Abstract:
A layer system is provided having a substrate, an outermost layer with a surface, at least one hole at least in the outermost layer, wherein in the vicinity around the hole or directly adjoining a boundary line of the hole, at least one, not closed extending recess is present in the surface of the outermost layer in which the recess is curved and its ends face each other. Through the use of depressions in a layer, spalling within the interfaces through the layers is prevented.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is the US National Stage of International Application No. PCT/EP2013/073001 filed Nov. 5, 2013, and claims the benefit thereof. The International Application claims the benefit of European Application No. EP12192931 filed Nov. 16, 2012. All of the applications are incorporated by reference herein in their entirety. 
     
    
     FIELD OF INVENTION 
       [0002]    The invention relates to the modification of a surface around a hole in a layer system. 
       BACKGROUND OF INVENTION 
       [0003]    Layer systems are used in particular for components operated at high temperatures. These are in particular turbine blades or vanes with a metallic substrate, metallic bonding layer and ceramic thermal barrier layer. 
         [0004]    In addition, gas turbine components in particular are cooled by a cooling medium flowing out of a cooling hole in order to cool the component on the inside or else in order to protect the component on the outside against excessively hot gases. 
         [0005]    Holes of this type are often made after complete coating of the substrate, in which case the opening can then be a flaw or starting point for crack growth on its inner face. 
       SUMMARY OF INVENTION 
       [0006]    It is therefore an object of the invention to solve the problem mentioned above. 
         [0007]    The object is achieved by one or more recesses around the hole as claimed in the independent claim. 
         [0008]    The dependent claims list further advantageous measures which can be combined with one another, as desired, in order to achieve further advantages. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0009]      FIGS. 1, 2  show a cooling hole according to the prior art, 
           [0010]      FIGS. 3-5  show exemplary embodiments of the invention, 
           [0011]      FIG. 6  shows a turbine blade or vane, 
           [0012]      FIG. 7  shows a combustion chamber, 
           [0013]      FIG. 8  shows a list of superalloys. 
       
    
    
     DETAILED DESCRIPTION OF INVENTION 
       [0014]    The description and the figures represent merely exemplary embodiments of the invention. 
         [0015]      FIG. 1  shows a plan view of a surface  42  of a layer  10  ( FIG. 2 ) having a hole  13 , which here in particular is in the form of a film-cooling hole. 
         [0016]    There is a contour  45  around the hole  13  on the surface  42 . The film-cooling hole  13  can have a radial bore  16  with a symmetrical or asymmetrical cross section. Depending on the application or location on the combustion chamber brick  155  ( FIG. 9 ) or the turbine blade or vane  120 ,  130  ( FIG. 8 ), the film-cooling hole  13  is formed with a diffuser  19 . 
         [0017]    The diffuser  19  constitutes a widening of the bottom portion  16  of the hole  13  ( FIG. 2 ). 
         [0018]      FIG. 2  shows a cross section through a layer system  25 . The layer system  25  comprises a substrate  4 . 
         [0019]    The substrate  4  is advantageously metallic and very particularly comprises nickel-based or cobalt-based superalloys. 
         [0020]    In this respect, use is advantageously made of alloys as shown in  FIG. 8 . 
         [0021]    An outer ceramic layer  10 , which has the outermost surface  42 , is applied to the substrate  4  directly or on a metallic bonding layer  7 . 
         [0022]    A hole  13 , which can also have a diffuser  19  on the surface  42 , is present continuously through the layer system  25 , i.e. through the layers  7 ,  10  and the substrate  4 . 
         [0023]      FIG. 3  shows, in plan view, the diffuser  19  or the hole  13 .  48  indicates the point in the interface between the ceramic layer and the substrate or the adhesion-promoting layer, whence cracks along the interface can start. 
         [0024]    For that reason, there is a curved recess  51  which begins upstream of the hole  13 , in an overflow direction  60 , and whose ends  54 ,  57  face each other, in particular approximately centrally at the level of the hole  13  or of the diffuser  19 . 
         [0025]    The diffuser  19  is not encircled by the recess  51 . 
         [0026]    The recess  51  is curved, advantageously tong-shaped or O-shaped. As seen in the overflow direction  60 , the ends  54 ,  57  of the recess  51  also represent the end, as seen in the overflow direction  60 , of the recess  51 . 
         [0027]    In this case, the recess  51  is round, oval or curved and can either not touch the hole  13  or the diffuser  19  ( FIG. 3 ) or touch and merge into this hole  13 , as is shown in  FIG. 5 , i.e. the recess  51  is interrupted only by the hole  13  or is at a small distance  60 ′,  60 ″ from the opening of the hole  13 . In this context, small means &lt;10% of the length of the recess  51 . 
         [0028]      FIG. 4  shows a cross section through  FIG. 2  or, respectively, also through  FIG. 5 , in which  48  indicates the region of the interface between the ceramic layer and the substrate  4  or adhesion-promoting layer  7 , which region begins at the hole  13  and the recess  51  which is located downstream of the region  48  as seen in the flow direction. 
         [0029]    The recess  51  thus runs over the hole  13 . 
         [0030]    In the event of spalling of the TBC  10  proceeding from the cooling air hole  13 , the TBC will break off only as far as the recess  51  and a crack will not propagate beyond the recess  51  as seen in the flow direction. 
         [0031]    Accordingly, the recess  51  extends largely over the thickness of the outermost layer  10 . 
         [0032]      FIG. 6  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  and 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, 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. 
         [0043]    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. 
         [0044]    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. 
         [0045]    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). 
         [0046]    Processes of this type are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1. 
         [0047]    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. 
         [0048]    The density is advantageously 95% of the theoretical density. 
         [0049]    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). 
         [0050]    The layer advantageously 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 advantageous to use nickel-based protective layers, such as Ni-10Cr-12A1-0.6Y-3Re or Ni-12Co-21Cr-11Al-0.4Y-2Re or Ni-25Co-17Cr-10Al-0.4Y-1.5Re. 
         [0051]    It is also possible for a thermal barrier layer, which is advantageously 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. 
         [0052]    The thermal barrier layer covers the entire MCrAlX layer. 
         [0053]    Columnar grains are produced in the thermal barrier layer by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD). 
         [0054]    Other coating processes are possible, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier layer 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 layer is therefore advantageously more porous than the MCrAlX layer. 
         [0055]    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. 
         [0056]    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). 
         [0057]      FIG. 9  shows a combustion chamber  110  of a gas turbine. 
         [0058]    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 . 
         [0059]    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 . 
         [0060]    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). 
         [0061]    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. 
         [0062]    A for example ceramic thermal barrier layer, 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. 
         [0063]    Columnar grains are produced in the thermal barrier layer by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD). 
         [0064]    Other coating processes are conceivable, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier layer may include grains that are porous or have micro-cracks or macro-cracks, in order to improve the resistance to thermal shocks. 
         [0065]    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). 
         [0066]    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. 
         [0067]    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 . 
         [0068]      FIG. 10  shows by way of example a partial longitudinal section through a gas turbine  100 . 
         [0069]    In its interior, the gas turbine  100  has a rotor  103  which is mounted such that it can rotate about an axis of rotation  102 , has a shaft  101 , and is also referred to as the turbine rotor. 
         [0070]    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 . 
         [0071]    The annular combustion chamber  110  is in communication with a for example annular hot gas duct  111 . There, by way of example, four successive turbine stages  112  form the turbine  108 . 
         [0072]    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 , a guide vane row  115  is followed in the hot gas duct  111  by a row  125  formed from rotor blades  120 . 
         [0073]    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 on the rotor  103 , for example by means of a turbine disk  133 . 
         [0074]    A generator (not shown) is coupled to the rotor  103 . 
         [0075]    While the gas turbine  100  is operating, air  135  is drawn in through the intake housing  104  and compressed by the compressor  105 . The compressed air provided at the turbine end of the compressor  105  is passed to the burners  107 , where it is mixed with a fuel. The mixture is then burnt in the combustion chamber  110 , forming the working medium  113 . From there, the working medium  113  flows along the hot gas duct  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. 
         [0076]    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. 
         [0077]    To be able to withstand the temperatures which prevail there, they can be cooled by means of a coolant. 
         [0078]    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). 
         [0079]    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 . 
         [0080]    Superalloys of this type are known for example from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 Al, WO 99/67435 or WO 00/44949. 
         [0081]    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. 
         [0082]    A thermal barrier layer, 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. 
         [0083]    Columnar grains are produced in the thermal barrier layer by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD). 
         [0084]    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 .