Abstract:
Turbine components are often shipped individually and are not shipped assembled into a turbine. To this end, the turbine blade has to be protected of external stresses and external damage. This is done by an easily removable protective coating that easily evaporates during the first operation of the newly produced or restored component, so that the protective coating does not have to be removed in an additional operational step before installation.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is the US National Stage of International Application No. PCT/EP2008/050556, filed Feb. 4, 2010 and claims the benefit thereof. The International Application claims the benefits of European. Patent Office application No. 09001524.9 EP filed Feb. 4, 2009. All of the applications are incorporated by reference herein in their entirety. 
     
    
     FIELD OF INVENTION 
       [0002]    The invention relates to a turbine component with an easily removable protective coating, a set of turbine components, a turbine and a method for protecting a component. 
       BACKGROUND OF INVENTION 
       [0003]    Turbine blades are often provided with metallic or ceramic protective layers for protection from oxidation or corrosion and from excessive introduction of heat, and are either shipped while fitted in a turbine or, in case of doubt, are shipped individually or multiply to allow them to be newly fitted again in situ in a plant. 
         [0004]    Similarly, turbine blades have film cooling holes, which are necessary since the cooling makes a higher operating temperature of the turbine blade possible. 
         [0005]    During transit, it may happen that the ceramic layer becomes scratched, and this may cause a crack if there is thermal stress. Similarly, the film cooling holes may be clogged by dirt and prevent the emergence of cooling air during operation. 
       SUMMARY OF INVENTION 
       [0006]    It is therefore the object of the invention to solve the aforementioned problems. 
         [0007]    The invention is achieved by a turbine component with an easily removable layer as claimed in the claims, turbine components as claimed in the claims, a set of turbine components as claimed in the claims, a turbine as claimed in the claims and a method as claimed in the claims. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0008]    Further advantageous measures that can be combined with one another as desired in order to achieve further advantages are listed in the subclaims. 
           [0009]      FIGS. 1 ,  2 ,  3  and  4  show exemplary embodiments of a turbine blade, 
           [0010]      FIG. 5  shows a gas turbine, 
           [0011]      FIG. 6  shows a turbine blade and 
           [0012]      FIG. 7  shows a combustion chamber, 
           [0013]      FIG. 8  shows a list of superalloys. 
       
    
    
       [0014]    The examples listed in the figures and in the description only represent exemplary embodiments of the invention. 
       DETAILED DESCRIPTION OF INVENTION 
       [0015]      FIG. 1  shows a turbine component  1 ,  120 ,  130 ,  155  with an outer hole  7 , which is adjacent an outer surface  13  of a substrate  4 . The invention is not restricted to turbine components. 
         [0016]    An outer hole  7  means a hole in an outer wall of a hollow turbine component  1 ,  120 ,  130 ,  155 . 
         [0017]    The hole  7  is preferably a through-hole  7 , that is to say a film cooling hole in the case of a turbine blade  120 ,  130  ( FIG. 6 ) or a combustion chamber element  155  ( FIG. 7 ). 
         [0018]    The part that is identified by the reference numeral  4  is the substrate  4  of a superalloy ( FIG. 8 ) and preferably also has metallic and/or ceramic protective coatings  16  ( FIGS. 3 and 4 ), which are not represented in any more specific detail in  FIGS. 1 and 2 . 
         [0019]    The turbine component  1 ,  120 ,  130 ,  155  is used at high operating temperatures, at least 700° C., in particular at least 850° C. 
         [0020]    A final, further, outermost layer  10  is applied on the surface  13  of the substrate  4  or the metallic coating  16  (MCrAlY coating) or the ceramic coating  16 . 
         [0021]    The outermost coating  10  can be easily removed at removal temperatures well below the operating temperature of the component  1 ,  120 ,  130 ,  135  and preferably consists of an organic material, in particular of a polymer. 
         [0022]    The high-temperature-resistant polymers are known from the prior art, and so too is the coating of the component  1 ,  120 ,  130 ,  155  with the polymer. Coming into consideration as polymers are polyamides (Aurum), PEEK or PEK (polyether ketones). 
         [0023]    The protective coating  10  may preferably contain at least one, particularly only one, dye (preferably inorganic material). 
         [0024]    The coating  10  may preferably leave the film cooling hole  7  open ( FIGS. 1 and 3 ) or preferably also cover the opening partially, largely or entirely, as represented in  FIGS. 2 and 4 . 
         [0025]    If the hole  7  is narrowed, is also possible to prevent coarse dust particles from penetrating further into the hole  7 . 
         [0026]    The component  1 ,  120 ,  130 ,  155  is fitted in a device, preferably a gas turbine  100 , while the coating  10  is still present on the component  120 ,  130 ,  155  as shown in  FIGS. 1 ,  2  and  3  or  FIG. 4 . 
         [0027]    As a result of the lower removal temperature during commissioning (preferably start-up, test operation, . . . ) in comparison with the maximum operating temperatures of the gas turbine  100 , at the lower removal temperatures the protective coating  10  is thermally removed or vaporized by evaporation and burning or a similar chemical process and then exposes the film cooling hole  7  or removes itself from the surface of the component  1 ,  120 ,  130 ,  155 . When the newly fitted component  1 ,  120 ,  130  is used for the first time, cooling is initially not yet necessary, so that it is quite acceptable for the cooling hole  7  still to be covered by the protective coating  10 . 
         [0028]    The operating temperature for a gas turbine  100  is ≧800° C., in particular ≧1000° C. The protective layer  10  evaporates, burns or sublimates within the turbine  100 , preferably at least at 100° C., in particular ≧200° C., in particular at at least 300° C. 
         [0029]    The difference between these two temperatures (operating temperature and removal temperature of the layer  10 ) is preferably at least 500° C. 
         [0030]    If the hole  7  is covered by the layer  10  ( FIGS. 2 and 4 ) or narrowed ( FIG. 1 ), no dirt can penetrate into the hole  7  and temporarily or permanently clog it or constrict it (protection while in transit). 
         [0031]    If the color of the layer  10  is different at one point, this is an indication of possible damage, and the component  120 ,  130 ,  150  can be examined at this point. 
         [0032]    The turbine blades  120 ,  130  of the first stage of the turbine  100  may preferably be of a different color than the turbine blades  120 ,  130  of the second stage of the turbine  100  for better differentiation. 
         [0033]    Similarly, refurbished and new turbine blades  120 ,  130 , preferably of the same turbine stage, may be of different colors. 
         [0034]    Similarly, moving and stationary blades  120 ,  130  of one turbine stage of a turbine  100  may be of different colors. 
         [0035]    Similarly, moving and stationary blades of one turbine stage but of different turbines  100  or types of turbine may be of different colors. 
         [0036]    The color does not have to be monochrome. 
         [0037]    Protective coatings  10  may also be applied in the case of steam turbines. 
         [0038]      FIG. 5  shows a gas turbine  100  by way of example in a longitudinal partial section. The gas turbine  100  has in the interior a rotor  103  with a shaft  101 , which is rotatably mounted about an axis of rotation  102  and is also referred to as a turbine runner. 
         [0039]    Following one another along the rotor  103  are an intake housing  104 , a compressor  105 , a combustion chamber  110 , for example toroidal, in particular an annular combustion chamber, with a number of coaxially arranged burners  107 , a turbine  108  and the exhaust housing  109 . 
         [0040]    The annular combustion chamber  110  communicates with a hot gas duct  111 , for example of an annular form. There, the turbine  108  is formed by four successive turbine stages  112 , for example. 
         [0041]    Each turbine stage  112  is formed, for example, by two blade rings. As seen in the direction of flow of a working medium  113 , a row of stationary blades  115  is followed in the hot gas duct  111  by a row  125  formed by moving blades  120 . 
         [0042]    The stationary blades  130  are in this case fastened to an inner housing  138  of a stator  143 , whereas the moving blades  120  of a row  125  are attached to the rotor  103 , for example by means of a turbine disk  133 . 
         [0043]    Coupled to the rotor  103  is a generator or a machine (not represented). 
         [0044]    During the operation of the gas turbine  100 , air  135  is sucked in by the compressor  105  through the intake housing  104  and compressed. The compressed air provided at the end of the compressor  105  on the turbine side is passed to the burners  107  and mixed there with a fuel. The mixture is then burned in the combustion chamber  110  to form the working medium  113 . From there, the working medium  113  flows along the hot gas duct  111  past the stationary blades  130  and the moving blades  120 . At the moving blades  120 , the working medium  113  expands, transferring momentum, so that the moving blades  120  drive the rotor  103  and the latter drives the machine coupled to it. 
         [0045]    The components that are exposed to the hot working medium  113  are subjected to thermal loads during the operation of the gas turbine  100 . The stationary blades  130  and moving blades  120  of the first turbine stage  112 , as seen in the direction of flow of the working medium  113 , are thermally loaded the most, along with the heat shielding elements lining the annular combustion chamber  110 . 
         [0046]    In order to withstand the temperatures prevailing there, these may be cooled by means of a coolant. 
         [0047]    Similarly, substrates of the components may have a directional structure, i.e. they are monocrystalline (SX structure) or only have longitudinally directed grains (DS structure). 
         [0048]    Iron-, nickel- or cobalt-based superalloys are used for example as the material for the components, in particular for the turbine blade  120 ,  130  and components of the combustion chamber  110 . 
         [0049]    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. 
         [0050]    Similarly, the blades  120 ,  130  may have coatings against corrosion (MCrAlX; M is at least one element of the group comprising iron (Fe), cobalt (Co) and nickel (Ni), X is an active element and represents yttrium (Y) and/or silicon, scandium (Sc) and/or at least one element of the rare earths, or hafnium). 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. 
         [0051]    A thermal barrier coating, which consists for example of ZrO 2 , Y 2 O 3 —ZrO 2 , i.e. is unstabilized, partially stabilized or completely stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide, may also be present on the MCrAlX. 
         [0052]    Columnar grains are produced in the thermal barrier coating by suitable coating methods, such as for example electron-beam physical vapor deposition (EB-PVD). 
         [0053]    The stationary blade  130  has a stationary blade root (not represented here), facing the inner housing  138  of the turbine  108 , and a stationary blade head, at the opposite end from the stationary blade root. The stationary blade head faces the rotor  103  and is fixed to a fastening ring  140  of the stator  143 . 
         [0054]      FIG. 6  shows in a perspective view a moving blade  120  or stationary blade  130  of a turbomachine, which extends along a longitudinal axis  121 . 
         [0055]    The turbomachine may be a gas turbine of an aircraft or of a power plant for generating electricity, a steam turbine or a compressor. 
         [0056]    The blade  120 ,  130  has, following one after the other along the longitudinal axis  121 , a fastening region  400 , an adjoining blade platform  403  and also a blade airfoil  406  and a blade tip  415 . 
         [0057]    As a stationary blade  130 , the blade  130  may have a further platform at its blade tip  415  (not represented). 
         [0058]    In the fastening region  400  there is formed a blade root  183 , which serves for the fastening of the moving blades  120 ,  130  to a shaft or a disk (not represented). 
         [0059]    The blade root  183  is designed for example as a hammer head. Other designs as a firtree or dovetail root are possible. 
         [0060]    The blade  120 ,  130  has for a medium which flows past the blade airfoil  406  a leading edge  409  and a trailing edge  412 . 
         [0061]    In the case of conventional blades  120 ,  130 , solid metallic materials, in particular superalloys, are used for example in all the regions  400 ,  403 ,  406  of the blade  120 ,  130 . 
         [0062]    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. 
         [0063]    The blade  120 ,  130  may in this case be produced by a casting method, also by means of directional solidification, by a forging method, by a milling method or combinations of these. 
         [0064]    Workpieces with a monocrystalline structure or structures are used as components for machines which are exposed to high mechanical, thermal and/or chemical loads during operation. 
         [0065]    The production of monocrystalline workpieces of this type takes place for example by directional solidification from the melt. This involves casting methods in which the liquid metallic alloy solidifies to form the monocrystalline structure, i.e. to form the monocrystalline workpiece, or in a directional manner. 
         [0066]    Dendritic crystals are thereby oriented along the thermal flow and form either a columnar grain structure (i.e. grains which extend over the entire length of the workpiece and are commonly referred to here as directionally solidified) or a monocrystalline structure, i.e. the entire workpiece comprises a single crystal. In these methods, the transition to globulitic (polycrystalline) solidification must be avoided, since undirected growth necessarily causes the formation of transversal and longitudinal grain boundaries, which nullify the good properties of the directionally solidified or monocrystalline component. 
         [0067]    While reference is being made generally to solidified structures, this is intended to mean both monocrystals, which have no grain boundaries or at most small-angle grain boundaries, and columnar crystal structures, which indeed have grain boundaries extending in the longitudinal direction but no transversal grain boundaries. These second-mentioned crystalline structures are also referred to as directionally solidified structures. 
         [0068]    Such methods are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1. 
         [0069]    Similarly, the blades  120 ,  130  may have coatings against corrosion or oxidation, for example (MCrAlX; M is at least one element of the group comprising iron (Fe), cobalt (Co) and nickel (Ni), X is an active element and represents yttrium (Y) and/or silicon and/or at least one element of the rare earths, or hafnium (HO). 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. 
         [0070]    The density is preferably 95% of the theoretical density. 
         [0071]    A protective aluminum oxide layer (TGO=thermal grown oxide layer) forms on the MCrAlX layer (as an intermediate layer or as the outermost layer). 
         [0072]    The composition of the layer preferably comprises Co-30Ni-28Cr-8Al-0.6Y-0.7Si or Co-28Ni-24Cr-10Al-0.6Y. Apart from these cobalt-based protective coatings, nickel-based protective coatings are also preferably used, 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. 
         [0073]    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. is unstabilized, partially stabilized or completely stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide, may also be present on the MCrAlX. 
         [0074]    The thermal barrier coating covers the entire MCrAlX layer. 
         [0075]    Columnar grains are produced in the thermal barrier coating by suitable coating methods, such as for example electron-beam physical vapor deposition (EB-PVD). 
         [0076]    Other coating methods are conceivable, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may have grains which are porous, are provided with microcracks or are provided with macrocracks for better thermal shock resistance. The thermal barrier coating is therefore preferably more porous than the MCrAlX layer. 
         [0077]    Refurbishment means that components  120 ,  130  may have to be freed of protective layers after use (for example by sandblasting). This is followed by removal of the corrosion and/or oxidation layers or products. If need be, cracks in the component  120 ,  130  are then also repaired. This is followed by recoating of the component  120 ,  130  and renewed use of the component  120 ,  130 . 
         [0078]    The blade  120 ,  130  may be hollow or be of a solid form. If the blade  120 ,  130  is to be cooled, it is hollow and may also have film cooling holes  418  (indicated by dashed lines). 
         [0079]      FIG. 7  shows a combustion chamber  110  of a gas turbine. The combustion chamber  110  is designed for example as what is known as an annular combustion chamber, in which a multiplicity of burners  107 , which produce flames  156  and are arranged in the circumferential direction around an axis of rotation  102 , open out into a common combustion chamber space  154 . For this purpose, the combustion chamber  110  is designed as a whole as an annular structure, which is positioned around the axis of rotation  102 . 
         [0080]    To achieve a comparatively high efficiency, the combustion chamber  110  is designed for a comparatively high temperature of the working medium M of approximately 1000° C. to 1600° C. To permit a comparatively long operating time even with these operating parameters that are unfavorable for the materials, the combustion chamber wall  153  is provided on its side facing the working medium M with an inner lining formed by heat shielding elements  155 . 
         [0081]    Each heat shielding element  155  of an alloy is provided on the working medium side with a particularly heat-resistant protective layer (MCrAlX layer and/or ceramic coating) or is produced from material that is resistant to high temperature (solid ceramic bricks). 
         [0082]    These protective layers may be similar to the turbine blades, meaning for example MCrAlX: M is at least one element of the group comprising iron (Fe), cobalt (Co) and nickel (Ni), X is an active element and represents yttrium (Y) and/or silicon and/or at least one element of the rare earths, 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. 
         [0083]    A thermal barrier coating which is for example a ceramic thermal barrier coating and consists for example of ZrO 2 , Y 2 O 3 —ZrO 2 , i.e. is unstabilized, partially stabilized or completely stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide, may also be present on the MCrAlX. 
         [0084]    Columnar grains are produced in the thermal barrier coating by suitable coating methods, such as for example electron-beam physical vapor deposition (EB-PVD). 
         [0085]    Other coating methods are conceivable, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may have grains which are porous, are provided with microcracks or are provided with macrocracks for better thermal shock resistance. 
         [0086]    Refurbishment means that heat shielding elements  155  may have to be freed of protective layers after use (for example by sandblasting). This is followed by removal of the corrosion and/or oxidation layers or products. If need be, cracks in the heat shielding element  155  are then also repaired. This is followed by recoating of the heat shielding elements  155  and renewed use of the heat shielding elements  155 . 
         [0087]    On account of the high temperatures in the interior of the combustion chamber  110 , a cooling system may also be provided for the heat shielding elements  155  or for their holding elements. The heat shielding elements  155  are for example hollow and, if need be, also have cooling holes (not represented) opening out into the combustion chamber space  154 .