Patent Abstract:
Bonding agent layers are often used in heat insulation layers in order to improve the bonding of an outer ceramic layer to a metal substrate. A process is provided wherein a MCrAlX or MCrAl alloy is applied to a substrate whereby an outer layer region within the layer is produced using a heat treatment.

Full Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is the US National Stage of International Application No. PCT/EP2009/054127, filed Apr. 7, 2009 and claims the benefit thereof. The International Application claims the benefits of European Patent Office application No. 08009023.6 EP filed May 15, 2008. All of the applications are incorporated by reference herein in their entirety. 
     
    
     FIELD OF INVENTION 
       [0002]    The invention relates to a process for producing a bonding layer and to a layer system according to the claims. 
       BACKGROUND OF INVENTION 
       [0003]    In thermal barrier coating systems, use is often made of a metallic bonding layer in order to improve the bond between the outer ceramic thermal barrier coating and the metallic substrate. Single-layer MCrAlX systems or even recently roughened two-layer MCrAlX systems are often used as bonding layers. In this case, the outer MCrAlX layer has a different structure, which contributes in particular to an improvement in resistance to oxidation and corrosion. Said second MCrAlX layer is applied separately, which represents an additional process step and also entails bonding problems. The desired phase of the outer layer cannot always be controlled precisely. 
       SUMMARY OF INVENTION 
       [0004]    It is therefore an object of the invention to solve the above-mentioned problem. 
         [0005]    The object is achieved by a process as claimed in the claims, in which only a single-layer system is applied but is converted into a two-layer system by a heat treatment, and by a layer system as claimed in the claims. 
         [0006]    The dependent claims each list further advantageous measures which can be combined with one another, as desired, in order to obtain further advantages. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0007]      FIG. 1  schematically shows the sequence of the process, 
           [0008]      FIGS. 2 ,  3  show exemplary uses of the layer system produced in this way, 
           [0009]      FIG. 4  shows a gas turbine, 
           [0010]      FIG. 5  shows a perspective view of a turbine blade or vane, 
           [0011]      FIG. 6  shows a perspective view of a combustion chamber, and 
           [0012]      FIG. 7  shows a list of superalloys. 
       
    
    
       [0013]    The figures and the description represent only exemplary embodiments. 
       DETAILED DESCRIPTION OF INVENTION 
       [0014]      FIG. 1  schematically shows the sequence of the process. 
         [0015]    The layer system  1  has a substrate  4  and a metallic layer  7  made of an MCrAlX or MCrAl alloy (M=Ni and/or Co). 
         [0016]    In particular in the case of the component  120 ,  130 ,  155  ( FIGS. 5 ,  6 ) of a gas turbine  100  ( FIG. 4 ), the substrate  4  consists of a superalloy according to  FIG. 7 . 
         [0017]    An MCrAl or MCrAlX layer  7  applied by APS, LPPS, VPS, HVOF or other coating processes is present on the nickel- or cobalt-based superalloy. 
         [0018]    X is preferably yttrium (X═Y) and M is preferably Ni and Co. 
         [0019]    According to the invention, only one coating operation of the layer  7  takes place with only one powder type. 
         [0020]    By virtue of a heat treatment (T) at 1000° C.-1200° C., preferably 1140° C.-1180° C., preferably in a vacuum, the chromium in the MCrAl or MCrAlX alloy evaporates, such that a different chemical composition (reduced chromium content) is present in the outermost layer region  8 ′. 
         [0021]    The duration of the heat treatment is two to eight hours. It is preferable for a different phase to also form, and it is very preferable for a β-NiAl layer to form. The heat treatment is preferably carried out for an accordingly long time. 
         [0022]    If appropriate, a second heat treatment which can be distinguished from the chromium evaporation is carried out, in order to carry out the phase transformations of Ni—Al, Ni—Al—Cr, Ni—Al—Co, Ni—Al—Cr—Co to β-NiAl. 
         [0023]    The layer  7 ′ which is changed in this way thus consists of an outer layer region  8 ′ with a reduced chromium content, preferably of a β-NiAl phase, and an unchanged lower layer region  8 , which has the same composition as the originally applied layer  7  but is thinner (thickness of  8 ′ thickness of  7 ′ or thickness ( 8 + 8 ′)=thickness ( 7 ) or thickness ( 8 + 8 ′)=thickness ( 7 ′)). 
         [0024]    This heat treatment has two advantages. 
         [0025]    On the one hand, a homogeneous single-phase structure is formed on the surface. On the other hand, a homogeneous oxide layer with very small spinel fractions and very small fractions of nickel and/or chromium oxides is formed at high temperatures. The oxide layer thus formed is the starting point for a further homogeneous, thermally grown oxide layer  10  (TGO) ( FIGS. 2 ,  3 ). 
         [0026]    For use as the layer system  1 , oxidation can be brought about intentionally or the oxide layer  10  forms during the application of a ceramic outer thermal barrier coating  13  ( FIG. 3 ). 
         [0027]    The layer  7 ′ may likewise be used as an overlay layer, i.e. it forms the outermost layer with the exception of the TGO layer  10  which forms thereon. 
         [0028]    The layer region  8 ′ has therefore not been applied by a second coating operation or not by the change of the powder (from MCrAl to NiAl powder) during the coating operation, and therefore also bonds well to the underlying layer region  8 . 
         [0029]      FIG. 4  shows, by way of example, a partial longitudinal section through a gas turbine  100 . 
         [0030]    In the interior, the gas turbine  100  has a rotor  103  with a shaft which is mounted such that it can rotate about an axis of rotation  102  and is also referred to as the turbine rotor. 
         [0031]    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 . 
         [0032]    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 . 
         [0033]    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 . 
         [0034]    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 . 
         [0035]    A generator (not shown) is coupled to the rotor  103 . 
         [0036]    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. 
         [0037]    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. 
         [0038]    To be able to withstand the temperatures which prevail there, they may be cooled by means of a coolant. 
         [0039]    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). 
         [0040]    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 . 
         [0041]    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. 
         [0042]    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 . 
         [0043]      FIG. 5  shows a perspective view of a rotor blade  120  or guide vane  130  of a turbomachine, which extends along a longitudinal axis  121 . 
         [0044]    The turbomachine may be a gas turbine of an aircraft or of a power plant for generating electricity, a steam turbine or a compressor. 
         [0045]    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  and a blade or vane tip  415 . 
         [0046]    As a guide vane  130 , the vane  130  may have a further platform (not shown) at its vane tip  415 . 
         [0047]    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 . 
         [0048]    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. 
         [0049]    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 . 
         [0050]    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 . 
         [0051]    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. 
         [0052]    The blade or vane  120 ,  130  may in this case be produced by a casting process, by means of directional solidification, by a forging process, by a milling process or combinations thereof. 
         [0053]    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. 
         [0054]    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. 
         [0055]    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. 
         [0056]    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). 
         [0057]    Processes of this type are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1. 
         [0058]    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. 
         [0059]    The density is preferably 95% of the theoretical density. 
         [0060]    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). 
         [0061]    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. 
         [0062]    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. 
         [0063]    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). 
         [0064]    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. 
         [0065]    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). 
         [0066]      FIG. 6  shows a combustion chamber  110  of the gas turbine  100 . 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 , 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 . 
         [0067]    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 . 
         [0068]    Moreover, a cooling system may 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) opening out into the combustion chamber space  154 . 
         [0069]    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). 
         [0070]    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. 
         [0071]    It is also possible for a, for example, ceramic thermal barrier coating to 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. 
         [0072]    Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD). 
         [0073]    Other coating processes are possible, e.g. 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. 
         [0074]    Refurbishment means that after they have been used, protective layers may have to be removed from turbine blades or vanes  120 ,  130  or 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 turbine blade or vane  120 ,  130  or in the heat shield element  155  are also repaired. This is followed by recoating of the turbine blades or vanes  120 ,  130  or heat shield elements  155 , after which the turbine blades or vanes  120 ,  130  or the heat shield elements  155  can be reused.

Technology Classification (CPC): 8