Abstract:
A method is provided for repairing a metallic turbine component which includes a thermal barrier coating system including a metallic bond coat and a ceramic top coat. The method includes: (a) removing the top coat using a mechanical process; (b) partially stripping the metallic bond coat from the component, such that substantially no material of the component is removed; (c) repairing at least one defect in the turbine component; (d) applying a new metallic bond coat to the turbine component; and (e) applying a new ceramic top coat over the metallic bond coat.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    This invention relates generally to the repair of gas turbine engine components and more particularly to the repair of turbine components incorporating ceramic thermal barrier coatings. 
         [0002]    Gas turbine engines include a “hot section” comprising a combustor and one or more downstream turbines. Components in the hot section are exposed during operation to a high temperature, corrosive gas stream that limits their effective service life. Accordingly, these components are typically fabricated from high temperature cobalt or nickel-based “superalloys” and are often coated with corrosion and/or heat resistant materials, and in particular with ceramic thermal barrier coatings (TBCs). 
         [0003]    Examples of such components include but are not limited to high pressure turbine blades and nozzles, and turbine blade shrouds. Despite the use of protective coatings, such components commonly develop defects such as cracks, damage, or material loss during service. 
         [0004]    These components, after engine operation, are commonly repaired by brazing processes such as activated diffusion healing (“ADH”), or by welding processes such as superalloy welding at elevated temperature (“SWET”). In order to achieve a successful repair, the components should be clean and free of oxides at the interface of the defect. 
         [0005]    One conventional process used to repair TBC coated components with service cracks is to completely strip the TBC coatings, including the ceramic top coat, and the metallic bond coat by a combination of grit blasting and chemical stripping. These two processes have a number of negative aspects, including low labor cost productivity, process variance and rework because of the manual nature of the processes, and component wall loss during chemical stripping which results in a high scrap rate. 
       BRIEF SUMMARY OF THE INVENTION 
       [0006]    These and other shortcomings of the prior art are addressed by the present invention, which provides a method for repairing a coated turbine component while preserving the material thickness of the component. 
         [0007]    According to an aspect of the invention, a method is provided for repairing a metallic turbine component which includes a thermal barrier coating system including a metallic bond coat and a ceramic top coat. The method includes: (a) removing the top coat using a mechanical process; (b) partially stripping the metallic bond coat from the component, such that substantially no material of the component is removed; (c) repairing at least one defect in the turbine component; (d) applying a new metallic bond coat to the turbine component; and (e) applying a new ceramic top coat over the metallic bond coat. 
         [0008]    According to another aspect of the invention, a repaired metallic turbine component includes: (a) a metallic body; (b) a bond coat applied to the body, comprising: (i) a first metallic layer of a nickel-based alloy; and (ii) a second metallic layer of a nickel-based alloy overlying the first layer; and (c) a ceramic top coat overlying the bond coat. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0009]    The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which: 
           [0010]      FIG. 1  is a perspective view of an service-run turbine nozzle having one or more defects therein; 
           [0011]      FIG. 2  is an enlarged cross-sectional view of a portion of a vane of the turbine nozzle of  FIG. 1 , showing a defect therein; 
           [0012]      FIG. 3  is a view of the vane of  FIG. 2  after removal of a TBC top coat; 
           [0013]      FIG. 4  is a view of the vane of  FIG. 3  after partial stripping of a TBC bond coat; 
           [0014]      FIG. 5  is a view of the vane of  FIG. 4  after a crack repair; 
           [0015]      FIG. 6  is a view of the vane of  FIG. 5  after a bond coat application; 
           [0016]      FIG. 7  is a view of the vane of  FIG. 6  after reapplication of an aluminide coating; and 
           [0017]      FIG. 8  is a view of the vane of  FIG. 7  after application of a new TBC top coat. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0018]    Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,  FIG. 1  illustrates an exemplary turbine nozzle segment  10 . A gas turbine engine will include a plurality of such segments  10  arranged in an annular array. The turbine nozzle segment  10  is merely an example of a coated metallic turbine component, and the repair methods described herein are equally applicable to other components, nonlimiting examples of which include combustor liners, rotating turbine blades, and turbine shrouds. 
         [0019]    The turbine nozzle  10  includes first and second nozzle vanes  12  disposed between an arcuate outer band  14  and an arcuate inner band  16 . The vanes  12  define airfoils configured so as to optimally direct the combustion gases to a turbine rotor (not shown) located downstream thereof. The outer and inner bands  14  and  16  define the outer and inner radial boundaries, respectively, of the gas flow through the nozzle segment  10 . The interior of the vanes  12  are mostly hollow and may include a number of internal cooling features of a known type, such as walls defining serpentine passages, ribs, turbulence promoters (“turbulators”), etc. The vanes  12  can have a plurality of conventional cooling holes  18  and trailing edge slots  20  formed therein. The presence of at least a specified minimum amount of material (i.e. minimum wall thickness) in the vanes  12  and the bands  14  and  16  is important to maintaining the structural integrity and aerodynamic performance of the turbine nozzle segment  10 . 
         [0020]    Such nozzle segments  10  and other hot section components commonly cast in one or more sections from a cobalt or nickel-based superalloy which has acceptable strength at the elevated temperatures of operation in a gas turbine engine (e.g. RENE 80, RENE 142, RENE N4, RENE N5, RENE N6). 
         [0021]    RENE N5 is one particularly commonly used alloy and has a nominal composition in weight percent of about 7.5 percent cobalt, about 7.0 percent chromium, about 1.5 percent molybdenum, about 5 percent tungsten, about 3 percent rhenium, about 6.5 percent tantalum, about 6.2 percent aluminum, about 0.15 percent hafnium, about 0.05 percent carbon, about 0.004 percent boron, about 0.01 percent yttrium, balance nickel and minor elements. 
         [0022]    The turbine nozzle segment  10  is coated with a TBC coating system of a known type.  FIG. 2  shows a portion of one of the vanes  12 . The coating system comprises a bond coat  22  and a top coat  24  of a ceramic material. The thicknesses of the various layers of the TBC coating are exaggerated for illustrative clarity. 
         [0023]    Examples of metallic TBC bond coats in wide use include alloys such as MCrAlX overlay coatings (where M is iron, cobalt and/or nickel, and X is yttrium or a rare earth element), and diffusion coatings that contain aluminum intermetallics, predominantly beta-phase nickel aluminide and platinum-modified nickel aluminides (PtA1). 
         [0024]    An example of an MCrAlX coating is commercially known as BC52 and has a nominal composition of, by weight, about 18% chromium, 10% cobalt, 6.5% aluminum, 2% rhenium, 6% tantalum, 0.5% hafnium, 0.3% yttrium, 1% silicon, 0.015% zirconium, 0.06% carbon and 0.015% boron, the balance nickel. BC52 and other bond coats may be applied by processes such as physical vapor deposition (PVD), particularly electron beam physical vapor deposition (EBPVD), and thermal spraying, particularly plasma spraying (air, low pressure (vacuum), or inert gas) and high velocity oxy-fuel spraying (HVOF). 
         [0025]    The outer portion of the bond coat  22  is overcoated by an aluminiding process to increase the Al content for improved environmental resistance while retaining appropriate surface roughness as an anchor for the top coat  24  and sealing porosity in the bond coat  22 . The complete bond coat  22  thus comprises a metallic layer  26  and an aluminide layer  28 . Various aluminiding processes are known, for example pack cementation, vapor atmosphere, local powder application, etc. 
         [0026]    The ceramic top coat  24  may comprise any suitable ceramic material alone or in combination with other materials. For example, it may comprise fully or partially stabilized yttria-stabilized zirconia and the like, as well as other low conductivity oxide coating materials known in the art. One suitable method for deposition is by electron beam physical vapor deposition (EB-PVD), although plasma spray deposition processes, such as air plasma spray (APS), also may be employed. 
         [0027]    During engine operation, the vanes  12  can experience damage such as might result from local gas stream over-temperature or foreign objects impacting thereon. By way of example, the vanes  12  are shown in  FIGS. 1 and 2  as having defects such as cracks “C”. As seen in  FIG. 2 , the cracks penetrate the TBC top coat  24 , the bond coat  22 , and the wall of the vane  12 . 
         [0028]    Using the vane  12  as a working example, TBC coated components may be repaired as follows, with reference to  FIGS. 3-7 . First, the top coat  24  is removed using a mechanical stripping method such as grit blasting. The mechanical stripping process parameters are selected so as to remove the top coat  24  and potentially part of the bond coat  22 , but not to penetrate through to any base material of the vane  12 . An example of a suitable stripping process is a light grit blast using about 138 kPa (20 psi) to about 414 kPa (60 psi) air pressure, with 120-240 grit aluminum oxide particles. The results are shown in  FIG. 3 . 
         [0029]    Next, the bond coat  22  is cleaned through a known fluoride ion cleaning (FIC) process, which is a high temperature gas-phase treatment of the nozzle segment  10  using hydrogen fluoride and hydrogen gas. The FIC process parameters are selected so as to remove the aluminide layer  28  of the bond coat  22 , but not to penetrate through the metallic layer  26  to the surface  30  of the vane  12 . An example of a suitable FIC process may consist of heating parts to a working temperature of about 1038° C. (1900° F.) to about 1093° C. (2000° F.), in a gas atmosphere of about 2-9% hydrogen fluoride/hydrogen with a soak time of about 2-8 hours at the working temperature. In other words, the FIC process is biased to leave some of the bond coat  22  remaining. This ensures that substantially no base metal attack occurs. In contrast, prior art stripping processes have focused on complete removal of the bond coat  22 , which inevitably leads to wall thickness loss in the vane  12 .  FIG. 4  shows the vane  12  after the FIC process A minimum amount of bond coat  22 , for example, about 0.03 mm (0.001 in.) bond coat thickness should be left intact on the part surface after the FIC process. Preferably about half the initial bond coat thickness is left intact. 
         [0030]    Next, crack repairs are made where necessary using conventional braze or welding processes, such as the above-noted ADH or SWET processes.  FIG. 5  shows a braze deposit “W” filling the crack C in the vane  12 . Contrary to conventional expectations, is has been found that satisfactory braze and weld repairs may be made even when part of the bond coat  22  is still in place. This unique result is achieved by using the FIC process which removes aluminum from the bond coat  22 . 
         [0031]    Once repairs are complete, the TBC system can be reapplied. First, a metallic flash coat is applied. For example, a layer  26 ′ of the above-mentioned BC52 material about 0.08 mm (3 mils) to about 0.13 mm (5 mils) in thickness may be applied (see  FIG. 6 ). Next, an aluminide coating  28 ′ is reapplied, as shown in  FIG. 7 . Finally, a ceramic TBC top coat  24 ′ is applied, as seen in  FIG. 8 . The completed vane  12  is then ready for return to service. 
         [0032]    Furnace cycle tests at high temperature, e.g. 1093° C. (2000° F.), of components repaired by the process described above have shown that the replaced TBC coating system is satisfactory for return to service in gas turbine engines. In fact, testing indicates that the repaired TBC system may survive more cycles at high temperature than an OEM TBC coating system. 
         [0033]    The foregoing has described a method for repairing gas turbine engine coated hot section components. While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention. Accordingly, the foregoing description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation.