Patent Publication Number: US-2019195080-A1

Title: Ceramic coating system and method

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This disclosure is a continuation-in-part of U.S. application Ser. No. 14/812,668, filed Jul. 29, 2015, which claims priority to U.S. Provisional Application No. 62/033,883, filed on Aug. 6, 2014 and is incorporated herein by reference. 
    
    
     BACKGROUND 
     A gas turbine engine typically includes a fan section, a compressor section, a combustor section, and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. 
     Components that are exposed to high temperatures during operation of the gas turbine engine typically require protective coatings. For example, components such as turbine blades, turbine vanes, blade outer air seals (BOAS), and compressor components may require at least one layer of coating for protection from the high temperatures. 
     Some BOAS for a turbine section include an abradable ceramic coating that contacts tips of the turbine blades such that the blades abrade the coating upon operation of the gas turbine engine. The abradable material allows for a minimum clearance between the BOAS and the turbine blades to reduce gas flow around the tips of the turbine blades to increase the efficiency of the gas turbine engine. Over time, internal stresses can develop in the protective coating to make the coating vulnerable to erosion and spalling. The BOAS may then need to be replaced or refurbished after a period of use. Therefore, there is a need to increase the longevity of protective coatings in gas turbine engines. 
     SUMMARY 
     A gas turbine engine article according to an example of the present disclosure includes a substrate that has at least one step. The at least one step has an undercut and a thermally insulating topcoat disposed on the substrate. The thermally insulating topcoat has at least one fault extending from the at least one step. 
     In a further embodiment of any of the foregoing embodiments, the at least one step includes, relative to an outer surface of the thermally insulating topcoat, a proximal surface, a distal surface, and a sidewall that joins the proximal surface and the distal surface, and the undercut is in the sidewall. 
     In a further embodiment of any of the foregoing embodiments, the sidewall defines a linear distance between the proximal surface and the distal surface, and the undercut defines a linear height of at least about 10% of the linear distance. 
     In a further embodiment of any of the foregoing embodiments, the sidewall defines a linear distance between the proximal surface and the distal surface, and the undercut defines a lateral undercut distance that is at least about 5% of the linear distance. 
     In a further embodiment of any of the foregoing embodiments, the at least one step includes, relative to an outer surface of the thermally insulating topcoat, a proximal surface, a distal surface, and a sidewall that joins the proximal surface and the distal surface, and the undercut is in the distal surface. 
     In a further embodiment of any of the foregoing embodiments, the sidewall defines a linear distance between the proximal surface and the distal surface, and the undercut has a linear height of less than about 50% of the linear distance. 
     In a further embodiment of any of the foregoing embodiments, the sidewall defines a diametric distance, and the undercut defines a lateral undercut distance that is less than about 50% of the diametric distance. 
     In a further embodiment of any of the foregoing embodiments, the at least one step includes, relative to an outer surface of the thermally insulating topcoat, a proximal surface, a distal surface, and a sidewall that joins the proximal surface and the distal surface, and the sidewall and the proximal surface meet at a 90° corner. 
     In a further embodiment of any of the foregoing embodiments, the at least one step is annular. 
     In a further embodiment of any of the foregoing embodiments, the at least one step includes a plurality of steps in a pattern. 
     In a further embodiment of any of the foregoing embodiments, the at least one fault is a microstructural discontinuity in the topcoat. 
     In a further embodiment of any of the foregoing embodiments, the fault extends to a surface of the thermally insulating topcoat. 
     A gas turbine engine according to an example of the present disclosure has a plurality of rotatable blades, and a seal arranged radially outwards of the plurality of rotatable blades. The seal has a substrate that has at least one step. The at least one step has an undercut, and a thermally insulating topcoat disposed on the substrate. The thermally insulating topcoat has at least one fault extending from the at least one step. 
     In a further embodiment of any of the foregoing embodiments, the at least one step includes, relative to an outer surface of the thermally insulating topcoat, a proximal surface, a distal surface, and a sidewall that joins the proximal surface and the distal surface, and the undercut is in the sidewall. 
     In a further embodiment of any of the foregoing embodiments, the at least one step includes, relative to an outer surface of the thermally insulating topcoat, a proximal surface, a distal surface, and a sidewall that joins the proximal surface and the distal surface, and the undercut is in the distal surface. 
     In a further embodiment of any of the foregoing embodiments, the at least one step is annular. 
     A method for fabricating a gas turbine engine article according to an example of the present disclosure includes forming at least one step in a substrate. The at least one step has an undercut, which deposits a thermally insulating topcoat on the substrate. The thermally insulating topcoat forms at least one fault during the depositing that extends from the at least one step. 
     In a further embodiment of any of the foregoing embodiments, the forming includes forming the at least one step and undercut using at least one of additive manufacturing, chemical milling, or mechanical milling. 
     In a further embodiment of any of the foregoing embodiments, the at least one step includes, relative to an outer surface of the thermally insulating topcoat, a proximal surface, a distal surface, and a sidewall that joins the proximal surface and the distal surface, and the undercut is formed in the sidewall. 
     In a further embodiment of any of the foregoing embodiments, the at least one step includes, relative to an outer surface of the thermally insulating topcoat, a proximal surface, a distal surface, and a sidewall that joins the proximal surface and the distal surface, and the undercut is formed in the distal surface. 
     The various features and advantages of this disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  illustrates an example gas turbine engine. 
         FIG. 2  illustrates a turbine section of the gas turbine engine of  FIG. 1 . 
         FIG. 3  illustrates an example portion of a turbine component. 
         FIG. 4  illustrates a perspective view of another example turbine component. 
         FIG. 5  illustrates another perspective view of the turbine component of  FIG. 4 . 
         FIG. 6  illustrates an example portion of the turbine component of  FIG. 4 . 
         FIG. 7A  illustrates a cross-section of a representative portion of another example gas turbine engine article. 
         FIG. 7B  illustrates a radial outward view of the gas turbine engine article of  FIG. 7A . 
         FIG. 8A  illustrates a cross-section of a representative portion of another example gas turbine engine article. 
         FIG. 8B  illustrates a radial outward view of the gas turbine engine article of  FIG. 8A . 
         FIG. 9  illustrates a comparative example of a deposition process using a substrate that has a step without an undercut. 
         FIG. 10  illustrates a comparative example of a deposition process using a substrate that has a step with an undercut. 
         FIG. 11  illustrates an example method for fabricating a gas turbine engine article. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct defined within a nacelle  15 , while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. 
     The exemplary engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The high speed spool  32  includes an outer shaft  50  that interconnects a second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     The core airflow is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core airflow path C. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture  48  is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low pressure turbine  46  has a pressure ratio that is greater than about five 5:1. Low pressure turbine  46  pressure ratio is pressure measured prior to inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)] 0.5 . The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second). 
       FIG. 2  illustrates a portion of the turbine section  28  of the gas turbine engine  20 . Turbine blades  60  receive a hot gas flow from the combustor section  26  ( FIG. 1 ). A blade outer air seal (BOAS) system  62  is located radially outward from the turbine blades  60 . The BOAS system  62  includes multiple seal members  64  circumferentially spaced around the axis A of the gas turbine engine  20 . Each seal member  64  is attached to a case  66  surrounding the turbine section by a support  68 . It is to be understood that the seal member  64  is only one example of an article within the gas turbine engine that may benefit from the examples disclosed herein. 
       FIG. 3  illustrates a portion of the seal member  64  having two circumferential sides  70  (one shown), a leading edge  72 , a trailing edge  74 , a radially outer side  76 , and a radially inner side  78  that is adjacent the hot gas flow and the turbine blade  60 . The term “radially” as used in this disclosure relates to the orientation of a particular side with reference to the axis A of the gas turbine engine  20 . 
     The seal member  64  includes a substrate  80 , a bond coat  82  covering a radially inner side of the substrate  80 , and a thermally insulating topcoat  84  covering a radially inner side of the bond coat  82 . In this example, the bond coat  82  covers the entire radially inner side of the substrate  80  and the thermally insulating topcoat  84  is a thermal barrier made of a ceramic material. The substrate  80  includes a slanted region  80   a  adjacent the leading edge  72  and a downstream portion  80   b  having a generally constant radial dimension. 
     The bond coat  82  includes a thicker region D1 adjacent the leading edge  72  and the trailing edge  74  and a thinner region D2 axially between the thicker regions D1. The thinner region D2 extends axially from upstream of the turbine blade  60  to downstream of the turbine blade  60 . 
     A step  86  is formed in the bond coat  82  between both of the thicker regions D1 and the thinner region D2. The step  86  extends in a radial and circumferential direction such that multiple BOAS systems  62  arranged together form a circumference around the axis A of the gas turbine engine  20  with the step  86  extending entirely around the circumference. 
     The step  86  includes a radially inner edge  88  having a radius R1 and a radially outer fillet  90  having a radius R2. In one example, the step  86  extends generally perpendicular to the axis A of the gas turbine engine  20 . In another example, the step  86  extends in a non-perpendicular direction such that the step forms an undercut. The step  86  extends for a radial thickness D3. 
     In one example, the sum of R1 and R2 equals less than or equal to 50% of the thickness of region D3. In another example, the sum of R1 and R2 equals less than or equal to 25% of the thickness of region D3. 
     The thermally insulating topcoat  84  includes a leading edge region  92  and a trailing edge region  94  having a thickness D4 and an axially central region  96  having a thickness D5. The central region  96  extends from axially upstream of the turbine blade  60  to axially downstream of the turbine blade  60 . The leading edge region  92  and the trailing edge region  94  are separated from the central region  96  by faults  98  extending radially through the thickness of the thermally insulating topcoat  84 . 
     The faults  98  extend from the steps  86  formed in the bond coat  82  and reduce internal stresses within the thermally insulating topcoat  84  that may occur from sintering of the thermal material at relatively high surface temperatures within the turbine section  28  during use of the gas turbine engine  20 . Although the central region  96  is separated from the trailing edge  74  by the trailing edge region  94 , the central region  96  could extend to the trailing edge  74 . 
     In one example, the thickness of region D1 is approximately 0.019 inches (0.483 mm), the thickness of region D4 is approximately 0.012 inches (0.305 mm), the thickness of region D2 is approximately 0.007 inches (0.178 mm), the thickness of region D3 is approximately 0.012 inches (0.305 mm) and the thickness of region D5 is approximately 0.025 inches (0.635 mm). In one example, at least one of the radius R1 and the radius R2 are approximately 0.003 inches (0.076 mm). In another example, at least one of the radius R1 and the radius R2 are less than 0.004 inches (0.102 mm). In yet another example, at least one of the radius R1 and the radius R2 are less than 0.005 inches (0.127 mm). 
     Depending on the composition of the thermally insulating topcoat  84 , surfaces temperatures of about 2500° F. (1370° C.) and higher may cause sintering. The sintering may result in partial melting, densification, and diffusional shrinkage of the thermally insulating topcoat  84 . The faults  98  provide pre-existing locations for releasing energy associated with the internal stresses (e.g., reducing shear and radial stresses). That is, the energy associated with the internal stresses may be dissipated in the faults  98  such that there is less energy available for causing delamination cracking between the thermally insulating topcoat  84  and the bond coat  82 . 
     The faults  98  may vary depending upon the process used to deposit the thermally insulating topcoat  84 . In one example, the faults  98  may be gaps between adjacent regions. In another example, the faults  98  may be considered to be microstructural discontinuities between the adjacent regions  92 ,  94 , and  96 . The faults  98  may also be planes of weakness in the thermally insulating topcoat  84  such that the regions  92 ,  94 , and  96  can thermally expand and contract without cracking the thermally insulating topcoat  84 . 
     The material selected for the substrate  80 , the bond coat  82 , and the thermally insulating topcoat  84  are not necessarily limited to any kind. In one example, the substrate  80  is made of a nickel based alloy and the thermally insulating topcoat  84  is an abradable ceramic material suited for providing a desired heat resistance. 
     The faults  98  in the thermally insulating topcoat  84  on the seal member  64  may be formed during application of the thermally insulating topcoat  84 . Once the bond coat  82  has been applied to the substrate  80 , the bond coat  82  is machined or ground to form the step  86  with the radially outer fillet  90  and the radially inner edge  88  having the desired radius R2 and R1, respectively. Alternatively, the step  86  is formed in the substrate  80  and the bond coat  82  is only applied to the radially inward facing portions of the substrate  80  excluding the step  86  in order to facilitate formation of the fault  98  along the step  86 . Therefore, the substrate  80  would include a first portion have a first thickness and a section portion having a second thickness different from the first thickness 
     The thermally insulating topcoat  84  is applied to the bond coat  82  and/or substrate  80  with a thermal spray process. The thermal spray process allows the thermally insulating topcoat  84  to build up discontinuously such that there is no bridging between the leading edge region  92 , the central region  96 , and the trailing edge region  94 . Because of the discontinuity created by the step  86 , the continued buildup of the thermally insulating topcoat  84  between the central region  96  and the leading and trailing regions  92  and  94  forms the faults  98 . The radially inner side  78  of the seal member  64  may be machined to remove unevenness introduced by the varying thickness associated with thermal spraying the step  86 . 
       FIGS. 4-6  illustrate another example seal member  164 . The seal member  164  is similar to the seal member  64  except where described below or shown in the Figures. The seal member  164  includes the substrate  80  covered by a bond coat  182 . The bond coat includes a leading edge portion  182   a  axially upstream of a step  186  and a trailing edge portion  182   b  axially downstream of the step  186 . The leading edge portion  182   a  and the trailing edge portion  182   b  include geometric features  185  formed in the bond coat  182 . In this example, the geometric features  185  are cylindrical. However, other shapes such as elliptical or rectangular rods could be formed in the bond coat  182 . Alternatively, the geometric features  185  could be formed in the substrate  80  with the radially inner surface of the substrate  80  being covered with the bond coat  182 . 
     The thermally insulating topcoat  84  can be applied as discussed above. However, when the thermally insulating topcoat  84  is applied over the geometric features  185 , faults  199  will form in the thermally insulating topcoat  184  in addition to a fault  198  formed radially inward from the step  186 . The faults  198  and  199  form in a similar fashion as the faults  98  described above. 
       FIG. 7A  illustrates a cross-section of a representative portion of a seal member  264 . In this example, the seal member  264  includes a substrate  265  that has one or more steps  267 . Although only one representative step  267  is shown, the substrate  265  may include a plurality of such steps  267 . For example, the steps  267  may be provided in a pattern, similar to the pattern shown in  FIGS. 4 and 5 . The substrate  265  can be metallic, ceramic, or a combination thereof, and may be, or may include, a bond coat. 
     The step  267  includes undercut  269 , which as discussed below, facilitates the formation of the faults  298  in the overlying thermally insulating topcoat  284 . As used herein, the term “undercut” refers to a recessed region. Although an “undercut” may be formed by a cutting action, the term does not necessarily imply formation by cutting action. 
       FIG. 7B  shows a radially outward view of the portion of the seal member  264  but without the thermally insulating topcoat  284  (i.e. a view from the engine central longitudinal axis A). As shown, the step  267  in this example is annular. Although shown as a circular annulus, the step  267  may alternatively be elliptical, rectangular, toroidal, or other geometric shape. Likewise, the undercut  269  is also annular. 
     Relative to an outer surface  284   a  of the thermally insulating topcoat  284 , the step  267  includes ( FIG. 7A ) a proximal surface  271 , a distal surface  273 , and a sidewall  275  that joins the proximal surface  271  and the distal surface  273 . The proximal surface  271  and the distal surface  273  face radially inwardly, toward the core flow path and engine central axis A. The sidewall  275  and the proximal surface  271  meet at a 90° corner  277 . In this example, the undercut  269  is in the sidewall  275 . The undercut  269  is thus a recessed region of the sidewall  275 . 
       FIG. 8A  shows a modified version of the seal member  264 . In this example, rather than the undercut  269  in the sidewall  275 , there is an undercut  369  in the distal surface  273 . The undercut  369  is thus a recessed region of the distal surface  273 . 
     The undercuts  269 / 369  facilitate formation of the faults  298  that extend from the step  267  by avoiding or reducing the potential for bridging of the thermally insulating topcoat  284  during spray deposition (e.g., thermal spray) of the topcoat  284 . To illustrate,  FIGS. 9 and 10  each show three progressions through a spray deposition process. In  FIG. 9  the step, S 1 , does not have an undercut. In the first progression at P1 on the left-hand side, the coating, C, begins to build up on the surfaces around the step S 1  and at the bottom of the step. In the middle progression at P2, due to deflection of the sprayed coating material, the coating builds-up more rapidly along the sidewall of the step S 1  (i.e., the coating builds in a non-planar manner). In the last progression at P3 on the right-hand side, the coating build-up along the sidewall of the step S 1  has caused bridging of the coating across the step in region Z. The faults, F, thus do not extend into the region Z. The bridging limits formation of the faults F and, in turn, limits the thickness, t1, of the coating that can be produced with faults F. For ceramic materials, such as stabilized zirconia, that are used for thermally insulating topcoats, this thickness is less than the depth of the step. 
     In  FIG. 10  the step, S 2 , has an undercut (i.e. undercut  269 ). In the first progression at P1 on the left-hand side, the coating, C, begins to build up on the surfaces around the step S 2  and at the bottom of the step S 2 . In the middle progression at P3, the undercut allows deflected coating material to spread laterally. Thus, the coating does not build-up along the sidewall (i.e., builds in a planar manner). In the last progression at P3 on the right-hand side, due to avoidance of build-up along the sidewall there is no bridging, and the faults, F, thus extend to the surface or very near surface of the coating. The elimination or reduction in the potential for bridging permits a greater thickness, t2, of the coating that can be produced with faults. For ceramic materials, such as stabilized zirconia, that are used for thermally insulating topcoats, this greater thickness may be equal to or greater than the depth of the step. 
     Similar to the undercut  269 , the undercut  369  eliminates or reduces the potential for bridging. However, rather than permitting the coating to deflect and spread laterally during spray deposition, the undercut  369  permits the coating to spread in the depth direction such that the coating does not build-up along the sidewall. 
     The 90° corner  277  may also facilitate formation of the faults  298 . For instance a highly rounded edge would provide less of a distinct change in depth at the step and thus contribute to bridging across the step. However, the 90° corner  277  provides a distinct change in depth and the step  267  and thus facilitates the formation of the thermally insulating topcoat  284  in a planar manner, which in turn facilitates formation of the faults  298  through the full thickness of the topcoat  284 . 
     The undercuts  269 / 369  may be configured in size to more effectively facilitate the elimination or reduction in the potential for coating build-up and bridging. For example, referring again to  FIG. 7A , the sidewall  275  defines a linear (radial) distance, LD, between the proximal surface  271  and the distal surface  273 . For example, the linear distance is along a direction that is perpendicular to the engine central longitudinal axis A. In one example, the undercut  269  comprises at least 10%, represented at linear height LD1-1, of the linear distance LD and may be up to 100% of the linear distance LD. In a further example, LD1-1 is 10% to 100% of the linear distance or 25% to 50% of the linear distance. The undercut  269  also defines a lateral undercut distance (perpendicular width to LD1-1), represented at LD1-2, which is at least 5% of the linear distance LD and may be up to 100% of the linear distance LD. In a further example, LD1-2 is 10% to 100% of the linear distance or 10% to 25% of the linear distance. In the context of the function of the undercut  269 , LD1-1 and LD1-2 thus represent a minimum relative size of the undercut  269  to more effectively allow spread of the coating material during coating deposition, relative to a given depth of the step  267 . For example, the sum of the percentages for LD1-1 and LD1-2 may be at least 20%. In a further example, the sum of the percentages for LD1-1 and LD1-2 is at least 20%, LD1-1 is at least 15% and LD1-2 is at least 5%. In further examples, the sum of the percentages for LD1-1 and LD1-2 is at least 75%, at least 100%, or at least 150%, which each provide more space for spread of the coating material during coating deposition. In one further example, the sum of the percentages for LD1-1 and LD1-2 is at least 150% and the angle of the undercut  269  is approximately 45°. 
     Similarly, the size of the undercut  369  ( FIG. 8A ) may comprise less than about 50%, represented at LD2-1, of the linear distance LD and a lateral undercut distance, represented at LD2-2, may comprise less than about 50% of a diametric distance DD of the of the step  267 . In a further example, LD2-1 is the smaller of, or is less than the smaller of, about 25% of the linear distance LD or about 25% of the diametric distance DD. In a further example, LD2-2 is about 25% of the diametric distance DD. In any of the above examples, the undercut  369  may also define a straight taper from a central region to the sidewall  275 . In an additional example, the sum of the percentages for LD2-1 and LD2-2 may be at least 20% % and less than approximately 100% with respect to the linear distance LD, the diametric distance DD, or combinations. 
       FIG. 11  illustrates an example method  401  of fabricating a gas turbine engine article, such as the seal members disclosed herein. In this example, the method  401  includes a forming process  403  and a deposition process  405 . The forming process step includes forming one or more steps with an undercut in a substrate, such as forming the step  267  and undercut  269  or  369  in substrate  265 . The deposition process  405  includes depositing a thermally insulating topcoat on the substrate, such as the thermally insulating topcoat  284 . As discussed herein, one or more faults form in the topcoat during deposition and extend from the step in the substrate. 
     The substrate and one or more steps with an undercut can be formed using one or more of several different processing techniques. For example, one cost effective processing technique includes forming the substrate and the one or more steps using additive manufacturing. Direct metal laser sintering and electron-beam melting are non-limiting examples of additive manufacturing techniques. In additive manufacturing a powdered material is fed to a machine, which may provide a vacuum, for example. The machine deposits multiple layers of the powdered material onto one another. At each iteration of layer deposition, the layer is selectively consolidated with reference to Computer-Aided Design data of the component being formed. Other layers or portions of layers corresponding to negative features, such as cavities or openings, are not joined and thus remain as a powdered material. The unjoined powder material may later be removed using blown air, for example. The additive manufacturing technique may be used to make the step  267  and undercut  269  or  369 . 
     Another processing technique includes forming an undercut using chemical milling. In this example, a substrate is provided that initially has a step without an undercut. A chemical, such as an acid etchant, is used to form the undercut. Other areas of the substrate may be masked off. Such a chemical milling technique may be used to make the undercut  269  or  369 . 
     Another processing technique includes forming a step and an undercut using laser ablation milling. In this example, a substrate is provided that initially has a step without an undercut. A high frequency pulsed laser beam, is used to form the undercut. Such a chemical milling technique may be used to make the undercut  269  or  369 . 
     Another processing technique includes forming a step and an undercut using mechanical milling. In this example, a substrate is provided that initially has no step. A tool, such as a drill bit or other cutting tool, is used to form the step and the undercut. The tool has a concave tip or other such configuration that forms the undercut. Such a mechanical milling technique may be used to make the step  267  and undercut  369 . 
     Any of the above processing techniques can additionally include formation of the 90° corner  277 . For example, the 90° corner  277  may be formed during formation of the step  267  in an additive manufacturing or milling technique. Additionally or alternatively, the 90° corner  277  may be formed by grinding down the surface of the substrate  265 . Thus, if the edge between the proximal surface  271  and the sidewall  275  is initially rounded, the rounded portion can be removed by grinding to produce the 90° corner  277 . 
     Although the different non-limiting embodiments are illustrated as having specific components, the embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
     It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed and illustrated in these exemplary embodiments, other arrangements could also benefit from the teachings of this disclosure. 
     The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would understand that certain modifications could come within the scope of this disclosure. For these reasons, the following claim should be studied to determine the true scope and content of this disclosure.