Turbo machine airfoil having a variable thickness thermal insulation coating

Abstract

To provide an airfoil for turbomachines such as stator vanes or rotor blades thereof.SOLUTION: An airfoil (100) includes a root (102) and a tip (104), which define a span (128) of the airfoil (100) therebetween. The airfoil (100) also includes a leading edge (124) and a trailing edge (126) downstream of the leading edge (124) along a flow direction. The leading edge (124) and the trailing edge (126) each extend across the span (128) of the airfoil (100) from the root (102) to the tip (104). The airfoil (100) further includes a pressure side surface (120) and a suction side surface (122). The airfoil (100) also includes a thermal barrier coating (108) on the pressure side surface (120) and the suction side surface (122). The thermal barrier coating (108) includes a base layer (110) and a top coat (112). A thickness of the base layer (110) varies across each of the pressure side surface (120) and the suction side surface (122) with a maximum thickness of the base layer (110) at the leading edge (124).SELECTED DRAWING: Figure 2

Description

Technical Field

[0001] The present disclosure generally relates to turbomachinery. More particularly, the present disclosure relates to airfoils for turbomachinery such as stator vanes or rotor blades thereof.

Background Art

[0002] A gas turbine engine generally includes a compressor section, a combustion section, a turbine section, and an exhaust section. The compressor section gradually increases the pressure of the working fluid entering the gas turbine engine and supplies this compressed working fluid to the combustion section. The compressed working fluid and fuel (e.g., natural gas) are mixed within the combustion section and combust within the combustion chamber to produce high-pressure and high-temperature combustion gases. The combustion gases flow from the combustion section into the turbine section, where they expand to generate work. For example, the expansion of the combustion gases in the turbine section can rotate a rotor shaft connected to a generator to generate electricity. The combustion gases are then discharged from the gas turbine via the exhaust section.

[0003] The turbine section generally includes a plurality of stator vanes and a corresponding plurality of rotor blades. Each stator vane and each rotor blade includes an airfoil positioned within the flow of combustion gases, and thus the airfoils are referred to as hot gas path components. The airfoils of each stator vane and each rotor blade typically extend radially outward from a platform such as an inner platform in the case of a stator vane. The airfoil of each stator vane extends to an outer platform at the radially outer end of the stator vane airfoil. The airfoil of each rotor blade extends to a tip at the radially outer end of the rotor blade airfoil. A particular rotor blade can include a tip shroud coupled to the radially outer end of the airfoil. Fillets can be provided at each transition between the airfoil and the platform and / or at the transition between the airfoil and the tip shroud.

[0004] The airfoil can extend from a leading edge to a trailing edge downstream of the leading edge and can define aerodynamic surfaces therebetween, such as a pressure side surface and a suction side surface. Since the airfoil is a hot gas path component, the surfaces of the airfoil, such as the aerodynamic surfaces, are typically treated to enhance their resistance to the high temperature environment of the hot gas path. One such surface treatment is a thermal barrier coating. In conventional airfoils, each layer of the thermal barrier coating is substantially uniform across the aerodynamic surface in both the spanwise direction (i.e., from root to tip) and the flow direction (i.e., from leading edge to trailing edge), and for example, has a constant uniform thickness. However, the conditions present at various locations around the airfoil can vary, and the properties of the layers within the thermal barrier can also change. For example, one layer may be more robust against physical impact, while another layer may provide better temperature resistance.

[0005] Therefore, an airfoil for a turbomachine having a thermal barrier coating that provides robust physical properties to selected regions or portions of the airfoil would be useful. SUMMARY OF THE INVENTION

[0006] Aspects and advantages of the systems according to the present disclosure are described in part in the following description, or will become apparent from the description, or can be learned through practice of the technology.

[0007] According to one embodiment, an airfoil for a turbomachine is provided. The airfoil includes a root and a tip spaced radially outward from the root. The span of the airfoil is defined between the root and the tip. The airfoil also includes a leading edge extending across the span of the airfoil from the root to the tip, and a trailing edge downstream of the leading edge along the flow direction. The trailing edge also extends across the span of the airfoil from the root to the tip. The airfoil further includes a pressure side extending between the root and the tip and between the leading edge and the trailing edge, and a suction side extending between the root and the tip and between the leading edge and the trailing edge. The suction side faces the pressure side. The airfoil also includes a thermal barrier coating on the pressure side and the suction side. The thermal barrier coating includes a base layer and a top coat. The thickness of the base layer varies across each of the pressure side and the suction side with a maximum thickness of the base layer at the leading edge.

[0008] According to another embodiment, a turbomachine is provided. The turbomachine includes a compressor, a combustor disposed downstream of the compressor, and a turbine disposed downstream of the combustor. The turbine includes rotor blades and stator vanes. At least one of the rotor blades and the stator vanes includes an airfoil. The airfoil includes a root and a tip spaced radially outward from the root. The span of the airfoil is defined between the root and the tip. The airfoil also includes a leading edge extending across the span of the airfoil from the root to the tip, and a trailing edge downstream of the leading edge along the flow direction. The trailing edge also extends across the span of the airfoil from the root to the tip. The airfoil further includes a pressure side extending between the root and the tip and between the leading edge and the trailing edge, and a suction side extending between the root and the tip and between the leading edge and the trailing edge. The suction side faces the pressure side. The airfoil also includes a thermal barrier coating on the pressure side and the suction side. The thermal barrier coating includes a base layer and a top coat. The thickness of the base layer varies across each of the pressure side and the suction side with a maximum thickness of the base layer at the leading edge.

[0009] These and other features, aspects, and advantages of the present system will be better understood with reference to the following description and the appended claims. The accompanying drawings, which are incorporated herein and form a part of this specification, illustrate embodiments of the technology and, together with the description, serve to explain the principles of the technology.

[0010] A complete and enabling disclosure of the present system, including the best mode of making and using the system and method for those of ordinary skill in the art, is set forth in this specification with reference to the accompanying drawings.

Brief Description of the Drawings

[0011]

Figure 1

Figure 2

Figure 3

Figure 4

Figure 5

Figure 6

Figure 7

Figure 8

Figure 9

Best Mode for Carrying Out the Invention

[0012] The repeated use of reference numerals in this specification and the drawings is intended to represent the same or similar features or elements of the present technology.

[0013] Here, embodiments of the present system will be referred to in detail, one or more examples of which are shown in the accompanying drawings. Each example is provided for the purpose of explaining the present technology and is not intended to limit the present technology. In fact, it will be apparent to those skilled in the art that various modifications and changes are possible in the present technology without departing from the scope or spirit of the claimed technology. For example, features illustrated or described as part of one embodiment can be used in another embodiment to yield further embodiments. Accordingly, the present disclosure is intended to embrace such modifications and changes that fall within the scope of the appended claims and their equivalents.

[0014] The detailed description uses numerical and alphabetical reference signs to refer to the features of the drawings. Similar or like reference signs in the drawings and the description are used to refer to similar or like parts of the present technology. As used herein, the terms "first," "second," and "third" can be used interchangeably to distinguish one component from another and are not intended to indicate the position or importance of individual components.

[0015] As used herein, the terms "upstream" (or "forward") and "downstream" (or "rearward") refer to the relative directions with respect to the flow of fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction in which the fluid flows. The term "radially" refers to the relative direction substantially perpendicular to the axial centerline of a particular component, the term "axially" refers to the relative direction substantially parallel and / or coaxial with the axial centerline of a particular component, and the term "circumferentially" refers to the relative direction extending around the axial centerline of a particular component.

[0016] Approximating terms such as "generally" or "about" include values within a range of plus or minus 10 percent of the recited value. When used in the context of an angle or direction, such terms include a range of plus or minus 10 degrees of the recited angle or direction. For example, "generally perpendicular" includes directions within 10 degrees of perpendicular in either direction, e.g., clockwise or counterclockwise.

[0017] Industrial or land-based gas turbines are shown and described herein, but the systems shown and described herein are not limited to land-based and / or industrial gas turbines unless specifically recited in the claims. For example, the technology described herein can be used in any type of turbomachine including, but not limited to, steam turbines, aircraft gas turbines, or marine gas turbines.

[0018] Referring now to the drawings, like numerals throughout the figures indicate like elements, and FIG. 1 shows a schematic view of one embodiment of a turbomachine, which in the illustrated embodiment is a gas turbine 10. It should be understood that the gas turbine 10 of the present disclosure need not be a gas turbine engine and may be any suitable turbomachine such as a steam turbine engine or other suitable engine.

[0019] As shown, the gas turbine 10 generally includes an inlet section 12, a compressor section 14 disposed downstream of the inlet section 12, a plurality of combustors (not shown) within a combustor section 16 disposed downstream of the compressor section 14, a turbine section 18 disposed downstream of the combustor section 16, and an exhaust section 20 disposed downstream of the turbine section 18. Additionally, the gas turbine 10 can include one or more shafts 22 coupled between the compressor section 14 and the turbine section 18.

[0020] The compressor section 14 can generally include a plurality of rotor disks 24 (one of which is shown) and a plurality of rotor blades 26 that extend radially outward from each rotor disk 24 and are connected to each rotor disk 24. Next, each rotor disk 24 may be coupled to or form a part of a shaft 22 that extends through the compressor section 14.

[0021] The turbine section 18 can generally include a plurality of rotor disks 28 (one of which is shown) and a plurality of rotor blades 30 that extend radially outward from each rotor disk 28 and are connected to each rotor disk 28. Next, each rotor disk 28 may be coupled to or form a part of a shaft 22 that extends through the turbine section 18. The turbine section 18 further includes an outer casing 31 that circumferentially surrounds a part of the shaft 22 and the rotor blades 30, thereby at least partially defining a hot gas path 32 through the turbine section 18. The turbine section can also include a plurality of stator vanes 29 attached to the casing 31 within the hot gas path 32.

[0022] During operation, a working fluid such as air flows into the compressor section 14 through the inlet section 12, where the air is gradually compressed, thereby providing pressurized air to the combustors of the combustor section 16. The pressurized air is mixed with fuel and burned in each combustor to generate combustion gases 34. The combustion gases 34 flow from the combustor section 16 into the turbine section 18 through the hot gas path 32, where energy (kinetic energy and / or thermal energy) is transferred from the combustion gases 34 to the rotor blades 30, causing the shaft 22 to rotate. The mechanical rotational energy can then be used to power the compressor section 14 and / or generate electricity. The combustion gases 34 discharged from the turbine section 18 can then be exhausted from the gas turbine 10 via the exhaust section 20.

[0023] FIG. 2 shows a perspective view of an exemplary airfoil 100 that can be incorporated into the rotor blades 30 and / or stator vanes 29 of the turbine section 18 of the gas turbine 10. As shown in FIG. 2, the airfoil 100 can extend radially outward from a root 102 to a tip 104. The airfoil 100 includes a pressure side surface 120 and an opposing suction side surface 122 (FIG. 3). The pressure side surface 120 and the suction side surface 122 are joined or interconnected together at a leading edge 124 of the airfoil 100 that is oriented with respect to the flow of combustion gas 34 (FIG. 1). The pressure side surface 120 and the suction side surface 122 are also joined or interconnected together at a trailing edge 126 of the airfoil 100 that is spaced downstream from the leading edge 124. The pressure side surface 120 is generally concave, and the suction side surface 122 is generally convex.

[0024] Referring particularly to FIG. 2, the airfoil 100 defines a span 128 that extends from the root 102 to the tip 104. In particular, the root 102 is positioned at 0 percent (0%) of the span 128, and the tip 104 is positioned at 100 percent (100%) of the span 128.

[0025] FIGS. 3-9 show various cross-sectional views of the exemplary airfoil 100. It should be noted that each of the cross-sectional views of FIGS. 3-9 is a constant-span cross-section. For example, FIG. 3 may be taken at approximately 50 percent (50%) of the span 128, and the entire cross-section through the airfoil 100 shown in FIG. 3 is at the same position along the span 128, e.g., at approximately 50 percent (50%) of the span 128.

[0026] For example, as seen in FIG. 3, the airfoil 100 defines a camber line 118. More specifically, the camber line 118 extends from the leading edge 124 to the trailing edge 126. The camber line 118 is also positioned between and equidistant from the pressure side surface 120 and the suction side surface 122. Also, as generally seen in FIG. 3, a thermal barrier coating 108 can be provided on the outermost surface of the airfoil 100. For example, as shown in FIG. 3, the thermal barrier coating 108 may be provided on each of the pressure side surface 120 and the suction side surface 122.

[0027] Figure 4 shows an enlarged view of a part of the airfoil 100 including the leading edge portion of the airfoil 100, for example, the leading edge 124 and the portion of the airfoil 100 adjacent thereto. Figure 5 shows an enlarged view of a part of the airfoil 100 in the central forward portion of the airfoil 100, for example, behind the leading edge 124 and in front of the midpoint of the airfoil 100, the midpoint being defined along the direction of the flow of combustion gas. Figure 6 shows an enlarged view of a part of the airfoil 100 in the central rearward portion of the airfoil 100, for example, behind the midpoint of the airfoil 100 and in front of the trailing edge 126. Figure 7 shows an enlarged view of a part of the airfoil 100 including the trailing edge portion of the airfoil 100, for example, the trailing edge 126 and the portion of the airfoil 100 adjacent thereto, according to one or more embodiments. Figures 8 and 9 show further views of the trailing edge portion of the airfoil according to one or more additional embodiments.

[0028] As seen in FIGS. 3-9, the base or core material of the airfoil 100 can include a substrate 106 to which a thermal barrier coating 108 is applied. For example, the substrate 106 can be a metallic material such as an alloy containing iron and nickel or cobalt, such as a high-temperature steel, a superalloy, and / or other suitable metal alloys, or can include them. The thermal barrier coating 108 can include, for example, a ceramic material. The thermal barrier coating 108 is generally formed on the outer surface 107 of the substrate 106. In various embodiments, the airfoil 100 can include the thermal barrier coating 108 directly on the outer surface 107, or can include a bond coat 114 formed directly on the outer surface 107 of the substrate 106. In embodiments where the bond coat 114 is provided, the thermal barrier coating 108 can be formed directly on the bond coat 114.

[0029] As seen in FIGS. 4-9, the thermal barrier coating 108 can include multiple layers. For example, the thermal barrier coating 108 can include a base layer 110 closer to the substrate 106 and a top coat 112 outside and / or on the base layer 110.

[0030] The thickness of the base layer 110 may vary across each of the positive-pressure side surface 120 and the negative-pressure side surface 122. For example, the thickness of the base layer 110 may decrease across each of the positive-pressure side surface 120 and the negative-pressure side surface 122. In at least some embodiments, the thickness of the base layer 110 may decrease from a maximum at the leading edge 124 to a minimum at the trailing edge 126. Providing the maximum thickness of the base layer 110 at and around the leading edge 124 can advantageously provide improved physical resistance to the leading edge portion of the airfoil 100, which can be advantageous when the leading edge portion of the airfoil 100 is more likely to receive physical impact and / or erosion than the downstream portion of the airfoil 100.

[0031] Thus, in various embodiments, the thickness of the base layer 110 at the leading edge 124 may be greater than the thickness of the top coat 112 at the leading edge 124. In some embodiments, the thickness of the base layer 110 may continuously taper from the leading edge 124 to the trailing edge 126 across each of the positive-pressure side surface 120 and the negative-pressure side surface 122. In some embodiments, the thickness of the base layer 110 at the trailing edge 126 may be less than the thickness of the top coat 112 at the trailing edge 126.

[0032] In some embodiments, the thermal barrier coating 108 may be composed only of the base layer 110 at the leading edge 124. For example, the ratio of the base layer 110 to the top coat 112 may be 100:0 at the leading edge portion, and the top coat 112 may not be present at the leading edge 124. In alternative embodiments, the thermal barrier coating 108 may mainly include the base layer 110 at and around the leading edge, such as a ratio of the base layer 110 to the top coat 112 of about 90:10, for example about 80:20.

[0033] In some embodiments, for example, as shown in FIG. 7, the thermal barrier coating 108 may not be provided at the trailing edge 126. For example, in such embodiments, both the base layer 110 and the top coat 112 may taper to zero thickness at or near the trailing edge 126, as shown in FIG. 7.

[0034] In other embodiments, the thermal barrier coating 108 may be continuous across the trailing edge 126. For example, in some embodiments, the proportion of the layer of the thermal barrier coating 108 at the trailing edge 126 may be the same as or similar to that at the leading edge 124 (e.g., as shown in FIG. 4). In such embodiments, for example, as shown in FIG. 8, the top coat 112 may taper in thickness as it approaches or at the trailing edge 126, such that the ratio of the base layer 110 to the top coat 112 may be 100:0 at the trailing edge portion and the top coat 112 may not be present at the trailing edge 126. Additionally, similar to the leading edge 124 as described above, the thermal barrier coating 108 may mainly include the base layer 110 at and around the trailing edge 126, such as a ratio of the base layer 110 to the top coat 112 of about 90:10, for example about 80:20.

[0035] In some embodiments, for example, as shown in FIG. 9, both layers of the thermal barrier coating 108, such as the base layer 110 and the top coat 112, can wrap around the trailing edge 126 with a substantially constant thickness around the trailing edge 126.

[0036] Referring back to FIG. 2, in some embodiments, the leading edge portion and / or region or part of the airfoil 100 where the thickness of the base layer 110 is maximum can include the high-impact zone 130 on the airfoil 100. It should be noted that the high-impact zone 130 refers to a relatively high probability of physical impact and / or erosion within the depicted region 130 relative to the rest of the airfoil 100, particularly the rest of its aerodynamic surfaces 120 and 122. Additionally, it should be understood that the high-impact zone 130 is substantially symmetric with respect to the leading edge 124 and extends along the length of the airfoil 100 at substantially the same distance on both the positive pressure side 120 (shown in FIG. 2) and the negative pressure side 122.

[0037] As described above, the length of the airfoil portion 100 is defined along the flow direction. In at least some embodiments, the leading edge portion may have the same extent as the high impact zone 130. For example, the maximum thickness of the base layer 110 may be provided at the leading edge 124 and throughout the high impact zone 130. As moving from the leading edge 124 to the trailing edge 126, the thickness of the base layer 110 may be maximum in the region 130, may decrease in the first intermediate zone 132, and may decrease again in the second intermediate zone 134. As a result, the thickness of the base layer 110 may be smaller in the second intermediate zone 134 than in the first intermediate zone 132. Also, the thickness of the base layer 110 can decrease again in the downstream or rear zone 136. In some embodiments, the thickness of the base layer 110 may be minimum within the region of the trailing edge 126 and / or the rear zone shown by 136 in FIG. 2.

[0038] In some embodiments, the minimum thickness of the base layer 110 may occupy 70 percent (70%) or less, for example about 60 percent (60%) or less, for example about 50 percent (50%) or less, for example about 40 percent (40%) or less, or for example about 30 percent (30%) or less of the total thermal insulation coating 108.

[0039] Referring further to FIG. 2, in some embodiments, the thickness of the base layer 110 may also vary across the span 128 of the airfoil 100. For example, the ratio of the base layer 110 to the top coat 112 within the thermal barrier coating 108 may be a base ratio over the entire length of the airfoil 100 at the root 102, e.g., the minimum thickness of the base layer 110, and the thickness of the base layer 110 may increase by about 10 percent (10%) to about 20 percent (20%) of the span 128. That is, the thermal barrier coating 108 may transition from the minimum thickness of the base layer 110 in the rear zone 136 to one of the intermediate zones 134 or 132 within a zone extending from about 5 percent (5%) or more of the span 128, e.g., about 10 percent (10%) of the span 128, e.g., about 20 percent (20%) of the span 128, e.g., about 30 percent (30%) of the span 128, and / or from about 5 percent (5%) to about 30 percent (30%) of the span 128. In various embodiments, there may be an increase in the second span direction in the thickness of the base layer 110, e.g., a transition from the second intermediate zone 134 to the first intermediate zone 132, which may occur within a zone extending from about 10 percent (10%) or more of the span 128, e.g., about 20 percent (20%) of the span 128, e.g., about 30 percent (30%) of the span 128, e.g., about 40 percent (40%) of the span 128, and / or from about 10 percent (10%) to about 40 percent (40%) of the span 128.

[0040] In some embodiments, the thickness of the base layer 110 at the leading edge 124 may vary across the span 128 of the airfoil 100 and may be constant across the span 128 at other portions of the airfoil 100, such as the trailing edge 126. In such embodiments, the maximum thickness of the base layer 110 may be defined at the leading edge 124, particularly at approximately the mid-span point (e.g., about 50 percent of the span) on the leading edge 124. For example, the maximum thickness of the base layer 110 may be provided outwardly from about 40 percent of the span 128, e.g., starting from 40 percent of the span and continuing to 100 percent of the span, outwardly from, e.g., about 50 percent of the span, or, e.g., outwardly from about 60 percent of the span.

[0041] Furthermore, it should be understood that additional embodiments can include more or fewer transition portions in one or both of the spanwise and flow directions, e.g., along the length of the airfoil 100. For example, some embodiments may include only one intermediate zone or three, four, or more intermediate zones. One or more transitions in the thickness of the base layer 110 generally taper gradually, and thus the lines indicating the various zones in FIG. 2 are for illustrative purposes only and should not be understood as sharp boundaries between regions of various thicknesses of the base layer 110. Additionally, some embodiments may include a continuous change in the thickness of the base layer 110 and thus there is a substantially infinite or constant transition in the thickness of the base layer 110. For example, when the thermal barrier coating 108 is continuous around the airfoil 100, as shown, e.g., in FIG. 3, the various thicknesses of the base layer 110 can thus form a substantially infinite loop around the periphery of the airfoil 100.

[0042] In at least some embodiments, the total thickness of the thermal barrier coating 108 may be the same over a majority of the airfoil 100, such as at least about 75 percent (75%) of the length of the airfoil 100 from the leading edge 124, for example, if the length of the airfoil 100 is defined along the flow direction. For example, the overall thickness of the thermal barrier coating 108 may be at least about 85 percent (85%) of the length of the airfoil 100, such as at least about 90 percent (90%) of the length of the airfoil 100, such as at least about 95 percent (95%) of the length of the airfoil 100, such as about 98 percent (98%) or more of the length of the airfoil 100. In some embodiments, the total thickness of the thermal barrier coating 108 may be constant over the entire length of the airfoil 100 from the leading edge 124 to the trailing edge 126. Providing a uniform and constant overall thickness to the thermal barrier coating 108 over all or a majority of the length of the airfoil 100 can advantageously improve the balance of the airfoil 100, such as the mass distribution, and also enable improved resistance to physical shock and / or erosion at the leading edge 124 and / or leading edge portion having the relatively thick base layer 110, as described above.

[0043] This specification uses examples to disclose the invention, including the best mode, and to enable any person skilled in the art to practice the invention, including making and using any device or system and performing any incorporated method. The patentable scope of the invention 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 include 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.

Explanation of Reference Numerals

[0044] 10 Gas turbine 12 Inlet section 14 Compressor section 16 Combustor section 18 Turbine section 20 Exhaust section 22 Shaft 24 Rotor disk 26 Rotor blade 28 Rotor disk 29 Stator vane 30 Rotor blade 31 Outer casing 32 High-temperature gas path 34 Combustion gas 100 Airfoil 102 Root 104 Tip 106 Substrate 107 Outer surface 108 Thermal barrier coating 110 Base layer 112 Top coat 114 Bond coat 118 Camber line 120 Positive pressure side 122 Negative pressure side 124 Leading edge 126 Trailing edge 128 Span 130 High-impact zone, region 132 First intermediate zone 134 Second intermediate zone 136 Rear zone

Claims

**Claim 1** An airfoil (100) for a turbomachine, the airfoil (100) having a root (102), a tip (104) spaced radially outward from the root (102), the root (102) and the tip (104) defining a span (128) of the airfoil (100) therebetween, a leading edge (124) extending across the span (128) of the airfoil (100) from the root (102) to the tip (104), a trailing edge (126) downstream of the leading edge (124) along the flow direction, the trailing edge (126) extending across the span (128) of the airfoil (100) from the root (102) to the tip (104), a pressure side surface (120) extending between the root (102) and the tip (104) and between the leading edge (124) and the trailing edge (126), a suction side surface (122) extending between the root (102) and the tip (104) and between the leading edge (124) and the trailing edge (126), the suction side surface (122) being opposite to the pressure side surface (120), and a thermal barrier coating (108) on the pressure side surface (120) and the suction side surface (122), the thermal barrier coating (108) comprising a base layer (110) and a top coat (112), the thickness of the base layer (110) decreasing in a constant transition from the leading edge (124) to the trailing edge (126) on each of the pressure side surface (120) and the suction side surface (122) to be maximum at the leading edge (124), and the thickness of the top coat (112) increasing from the leading edge (124) towards the trailing edge (126) on each of the pressure side surface (120) and the suction side surface (122). An airfoil (100). **Claim 2** The airfoil (100) according to claim 1, wherein the thickness of the base layer (110) at the leading edge (124) is greater than the thickness of the top coat (112) at the leading edge (124). **Claim 3** The airfoil (100) according to claim 1, wherein the thickness of the base layer (110) proximate to the trailing edge (126) is less than the thickness of the top coat (112) proximate to the trailing edge (126). **Claim 4** The airfoil (100) according to claim 1, wherein the total thickness of the heat shield coating (108) is constant along the direction from the leading edge (124) to the trailing edge (126) over most of the airfoil (100) on each of the positive pressure side surface (120) and the negative pressure side surface (122).

5. The airfoil (100) according to claim 1, wherein the thickness of the base layer (110) varies across the span (128) of the airfoil (100).

6. The airfoil (100) according to claim 1, further comprising a substrate (106), wherein the heat shield coating (108) is formed on the outer surface (107) of the substrate (106).

7. The airfoil (100) according to claim 6, further comprising a bond coat (114) formed directly on the outer surface (107) of the substrate (106), wherein the heat shield coating (108) is formed directly on the bond coat (114).

8. The airfoil (100) according to claim 6, wherein the substrate (106) includes a metallic material and the heat shield coating (108) includes a ceramic material.

9. A turbomachine, wherein the turbomachine includes a compressor, a combustor disposed downstream of the compressor, and a turbine disposed downstream of the combustor and the turbine includes rotor blades (30) and stator vanes (29), and at least one of the rotor blades (30) and the stator vanes (29) includes the airfoil (100) according to any one of claims 1 to 8.

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