Patent Publication Number: US-2009229101-A1

Title: Method of Repairing a Component, and a Component

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is the US National Stage of International Application No. PCT/EP2007/051475 filed Feb. 15, 2007, and claims the benefit thereof. The International Application claims the benefits of European application No. 06010253.0 EP filed May 18, 2006. Both applications are incorporated by reference herein in their entirety. 
    
    
     FIELD OF INVENTION 
     The invention relates to a method for repairing a component and to a component. 
     BACKGROUND OF INVENTION 
     Newly produced components, for example cast components, or components after use, often have cracks which are repaired so that the component can be used again. 
     According to the prior art, material is excavated around the crack and it is filled with a braze or closed by welding. 
     So-called coupon brazing is also known from EP 0 868 253 B1. 
     The methods mentioned above, however, often cannot sufficiently prevent crack growth during reuse of the component. 
     SUMMARY OF INVENTION 
     It is therefore an object of the invention to overcome the problem mentioned above. 
     The object is achieved by a method and a component as claimed in the independent claims. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention will be explained in more detail with the aid of the figures. 
       The dependent claims list other advantageous measures, which may advantageously be combined with one another in any desired way. 
         FIG. 1  shows a component surface with a crack, 
         FIG. 2  shows a sectional representation of  FIG. 1  along the line II-II, 
         FIG. 3  shows a sectional representation of a hollow, 
         FIG. 4  shows a surface of a component with a hollow, 
         FIG. 5  shows the negative of the excavated crack surface (hollow), 
         FIG. 6  shows the use of rods in the further repair of the component, 
         FIG. 7  shows sectional representation of  FIG. 6  along the line VII-VII, 
         FIGS. 8 ,  9 ,  10 ,  11  show other exemplary embodiments of the invention, 
         FIGS. 12 ,  13  show another possibility for arranging rods in the hollow, 
         FIG. 14  shows a gas turbine, 
         FIG. 15  shows a turbine blade in perspective and 
         FIG. 16  shows a combustion chamber in perspective. 
     
    
    
     DETAILED DESCRIPTION OF INVENTION 
       FIG. 1  shows a surface  4  of a component  1 , which has a crack  7  with a length  1  along a crack direction  8 . The crack  7  may also extend in a curve, although in this case it also has an (averaged) direction  8 . 
     The component  1  is preferably a turbine blade  120 ,  130  ( FIG. 14 ) or a combustion chamber element  155  ( FIG. 15 ) of a turbine, for example an aircraft turbine, preferably a gas turbine  100  ( FIG. 13 ). 
     Particularly in the case of turbine components, a substrate  10  of the component  1  consists of a nickel- or cobalt-based superalloy. 
     The method of crack repair is not however restricted to such components, rather it also encompasses all components having cracks, cavities, indentations which are repaired by means of brazing or welding, and also those which consist of other materials. 
       FIG. 2  shows a sectional representation of  FIG. 1 , from which it may be seen that the crack  7  has a crack depth a in the substrate  10  of the component  1 , which has a thickness or wall thickness t (a&lt;&lt;t). 
     The crack  7  according to  FIGS. 1 ,  2  is excavated as in the prior art, so as to create a hollow  11  which according to the prior art comprises an indentation which is rectangular ( FIGS. 3 ,  4 ,  5 ) or triangular ( FIGS. 8 ,  9 ) in cross section. Other cross-sectional shapes may be envisaged. The hollow  11  is preferably of cuboid shape. 
     However, the hollow  11  additionally comprises at least three recesses  13 ,  13 ′,  16 ,  16 ′,  19 ,  19 ′. 
       FIG. 4  shows a plan view of such a hollow  11 . 
     The recesses  13 ,  13 ′, . . . may be arranged arbitrarily along the direction  8  around the circumference of the hollow  11 . The recesses  13 ,  13 ′, . . . are preferably arranged lying directly opposite in pairs ( FIGS. 4 ,  5 ,  6 ,  9 ,  10 ). 
     The distances between the recesses  13 ,  13 ′,  16 ,  16 ″, . . . , and in particular also from the start or end  17  of the hollow  11 , are preferably equidistant ( FIGS. 5 ,  6 ). In the exemplary embodiment according to  FIG. 4 , the recesses  13 ,  13 ′ are present at the end  17  of the hollow  11 . 
       FIG. 5  represents the negative of the excavated inner surface of the hollow  11  with recesses  13 ,  13 ′, . . . 
     Here the recesses  13 ,  13 ′ are not arranged at the end  17  of the hollow  11 , rather they lie at a distance from the end  17  of the hollow  11  along the direction  8 . 
     The excavated hollow  11  has a height a′, which is equal or essentially corresponds to the crack depth a according to  FIG. 3  (a≦a′). The hollow  11  has a length l′ along the crack direction  8 , which is equal or essentially corresponds to the length l of the crack  7  ( FIG. 1 ) (l≦l′). 
     Furthermore, the hollow  11  (without recesses) has a width b which is wide enough so that the crack  7  and attacked surfaces on the crack surface have been removed. 
     The recesses  13 ,  16 ,  19 , which preferably have recesses  13 ′,  16 ′,  19 ′ lying opposite, in particular perpendicularly to the crack direction  8 , substantially represent cuboids transversely to the direction  8  (see dashed indication) which have a length b+greater than the width b of the hollow  11 . 
     The recesses  13 ,  13 ′, . . . are for example of cuboid (or cubic) shape here and have a length f along the direction  8 , a depth g in the direction of the depth a′ and a width e in the direction of the width b. The width e of the recess is preferably much less than the width b of the hollow  11 , the length f is much less than the length l′ of the hollow  11  and the depth g is preferably much less than the depth a′ of the hollow  11 . 
     In  FIG. 5 , the recesses  13 ,  13 ′, . . . are of cubic or cuboid shape. The recesses  13 ,  13 ′, . . . may likewise have other shapes (round or triangular in cross section). 
     After the crack has been excavated according to  FIG. 5 , rods  22 ′,  22 ″,  22 ′″ are placed into the recesses  13 ,  13 ′, . . . ( FIG. 6 ), the melting temperature of these rods being higher than the melting temperature of the filler material, which constitutes in particular a braze  25 , with which the hollow  11  to be filled is then also filled ( FIG. 7 ). The hollow  11  may also be closed by welding. 
     The recesses  13 ,  13 ′ are preferably configured so that the rods  22 ′,  22 ″,  22 ′″ can be put into them from above, i.e. the recesses  13 ,  13 ′, . . . are open at the height of the surface  4 , and they do not protrude from the hollow  11  beyond the surface  4 . The depth g of the recesses  13 ,  13 ′, . . . is preferably configured here ( FIG. 5 ) so that the rod  22 ′,  22 ″,  22 ′″ is arranged close to the surface  4  of the component  1 . 
     The rods  22 ′,  22 ″ are preferably arranged in a plane, i.e. at one height. The height of the rods  22 ′,  22 ″ between one another, i.e. the depth g of the respective recesses  13 ,  13 ′, . . . , may also be selected arbitrarily so that there is a larger distance from the surface  4 . 
     The material of the rods  22 ′,  22 ″,  22 ′″ preferably consists of a ceramic material. 
     The effect of the rods is that a crack, which is often formed on the surface  4  and propagates through the braze, meets the ceramic rods  22 ′,  22 ″,  22 ′″ where it is prevented from growing further in them owing to the greater toughness of the rods  22 ′,  22 ″,  22 ′″. 
     The recesses  13 ,  13 ′,  16 ,  16 ′,  19 ,  19 ′ may likewise extend over the entire height a′ of the hollow  11 , as is represented in  FIG. 10 . Plates instead of rods are then preferably inserted into these recesses  13 ,  13 ′, . . . , the plates extending over the entire width b+. The plates may have the height a′, although they may also be of smaller size. 
     Instead of plates, fine- or coarse-latticed meshes may also be inserted so that the filler material or the braze can enclose the mesh during the brazing and the mesh acts as crack prevention over the entire depth of the hollow  11 . 
     Fiber bundles may also be inserted into the hollow  11  instead of the rods  22 ′,  22 ″,  22 ′″, and fiber mats may also be inserted instead of the plates. 
     A coupon may also be brazed or welded into the hollow  11 , the coupon having a shape according to  FIG. 5  while having smaller dimensions than the hollow  11 , so that it can be fitted into it. 
       FIG. 11  shows another exemplary embodiment of an arrangement of the rods  22 ′,  22 ″,  22 ′″ in the hollow  11 . 
     The rods  22 ′,  22 ″, . . . do not extend perpendicularly to the crack direction  8  or to the longitudinal direction of the hollow  11  here, rather they are arranged obliquely. The rods  22 ″,  22 ′″ may likewise cross over, i.e. either they are arranged at different heights in the hollow  11  or a cross in the form of an “X” is placed into the recesses  16 ,  16 ′,  19 ,  19 ′, in which case the rods  22 ″,  22 ′″ may form one part. 
     Further arrangement possibilities for the rods  22 ′,  22 ″ with respect to one another, or the combination of rods and fiber mats ( FIG. 10 ), may be envisaged. 
     Instead of the recesses, which adjoin the surface  4  of the component  1 , there may also be recesses  13 ,  13 ′, . . . below the surface  4  on the inner surfaces of the hollow  11 . The rods  22 , . . . must then be flexible enough to be bent ( FIG. 12 ) so that they can be inserted into the hollow  11 , in which case the rods must have a length &gt;b so that their two ends rest in the indentations  13 ,  13 ′ ( FIG. 13 ). 
       FIG. 14  shows a gas turbine  100  by way of example in a partial longitudinal section. 
     The gas turbine  100  internally comprises a rotor  103 , which will also be referred to as the turbine rotor, mounted so as to 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 formed for example by two blade rings. As seen in the flow direction of a working medium  113 , a guide vane row  115  is followed in the hot gas channel  111  by a row  125  formed by rotor blades  120 . 
     The guide vanes  130  are fastened on an inner housing  138  of a 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 and compressed by the compressor  105  through the intake manifold  104 . The compressed air provided at the end of the compressor  105  on the turbine side 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 heated the most. 
     In order to withstand the temperatures prevailing there, they may be cooled by means of a coolant. Substrates of the components may likewise comprise a directional structure, i.e. they are monocrystalline (SX structure) or comprise only longitudinally directed grains (DS structure). Iron-, nickel- or cobalt-based superalloys used as material for the components, in particular for the turbine blades  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; with respect to the chemical composition of the alloys, these documents are part of the disclosure. 
     The guide vane  130  comprises 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 . 
       FIG. 15  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. 
     The blade  120 ,  130  comprises, successively along the longitudinal axis  121 , a fastening region  400 , a blade platform  403  adjacent thereto as well as a blade surface  406  and a blade tip  415 . 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 fir-tree 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; with respect to the chemical composition of the alloy, these documents are part of the disclosure. 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 during operation are exposed to heavy mechanical, thermal and/or chemical loads. Such monocrystalline workpieces are manufactured, for example, by directional solidification from the melts. These are casting methods in which the liquid metal alloy is solidified to form a monocrystalline structure, i.e. to form the monocrystalline workpiece, or is directionally solidified. 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 these methods, since nondirectional growth will necessarily form transverse and longitudinal grain boundaries which negate the beneficial 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; with respect to the solidification method, these documents are part of the disclosure. 
     The blades  120 ,  130  may likewise have 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, or hafnium (Hf)). Such alloys 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, with respect to the chemical composition of the alloy, are intended to be part of this disclosure. The density may preferably be 95% of the theoretical density. A protective aluminum oxide layer (TGO=thermal grown oxide layer) is formed on the MCrAlX layer (as an interlayer or as the outermost layer). 
     On the MCrAlX, there may furthermore be a thermal barrier layer, which is preferably the outermost layer and consists for example of ZrO 2 , Y 2 O 3 —ZrO 2 , i.e. it is not stabilized or is partially or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide. The thermal barrier layer covers the entire MCrAlX layer. Rod-shaped grains are produced in the thermal barrier layer by suitable coating methods, for example electron beam deposition (EB-PVD). Other coating methods may be envisaged, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier layer may comprise porous, micro- or macro-cracked grains for better thermal shock resistance. The thermal barrier layer is thus preferably more porous than the MCrAlX layer. 
     The blade  120 ,  130  may be designed to be 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  (indicated by dashes). 
       FIG. 16  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 , which produce flames  156  and are arranged in the circumferential direction around a rotation axis  102 , open into a common combustion chamber space  154 . To this end, the combustion chamber  110  as a whole 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. 
     Owing to the high temperatures inside the combustion chamber  110 , a cooling system may also be provided for the heat shield elements  155  or for their retaining elements. The heat shield elements  155  are then hollow, for example, and optionally also have film cooling holes (not shown) opening into the combustion chamber space  154 . 
     Each heat shield element  155  made of an alloy is equipped with a particularly heat-resistant protective layer (MCrAlX layer and/or ceramic coating) on the working medium side, 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, or hafnium (Hf). Such alloys 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, with respect to the chemical composition of the alloy, are intended to be part of this disclosure. 
     On the MCrAlX, there may furthermore be an e.g. ceramic thermal barrier layer which consists for example of ZrO 2 , Y 2 O 3 —ZrO 2 , i.e. it is not stabilized or is partially or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide. 
     Rod-shaped grains are produced in the thermal barrier layer by suitable coating methods, for example electron beam deposition (EB-PVD). Other coating methods may be envisaged, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier layer may comprise porous, micro- or macro-cracked grains for better thermal shock resistance. 
     Refurbishment means that turbine blades  120 ,  130  and heat shield elements  155  may need to have protective layers taken off (for example by sandblasting) after their use. The corrosion and/or oxidation layers or products are then removed. Optionally, cracks in the turbine blade  120 ,  130  or the heat shield element  155  are also repaired by the method. The turbine blades  120 ,  130  or heat shield elements  155  are then recoated and the turbine blades  120 ,  130  or the heat shield elements  155  are used again.