FIG. 5  illustrates, by way of example, a steam turbine  300 ,  303  with a turbine shaft  309  extending along an axis of rotation  306 .  
         [0116]     The steam turbine has a high-pressure part-turbine  300  and an intermediate-pressure part-turbine  303 , each with an inner casing  312  and an outer casing  315  surrounding it. The high-pressure part-turbine  300  is, for example, of pot-type design. The intermediate-pressure part-turbine  303  is of two-flow design. It is also possible for the intermediate-pressure part-turbine  303  to be of single-flow design. Along the axis of rotation  306 , a bearing  318  is arranged between the high-pressure part-turbine  300  and the intermediate-pressure part-turbine  303 , the turbine shaft  309  having a bearing region  321  in the bearing  318 . The turbine shaft  309  is mounted on a further bearing  324  next to the high-pressure part-turbine  300 . In the region of this bearing  324 , the high-pressure part-turbine  300  has a shaft seal  345 . The turbine shaft  309  is sealed with respect to the outer casing  315  of the intermediate-pressure part-turbine  303  by two further shaft seals  345 . Between a high-pressure steam inflow region  348  and a steam outlet region  351 , the turbine shaft  309  in the high-pressure part-turbine  300  has the high-pressure rotor blading  354 ,  357 . This high-pressure rotor blading  354 ,  357 , together with the associated rotor blades (not shown in more detail), constitutes a first blading region  360 . The intermediate-pressure part-turbine  303  has a central steam inflow region  333 . Assigned to the steam inflow region  333 , the turbine shaft  309  has a radially symmetrical shaft shield  363 , a cover plate, on the one hand for dividing the flow of steam between the two flows of the intermediate-pressure part-turbine  303  and also for preventing direct contact between the hot steam and the turbine shaft  309 . In the intermediate-pressure part-turbine  303 , the turbine shaft  309  has a second blading region  366  comprising the intermediate-pressure rotor blades  354 ,  342 . The hot steam flowing through the second blading region  366  flows out of the intermediate-pressure part-turbine  303  from an outflow connection piece  369  to a low-pressure part-turbine (not shown) which is connected downstream in terms of flow.  
         [0117]     The turbine shaft  309  is composed of two turbine part-shafts  309   a  and  309   b , which are fixedly connected to one another in the region of the bearing  318 . The blades or vanes  354 ,  357 ,  366 , shafts  309  or other housing parts  333  may have above-described protective layers  7 ,  10  protecting against corrosion (MCrAlX; M═Fe, X═Y, Si, rare earths) and heat (thermal barrier coating, for example ZrO 2 , Y 2 O 4 -ZrO 2 ).