Abstract:
Conventionally coated components with film cooling holes are known, comprising a diffuser, extending through the layers into the substrate. According to the invention, the component is embodied such that the whole diffuser is largely arranged in the layer.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation of U.S. Ser. No. 11/918,302 filed on Oct. 11, 2007 now U.S. Pat. No. 8,157,526. This application is the US National Stage of International Application No. PCT/EP2006/060794, filed Mar. 16, 2006 and claims the benefit thereof. The International Application claims the benefits of European application No. 05007993.8 filed Apr. 12, 2005, all of the applications are incorporated by reference herein in their entirety. 
    
    
     FIELD OF INVENTION 
     The invention relates to a component having a film cooling hole according to the claims. 
     BACKGROUND OF THE INVENTION 
     Components for applications at high temperatures consist of a superalloy with additional protection against oxidation, corrosion and high temperatures. To this end, the substrate of the component comprises a corrosion protection layer on which, for example, an outer ceramic thermal insulation layer is also applied. 
     Through-holes, out of which a coolant flows on the outer surface and contributes to the film cooling, are also made in the substrate and the layers for additional cooling. The film cooling hole is widened in the vicinity of the outer surface to form a so-called diffuser. When newly producing a component having a film cooling hole, problems arise since the diffuser must be made both through the layers and for the most part in the substrate. During the refurbishment of components, the problem is that the through-hole is already present and the substrate needs to be recoated, so that coating material must subsequently be removed from the diffuser region in the through-hole. 
     U.S. Pat. No. 4,743,462 discloses a method for closing a film cooling hole, in which a plug consisting of a pin and a spherical head is inserted into the film cooling hole. A bell-shaped indentation is thereby produced inside the coating. The indentation does not serve as a diffuser, however, since it is symmetrically designed. 
     The functionality of the head furthermore consists in the material of the head evaporating during the coating. It is not therefore possible to produce accurate, reproducible indentations for a multiplicity of film cooling holes. 
     Similar symmetrical widening of a film cooling hole is disclosed in FIG. 3 of U.S. Pat. No. 6,573,474. 
     EP 1 350 860 A1 discloses a method for masking a film cooling hole. The material of the masking means is selected so that no coating material is deposited there during the subsequent coating. An accurate, reproducible shape of the indentations inside a layer cannot be produced in this case. Furthermore, a diffuser is not described here. 
     EP 1 091 090 A2 discloses a film cooling hole in which a groove is made in the layer, so that the groove extends along a plurality of film cooling holes. Neither the film cooling holes nor the groove have a diffuser region. 
     U.S. Pat. No. 5,941,686 discloses a layer system, in which the substrate is processed. A diffuser region is not disclosed. 
     EP 1 076 107 A1 discloses a method for masking film cooling holes in which a plug, which protrudes from the hole, is respectively produced in the film cooling hole. To this end air is blown through the film cooling hole in a first step and a coating is applied, a precursor for the plug to be produced subsequently being introduced into the film cooling hole and into the coating. That part of the plug which is arranged inside the temporary layer has its shape determined by how strongly a medium is blown through the film cooling hole and how the coating of the temporary layer is carried out. The shape of that part of the plug which protrudes from the hole is therefore not reproducible. 
     SUMMARY OF INVENTION 
     It is therefore an object of the invention to overcome this problem. 
     The object is achieved by a component as claimed in the claims. 
     Further advantageous measures, which may arbitrarily be combined with one another in an advantageous way, are listed in the dependent claims. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIGS. 1 to 6  show exemplary embodiments of a component according to the invention having a film cooling hole, 
         FIGS. 7 ,  8  show a plan view of a film cooling hole according to the invention, 
         FIGS. 9 to 13  show configurations of a film cooling hole, 
         FIG. 14  shows a turbine blade, 
         FIG. 15  shows a combustion chamber, 
         FIG. 16  shows a gas turbine. 
     
    
    
     DETAILED DESCRIPTION OF INVENTION 
       FIG. 1  shows a component  1 ,  120 ,  130 ,  138 ,  155  consisting of a substrate  4  and a single outer layer  7 . 
     Particularly for components  120 ,  130 ,  138 ,  155  for turbines, the substrate  4  is a superalloy based on iron, nickel and/or cobalt. The outer layer  7  is preferably a corrosion and/or oxidation layer based on an MCrAlX alloy ( FIG. 15 ). 
     It may however also be ceramic. 
     The substrate  4  and the layer  7  comprise at least one film cooling hole  28  which, on the side  22  which is hot under operational conditions of use, comprises a diffuser  13  which departs from the e.g. cylindrical, square or generally speaking symmetrical contour  49  of the lower part  24  of the film cooling hole  28  near a cooling reservoir  31  and increases in cross section. 
     The film cooling hole  28  thus consists of a lower part  24  and the outer diffuser  13 . The diffuser  13  has an outlet opening  58 , over which a hot gas flows in an overflow direction  37 . 
     The diffuser  13  is formed from an imaginary extension  12  of the contour  49  as far as the surface  25  and an appendage  14  ( FIG. 2 ), which adjoins one or more side faces of the extension  12 . 
     In the cross-sectional view of  FIG. 1 , the appendage  14  preferably has a wedge shape. 
     In the plane of the outer surface, the diffuser  13  thus does not have rotational symmetry, the centroid of the asymmetric shape being displaced in the overflow direction  37  from the centroid of the symmetric shape of the contour  49 . 
     Along the normal  27  to the outer surface  25 , that cross-sectional area of the film cooling hole  28  which is perpendicular to the normal  27  becomes greater, i.e. the diffuser  13  is fully or preferably partially designed with a funnel shape. 
     According to the invention, the diffuser  13  is arranged for the most part inside the single layer  7 , i.e. when the diffuser  13  extends with an overall length  19  into the depth along a normal  27  of the component  1  which is perpendicular to the outer surface  25  or perpendicular to the overflow direction  37 , then there is a substrate length  16  of the diffuser  13  which constitutes the proportion of the diffuser  13  in the substrate  4 . The substrate length  16  is designed to be significantly less than the overall length  19 . The overall coating thickness  26  (here that of the layer  7 ) forms the remaining part of the overall length  19  of the diffuser  13 . The coating thickness  26  is at least 50%, preferably at least 60% or at least 70%, in particular 80% or 90% of the overall length  19 . 
     As an alternative, the diffuser  13  may be arranged entirely in the single layer  7  ( FIG. 3 , layer thickness  26 =overall length  19 ). 
     In  FIG. 4 , there are two layers on the substrate  4 . 
     These are in turn a corrosion and oxidation protection layer  7 , on which an outer ceramic thermal insulation layer  10  is also applied. 
     As in  FIG. 1 , there are lengths  16 ,  19  of the diffuser  13 , the layer thickness  26  again constituting at least 50%, 60% or in particular 70%, in particular 80% or 90% of the overall length  19 . 
     The diffuser  13  may likewise be arranged entirely in the two layers  7 ,  10  ( FIG. 5 ). 
     Correspondingly as for the two layers according to  FIGS. 4 ,  5 , this also applies for three or more layers. 
     The fact that the diffuser  13  is arranged for the most part or entirely in the layers  7 ,  10  provides advantages for refurbishing the component  1 , for example in respect of laser erosion or removal of material, above the lower part  24 , which covers the outlet opening  58  after recoating of the component  1 , specifically in that the laser or other coating apparatus only needs to be adjusted for the material of the layers  7 ,  10  and processing of the other material, i.e. that of the substrate  4 , does not need to be taken into account. 
       FIG. 6  shows a cross section through a component  1  having a film cooling hole  28 . 
     The substrate  4  comprises an outer surface  43 , on which the at least one layer  7 , is applied. 
     The diffuser  13  is for example arranged for the most part (according to  FIGS. 1 ,  3 ,  4 ,  5 ) in the layer  7 ,  10 , although it may also exist entirely in the substrate  4  or for the most part in the substrate  4 . 
     The lower part  24  of the film cooling hole  28  comprises for example a symmetry line  46  in longitudinal section. 
     The symmetry line  46  also constitutes for example an outflow direction  46  for a coolant, which flows through the cooling hole  28 . 
     A contour line  47 , which extends parallel to the symmetry line  46  on the inner side of the film cooling hole  28  or represents a projection of the symmetry line  46  onto the inner side of the lower part  24  of the film cooling hole  28 , makes an acute angle α1 with the outer surface  43 , which is in particular 30°+/−10%. The film cooling hole  28  is thus inclined in the overflow direction  37 . 
     The edge length a 28  ( FIG. 8 ) or the diameter φ 28  of the film cooling hole  28  is for example about 0.62 mm or 0.7 mm for a rotor blade and about 0.71 mm or 0.8 mm for guide vanes. 
     The contour line  47 , which preferably extends parallel to the outflow direction  46  along the contour  49  of the lower part  24 , makes an acute angle α2 with a diffuser line  48  which extends on the inner face  50  of the appendage  14  of the diffuser  13 , and which represents a projection of the overflow direction  37  onto the inner face  50  of the appendage  14  of the diffuser  13 . 
     The angle α2 is in particular 10°+/−10%. 
     Along the symmetry line  46 , the lower part  24  has a constant cross section which comprises in particular n-fold rotational symmetry (square, rectangular, round, oval, . . . ). 
     The diffuser  13  is created by the cross-sectional area of the film cooling hole  28  widening, i.e. being designed with a funnel shape in cross section. The appendage  14  to the contour  49  does not necessarily extend entirely around the outlet opening  58  of the film cooling hole  28 , rather only partially, in particular over half or less of the circumference of the outlet opening  58 . 
     The diffuser  13  is preferably arranged only—as seen in the overflow direction  37  of the hot gas  22 —in the rear region of the opening  58  ( FIG. 7 ). Side lines  38  of the diffuser  13  or of the appendage  14  extend for example parallel to the overflow direction  37  in plan view ( FIG. 7 ). 
     The overall layer thickness of the at least one layer  7 ,  10  is from about 400 μm to 700 μm, in particular 600 μm. 
       FIG. 8  shows another configuration of the film cooling hole  28  and a plan view of the diffuser  13  in the plane of the outer surface  25  of the layer system or component  1 . 
     The appendage  14  has, for example, a trapezoidal shape in the plane of the outer surface  25 . 
     In the plane of the surface  25 , the appendage  14  of the diffuser  13  has a longitudinal length l 1  of preferably about 3 mm in the overflow direction  37 . 
     The greatest width i.e. the greatest transverse length l 2  of the diffuser  13  in the surface, i.e. measured perpendicularly to the overflow direction  37 , preferably has a size of 2+−0.2 mm for rotor blades and a size of 4+−0.2 mm for guide vanes, and is at most 8 mm. 
     In the exemplary embodiment of  FIG. 8 , the widening of the diffuser  13  begins on a widening front edge  62 , i.e. at the appendage  14 , and widens in the overflow direction  37 . 
     The overflow direction  37  makes an acute angle α3, in particular 10°+/−10%, with a lateral delimiting line  38  of the appendage  14  in the plane of the outer surface  25 . 
     The diffuser  13  preferably widens departing from the contour  49  of the lower part  24 , which is for example symmetrical with respect to two mutually perpendicular axes, transversely to the flow direction  37  in each case by an angle α3, which is in particular 10°+/−10%, in which case the widening already begins on a leading edge  61  (as seen in the overflow direction  37 ) of the film cooling hole  28  and extends as far as the trailing edge  64 . 
     The diffuser  13  therefore has a trapezoidal cross section in the plane of the surface  25  ( FIG. 9 ). 
     The diffuser  13  is produced by a material erosion method, for example electron bombardment or laser irradiation. Only in this way can a multiplicity of cooling holes be produced accurately and reproduced. 
       FIGS. 10 ,  11 ,  12  and  13  show various contours of the film cooling hole  28 . 
     The lower part  24  of the film cooling hole  28  is designed to be cuboid here, merely by way of example, although it may also have a round or oval cross-sectional shape. 
     The diffuser  13  in  FIG. 10  is lengthened for example only in the overflow direction  37 , so that the cross section of the outlet opening  58  is greater than the cross section of the lower part  24 . The film cooling hole  28  thus corresponds to the film cooling hole according to  FIG. 2 ,  6  or  7 . 
     Based on  FIG. 10 ,  FIG. 11  represents a film cooling hole  28  which is also widened in the overflow direction  37  transversely to the overflow direction  37 , i.e. it corresponds to  FIG. 8 . 
     The diffuser  13  in  FIG. 12  is lengthened for example only transversely to the overflow direction  37 , so that here again the cross section of the outlet opening  58  is greater than the cross section of the lower part  24 . 
     The film cooling hole  28  consists for example of a cuboid lower part  24 , which is adjoined by a diffuser  13  in the form of a hexahedron with parallel trapezoidal side faces. 
     The diffuser  13  in  FIG. 13  is widened both only in the overflow direction  37  and in both directions transversely to the overflow direction  37 . 
       FIGS. 6 ,  7 ,  8 ,  9 ,  10 ,  11  and  13  respectively show that the diffuser  13  is for the most part arranged behind the outlet opening  58 , as seen in the overflow direction  37 . 
     This means that the diffuser  13  is formed by an asymmetric widening as seen in the overflow direction  37 . Uniform widening of the cross section of the lower part  24  of the film cooling hole  28  at the level of the outer surface  25  is not desired. 
     It can be seen clearly in  FIG. 6 , and is correspondingly described, that the appendage  14  represents a widening of the cross section in the overflow direction  37  so that the diffuser is formed. This is also shown by the plan view of  FIG. 6  according to  FIG. 7 . 
     In  FIG. 8 , the widening of the aperture of the cross section of the film cooling hole in the overflow direction  37  begins from the line  62 . 
     In  FIG. 9 , the widening of the diffuser  13  already begins on the leading edge  61  as seen in the overflow direction  37 . 
     Widening of the cross section of the film cooling hole  28  at the level of the outer surface  25  against the flow direction  37 , i.e. before the leading edge  61 , is not present or is present only to a small extent compared with the widening of the cross section in the overflow direction  37 . 
       FIG. 14  shows a perspective view of a rotor blade  120  or guide vane  130  of a turbomachine, which extends along a longitudinal axis  121 . 
     The turbomachine may be a gas turbine of an aircraft or of a power plant for electricity generation, a steam turbine or a compressor. 
     Successively along the longitudinal axis  121 , the blade  120 ,  130  comprises a fastening region  400 , a blade platform  403  adjacent thereto and a blade surface  406 . 
     As a guide vane  130 , the vane  130  may have a further platform (not shown) at its vane tip  415 . 
     A blade root  183 , which is used to fasten the rotor blades  120 ,  130  on a shaft or a disk (not shown), is formed in the fastening region  400 . 
     The blade root  183  is configured, for example, as a hammerhead. Other configurations as a firtree or dovetail root are possible. 
     The blade  120 ,  130  comprises a leading edge  409  and a trailing edge  412  for a medium which flows past the blade surface  406 . 
     In conventional blades  120 ,  130 , for example, solid metallic materials, in particular superalloys, are used in all regions  400 ,  403 ,  406  of the blade  120 ,  130 . 
     Such superalloys 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 are part of the disclosure in respect of the chemical composition of the alloy. 
     The blades  120 ,  130  may in this case be manufactured by a casting method, also by means of directional solidification, by a forging method, by a machining method or combinations thereof. 
     Workpieces with a monocrystalline structure or structures are used as components for machines which are exposed to heavy mechanical, thermal and/or chemical loads during operation. 
     Such monocrystalline workpieces are manufactured, for example, by directional solidification from the melt. These are casting methods in which the liquid metal alloy is solidified to form a monocrystalline structure, i.e. to form the monocrystalline workpieces, or directionally. 
     Dendritic crystals are in this case aligned along the heat flux and form either a rod-crystalline grain structure (columnar, i.e. grains which extend over the entire length of the workpiece and in this case, according to general terminology usage, are referred to as directionally solidified) or a monocrystalline structure, i.e. the entire workpiece consists of a single crystal. It is necessary to avoid the transition to globulitic (polycrystalline) solidification in this method, since nondirectional growth will necessarily form transverse and longitudinal grain boundaries which negate the good properties of the directionally solidified or monocrystalline component. 
     When directionally solidified structures are referred to in general, this is intended to mean both single crystals which have no grain boundaries or at most small-angle grain boundaries, and also rod-crystal structures which, although they do have grain boundaries extending in the longitudinal direction, do not have any transverse grain boundaries. These latter crystalline structures are also referred to as directionally solidified structures. 
     Such methods are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1; these documents are part of the disclosure. 
     The blades  120 ,  130  may likewise comprise coatings against corrosion or oxidation, for example (MCrAlX; M is at least one element from the group 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, for example hafnium (Hf)). Such alloys are known, for example, 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 be part of this disclosure in respect of the chemical composition of the alloy. 
     On the MCrAlX, there may also be a thermal insulation layer which consists for example of ZrO 2 , Y 2 O 3 —ZrO 2 , i.e. it is non-stabilized or partially or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide. 
     Rod-shaped grains are generated in the thermal insulation layer by suitable coating methods, for example electron beam deposition (EB-PVD). 
     Refurbishment means that components  120 ,  130  may need to have protective layers removed from them after their use (for example by sandblasting). Corrosion and/or oxidation layers or products are then removed. Optionally, cracks in the component  120 ,  130  will also be repaired. The component  120 ,  130  is then recoated and the component  120 ,  130  is used again. 
     The blade  120 ,  130  may be designed to be a hollow or solid. If the blade  120 ,  130  is intended to be cooled, it will be hollow and optionally also comprise film cooling holes  418  (represented by dashes). 
       FIG. 15  shows a combustion chamber  110  of a gas turbine  100 . 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 a rotation axis  102 , which produce flames  156 , open into a common combustion chamber space  154 . To this end, the combustion chamber  110  in its entirety is designed as an annular structure which is positioned around the rotation axis  102 . 
     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  made of an alloy is equipped with a particularly heat-resistant protective layer on the working medium side (MCrAlX layer and/or ceramic coating), or is made of refractory material (solid ceramic blocks). 
     These protective layers may be similar to the turbine blades, i.e. for example MCrAlX means: M is at least one element from the group 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, for example hafnium (Hf). Such alloys are known, for example, 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 be part of this disclosure in respect of the chemical composition of the alloy. 
     On the MCrAlX, there may also be an e.g. ceramic thermal insulation layer which consists for example of ZrO 2 , Y 2 O 3 —ZrO 2 , i.e. it is non-stabilized or partially or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide. 
     Rod-shaped grains are generated in the thermal insulation layer by suitable coating methods, for example electron beam deposition (EB-PVD). 
     Refurbishment means that heat shield elements  155  may need to have protective layers removed from them after their use (for example by sandblasting). Corrosion and/or oxidation layers or products are then removed. Optionally, cracks in the heat shield element  155  will also be repaired. The heat shield elements  155  are then recoated and the heat shield elements  155  are used again. 
     Owing to the high temperatures inside the combustion chamber  110 , a cooling system is also provided for the heat shield elements  155  or their holding elements. The heat shield elements  155  are then for example hollow and optionally also comprise cooling holes (not shown) opening into the combustion chamber space  154 . 
       FIG. 16  shows by way of example a gas turbine  100  in a longitudinal partial section. 
     The gas turbine  100  internally comprises a rotor  103 , or turbine rotor, mounted so that it can rotate about a rotation axis  102  and having a shaft  101 . 
     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, having a plurality of burners  107  arranged coaxially, a turbine  108  and the exhaust manifold  109 . 
     The ring combustion chamber  110  communicates with an e.g. annular hot gas channel  111 . There, for example, four successively connected turbine stages  112  form the turbine  108 . 
     Each turbine stage  112  is for example foamed by two blade rings. As seen in the flow direction of a working medium  113 , a row  125  formed by rotor blades  120  follows in the hot gas channel  111  of a guide vane row  115 . 
     The guide vanes  130  are fastened on the 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). 
     During operation of the gas turbine  100 , air  135  is taken in by the compressor  105  through the intake manifold  104  and compressed. 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 . 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. 
     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  110 , 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 thermally loaded most greatly. 
     In order to withstand the temperatures prevailing there, they may be cooled by means of a coolant. 
     The substrates may likewise comprise a directional structure, i.e. they are monocrystalline (SX structure) or comprise only longitudinally directed grains (DS). 
     Iron-, nickel- or cobalt-based superalloys, for example, are used as material for the components, in particular for the turbine blades and vanes  120 ,  130  and components of the combustion chamber  110 . 
     Such superalloys 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 are part of the disclosure in respect of the chemical composition of the alloy. 
     The blades and vanes  120 ,  130  may likewise comprise coatings against corrosion (MCrAlX; M is at least one element in the group 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). Such alloys are known, for example, 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 be part of this disclosure in respect of the chemical composition of the alloy. 
     On the MCrAlX, there may also be a thermal insulation layer, which consists for example of ZrO 2 , Y 2 O 3 —ZrO 2 , i.e. it is non-stabilized or partially or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide. 
     Rod-shaped grains are generated in the thermal insulation layer by suitable coating methods, for example electron beam deposition (EB-PVD). 
     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 .