Patent Publication Number: US-2012034491-A1

Title: Cmas resistant tbc coating

Description:
BACKGROUND 
     The present disclosure is directed to a coating system for a turbine engine component which has an activated CMAS resistant reactive layer deposited over a thermal barrier coating and a method of forming same. 
     The durability and the maximum temperature capability of a thermal barrier coating (TBC) system used in gas turbine engines is often limited by deposits of calcium-alumino-silicate (CMAS). These deposits melt and wet the material, typically yttria-stabilized zirconia, used as the thermal barrier coating, causing it to be drawn by capillarity into all of the open void space. Upon cooling, when the CMAS solidifies, the penetrated layer develops a high modulus. Since the thermal barrier coatings rely on spatially configured voids to achieve strain tolerance with the superalloy substrate, those regions penetrated by the CMAS can be detrimental, causing the thermal barrier coating to be susceptible to extensive spallation when subjected to subsequent thermal cycles. Thermal barrier coating spallation can lead to a drastic reduction in the turbine engine component durability and to a direct attack on the underlying substrate. 
     A common approach to deal with this problem is to deposit an extra layer over the thermal barrier coating. However, this extra layer is not activated prior to the introduction of the component into service. As a result, it does not become activated until CMAS is encountered in service. CMAS comes with a variety of chemical compositions depending upon its geographical origin. Consequently, the effectiveness of this extra layer however is unknown and can be less than desirable. 
     SUMMARY 
     In this disclosure, the CMAS risk is mitigated and the durability of a thermal barrier coating can be increased by creating an activated reactive layer with known CMAS reaction kinetics on thermal barrier coating, prior to service. 
     In accordance with the instant disclosure, there is described a process for forming a coating system on a turbine engine component which broadly comprises the steps of: providing a substrate; depositing a thermal barrier coating on the substrate; depositing a reactive with known CMAS reaction kinetics on the thermal barrier coating; and activating the reactive layer prior to the component being placed in service. 
     Further, in accordance with the present disclosure, there is provided a turbine engine component which broadly comprises a substrate; a thermal barrier coating deposited on the substrate; a reactive layer deposited on the thermal barrier coating, which reactive layer has known CMAS reaction kinetics and is activated prior to the turbine engine component entering into service. 
     Other details of the process and the turbine engine component are set forth in the following detailed description and the accompanying drawings, wherein like reference numerals depict like elements. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIGS. 1 and 2  are schematic representations of a turbine engine component having a coating system with an outer CMAS reactive layer; and 
         FIGS. 3 and 4  are schematic representations of a turbine engine component having an alternative coating system with an outer CMAS reactive layer. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT(S) 
     Referring now to  FIGS. 1 and 2 , there is shown a turbine engine component  10 , such as a combustor panel or a turbine blade. The component  10  has a substrate  12  which may be formed from any suitable material known in the art, such as a ceramic composite material, a nickel based superalloy, a cobalt based alloy, or a titanium based alloy. 
     If desired, an optional bond coat  14  may be deposited on a surface of the substrate  12 . The bond coat  14  may be deposited using any suitable technique known in the art and may comprise a material selected from the group consisting of aluminides, platinum alumindes, and MCrAlYs, where M is selected from the group consisting of nickel, cobalt, iron, and mixtures thereof. 
     Deposited on a surface of the bond coat  14 , when the bond coat  14  is present, is a thermal barrier coating  16 . If the bond coat  14  is not present, the thermal barrier coating  16  may be deposited directly on a surface of the substrate  12 . The thermal barrier coating  16  may be formed from any suitable material known in the art. For example, the thermal barrier coating  16  can be a yttria stabilized zirconia deposited using an APS technique or EBPVD technique. The thermal barrier coating  16  may contain crystallization promoting elements, such as La 2 Zr 2 O 7 , Gd 2 Zr 2 O 7 , Al 2 O 3 , TiO 2 , ZrO 2  and mixtures thereof. One suitable thermal barrier coating composition could be an APS deposited coating having a composition of (YSZ+A+B), where YSZ is yttria partially stabilized zirconia, and A and B are at least two of the crystallization promoting elements. Crystallization is a conversion of glass to fine-grained crystalline “glass-ceramics”. 
     If desired, a layer  20  may be deposited between the bond coat  14  and the layer  16 . The layer  20  may be a conventional thermal barrier coating with YSZ or any suitable thermal barrier coating system. 
     Thereafter, an exterior CMAS resistant layer  18  may be deposited on the thermal barrier coating  16 . The layer  18  may be a thin film of up to approximately 50 μm thick chemically conditioned CMAS material containing at least one reactive element, which can be selected from the group consisting of Gd 2 Zr 2 O 7  and TiO 2 . 
     To create the CMAS layer  18 , a powder mixture is prepared, having a nominal composition in Table 1. 
                     TABLE 1                  Nominal composition of the conditioned CMAS in mol. %                                         SiO 2     CaO   MgO   Al 2 O 3     Na 2 O   K 2 O   Fe 2 O 3                                           45   33   9   13                    
The chemical composition provides a powder mixture, which after thermo-chemical reaction with the thermal barrier coating layer  16 , has a melting temperature significantly higher, at least 50 degrees Fahrenheit higher, than the melting temperature of the CMAS encountered in field engine applications.
 
     The CMAS resistant layer  18  may be deposited using a solution-precursor plasma spray (SPPS) or a conventional APS procedure. The spray method of SPPS is identical to APS deposition with the substitution of a solution atomizer for the solid powder feeder. Plasma torch operating power and gas flow rates are in the same range as used for depositing APS coatings. 
     If desired, the top CMAS layer  18  may have graded characteristics from the interface with the thermal barrier coating  16  to the outer surface of the layer  18  for better adherence and strain compatibility. 
     After deposition, the deposited film of CMAS is subjected to a heat treatment at 2100° F.-2200° F. for up to 24 hours to form an active reaction layer. In this way, the CMAS layer is activated before the layer comes into contact with CMAS during service. 
     A reactive layer  18  formed as set forth hereinabove has CMAS reaction kinetics. The heat treatment turns the CMAS layer into CMAS glass, which reacts with the layer  16  underneath with nucleating agent, transitions to form globular anorthide phase, or to form a cubic garnet type crystal structure, depending upon the nucleating agents that are used. 
     The actively formed reaction layer  18  over the thermal barrier coating  16  has a melting temperature, which is significantly higher than the common CMAS encountered in the field. Upon contact with the common CMAS in a high temperature environment, the common CMAS with lower melting temperature will not readily dissolve into the reaction layer. Even if it dissolves into the reaction layer at higher temperature, it will recrystallize quickly within the reaction layer, which prohibits further CMAS penetration into the thermal barrier coating layer. Protecting the thermal barrier coating from spallation can significantly increase the durability of the turbine engine component  10  and thus save maintenance costs. The process described herein is simple and should be effective in protecting the turbine engine component by reducing and/or eliminating CMAS attacks. 
     Referring now to  FIGS. 3 and 4 , there is shown an alternative coating system deposited on the substrate  12 . The coating system may have an optional bond coat  14 , a thermal barrier coating  20  and an outer engineered CMAS layer  22  with at least one recrystallization agent discussed hereinbefore. As shown in  FIG. 4 , a layer  24  can be provided between the bond coat  14  and the thermal barrier coating  20 . The layer  22  in each embodiment can be applied by APS or SPPS. 
     It is apparent that there has been provided in accordance with the instant disclosure a CMAS resistant TBC coating. While a specific embodiment of the coating has been described herein, other unforeseeable alternatives, variations, and modifications may become apparent to those skilled in the art having read the foregoing description. Accordingly, it is intended to embrace those alternatives, modifications, and variations which fall within the broad scope of the appended claims.