Abstract:
During the coating of a component including cooling-air holes, it is necessary, when remachining the coated cooling-air holes, to determine the correct alignment of the cooling-air row. A process is provided in that at least two cooling-air holes are kept clear by a masking material during the coating and the alignment of the cooling-air row is effected on the basis of these two unclosed cooling-air hole.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims priority of European Patent Office application No. 09014306.6 EP filed Nov. 16, 2009, which is incorporated by reference herein in its entirety. 
       FIELD OF INVENTION 
       [0002]    The invention relates to a process for coating a component which has holes, wherein not all the holes are closed during the coating, and to a process for opening said holes. 
       BACKGROUND OF INVENTION 
       [0003]    Components such as, for example, turbine blades or vanes of a gas turbine have cooling-air bores. Cooling-air bores of this type are made directly in the substrate before the further application of layers during the production of new components, or are already present in a component to be refurbished, in the case of which the coatings are replaced. Both cases are subject to the problem of “coat down”, i.e. the coating material penetrates into the cooling-air bores and undesirably closes them. 
         [0004]    There are processes for closing the cooling-air bores where complete coverage, in particular also during the coating process, is not always ensured. 
         [0005]    The subsequent removal of the “coat down” from the cooling-air bores is subject to the problem of correct alignment of the cooling-air row in the apparatus by means of which the coating is removed from the cooling-air bores, since the alignment differs from a desired geometry. 
       SUMMARY OF INVENTION 
       [0006]    It is therefore an object of the invention to solve the above-mentioned problem. 
         [0007]    The object is achieved by a process as claimed in the claims and by a process as claimed in the claims. 
         [0008]    The dependent claims list further advantageous measures which can be combined with one another, as desired, in order to obtain further advantages. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0009]      FIG. 1  shows a component having a row of cooling-air bores, 
           [0010]      FIGS. 2 ,  3 ,  4 ,  5 ,  6  show components for the coating process according to the invention, 
           [0011]      FIG. 7  shows a gas turbine, 
           [0012]      FIG. 8  shows a turbine blade or vane, 
           [0013]      FIG. 9  shows a combustion chamber, and 
           [0014]      FIG. 10  shows a list of superalloys. 
       
    
    
       [0015]    The drawings and the description which follows represent only exemplary embodiments of the invention. 
       DETAILED DESCRIPTION OF INVENTION 
       [0016]      FIG. 1  shows a component  1 ,  120 ,  130 ,  155  which, as an example of holes, has cooling-air bores  4  ( FIG. 1 ),  13 ′,  13 ″, . . . ,  15 ′,  15 ″ ( FIGS. 2-4 ) in a cooling-air row or row  10  of holes. 
         [0017]    The component  120 ,  130 ,  155  may have a plurality of such cooling-air rows  10 ′,  10 ″,  10 ″′ ( FIG. 5 ). 
         [0018]    A cooling-air row runs along a straight line ( FIG. 1 ) or a slightly bent straight line ( FIG. 5 ) or has a non-curved profile (not shown). 
         [0019]    By way of example, the component  1  in this case is a turbine blade or vane  120 ,  130  of a gas turbine  100  ( FIG. 7 ), the base material (substrate) of which contains a superalloy, in particular a nickel-based superalloy ( FIG. 10 ). 
         [0020]    The main blade or vane part  406  of the component  120 ,  130  has at least one row  10 ,  10 ′,  10 ″, . . . of cooling-air holes  4 ,  13 ′,  13 ″,  15 ′,  15 ″. 
         [0021]    In  FIG. 1 , only one cooling-air row  10  is shown (non-limiting). 
         [0022]    The cooling-air holes  4  in the cooling-air row  10  lie on a line  7 , in particular a straight line  7 , which is at a specific distance from the trailing or leading edge  412 ,  409 , or the holes  4  are aligned specifically with respect to one another. 
         [0023]    During the production of the holes  4  for producing the cooling-air row  10  in the component  1 ,  120 ,  130 ,  155 , undesired displacement of the dashed lines  412 ,  7  in relation to one another can occur. This is tolerated and represents no limitation whatsoever for the first use of the component  120 ,  130 . 
         [0024]    If, however, material has to be removed from the cooling-air holes  4 , for example during recoating or during refurbishment, the machining system, in particular that of a laser, assumes that line  412  and line  7  are ideally aligned. 
         [0025]    Since this is sometimes not the case, the machining system has to be reprogrammed in order to take the real situation into consideration. 
         [0026]    This is done by virtue of the fact that at least two cooling-air holes  13 ′,  13 ″, . . . in a cooling-air row  10  of cooling-air holes  4  are masked and closed before the coating. This can be effected in various ways, as described in the prior art. 
         [0027]    In this case, it is preferable for up to 50%, in particular up to 40%, of the cooling-air holes to be closed. 
         [0028]    The substrate  4  is then coated. This coating may be a metallic bonding layer (in particular MCrAlY) or a metallic bonding layer and a ceramic layer on top of the latter or only a single-ply, two-ply or multi-ply ceramic layer. 
         [0029]    Irrespective of whether one layer or two layers are applied, the unclosed cooling-air bores  15 ′,  15 ″ are coated and at least partially or completely closed. This coating material in the holes  15 ′,  15 ″ has to be removed after the coating process. However, the previously closed cooling-air holes  13 ′,  13 ″ have no or scarcely any coating material within the holes and do not have to be remachined. 
         [0030]    Before the remachining of the closed cooling-air bores  15 ′,  15 ″, . . . , the masking material is removed from the cooling-air bores  13 ′,  13 ″, . . . , in particular these bores are opened manually. An imaginary straight line  7  is then placed through the uncoated cooling holes. The machining system (not shown in more detail) detects the unclosed holes  13 ′,  13 ″ and is aligned with respect to a line  7  between the cooling-air bores  13 ′,  13 ″′, such that the machining system or, in particular, the laser knows how the line  7  of the cooling-air row  10  is defined independently with respect to the alignment of the line  412 . 
         [0031]    Therefore, it is always necessary to close and reopen at least two cooling-air bores  13 ′,  13 ″, so that a line  7  or a line with a known profile (length, angle of curvature) can be formed. It is immaterial which cooling-air holes are closed. These can be an upper and a central cooling-air hole or the topmost and bottommost cooling-air holes ( FIG. 4 ), or these can be two central cooling-air holes ( FIG. 3 ). 
         [0032]    These are preferably the topmost and the bottommost holes ( FIG. 6 ). 
         [0033]    At least one cooling-air hole  15 ′,  15 ″, in particular at least a plurality of unclosed cooling-air holes  15 ′,  15 ″, is or are preferably present between the at least two closed cooling-air holes  13 ′,  13 ″. 
         [0034]    It is likewise possible for two adjacent cooling-air bores  13 ′,  13 ″ to be closed, where a number of unclosed cooling-air bores  15 ′,  15 ″ then follows in the row  10  and then, in turn, one or two masked cooling-air bores  13 ′,  13 ″, which are closed before the coating. 
         [0035]      FIG. 5  shows a further exemplary embodiment of the invention. 
         [0036]    In this case, a cooling-air row  10  consists of a plurality of cooling-air portions  10 ′,  10 ″,  10 ′″, which are aligned differently with respect to the leading edge  409  or lie on a straight line and have different properties with respect to their diffuser geometry (each portion is virtually a cooling-air row). It is thus also possible, for example, for portions  10 ′,  10 ″,  10 ″′ to be present on a preferably straight line  10 , in which case the cooling-air hole in the respective portion  10 ′,  10 ″,  10 ″′ has a different size or a different orientation/diffuser geometry with respect to the overflow direction  19  or longitudinal axis  22  of the main blade or vane part  406 . 
         [0037]    For the preparation for the coating, a comparable situation applies to the procedure according to  FIG. 2 ,  3 ,  4  or  6 , i.e. here the respective pattern and specifications of  FIG. 2 ,  3 ,  4  or  6  are employed for each portion  10 ′,  10 ″,  10 ″′, i.e. at least two cooling-air holes  13 ′,  13 ″ of a portion  10 ′,  10 ″,  10 ″′ are closed, and are then opened after the coating. 
         [0038]    Within a cooling-air portion  10 ′,  10 ″,  10 ″′, the cooling-air holes  4  are preferably identical. The cooling-air portions  10 ′,  10 ″,  10 ″′ virtually represent a cooling-air row. 
         [0039]      FIG. 7  shows, by way of example, a partial longitudinal section through a gas turbine  100 . 
         [0040]    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. 
         [0041]    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 . 
         [0042]    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 . 
         [0043]    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 . 
         [0044]    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 . 
         [0045]    A generator (not shown) is coupled to the rotor  103 . 
         [0046]    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. 
         [0047]    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. 
         [0048]    To be able to withstand the temperatures which prevail there, they may be cooled by means of a coolant. 
         [0049]    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). 
         [0050]    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 . 
         [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. 
         [0052]    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 . 
         [0053]      FIG. 8  shows a perspective view of a rotor blade  120  or guide vane  130  of a turbomachine, which extends along a longitudinal axis  121 . 
         [0054]    The turbomachine may be a gas turbine of an aircraft or of a power plant for generating electricity, a steam turbine or a compressor. 
         [0055]    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 . 
         [0056]    As a guide vane  130 , the vane  130  may have a further platform (not shown) at its vane tip  415 . 
         [0057]    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 . 
         [0058]    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. 
         [0059]    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 . 
         [0060]    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 . 
         [0061]    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. 
         [0062]    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. 
         [0063]    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. 
         [0064]    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. 
         [0065]    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. 
         [0066]    Where the text refers in general teens 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). 
         [0067]    Processes of this type are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1. 
         [0068]    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, which are intended to form part of this disclosure with regard to the chemical composition of the alloy. 
         [0069]    The density is preferably 95% of the theoretical density. 
         [0070]    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). 
         [0071]    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. 
         [0072]    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. 
         [0073]    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). 
         [0074]    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. 
         [0075]    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). 
         [0076]      FIG. 9  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 . 
         [0077]    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 . 
         [0078]    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 . 
         [0079]    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). 
         [0080]    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, which are intended to form part of this disclosure with regard to the chemical composition of the alloy. 
         [0081]    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. 
         [0082]    Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD). 
         [0083]    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. 
         [0084]    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. 
         [0085]    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.