Patent Abstract:
A method for coating a component of a turbo-machine. The method allows arranging a turbine rotor in the turbo-machine and introducing a coating material into the interior of the turbo-machine such that the rotor is coated. The rotor is rotated while it is being coated.

Full Description:
CROSS REFERENCE TO RELATED APPLICATION 
     This application claims priority of the European application No. 04028484.6 EP filed Dec. 01, 2004, which is incorporated by reference herein in its entirety. 
     FIELD OF THE INVENTION 
     The invention relates to a process for coating a component in accordance with the preamble of the claims. 
     BACKGROUND OF THE INVENTION 
     Components are often provided with layers in order to achieve a certain function, such as for example resistance to corrosion, oxidation and/or heat (thermal barrier). In this case, the coated components may also be rotating components, such as for example blades of a compressor or a turbine of a gas turbine installation which have layers protecting against erosion or heat. In the case of large machines, a rotor is composed of a large number of individual parts (a plurality of disks each having a plurality of turbine blades), which are coated individually or in groups, and consequently it takes a long time for all the individual parts to be coated. 
     JP 06 099 125 A discloses a coating apparatus in which the substrate is coated during rotation. 
     JP 06 219 762 A discloses a coating method in which a circular cutting tool is coated while it is rotating. 
     U.S. Pat. No. 5,897,921 discloses a process for applying a thermal barrier coating. 
     U.S. Pat. No. 6,585,569 B2 discloses a process for cleaning a compressor of a gas turbine in which dry ice is introduced into the turbine. 
     U.S. Pat. No. 6,180,262 discloses a process in which a plurality of dismantled turbine blades of a rotor are coated all at once. 
     Therefore, it is an object of the invention to overcome the above problem. 
     SUMMARY OF THE INVENTION 
     This object is achieved by the process as claimed. 
     The subclaims list further advantageous measures which can be combined with one another in any advantageous way. 
     The maintenance time for a turbine is considerably shortened by the process according to the invention, since, for example in the case of a gas turbine, there is no need to wait for the turbine and turbine housing to have cooled, and/or there is no need for forced cooling, and since it is not necessary to remove all the feed lines and outer parts of the housing. Also, the turbine does not have to be reassembled and started up again. This considerably reduces the maintenance time and the associated downtime for the operator of a turbine and obviates the need for heavy machines required to lift the housing parts. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       In the drawing: 
         FIG. 1  diagrammatically depicts the process according to the invention, 
         FIG. 2  shows a turbine blade or vane, 
         FIG. 3  shows a combustion chamber, 
         FIG. 4  shows a gas turbine, 
         FIG. 5  shows a steam turbine. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       FIG. 1  shows an apparatus  1  which includes a turbomachine with a rotor  10 . 
     The rotor  10  is to be coated within the interior of a housing  4 . 
       FIG. 1  diagrammatically depicts how the process according to the invention is to be carried out in order to coat the rotor  10 . The rotor  10  comprises a plurality of parts  7 ,  120 . In the case of a rotor  103  for a turbine  100 ,  300 ,  303  or of a compressor  105 , the rotor  10 ,  103  has, for example, at least one disk  11 ,  11 ′  133  ( FIG. 4 ), on which a plurality of turbine blades  120 ,  354  ( FIG. 2 ,  4 ,  5 ) are arranged in a radial orientation distributed over the circumference. 
     The rotor  10  may comprise a plurality of disks  11 ,  11 ′ each having a plurality of turbine blades  120 ,  354 . A row  125  comprises, for example, a disk  11 ,  133  ( FIG. 4 ) with turbine rotor blades  120 . 
     The rotor  10 ,  103  is still arranged in its housing  4 ,  138  ( FIG. 4 ) and can be rotated about a rotatable axis  16 ,  102  ( FIG. 4 ). 
     The housing  4 ,  138  is, for example, the housing of a turbine  100 ,  300 ,  303  or of a compressor  105 , in which the rotor  10 ,  103  is operated. 
     According to the invention, coating material  13  is applied while the rotor  10 ,  103  is therefore still in its installed position. 
     During the coating operation, the rotor  10  can also rotate about the axis of rotation  16 ,  102 . 
     The coating material  13  may be in gas form, as in known from a PVD or CVD installation, and is then deposited on the surfaces of the parts  7 ,  120  that are to be coated. 
     The coating material  13  may also be applied to the parts  7  in liquid form, in particular finely dispersed in the air. 
     By way of example, it is also possible to apply a slip which contains a binder and powder particles, which then produce the definitive layer. In this case, the carrier medium can also be evaporated and the binder burnt out when the coating material  13  has been applied. Then, for example by further increasing the temperature, the powder particles are sintered together, so as to produce a fixed coating. The increase in temperature can be achieved by applying a flame to the components  7 ,  120  or, for example, by passing a hot gas or steam through the hollow turbine blades  120 , which leads to heating of the coated component  7 . 
     The coating of the rotor  10  can also be carried out while the rotor is operating. In this context, the term “operating” means that the rotor  10  is being used as intended. This means that the rotor of a compressor is compressing air, while a gas (steam, hot gas) is being expanded and performing work in the rotor of a turbine. Where the text refers to the rotor “rotating”, this does not necessarily mean that the rotor is operating. The rotor  10  is in this case, for example, part of a compressor  105 , for example, a gas turbine  100 , in which case the coating material is added as an additive to the air that is to be compressed. 
     It is also possible for the rotor  10  to be a rotor  103  of a gas turbine  100 , in which case the coating material  13  is added to the hot gas in operation during a reduction or increase in power of the gas turbine  100 , in order to coat the rotor blades  120  and/or guide vanes  130  of the rotor. 
     Furthermore, in operation the temperature of the gas can be deliberately matched to the required thermal boundary conditions for the respective coating process by more or less fuel being burnt and less compressed air being fed from the compressor to the turbine or by the temperature of the steam being controlled. 
     The coating material  13  may contain a metal halide (AlF 3 , AICl 3 , CrF, . . . ) which is in gas form or in the form of powder particles. 
     It is also possible for particles  13  in powder form (e.g. ceramics, hard metals) to be applied to the components  7 ,  120  that are to be coated, and these particles are then embedded on the region of the rotors  10 ,  103  which is close to the surface, if the surface or a layer in which the particles can be embedded is soft enough (for example by heating). These are, for example, coarser particles which are intended, for example, to increase the resistance to erosion of the component  7 ,  120 . 
     The coating material  13  may, for example, be metallic (MCrAlX) or vitreous (compressor blade). 
     The coating process according to the invention can also be used to repair damaged blades or vanes of the compressor  105  or of the turbine  100 ,  300 ,  303 . In this case, the material can be selected in such a way that it is preferentially deposited on the damaged areas. 
     If the rotor  10  is a rotor  103  of a gas turbine  100 , the coating material  13  can, for example, be introduced into the combustion chamber  110  with the fuel, and the combustion of the fuel can heat the coating material  13 , so that it is deposited on the components  7 ,  120  in a similar way to in the plasma spraying process. 
     It is also possible for the blades or vanes of a compressor, in particular of a compressor  105  of a turbine  100 , to be coated. 
     When air is being compressed in the compressor, water precipitates and can form an electrolyte in combination with other elements contained in the air, which can lead to corrosion and erosion at the compressor blades or vanes. To prevent the corrosion and erosion, it is possible for compressor blades or vanes to be provided with coatings. A coating of this type comprises, for example, a basecoat and a topcoat. A suitable basecoat is in particular a coating which comprises an inorganic binder composed of chromium phosphate compounds and contains, for example, spherical aluminum particles. Coatings of this type are disclosed in EP 0 142 418 B1 or in EP 0 905 279 A1, with the layer composition and layer structure of these patents forming part of the present disclosure. The topcoat used may, for example, be water-based chromium phosphate compounds with inert fillers and colored pigmentations. 
     The same procedure can also be adopted for internal coating of a turbine  100  or a compressor  105 . The turbine  100  comprises guide vanes  130  and a rotor  103  which has the rotor blades  120 . 
     The coating material  13 , as described above, is introduced into the turbine  100 , with the coating material  13  being deposited both on the rotor  103  and on the guide vanes  130 . The rotor  103  with the rotor blades  120  can in the process also rotate. 
     In this context, it is also possible to coat housing parts  4 ,  138  and blades and vanes  120 ,  130  or just the housing  4 ,  138 . 
     This is done by controlled setting of temperature differences between housing  4 ,  138  and blades and vanes  120 ,  130 . The hollow rotor blades  120  and the hollow guide vanes  130  have separate feeds leading into their cavity for supplying a medium, so that the rotor blades  120  can be heated while the guide vanes  130  are not, or vice versa. The appropriate growth conditions for a layer to grow on a substrate (blades or vanes) are only established at a certain elevated temperature, or only this elevated temperature makes it possible to ensure that the layer will not flake off a substrate which is too cold. 
     A combustion chamber  110  may likewise be a component of an apparatus  1 , i.e. a gas turbine  100 , which is to be coated. In this case too, the coating material is fed into the combustion chamber  110  from the outside. As with the turbine, this can take place during operation. In this case, the coating material can be supplied via the burner  107  and is then deposited on the heat shield elements  155 , which have been suitably “temperature-controlled”. 
     Another way of feeding the coating material to the combustion chamber or the turbine for the coating operation is for the coating material to be added to the flow medium within the turbine  100  at the compressor outlet. 
       FIG. 2  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 generating electricity, a steam turbine or a compressor. 
     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 . 
     As a guide vane  130 , the vane  130  may have a further platform (not shown) at its vane tip  415 . 
     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 . 
     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. 
     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 . 
     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  120 ,  130 . 
     Superalloys of this type are known, for example, from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949; these documents form part of the disclosure. The blade or vane  120 ,  130  may in this case be produced by a casting process, also by means of directional solidification, by a forging process, by a milling process or combinations thereof. 
     Work pieces 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. 
     Single-crystal work pieces 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 work piece, or solidifies directionally. 
     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 work piece and are referred to here, in accordance with the language customarily used, as directionally solidified) or a single-crystal structure, i.e. the entire work piece 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. 
     Where the text refers in general terms to directionally solidified micro structures, 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 micro structures (directionally solidified structures). 
     Processes of this type are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1; these documents form part of the disclosure. 
     The blades or vanes  120 ,  130  may likewise have coatings protecting against corrosion or oxidation (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 represents yttrium (Y) and/or silicon (Si) 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 the present disclosure. 
     It is also possible for there to be a thermal barrier coating, consisting for example of ZrO 2 , Y 2 O 4 —ZrO 2 , i.e. unstabilized, partially stabilized or completely stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide, to be present on the MCrAlX. Columnar grains are produced in the thermal barrier coating by means of suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD). 
     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 by soldering or welding. This is followed by recoating of the component  120 ,  130 , after which the component  120 ,  130  can be reused. 
     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). 
       FIG. 3  shows a combustion chamber  110  of a gas turbine. The combustion chamber  110  is configured, for example, as what is known as an annular combustion chamber, in which a multiplicity of burners  107  arranged circumferentially around the axis of rotation  102  open out into a common combustion chamber space. For this purpose, the combustion chamber  110  overall is of annular configuration positioned around the axis of rotation  102 . 
     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 . 
     On the working medium side, each heat shield element  155  is equipped with a particularly heat-resistant protective layer or is made from material that is able to withstand high temperatures. These may be solid ceramic bricks or alloys with MCrAlX and/or ceramic coatings. The materials of the combustion chamber wall and their coatings may be similar to the turbine blades or vanes. 
     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 . 
       FIG. 4  shows, by way of example, a partial longitudinal section through a gas turbine  100 . 
     In the interior, the gas turbine  100  has a rotor  103  which is mounted such that it can rotate about an axis of rotation  102  and is also referred to as the turbine rotor. 
     An intake housing  104 , a compressor  105 , a, for example, toroidal combustion chamber  110 , in particular an annular combustion chamber  106 , with a plurality of coaxially arranged burners  107 , a turbine  108  and the exhaust-gas housing  109  follow one another along the rotor  103 . 
     The annular combustion chamber  106  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 . 
     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 . 
     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 . A generator (not shown) is coupled to the rotor  103 . 
     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. 
     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, as seen in the direction of flow of the working medium  113 , together with the heat shield bricks which line the combustion chamber  106 , are subject to the highest thermal stresses. 
     To be able to withstand the temperatures which prevail there, they have to be cooled by means of a coolant. 
     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). 
     By way of example, iron-base, nickel-base or cobalt-base 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 . 
     Superalloys of this type are known, for example, from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949; these documents form part of the disclosure. 
     The blades or vanes  120 ,  130  may also have coatings which protect 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 represents yttrium (Y) and/or silicon 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, which are intended to form part of the present disclosure. 
     A thermal barrier coating, consisting for example of ZrO 2 , Y 2 O 4 —ZrO 2 , i.e. unstabilized, partially stabilized or completely stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide, may also be present on the MCrAlX. Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD). 
     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 . 
       FIG. 5  illustrates, by way of example, a steam turbine  300 ,  303  with a turbine shaft extending along an axis of rotation  306 . 
     The steam turbine has a high-pressure partial turbine  300  and a medium-pressure partial turbine  303 , each with an inner housing  312  and an outer housing  315  surrounding it. 
     The high-pressure partial turbine  300  is, for example, of pot-like configuration. 
     The medium-pressure partial turbine  303  is of two-flow design. 
     It is also possible for the medium-pressure partial turbine  303  to be of single-flow design. 
     A bearing  318  is arranged between the high-pressure partial turbine  300  and the medium-pressure partial turbine  303  along the axis of rotation  306 , the turbine shaft  309  having a bearing region  321  in the bearing  318 . The turbine shaft  309  is mounted on a further bearing  324  next to the high-pressure partial turbine  300 . In the region of this bearing  324 , the high-pressure partial turbine  300  has a shaft seal  345 . The turbine shaft  309  is sealed by two further shaft seals  345  with respect to the outer housing  315  of the medium-pressure partial turbine  303 . Between a high-pressure steam inlet region  348  and a steam outlet region  351 , the turbine shaft  309  has the high-pressure rotor blading  354 ,  357  in the high-pressure partial turbine  300 . This high-pressure rotor blading  354 ,  357 , together with the associated rotor blades (not shown in more detail), represents a first blading region  360 . The medium-pressure partial turbine  303  has a central steam inlet region  333 . Assigned to the steam inlet region  333  the turbine shaft  309  has a radially symmetrical shaft shield, a covering plate, on the one hand for dividing the flow of steam into the two flows for the medium-pressure partial turbine  303  and for preventing direct contact between the hot steam and the turbine shaft  309 . In the medium-pressure partial turbine  303 , the turbine shaft  309  has a second blading region  366  comprising the medium-pressure rotor blades  354 ,  342 . The hot steam which flows through the second blading region  366  flows out of the medium-pressure partial turbine  303  from an outlet connection piece  369  to a low-pressure partial turbine, which is connected downstream in terms of flow but is not illustrated.

Technology Classification (CPC): 5