Patent Publication Number: US-11034842-B2

Title: Coating for improved surface finish

Description:
STATEMENT OF GOVERNMENT RIGHTS 
     The invention described in the present disclosure was made with the support of the U.S. Government under contract number DE-FE0024006, which was awarded by the Department of Energy. The U.S. Government has certain rights in this invention. 
    
    
     BACKGROUND 
     The present subject matter relates generally to coatings, and more specifically to coatings for gas turbine engines. 
     As the demand for more efficient gas turbine engines drives internal operating temperatures higher, the transition to higher temperature materials has driven the transition from metallic nickel-based superalloys to ceramic matrix composites (CMC), which offer both mechanical strength as well as resistance to high temperatures. However, CMCs may have higher production costs and/or longer manufacturing cycle times compared to metallic nickel-based superalloys. In addition, the geometries and/or topologies in which CMC components can be formed may be limited compared to metallic nickel-based superalloys. 
     The orthotropic nature of CMCs can lead to porosity and inconsistent machined surface finish. Porosity in a CMC sealing surfaces (i.e., in applications where CMCs are installed within or proximate gas turbine hot gas paths (HGP)) may allow cooling flow leakage past the seals, thereby resulting in efficiency losses. In addition, inconsistent machining, due to, for example varying speeds of material removal of ceramic matrices and/or fibers, can lead to ripples or rough surfaces in finished CMC components. Rough surfaces may create undulations and gaps between the CMC seals and sealing surfaces for cooling flow to leak through, again leading to efficiency losses in gas turbine engines. 
     BRIEF DESCRIPTION OF THE EMBODIMENTS 
     Aspects of the present embodiments are summarized below. These embodiments are not intended to limit the scope of the present claimed embodiments, but rather, these embodiments are intended only to provide a brief summary of possible forms of the embodiments. Furthermore, the embodiments may encompass a variety of forms that may be similar to or different from the embodiments set forth below, commensurate with the scope of the claims. 
     In one aspect, a coating includes: at least 34.9 percent by mass silicon dioxide; at least 9.1 percent by mass aluminum oxide; and at least 16.1 percent by mass yttrium oxide. 
     In another aspect, a coating includes: at least 9.8 percent by barium oxide; at least 5.2 percent by mass aluminum oxide; and at least 40.3 percent by mass silicon dioxide. 
     In another aspect, a coating system includes a S-A-Y material composition including: from about 49 percent to about 59 percent by mass silicon dioxide; from about 13 percent to about 23 percent by mass aluminum oxide; and from about 23 percent to about 33 percent by mass yttrium oxide. The coating system includes a B-A-S material composition including: from about 15 percent to about 25 percent by mass barium oxide; from about 8 to about 18 percent by mass aluminum oxide; and from about 62 percent to about 72 percent by mass silicon dioxide. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       These and other features, aspects, and advantages of the present disclosure will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein: 
         FIG. 1  is a side schematic representation of a gas turbine engine; 
         FIG. 2  is a side schematic representation of a turbine airfoil; 
         FIG. 3  is summary of coating composition test points; 
         FIG. 4  is summary of coating composition that were selected for further study; and 
         FIG. 5  is a method of forming a CMC component with a coating disposed thereon, according to aspects of the present embodiments. 
     
    
    
     Unless otherwise indicated, the drawings provided herein are meant to illustrate features of embodiments of the disclosure. These features are believed to be applicable in a wide variety of systems comprising one or more embodiments of the disclosure. As such, the drawings are not meant to include all conventional features known by those of ordinary skill in the art to be required for the practice of the embodiments disclosed herein. 
     DETAILED DESCRIPTION 
     In the following specification and the claims, reference will be made to a number of terms, which shall be defined to have the following meanings. 
     The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise. 
     “Optional” or “optionally” means that the subsequently described event or circumstance may or may not occur, and that the description includes instances where the event occurs and instances where it does not. 
     Approximating language, as used herein throughout the specification and claims, may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about” and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value. Here and throughout the specification and claims, range limitations may be combined and/or interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. 
     As used herein, the term “axial” refers to a direction aligned with a central axis or shaft of a gas turbine. 
     As used herein, the term “circumferential” refers to a direction or directions around (and tangential to) the outer circumference of the gas turbine, or for example the circle defined by the swept area of the rotor of the gas turbine. As used herein, the terms “circumferential” and “tangential” may be synonymous. 
     As used herein, the term “radial” refers to a direction moving outwardly away from the central axis of the gas turbine. A “radially inward” direction is aligned toward the central axis moving toward decreasing radii. A “radially outward” direction is aligned away from the central axis moving toward increasing radii. 
     Referring now to the drawings, wherein like numerals refer to like components,  FIG. 1  illustrates an example of a gas turbine  10  which may incorporate various aspects of the embodiments disclosed herein. As shown, the gas turbine  10  generally includes a compressor section  12  having an inlet  14  disposed at an upstream end of the gas turbine  10 , and a casing  16  that at least partially surrounds the compressor section  12 . The gas turbine  10  further includes a combustion section  18  having at least one combustor  20  downstream from the compressor section  12 , and a turbine section  22  downstream from the combustion section  18 . As shown, the combustion section  18  may include a plurality of the combustors  20 . A shaft  24  extends axially through the gas turbine  10 .  FIG. 1  illustrates the radial  94 , axial  92  and circumferential directions  90 . 
     Referring still to  FIG. 1 , the gas turbine  10  may include a transition piece  52  disposed between a downstream end of the combustor  20  and an upstream end of the turbine section  22 . The combustor  20  may include a combustor liner  54  defining the boundaries of the combustor  20 . The combustion section  18  may include a plurality of substantially cylindrical “can-style” combustors  20  circumferentially spaced around the gas turbine  10 , in which case the combustor liners  54  may also be substantially cylindrical. In other embodiments, the combustion section  18  may include an annular combustor  20 , in which case the combustor liner  54  may include both an inner liner (not shown) defining a radially inner boundary of the annular combustor  20 , as well as an outer liner (not shown) defining a radially outer boundary of the annular combustor  20 . The gas turbine  10  may also include a stage one nozzle  56  located in the turbine section  22  and disposed axially aft of the transition piece  52 , as well as a second stage nozzle  58  disposed downstream of a first stage turbine rotor  36 . The gas turbine  10  may also include one or more flow path ducts  60  defining a radially outer boundary of a turbine gas path at axial locations between rotors and stators (i.e., blades  36  and nozzles  34 ). A gas turbine “hot section” may include both the combustor section  18  and components thereof, as well as the turbine section  22  and components thereof. 
     In operation, air  26  is drawn into the inlet  14  of the compressor section  12  and is progressively compressed to provide compressed air  28  to the combustion section  18 . The compressed air  28  flows into the combustion section  18  and is mixed with fuel in the combustor  20  to form a combustible mixture. The combustible mixture is burned in the combustor  20 , thereby generating a hot gas  30  that flows from the combustor  20  across a first stage  32  of turbine nozzles  34  and into the turbine section  22 . The turbine section generally includes one or more rows of rotor blades  36  axially separated by an adjacent row of the turbine nozzles  34 . The rotor blades  36  are coupled to the rotor shaft  24  via a rotor disk. The rotor shaft  24  rotates about an engine centerline CL. A turbine casing  38  at least partially encases the rotor blades  36  and the turbine nozzles  34 . Each or some of the rows of rotor blades  36  may be concentrically surrounded by a shroud block assembly  40  that is disposed within the turbine casing  38 . The hot gas  30  rapidly expands as it flows through the turbine section  22 . Thermal and/or kinetic energy is transferred from the hot gas  30  to each stage of the rotor blades  36 , thereby causing the shaft  24  to rotate and produce mechanical work. The shaft  24  may be coupled to a load such as a generator (not shown) so as to produce electricity. In addition, or in the alternative, the shaft  24  may be used to drive the compressor section  12  of the gas turbine. 
       FIG. 2  illustrates an enlarged side view of a portion of the turbine section  22  including an exemplary rotor blade  36  and a portion of a shroud block assembly  40  according to various embodiments of the present disclosure. The turbine rotor blade or airfoil  36 , extends from an axially forward leading edge  44  to an axially aft trailing edge  46 , and from a radially inward root  48  to a radially outer tip  42 . The airfoil  36  includes a platform  50  defining a radially inner boundary of a hot gas path. As shown in  FIG. 2 , the shroud block assembly  40  generally extends in a radial direction  94  outward from the airfoil  36  between the turbine casing  38  (not shown) and a tip portion  42  of the rotor blade  36 . The shroud block assembly  40  generally includes mounting hardware  62  for securing a plurality of shroud block segments  100  to the shroud block assembly  40 . The plurality of shroud block segments  100  may be arranged circumferentially  90  in an annular array around the rotor blades  36  within the turbine casing  38  (not shown). 
     Still referring to  FIG. 2 , each shroud block segment  100  may include a slash face  64  forming a circumferential interface with an adjacent shroud block segment  100 . Stated otherwise, the plurality of shroud block segments  100  may be arranged circumferentially  90 , and the slash face  64  of each shroud block segment  100  may contact and/or be adjacent to the slash face  64  of an adjacent shroud block segment  100 . Each shroud block segment  100  may also include a shroud segment forward edge  68 , a shroud segment aft edge  66 , and a shroud hot gas surface  76 . The shroud segment forward edge  68  and the shroud segment aft edge  66  are located on the axially forward and aft ends respectively of each shroud block segment  100 , while the shroud hot gas surface  76  is disposed at a radially inward surface of each shroud block segment  100  and forms the radially outer boundary of the turbine hot gas path. The platform  50  may also include a platform forward edge  72 , a platform aft edge  70 , and a platform circumferential edge  74  disposed at the axially forward, axially aft and circumferential ends, respectively, of each platform  50 . Coatings of the embodiments disclosed herein may be disposed at one or more of the transition piece  52 , the combustor liners  54 , the first stage nozzle  56 , the second stage nozzle  58 , other turbine nozzles  34 , flow-path ducts  60 , the slash face  64 , the shroud segment forward edge  68 , the shroud segment aft edge  66 , the shroud hot gas surface  76 , the platform forward edge  72 , the platform aft edge  70 , the platform circumferential edge  74 , as well as on other seal surfaces of CMC components. 
       FIGS. 3 and 4  summarize a series of test points that were undertaken to quantify various material properties of the coatings of the embodiments disclosed herein. In each case, a coating with a thickness between about 0.01 inches and about 0.07 inches was applied to a crucible using various methods. In one or more cases, a coating with a thickness between about 0.02 inches and about 0.05 inches was used. Any crucible with suitable temperature resistance and surface qualities may be used. For example, the crucible may include one or more graphite crucibles, ceramic rods with rounded edges, machined ceramic surfaces, and/or other suitable crucibles. The coating may be applied to the crucible via a tape, mold, air spray, manually brushed on, applied as a powder and subsequently crystallized, and/or via other suitable techniques. In addition, various modifications may be made to the application process to arrive at the desired coating thickness, density, etc. For example, the flow rate at which the coating is applied via an air spray application process may be adjusted. In addition, the rate at which the crucible is translated under the application equipment (and/or the rate at which the application equipment is moved over the crucible) may be adjusted. The height from which the coating is applied as well as the pressure at which the coating is applied may both also be adjusted. 
       FIG. 3  illustrates the processing temperatures  356  at which a number of material systems  352  including specific compositions  354  where tested, as well as the resulting coefficients of thermal expansion (CTE)  358  and softening points  360 . Test points  302 ,  304 ,  306 , and  308  were all performed on a B-A-S material system including compositions with various mass percentages of barium oxide (i.e., BaO and “B” in “B-A-S”), aluminum oxide (i.e., Al2O3 and “A” in “B-A-S”) and silicon dioxide (i.e., SiO2 and “S” in “B-A-S”). For example, test point  302  includes 33 percent by mass silicon dioxide, 61.5 percent by mass barium oxide, and 5.5 percent by mass aluminum oxide, while test point  304  includes 40.3 percent by mass silicon dioxide, 51.3 percent by mass barium oxide, and 8.4 percent by mass aluminum oxide, as indicated by the composition percentages listed in the composition column  354  of  FIG. 3 . Test point  302 , which was performed at a processing temperature  356  of 1300° C., resulted in a CTE  358  of 12.4 (×10{circumflex over ( )}−6/(° C.)) and a softening point  360  greater than 1200° C. Although test point  302  resulted in a favorable softening point  360  greater than 1200° C., the CTE  358  was higher than desirable. Test point  304 , which was also performed at a processing temperature  356  of 1300° C. but on a different B-A-S composition than test point  302 , resulted in a CTE  358  of 8.9 (×10{circumflex over ( )}−6/(° C.)) and a softening point  360  of 815° C. Thus, test point  304  resulted in neither a desired CTE  358  nor a desired softening point  360 . Both test points  306  and  308 , which were performed on various B-A-S compositions at 1400° C. and 1300° C. respectively resulted in favorable CTE  358  (i.e., equal to or below about 8.0 (×10{circumflex over ( )}−6/(° C.))), but lower than desirable softening points  360 . A CTE  358  was not quantified for test point  308 . 
     Referring still to  FIG. 3 , test points  310 ,  312 , and  314  were performed on a B-A-S/SiC material system including compositions by mass of B-A-S and SiC (silicon carbide) as shown in the composition column  354 . The B-A-S material system that was used in each of test points  310 ,  312 , and  314  included 66.7 percent by mass silicon dioxide, 20 percent by mass barium oxide, and 13.3 percent by mass aluminum oxide (i.e., the same B-A-S material system as test point  306 ). Test point  310  was performed on a composition of 98% B-A-S and 2% SiC while test point  312  was performed on a coating with a composition of 80% B-A-S and 20% SiC. Each of test points  310 ,  312 , and  314  were performed at a processing temperature  356  of 1400° C. and resulted in favorable CTE  358  (i.e., equal to or below about 8.0 (×10{circumflex over ( )}−6/(° C.))), but lower than desirable softening points  360 . In addition, test points  312  and  314  resulted in porous coatings. 
     Still referring to  FIG. 3 , test points  316 ,  318 , and  320  were each performed on a S-A-Y material system including 54 percent by mass silicon dioxide (i.e., SiO2 and “S” in “S-A-Y”), 18.07 percent by mass aluminum oxide (i.e., Al2O3 and “A” in “S-A-Y”), and 27.93 percent by mass yttrium oxide (i.e., Y2O3 and “Y” in “S-A-Y”). Test points  316 ,  318 , and  320  were performed at processing temperatures  356  of 1300° C., 1350° C., and 1400° C., respectively and each resulted in favorable CTE  358  (i.e., equal to or below about 8.0 (×10{circumflex over ( )}−6/(° C.))) and favorable softening points  360  (i.e., greater than 1200° C.). However, test point  316  resulted in a porous coating. 
     Referring still to  FIG. 3 , test points  322 ,  324 ,  326 ,  328 ,  330 , and  332  were each performed on coating compositions include a mixture of S-A-Y and B2O3 material systems. The S-A-Y material system of each of test points  322 ,  324 ,  326 ,  328 ,  330 , and  332  includes 54 percent by mass silicon dioxide (i.e., SiO2 and “S” in “SAY”), 18.07 percent by mass aluminum oxide (i.e., Al2O3 and “A” in “S-A-Y”), and 27.93 percent by mass yttrium oxide (i.e., Y2O3 and “Y” in “S-A-Y”), similar to the S-A-Y material system of test points  316 ,  318 , and  320 . Each of test points  322 ,  324 ,  326 ,  328 ,  330 , and  332  also include boron oxide (i.e., B2O3). For example, test point  322  includes 97 percent by mass S-A-Y material and 3 percent by mass boron oxide while test points  324 ,  326 , and  328  each include 95 percent by mass S-A-Y material and 5 percent by mass boron oxide. Each of test points  322 ,  324 ,  326 ,  328 ,  330 , and  332  resulted in a favorable CTE  358  (i.e., equal to or below about 8.0 (×10{circumflex over ( )}−6/(° C.))), but only test points  322 ,  324 ,  326 , and  328  resulted in favorable softening points  360  (i.e., greater than 1200° C.). In addition, test points  322 ,  324 , and  326  all resulted in porous coatings. 
     Still referring to  FIG. 3 , test points  334 ,  336 ,  338 ,  340 ,  342 ,  344 ,  346 ,  348 , and  350  were each performed on coating compositions including a mixture of S-A-Y and B-A-S material systems, in different ratios and at different processing temperatures  356 . Each of test points  334 ,  336 ,  338 ,  340 ,  342 ,  344 ,  346 ,  348 , and  350  include a S-A-Y material system including 54 percent by mass silicon dioxide (i.e., SiO2 and “S” in “S-A-Y”), 18.07 percent by mass aluminum oxide (i.e., Al2O3 and “A” in “SAY”), and 27.93 percent by mass yttrium oxide (i.e., Y2O3 and “Y” in “S-A-Y”), similar to the S-A-Y material system of test points  316 - 332 . Each of test points  334 ,  336 ,  338 ,  340 ,  342 ,  344 ,  346 ,  348 , and  350  include a B-A-S material system including 66.7 percent by mass silicon dioxide, 20 percent by mass barium oxide, and 13.3 percent by mass aluminum oxide (i.e., the same B-A-S material system as test points  306 ,  310 ,  312 , and  314 ). Each of test points  334 ,  336 ,  338 ,  340 ,  342 ,  344 ,  346 ,  348 , and  350  resulted in a favorable CTE  358  (i.e., equal to or below about 8.0 (×10{circumflex over ( )}−6/(° C.))). Test points  334 ,  336 ,  338 ,  340 ,  342 ,  346 , and  348  (i.e., all but test points  344  and  350 ) resulted in favorable softening points  360  (i.e., greater than 1200° C.). In addition, test points  334 ,  336  and  340  resulted in porous coatings. 
       FIG. 4  illustrates a summary of each of the non-porous coatings that include favorable CTE  358  and softening temperatures  360  from  FIG. 3 . As such, the coatings resulting from test points  318 ,  320 ,  328 ,  338 ,  342 ,  346 , and  348  are summarized included in  FIG. 4 . Coatings resulting from these seven test points were further studied to understand surface roughness. The coatings resulting from test points  318 ,  320 ,  328 ,  338 , and  346  were found to have rough surfaces while the coatings resulting from test points  342  and  348  were found to have smooth surfaces. The surface finish was quantified in terms of roughness average, Ra, with units in micro-inches (inches×10{circumflex over ( )}−6). Roughness Average, Ra is calculated as the average of a surface&#39;s measured microscopic peaks and valleys, Rough surfaces may include surfaces with a roughness average of above about 150 Ra, or from about 150 Ra to about 250 Ra. Smooth surfaces may include surfaces with a roughness average of below about 150 Ra, or below about 100 Ra. In other embodiments, smooth surfaces may include surfaces with a roughness average between about 10 Ra and about 80 Ra. In other embodiments, smooth surfaces may include surfaces with a roughness average between about 15 Ra and about 60 Ra. In other embodiments, smooth surfaces may include surfaces with a roughness average between about 20 Ra and about 50 Ra. In other embodiments, smooth surfaces may include surfaces with a roughness average between about 30 Ra and about 40 Ra. Smooth surfaces may include surfaces with a roughness average between about 10 Ra and about 150 Ra, and all sub-ranges therebetween. 
     The compositions of the material systems included in  FIGS. 3 and 4  may vary around the exact constituents shown. For example, the S-A-Y composition including approximately 54% S, 18% A and 28% Y (i.e., test points  316 ,  318 , and  320 ) may include tolerance bands of 2%, 3%, and even 5% around each of the constituents. For example, the S-A-Y composition may include from about 49% to about 59% S, from about 13% to about 23% A, and from about 23% to about 33% Y, assuming 5% tolerance bands. Similarly, the S-A-Y composition may include from about 51% to about 57% S, from about 15% to about 21% A, and from about 25% to about 31% Y, assuming 3% tolerance bands. Similarly, the S-A-Y composition may include from about 52% to about 56% S, from about 16% to about 20% A, and from about 26% to about 30% Y, assuming 2% tolerance bands. Therefore, for the test points that include a mix of S-A-Y and B-A-S material systems (i.e., test points  334 - 350 ) which includes as little as 70% of S-A-Y material system in the overall composition, the overall composition may include as little as 70% of the lower limits of each of the ranges resulting from 5% tolerance bands. For example, the S-A-Y:B-A-S material systems in a 70:30 ratio may include as little as about 34.3 percent by mass S (i.e. silicon dioxide), about 9.1 percent by mass aluminum oxide, and about 16.1 percent by mass yttrium oxide, when accounting for the constituents of the S-A-Y material system. When also accounting for the B-A-S material system, in a 90:10 SAY:B-A-S ratio, the silicon dioxide may increase from about 34.3 percent to about 40.5 percent (i.e., an addition of 6.2% of the overall composition due to the silicon dioxide in the B-A-S material system, while the aluminum oxide may increase from about 9.1 to about 9.9 percent (i.e., an addition of about 0.8% of the overall composition due to the aluminum oxide in the B-A-S material system). 
     The B-A-S material composition of test point  306  was repeated again for test points  310 - 314  and  334 - 350  due to the low CTE  358 . This material composition included 66.7% S, 20% B, and 13.3% A. When surrounded by 5% tolerance bands, these percentages range from about 62% to about 72% S, about 15% to about 25% B, and about 8% to about 18% A. When surrounded by 3% tolerance bands, these percentages range from about 64% to about 70% S, about 17% to about 23% B, and about 10% to about 16% A. When surrounded by 2% tolerance bands, these percentages range from about 65% to about 79% S, about 18% to about 22% B, and about 11% to about 15% A. 
       FIG. 5  illustrates a method  500  of forming CMC components coating with coatings according to the embodiments disclosed herein. At step  502 , the method includes providing a CMC component. The CMC component may include a finished CMC component, a CMC component that is in a green state, and/or a CMC component that has been solidified but that may require one or more post-processing steps such as heat treat and/or machining. At step  504 , the method may include performing an EDM (electro discharge machining) process on the CMC component. At step  506 , the method  500  may include performing a grinding process on the CMC component. Steps  504  and  506  may be performed in any order. In some embodiments, both steps  504  and  506  are performed. In some embodiments, neither step  504  nor step  506  is performed. In some embodiments, only one of steps  504  and  506  is performed. At step  508 , the method  500  may include depositing a coating on the CMC component. The coating may be deposited on the CMC component via spray, plasma spray, and/or air spray, tape, manual brushing, and/or via powder which may be subsequently crystallized and/or reacted onto the CMC component. At step  510 , the method  500  may include heat treating the coated CMC component. At step  512 , the method  500  may include grinding the coated CMC component. At step  514 , the method  500  may include honing the coated CMC component. At step  516 , the method  500  may include extrude honing the coated CMC component. At step  518 , the method  500  may include polishing the coated CMC component. In some embodiments, other steps may be performed. In some embodiments, one or more of steps  502  through  518  may be omitted. In some embodiments, one or more of steps  502  through  518  may be performed in a different order that what is illustrated in  FIG. 5 . 
     Coatings and/or sealants of the embodiments disclosed herein may be disposed at each of the locations discussed above in which a CMC surface and/or component defines a flow-path boundary and/or defines an interfacing surface with an adjacent component (i.e., thereby forming a seal). Stated otherwise, it may be desirable to dispose the coating of the present embodiments on any CMC boundary and/or interfacing surface on which an enhanced seal and/or improved surface finish is desired. The surfaces on which the coatings may be disposed may include one or more of the transition piece  52 , the combustor liners  54 , the first stage nozzle  56 , the second stage nozzle  58 , other turbine nozzles  34 , flow-path ducts  60 , one or more slash faces  64 , one or more shroud segment forward edges  68 , one or more shroud segment aft edges  66 , one or more shroud hot gas surfaces  76 , a platform forward edge  72 , a platform aft edge  70 , a platform circumferential edge  74 , as well as on other seal surfaces of CMC components. 
     Each of the coatings according to the embodiments disclosed herein may include other oxides other than those listed above. For example, boron oxide may include boron dioxide, boron trioxide, boron monoxide, and/or boron suboxide. Similarly, aluminum oxide may include aluminum (I) oxide, aluminum (II) oxide, and/or aluminum (III) oxide. Each of the coatings according to the embodiments disclosed herein may include a glassy surface that remains smooth in operation. By contrast, other conventional EBCs (environmental barrier coating) and/or TBCs (thermal barrier coatings) may be brittle and may chip away over time, when exposed to gas turbine internal operating temperatures. Each of the coatings according to the embodiments disclosed herein may include a CTE (coefficient of thermal expansion) that is substantially similar to that of the CMC substrate on which they are disposed. For example, in some embodiments, the CTE of the coatings disclosed herein are within about 50% of the CTE of the CMC substrate on which they are disposed. In other embodiments, the CTE of the coatings disclosed herein are within about 20% of the CTE of the CMC substrate on which they are disposed. In other embodiments, the CTE of the coatings disclosed herein are within about 10% of the CTE of the CMC substrate on which they are disposed. In other embodiments, both the coatings disclosed herein as well as the CMC substrate on which they are disposed include a CTE that is equal to or less than about 8.0 (×10{circumflex over ( )}−6/(° C.)). With conventional coatings, undulations may form on seal surfaces which may allow undesired leakages to flow past. By including a CTE that approximately matches the CMC substrate on which they are disposed, by including a high softening temperature, and by remaining smooth when exposed to internal gas turbine operating temperatures, the coatings disclosed herein may provide a smooth sealing surface substantially free from undulations, thereby resulting in enhanced sealing. Stated otherwise, the coatings disclosed herein may prevent mating and/or interfacing CMC seal surfaces from developing undulations and/or gaps, which may lead to increases in undesired leakage flows. 
     Although specific features of various embodiments of the present disclosure may be shown in some drawings and not in others, this is for convenience only. In accordance with the principles of the present disclosure, any feature of a drawing may be referenced and/or claimed in combination with any feature of any other drawing. 
     This written description uses examples to disclose the embodiments of the present disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the embodiments described herein is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.