Turbine blade outermost polishing layer using gradient ceramic

Abstract

The present invention provides a turbine blade comprising an airfoil body having the outermost radial surface relative to the turbine rotor, a bond coat on the outermost radial surface, and an outermost polished layer on the bond coat. [Solution] The blade comprises an airfoil body having the radially outermost surface relative to the turbine rotor, a bond coat on the radially outermost surface, and an outermost polished layer on the bond coat. The outermost polished layer may include a gradient oxide ceramic layer or a gradient metal ceramic layer. In either case, vacuum heat treatment of the outermost polished layer hardens the layer, increasing the friction ratio, improving the clearance between blade tips, resulting in a smaller tip gap and less leakage compared to conventional systems. The heat-treated polished layer is also resistant to oxidation at high temperatures exceeding 1090°C (approximately 2000°F), such as those caused by the increasing use of high-temperature fuels such as hydrogen. Vacuum heat treatment also results in improved thermal cycling tests.

Description

Technical Field

[0001] The present disclosure generally relates to coatings. More specifically, the present disclosure relates to an outermost polishing layer of a turbine blade using ceramics.

Background Art

[0002] In a gas turbine, air is pressurized by a compressor and used to burn fuel in a combustor to generate a flow of high-temperature combustion gas, and such gas flows downstream through one or more turbines from which energy can be extracted. According to such a turbine, generally, a row of turbine blades arranged at intervals in the circumferential direction extends radially outward from a supporting rotor disk. Each blade typically includes a dovetail that enables assembly and disassembly of the blade in a corresponding dovetail slot in the rotor disk, and an airfoil body that extends radially outward from the dovetail.

[0003] The airfoil body has a generally concave positive-pressure sidewall and a generally convex negative-pressure sidewall that extend axially between a corresponding leading edge and trailing edge and radially between a root and a tip. In a turbine engine, a shroud is a ring of material that surrounds the rotating blades. The shroud may be stationary and formed of a ceramic matrix composite (CMC) protected by an environmental barrier coating (EBC) to avoid oxidation and erosion in the presence of a high-temperature gas flow. Alternatively, the shroud may include a metallic component protected by a thermal barrier coating (TBC) to avoid oxidation and erosion in the presence of a high-temperature air flow.

[0004] The performance and efficiency of a turbine can be improved by reducing the space between the radial outermost surface of the rotating turbine blades, i.e., the tip and the stationary shroud, thereby restricting the flow of gas over or around the blade tip, and potentially bypassing the blade. For example, blades can be configured so that their tips fit close to the shroud during engine operation. Therefore, generating and maintaining a small tip clearance is particularly desirable for efficiency purposes. During engine operation, the blade tips may rub against the shroud, thereby increasing the gap, resulting in loss of efficiency, and potentially damaging or destroying the blade set. [Overview of the project]

[0005] To reduce efficiency losses, a polished layer can be provided on the tip to help form a tight contact tolerance with the shroud. In one example, the polished layer is formed as a continuous but rough ceramic layer, which is typically very hard. Some current techniques use polished layers containing cubic zirconia (cZ) and hafnia (HfO2). Other current techniques use a polished layer containing cubic boron nitride (cBN) in a soft nickel boron silicon (NiBSi) matrix laser-clad on the radial outermost surface of the blade. Hard ceramic materials provide a cutter blade effect from the radial outermost surface of the airfoil blade to the shroud. The use of the above ceramic materials has several drawbacks. For example, the ceramic forming process, such as for cBN and NiBSi matrices, has variables that affect quality and are difficult to control. Another drawback of the listed ceramic materials, particularly cBN materials, is that they tend to oxidize at temperatures above the high temperatures currently in which turbine engines operate, e.g., 1090°C (approximately 2000°F), due to the increasing use of higher-temperature fuels such as hydrogen. Oxidation reduces the hardness of the ceramic and its ability to cut into the shroud and form a seal. Oxidation can also occur during manufacturing, and the rub-in process must be performed quickly during commissioning to avoid damage to the polished layer.

[0006] All aspects, examples, and features described below can be combined in any technically possible way.

[0007] One aspect of the present disclosure is a turbine blade comprising an airfoil body having a radially outermost surface with respect to a turbine rotor, a bond coat on the radially outermost surface, and an outermost polished layer on the bond coat, the outermost polished layer comprising a first layer on the bond coat comprising 100 wt% of a first ceramic and 0 wt% of a second ceramic different from the first ceramic, the first ceramic comprising a first layer consisting of 8 wt% of yttria-stabilized zirconia (8YSZ), and 65-85 wt% of the first ceramic and 15-35 wt The invention provides a turbine blade comprising an outermost polished layer that includes a gradient oxide ceramic layer comprising: a second layer on a first layer containing a second ceramic by weight; a third layer on the second layer containing 40-60 wt% of the first ceramic and 40-60 wt% of the second ceramic; a fourth layer on the third layer containing 5-25 wt% of the first ceramic and 75-95 wt% of the second ceramic; and a fifth layer on the fourth layer containing 0 wt% of the first ceramic and 100 wt% of the second ceramic.

[0008] Another aspect of the present disclosure includes any of the aforementioned aspects, wherein the bond coat comprises at least one of cobalt-nickel-chromium-aluminum-yttrium (CoNiCrAlY), nickel-chromium-aluminum-yttrium (NiCrAlY), nickel-cobalt-chromium-aluminum-yttrium (NiCoCrAlY), and cobalt-chromium-aluminum-yttrium (CoCrAlY).

[0009] Another aspect of the present disclosure includes any of the preceding aspects, wherein the second ceramic comprises one of pure alumina (Al2O3) or zirconia-reinforced alumina (ZTA).

[0010] Another aspect of the present disclosure includes any of the preceding aspects, wherein the bond coat has a thickness of 85 to 125 micrometers (μm), each of the first to fourth layers has a thickness of 55 to 75 μm, and the fifth layer has a thickness of 100 to 125 μm.

[0011] Another aspect of the present disclosure includes any of the preceding aspects, wherein the outermost polished layer is formed using thermal plasma spraying and post-formation vacuum heat treatment, which includes exposing the gradient oxide ceramic layer to a temperature above 1000°C for a period of at least 1 hour, and then cooling the gradient oxide ceramic layer.

[0012] Another aspect of the present disclosure includes a turbine blade comprising: an airfoil body having a radially outermost surface with respect to a turbine rotor; a bond coat on the radially outermost surface; and an outermost polished layer on the bond coat, the outermost polished layer comprising a gradient metal-ceramic layer which includes a first layer on the bond coat made of a first ceramic made of 8 wt% yttria-stabilized zirconia (8YSZ) having a porosity of 5 to 25%; a second layer on the first layer comprising 65 to 85 wt% metal and 15 to 35 wt% second ceramic; a third layer on the second layer comprising 40 to 60 wt% metal and 40 to 60 wt% second ceramic; a fourth layer on the third layer comprising 5 to 25 wt% metal and 75 to 95 wt% second ceramic; and a fifth layer on the fourth layer comprising 0 wt% metal and 100 wt% second ceramic.

[0013] Another aspect of the present disclosure includes any of the aforementioned aspects, wherein the second ceramic, unlike the first ceramic, includes either pure alumina (Al2O3) or zirconia-reinforced alumina (ZTA).

[0014] Another aspect of the present disclosure includes any of the aforementioned aspects, wherein the second ceramic includes 8YSZ without porosity.

[0015] Another aspect of the present disclosure includes any of the aforementioned aspects, wherein the bond coat comprises M-chromium-aluminum-yttrium (MCrAlY), and the metal comprises at least one of MCrAlY, nickel-boron-silicon (NiBSi) alloy, nickel-chromium-iron-molybdenum (NiCrFeMo) alloy, and cobalt-chromium-molybdenum (CoCrMo).

[0016] Another aspect of the present disclosure includes any of the preceding aspects, wherein the bond coat has a thickness of 85 to 125 micrometers (μm), each of the first to fourth layers has a thickness of 55 to 75 μm, and the fifth layer has a thickness of 100 to 125 μm.

[0017] Another aspect of the present disclosure includes any of the preceding aspects, wherein the outermost polished layer is formed using thermal plasma spraying and post-formation vacuum heat treatment, which includes exposing the gradient metal ceramic layer to a temperature above 1000°C for a period of at least 1 hour, and then cooling the gradient metal oxide layer.

[0018] Another aspect of the present disclosure includes a method for forming an outermost polished layer for the radial outermost surface of a turbine blade, comprising: forming a bond coat on the radial outermost surface of the turbine blade; forming one of a gradient oxide ceramic layer and a gradient metal ceramic layer on the bond coat on the radial outermost surface of the turbine blade; and heat-treating one of the gradient oxide ceramic layer and the gradient metal ceramic layer on the radial outermost surface of the turbine blade.

[0019] Another aspect of the present disclosure includes any of the preceding aspects, wherein the vacuum heat treatment includes exposing the outermost polished layer to a temperature above 1000°C for a period of at least 1 hour, and then cooling the outermost polished layer.

[0020] Another aspect of the present disclosure includes any of the preceding aspects, wherein the vacuum heat treatment includes raising the temperature of the outermost polished layer to a temperature above 1200°C for a period of at least 1.5 hours, and then cooling the outermost polished layer.

[0021] Another aspect of the present disclosure includes any of the aforementioned aspects, wherein the heat treatment includes exposing the outermost polished layer to a temperature of 750°C to 1000°C in air for a period of 30 minutes to 4 hours.

[0022] Another aspect of the present disclosure includes any of the aforementioned aspects, wherein the bond coat comprises at least one of cobalt-nickel-chromium-aluminum-yttrium (CoNiCrAlY), nickel-chromium-aluminum-yttrium (NiCrAlY), nickel-cobalt-chromium-aluminum-yttrium (NiCoCrAlY), and cobalt-chromium-aluminum-yttrium (CoCrAlY).

[0023] Another aspect of the present disclosure includes any of the preceding aspects, wherein the outermost polished layer is a first layer on a bond coat comprising 100 wt% of a first ceramic and 0 wt% of a second ceramic different from the first ceramic, the first ceramic comprising a gradient oxide ceramic layer comprising a first layer consisting of 8 wt% yttria-stabilized zirconia (8YSZ), a second layer on the first layer comprising 65-85 wt% of the first ceramic and 15-35 wt% of the second ceramic, a third layer on the second layer comprising 40-60 wt% of the first ceramic and 40-60 wt% of the second ceramic, a fourth layer on the third layer comprising 5-25 wt% of the first ceramic and 75-95 wt% of the second ceramic, and a fifth layer on the fourth layer comprising 0 wt% of the first ceramic and 100 wt% of the second ceramic.

[0024] Another aspect of the present disclosure includes any of the preceding aspects, wherein the second ceramic comprises one of pure alumina (Al2O3) or zirconia-reinforced alumina (ZTA).

[0025] Another aspect of the present disclosure includes any of the preceding aspects, wherein the bond coat has a thickness of 85 to 125 micrometers (μm), each of the first to fourth layers has a thickness of 55 to 75 μm, and the fifth layer has a thickness of 100 to 125 μm.

[0026] Another aspect of the present disclosure includes any of the foregoing aspects, and the outermost polishing layer comprises a first layer on the bond coat, which consists of a first ceramic made of 8 wt% yttria stabilized zirconia (8YSZ) having a porosity of 5 - 25%, a second layer on the first layer, which comprises 65 - 85 wt% metal and 15 - 35 wt% second ceramic, a third layer on the second layer, which comprises 40 - 60 wt% metal and 40 - 60 wt% second ceramic, a fourth layer on the third layer, which comprises 5 - 25 wt% metal and 75 - 95 wt% second ceramic, and a fifth layer on the fourth layer, which comprises 0 wt% metal and 100 wt% second ceramic, the inclined metal ceramic layer including the above layers.

[0027] Another aspect of the present disclosure includes any of the foregoing aspects, and the second ceramic is different from the first ceramic and includes one of pure alumina (Al2O3) or zirconia toughened alumina (ZTA).

[0028] Another aspect of the present disclosure includes any of the foregoing aspects, and the second ceramic includes non-porous 8YSZ.

[0029] Another aspect of the present disclosure includes any of the foregoing aspects, the bond coat includes M-chromium-aluminum-yttrium (MCrAlY), and the metal includes at least one of MCrAlY, nickel boron silicon (NiBSi) alloy, nickel-chromium-iron-molybdenum (NiCrFeMo) alloy, and cobalt-chromium-molybdenum (CoCrMo).

[0030] Another aspect of the present disclosure includes any of the foregoing aspects, the bond coat has a thickness of 85 - 125 micrometers (μm), each of the first to fourth layers has a thickness of 55 - 75 μm, and the fifth layer has a thickness of 100 - 125 μm.

[0031] Another aspect of the present disclosure includes any of the preceding aspects, wherein bond coat formation includes forming a first bond coat layer by one of a high-speed oxygen-fuel (HVOF) thermal spraying process and a low-temperature thermal spraying process, and forming a second bond coat layer by one of an HVOF thermal spraying process, a flame spraying process, and an air-plasma spraying process.

[0032] Two or more embodiments described in this disclosure, including those described in this summary section, can be combined to form embodiments not specifically described herein. That is, all embodiments described herein can be combined with one another.

[0033] Details of one or more embodiments are described in the accompanying drawings and the following description. Other features, purposes, and advantages will become apparent from the description and drawings, as well as the claims.

[0034] These and other features of the Disclosure will be more readily apparent from the following detailed description of various embodiments of the Disclosure, in conjunction with the accompanying drawings illustrating various embodiments of the Disclosure. [Brief explanation of the drawing]

[0035] [Figure 1] This is a schematic diagram of an example of a turbomachinery system. [Figure 2] This is a perspective view of an exemplary turbine component in the form of a turbine blade assembly, including a rotor disk, turbine blades, and a stationary shroud. [Figure 3] This is a perspective view of the radially outermost surface or tip of a turbine blade according to an embodiment of the present disclosure. [Figure 4A] This is a cross-sectional view of the outermost polished layer for a turbine blade according to an embodiment of the present disclosure. [Figure 4B] This is a cross-sectional view of the outermost polished layer for a turbine blade according to an embodiment of the present disclosure. [Figure 5] This is a cross-sectional view of a method for forming the outermost polished layer according to an embodiment of the present disclosure. [Figure 6] This is a cross-sectional view of a method for forming the outermost polished layer according to an embodiment of the present disclosure. [Figure 7] This is a cross-sectional view of a method for forming the outermost polished layer according to an embodiment of the present disclosure. [Figure 8] This is a cross-sectional view of a method for forming the outermost polished layer according to an embodiment of the present disclosure. [Figure 9] This figure shows a graph of vacuum heat treatment according to an embodiment of the present disclosure. [Figure 10] This figure shows a graph representation of an exemplary thermal cycling test according to an embodiment of the present disclosure.

[0036] Please note that the drawings in this disclosure are not necessarily to scale. The drawings are intended to illustrate only typical embodiments of this disclosure and should not be considered to limit the scope of this disclosure. In the drawings, similar reference numerals represent similar elements across drawings. [Modes for carrying out the invention]

[0037] As a first issue, in order to clearly describe this disclosure, it is necessary to select specific technical terms when referring to and describing relevant mechanical components within turbomachinery and related to turbine blades. Where this is done, common industrial terminology will be used and adopted in a manner that is consistent with its accepted meaning, whenever possible. Unless otherwise stated, such terminology should be given a broad interpretation consistent with the context of this application and the appended claims. Those skilled in the art will understand that, in many cases, certain components may be referred to using several different or overlapping terms. What may be described herein as a single part may comprise multiple components and be referred to as consisting of multiple components in another context. Or, what may be described herein as comprising multiple components may be referred to elsewhere as a single part.

[0038] In addition, several descriptive terms may be used regularly in this specification, and it will be useful to define these terms at the beginning of this section. These terms and their definitions are as follows, unless otherwise noted: As used herein, “downstream” and “upstream” are terms indicating direction relative to the flow of combustion gases through a turbine engine, or, for example, the flow of air through a combustor, or the flow of working fluid such as coolant through or by one of the turbine components. The term “downstream” corresponds to the direction of fluid flow, and the term “upstream” refers to the direction opposite to the flow. The terms “forward” and “backward” refer to direction, unless further specified, with “forward” referring to the upstream portion of the referenced part, i.e., the part closest to the compressor, and “backward” referring to the downstream portion of the referenced part, i.e., the part furthest from the compressor. Often, it is required to describe parts that are in different radial positions with respect to the central axis. The term “radial” refers to movement or position perpendicular to the axis. In such cases, if the first component is located closer to the axis than the second component, this specification states that the first component is “radially inward” or “inward” of the second component. On the other hand, if the first component is located further from the axis than the second component, this specification may state that the first component is “radially outward” or “outward” of the second component. The term “axial” refers to movement or position parallel to the axis. Finally, the term “circumferential” refers to movement or position about the axis. It will be understood that such terminology can be applied in relation to the central axis of the turbine.

[0039] When an element or layer is referred to as being “on top of,” “engaged,” “disengaged,” “connected,” or “joined” with respect to another element or layer, that element or layer may be directly on top of, engaged with, connected to, or joined to the other element or layer, or there may be an intervening element or layer. Conversely, when an element is referred to as being “directly on top of,” “directly engaged,” “directly connected,” or “directly joined” with respect to another element or layer, there may be no intervening element or layer. Other words used to describe relationships between elements should be interpreted similarly (e.g., “between” and “directly between,” “adjacent” and “directly adjacent,” etc.). As used herein, the term “and / or” includes any combination of one or more of the related enumerated items.

[0040] As described above, embodiments of the present disclosure provide a turbine blade comprising an airfoil body having a radially outermost surface relative to the turbine rotor, a bond coat on the radially outermost surface, and an outermost polished layer on the bond coat. The outermost polished layer may include a gradient oxide ceramic layer or a gradient metal ceramic layer. In either case, vacuum heat treatment of the outermost polished layer hardens the layer, increasing the rub ratio (e.g., 0.1 for cBN compared to the current approximately 0.4), improving the clearance between blade tips, resulting in a smaller tip gap and less leakage compared to conventional systems (e.g., an improvement of 380 micrometers (0.015 inches)). The heat-treated polished layer is also resistant to oxidation at high temperatures exceeding 1090°C (approximately 2000°F), caused by the increasing use of high-temperature fuels such as hydrogen. Vacuum heat treatment also results in improved thermal cycling tests.

[0041] Figure 1 is a schematic diagram of an exemplary turbomachinery system, such as a gas turbine system 100. The system 100 includes a compressor 102, a combustor 104, a turbine 106, a turbine rotor 108 (hereinafter referred to as "rotor 108"), and a fuel nozzle 110. In one embodiment, the system 100 may include multiple compressors 102, combustors 104, turbines 106, rotors 108, and fuel nozzles 110. The compressors 102 and turbines 106 are coupled by the rotor 108. The rotor 108 may be a single shaft or multiple shaft segments coupled to each other to form the rotor 108.

[0042] In one embodiment, the combustor 104 operates the engine using liquid and / or gaseous fuels such as natural gas or hydrogen-enriched synthesis gas. For example, a fuel nozzle 110 is in fluid communication with an air supply source and a fuel supply source 112. The fuel nozzle 110 generates an air-fuel mixture and discharges it into the combustor 104, thereby causing combustion to produce a hot, pressurized exhaust gas. The combustor 104 guides the hot, pressurized gas through a transition piece to the turbine nozzle (or "first-stage nozzle"), as well as to the buckets and other stages of the nozzle, thereby rotating the turbine 106. The rotation of the turbine 106 rotates the rotor 108, thereby compressing the air as it flows into the compressor 102. In one embodiment, but not limited to, hot gas path components including shrouds, diaphragms, nozzles, blades, and transition pieces are located within the turbine 106, where the hot gas flow across the components causes creep, oxidation, wear, and thermal fatigue of the turbine components. The efficiency of a gas turbine increases with increasing flame temperature in the turbine system 100. As flame temperature increases, the likelihood of heat-related damage to high-temperature gas path components such as turbine blades may increase. While the following explanation primarily focuses on the turbine blades of a gas turbine, the concepts described are not limited to gas turbines and can be applied to other forms of gas turbines, such as aircraft engines.

[0043] Figure 2 is a perspective view of an exemplary turbine blade 114 positioned on the turbine of a gas turbine system. It will be understood that the turbine 106 (Figure 1) is mounted downstream of the combustor 104 (Figure 1) and receives the hot combustion gases 116 from there. The turbine 106 (Figure 1), which is axisymmetric about the axial central axis, includes a rotor disk 117 and several circumferentially spaced turbine blades 114 (only one of which is shown) extending radially outward from the rotor disk 117 along the radial axis. The rotor disk 117 is coupled to the shaft or rotor 108 (Figure 1). An annular stationary blade shroud 120 is appropriately bonded to a stationary stator casing (not shown) and surrounds the turbine blades 114 so that a relatively small clearance or gap remains between them, limiting the leakage of combustion gases during operation.

[0044] Each turbine blade 114 includes a base 122 (also called root or dovetail) which can have any conventional form, such as an axial dovetail configured to fit into a corresponding dovetail slot around the rotor disk 117. The airfoil body 124 is integrally joined to the base 122 and extends radially or longitudinally outward from there. It will be recognized that the airfoil body 124 may be hollow and / or may contain cooling passages (not shown). The turbine blade 114 may also include an integral platform 126 positioned at the joint between the airfoil body 124 and the base 122 to define a portion of the radially inward flow path for the combustion gases 116. The turbine blade 114 can be formed in any conventional manner and will be understood to be typically a one-piece casting, an additively manufactured part, or an additively manufactured tip joined to a cast blade base section.

[0045] The airfoil body 124 may include positive pressure sidewalls 128 and negative pressure sidewalls 130 extending between the leading edge 132 and the trailing edge 134. More specifically, the airfoil body 124 preferably includes substantially concave positive pressure sidewalls 128 and substantially convex negative pressure sidewalls 130 that are opposite each other in the circumferential or transverse direction, respectively, and extending axially between the opposing leading edge 132 and trailing edge 134. The sidewalls 128 and 130 also extend radially from the platform 126 to the radially outer end 138, which may also be called the blade tip or simply the “tip”.

[0046] Figure 3 is a perspective view of an exemplary radially outer end 138 (hereinafter referred to as "tip 138" for brevity) at the radially outer end of the airfoil body 124, in which the outermost polishing layer 170 according to embodiments of the present disclosure can be employed. Generally, the tip 138 includes a tip rail 150, which may include any various cooling passages extending therein and / or through the radially outermost surface 160 of the tip rail 150. In some cases, the tip 138 may include a tip cavity 142, and the tip rail 150 surrounds at least a portion of the tip cavity 142. The tip 138 includes a tip plate 146 positioned opposite the base 122 (Figure 2) and defining an outward-facing tip end 148 between the positive pressure sidewall 128 and the negative pressure sidewall 130. The tip plate 146 is typically in contact with an internal cooling passage (not shown in detail) located within the airfoil body 124 and is defined between the positive pressure sidewall 128 and the negative pressure sidewall 130 of the airfoil body 124. The internal cooling cavity 156 is configured to supply coolant through the airfoil body 124, for example, radially. That is, coolant such as compressed air extracted from a compressor can be circulated through the internal cooling cavity 156 during operation. The internal cooling cavity 156 may include, but is not limited to, any currently known or future-developed coolant transport passages or circuits, including cooling passages, impingement sleeves or elements, connecting passages, cavities, pedestals, etc. The tip plate 146 may be integral with the turbine blade 114 or may be welded / brazed into place after the blade has been cast.

[0047] The blade tip 138 includes a tip rail 150 that provides certain performance advantages, such as reduced leak flow. Coinciding with the positive pressure sidewall 128 and negative pressure sidewall 130, the tip rail 150 can be described as including a positive pressure sidewall rail 152 and a negative pressure sidewall rail 154, respectively. Generally, the positive pressure sidewall rail 152 extends radially outward from the tip plate 146 and from the leading edge 132 to the trailing edge 134 of the airfoil body 124. As shown, the path of the positive pressure sidewall rail 152 is adjacent to or near the outer radial edge of the positive pressure sidewall 128 (i.e., around or near the tip plate 146 so as to align with the outer radial edge of the positive pressure sidewall 128). Similarly, as shown, the negative pressure sidewall rail 154 may extend radially outward from the tip plate 146 and from the leading edge 132 to the trailing edge 134 of the airfoil body 124. The path of the negative pressure sidewall rail 154 is adjacent to or near the outer radial edge of the negative pressure sidewall 130 (i.e., around or near the tip plate 146 so as to align with the outer radial edge of the negative pressure sidewall 130). The positive pressure sidewall rail 152 and / or negative pressure sidewall rail 154 can be described as having an inner rail surface 157 and an outer rail surface 159. The positive pressure sidewall rail 152 and / or negative pressure sidewall rail 154 can also be described as having a radially outermost surface 160, i.e., an end surface facing radially outward, between the inner rail surface 157 and the outer rail surface 159. Note here that the radial direction is relative to the turbine rotor 108 (Figure 1). It should be understood that the tip rail 150 does not necessarily have to be along the positive pressure or negative pressure sidewall rail. In other words, in alternative types of tips on which the present disclosure can be used, the tip rail 150 may move away from the edge of the tip plate 146 and may not extend to the trailing edge 134.

[0048] In this configuration, it will be understood that the tip rail 150 defines a tip cavity 142 at the tip 138 of the turbine blade 114. As those skilled in the art will understand, a tip 138 configured in this manner, i.e., having this type of tip cavity 142, is sometimes called a “skiler tip” or a tip having a “skiler pocket or cavity.” The height and width of the positive pressure sidewall rail 152 and / or negative pressure sidewall rail 154 (and therefore the depth of the tip cavity 142) can be varied according to best performance and the size of the overall turbine assembly. It will be understood that the tip plate 146 forms the floor of the tip cavity 142 (i.e., the inner radial boundary of the cavity), the tip rail 150 forms the sidewalls of the tip cavity 142, and the tip cavity 142 remains open through the outer radial surface. When installed in the turbine engine, the tip cavity 142 is closely bordered therefrom by an annular stationary turbine shroud 120 (see Figure 2) that is slightly radially offset from it. As further described herein, the radially outermost surface 160 of the tip 138, over which the outermost polishing layer 170 is located, may rub against the annular stationary turbine shroud 120 (Figure 2). In certain embodiments, the airfoil body 124 and the tip 138 may be cast, or they may be additively manufactured. In either case, the cooling passage 162 may be cast or additively manufactured simultaneously during the airfoil manufacturing process. The airfoil body 124 may include, but is not limited to, any currently known or future-developed material capable of withstanding the environment of the turbine 106 (Figure 1), such as a nickel-based superalloy. While a specific configuration of the tip 138 has been described herein, it will be recognized that other configurations including the radially outermost surface 160 are also possible and are considered to be within the scope of this disclosure.

[0049] Figures 4A to 4B show cross-sectional views of the outermost polishing layer 170 (hereinafter referred to as "polishing layer 170" for brevity) on the radial outermost surface 160 according to various embodiments of the present disclosure. Furthermore, Figures 5 to 8 show cross-sectional views of a method for forming the polishing layer 170 for the radial outermost surface 160 of the turbine blade 114. The method for forming the polishing layer 170 for the radial outermost surface 160 of the turbine blade 114 generally includes forming a bond coat 172 on the radial outermost surface 160 of the turbine blade 114, forming the polishing layer 170 as one of a gradient oxide ceramic layer 174 and a gradient metal ceramic layer 176 on the bond coat 172 on the radial outermost surface 160 of the turbine blade 114, and vacuum heat treatment of one of the gradient oxide ceramic layer 174 and the gradient metal ceramic layer 176 on the radial outermost surface 160 of the turbine blade 114.

[0050] The turbine blade 114 includes a bond coat 172 on its radially outermost surface 160. The bond coat 172 can include an M-chromium-aluminum-yttrium (MCrAlY) alloy, where M includes cobalt-nickel (more Co than Ni), nickel (Ni), nickel-cobalt (more Ni than Co), or cobalt (Co). Thus, the bond coat 172 can include at least one of cobalt-nickel-chromium-aluminum-yttrium (CoNiCrAlY), nickel-chromium-aluminum-yttrium (NiCrAlY), nickel-cobalt-chromium-aluminum-yttrium (NiCoCrAlY), and cobalt-chromium-aluminum-yttrium (CoCrAlY). Other forms of the bond coat 172 are also possible. Figure 5 shows the formation of the bond coat 172 on the radially outermost surface 160 of the turbine blade 114. The bond coat 172 can be formed using any currently known or future-developed technique, such as thermal spraying, thermal spraying, and / or plasma spraying, but is not limited to these. In certain embodiments, as shown only in Figures 4A and 5, the bond coat 172 may include two layers 172A and 172B, and the formation of the bond coat 172 may include forming the two layers 172A and 172B. For example, the first (dense) bond coat layer 172A may be formed by a high-speed oxygen fuel (HVOF) thermal spraying process or a low-temperature thermal spraying process, and the second (low-density, coarse) bond coat layer 172B may be formed by an HVOF thermal spraying process, a flame spraying process, or an air-plasma spraying process. The first bond coat layer 172A can have a thickness TB1 of 55 to 76 micrometers (μm) (Figure 4A), and the second bond coat layer 172B can have a thickness TB2 of 30 to 50 μm (Figure 4A). Therefore, the bond coat 172 can have an overall thickness T1 of 85 to 125 micrometers (μm) (approximately 0.003 to 0.0005 inches), although other thicknesses are also possible.

[0051] The turbine blade 114 also includes a polishing layer 170 on top of the bond coat 172. As shown in Figures 4A to 4B, the outermost polishing layer 170 can take various forms.

[0052] Referring to Figure 4A, in a particular embodiment, the polished layer 170 includes a gradient oxide ceramic layer 174 comprising several different layers having different concentrations of oxide ceramic. More specifically, the gradient oxide ceramic layer 174 may include a first layer 180A on a bond coat 172, comprising 100 wt percent (W%) of a first ceramic and 0 wt percent of a second ceramic different from the first ceramic. In this embodiment, the first ceramic is 8 wt percent yttria-stabilized zirconia (8YSZ), i.e., (ZrO2) 0.92 (Y2O3) 0.08 The first and second ceramics are solid, i.e., have 0% porosity.

[0053] When used herein, the weight percentage (W%) of materials is based on those present in the final product (and initial feedstock). However, it is recognized that, without limitation, mixtures, blends, and / or adhesives or carriers, such as polyvinyl acetate (PVA), may be used during the application process. Typically, these agents or carriers evaporate or dissipate during the application process described herein, but the final product may contain some small residual amounts of these agents or carriers, e.g., less than 1%.

[0054] The gradient oxide ceramic layer 174 may also include a second layer 180B on top of the first layer 180A. The second layer 180B may contain 65–85 wt% of the first ceramic (i.e., 8YSZ) and 15–35 wt% of the second ceramic (i.e., pure Al2O3 or ZTA). In certain embodiments, the second layer 180B may contain 75 wt% of the first ceramic and 25 wt% of the second ceramic. The gradient oxide ceramic layer 174 may also include a third layer 180C on top of the second layer 180B. The third layer 180C may contain 40–60 wt% of the first ceramic and 40–60 wt% of the second ceramic. In certain embodiments, the third layer 180C may contain 50 wt% of the first ceramic and 50 wt% of the second ceramic. The gradient oxide ceramic layer 174 may also include a fourth layer 180D on top of the third layer 180C. The fourth layer 180D may contain 5–25 wt% of a first ceramic and 75–95 wt% of a second ceramic. In certain embodiments, the third layer 180C may contain 15 wt% of a first ceramic and 85 wt% of a second ceramic. The gradient oxide ceramic layer 174 may also include a fifth (outermost) layer 180E on top of the fourth layer 180D. The fifth layer 180E may contain 0 wt% of a first ceramic (i.e., 8YSZ) and 100 wt% of a second ceramic (i.e., pure Al2O3 or ZTA). The fifth layer 180E may have a porosity of less than 2%, but a porosity of 5–10% is also possible. When used herein, "porosity" is the ratio of the open space volume to the total volume of the described structure, e.g., layer 180E. Typically, in this regard, porosity is described as the ratio of the volume of cavity space to the whole or total volume of the described structure. Cavity space is an empty area within a solid material, which may be referred to herein as “pores,” and may contain interconnection passages within the material of the described structure.As used herein, for the purpose of identifying its “total volume,” a three-dimensional boundary of a porous region or subregion can be identified by a change in porosity of more than 2% relative to an adjacent region or subregion, for example, occurring within layer 180E. In certain embodiments, each of the first to fourth layers 180A to D may have a thickness T2 of 55 to 75 μm (i.e., about 0.001 to 0.002 inches), and the fifth layer 180E may have a thickness T3 of 100 to 125 μm (about 0.0039 to 0.009 inches). Other thicknesses may also be possible.

[0055] As shown in Figure 7, each of layers 180A to E is formed sequentially or gradually on the bond coat 172 until the fifth layer 180E is formed to the desired thickness. Although five layers 180A to E are shown, it will be recognized that substantially any number of distinct layers may exist, as the gradient oxide ceramic layer 174 gradually changes the concentration of each of the first and second ceramics. The polished layer 170, i.e., the layer of the gradient oxide ceramic layer 174, can be formed using any currently known or hereafter developed form of thermal plasma spraying, e.g., atmospheric pressure plasma spraying, but is not limited to other processes such as laser cladding and sintering. By controlling the parameters of the formation process in a known manner, layers of different concentrations at the thicknesses described herein can be created.

[0056] In contrast to conventional processes, as shown in Figure 8, the formation of the polished layer 170, for example, the gradient oxide ceramic layer 174, includes a post-formation vacuum heat treatment (arrow 184). In certain embodiments, the heat treatment may include exposing the gradient oxide ceramic layer 174 to a temperature above 1000°C for a period of at least 1 hour, and then cooling the gradient oxide ceramic layer 174. In other embodiments, the vacuum heat treatment may include raising the temperature of the gradient oxide ceramic layer 174 to a temperature above 1200°C for a period of at least 1.5 hours, and then cooling the gradient oxide ceramic layer 174. Figure 9 shows a graphical representation of the vacuum heat treatment.

[0057] Referring to Figure 4B, in another embodiment, the polishing layer 170 may include a gradient metal-ceramic layer 176 comprising several different layers having different concentrations of metal and ceramic. More specifically, the gradient metal-ceramic layer 176 may include a first layer 182A on top of a bond coat 172 consisting of a third ceramic made of 8YSZ having a porosity of 5-25%. The gradient metal-ceramic layer 176 may also include a second layer 182B on top of the first layer 182A. The second layer 182B may contain 65-85 wt% metal and 15-35 wt% fourth ceramic. In a particular embodiment, the third layer 182C may contain 75 wt% metal and 25 wt% fourth ceramic. In a particular embodiment, the metal may include at least one of MCrAlY, nickel-boron-silicon (NiBSi) alloy, nickel-chromium-iron-molybdenum (NiCrFeMo) alloy, and cobalt-chromium-molybdenum (CoCrMo). As described herein, in MCrAlY alloys, M includes cobalt-nickel (more Co than Ni), nickel (Ni), nickel-cobalt (more Ni than Co), or cobalt (Co). Thus, in certain embodiments, the metal of the gradient metal ceramic layer 176 may include at least one of cobalt-nickel-chromium-aluminum-yttrium (CoNiCrAlY), nickel-chromium-aluminum-yttrium (NiCrAlY), nickel-cobalt-chromium-aluminum-yttrium (NiCoCrAlY), and cobalt-chromium-aluminum-yttrium (CoCrAlY). When NiCrFeMo alloy is used, the metal may include HASTELLOY® X, and when CoCrMo alloy is used, the metal may include TRIBALOY® T-800® and T-400®. In certain embodiments, the fourth ceramic may differ from the third ceramic and may include pure alumina (Al2O3) or zirconia-reinforced alumina (ZTA), as previously described herein.In another embodiment, the fourth ceramic may include a non-porous 8YSZ, i.e., it is solid in contrast to the porous 8YSZ of the first layer 182A of the gradient metal-ceramic layer 176.

[0058] The gradient metal-ceramic layer 176 may also include a third layer 182C on top of the second layer 182B. The third layer 182C may contain 40-60% by weight of metal and 40-60% by weight of a fourth ceramic. In certain embodiments, the third layer 182C may contain 50% by weight of metal and 50% by weight of a fourth ceramic. The gradient metal-ceramic layer 176 may also include a fourth layer 182D on top of the third layer 182C. The fourth layer 182D may contain 5-25% by weight of metal and 75-95% by weight of a fourth ceramic. In certain embodiments, the fourth layer 182D may contain 15% by weight of metal and 85% by weight of a fourth ceramic. The gradient metal-ceramic layer 176 may also include a fifth (outermost) layer 182E on top of the fourth layer 182D. The fifth layer 182E may contain 0 wt% metal, e.g., MCrAlY, and 100 wt% fourth ceramic, e.g., alumina or ZTA. In Figure 4B, the bond coat 172 can have a thickness T4 of 85–125 μm (approximately 0.003–0.005 inches). Furthermore, the first to fourth layers 182A–D can each have a thickness T5 of 55–75 μm (approximately 0.001–0.002 inches), and the fifth layer 182E can have a thickness T6 of 100–125 μm (approximately 0.0039–0.0049 inches). Other thicknesses are also possible.

[0059] As shown in Figure 7, each of layers 182A-E is formed sequentially or gradually on the bond coat 172 until the fifth layer 182E is formed to the desired thickness. Although five layers 182A-E are shown, it will be recognized that substantially any number of distinct layers may exist, as the gradient metal-ceramic layer 176 gradually changes the concentration of each of the metal and the fourth ceramic. The polished layer 170, i.e., the gradient metal-ceramic layer 176, can be formed using any currently known or hereafter developed form of thermal plasma spraying, e.g., atmospheric pressure plasma spraying, laser cladding, sintering, etc. In contrast to conventional processes, as shown in Figure 8, the method may include exposing the polished layer 170, i.e., the gradient metal-ceramic layer 176, to a vacuum heat treatment after formation (arrow 184). In certain embodiments, the heat treatment may include exposing the gradient metal-ceramic layer 176 to a temperature above 1000°C for a period of at least 1 hour, and then cooling the gradient metal-ceramic layer 176. In another embodiment, the vacuum heat treatment may include raising the temperature of the gradient metal-ceramic layer 176 to a temperature above 1200°C for a period of at least 1.5 hours, and then cooling the gradient metal-ceramic layer 176. As described above, Figure 9 shows a graphical representation of the vacuum heat treatment.

[0060] Furthermore, regarding heat treatment, regardless of the format of the polishing layer 170, the heat treatment results in hardening of the layer, increasing the friction ratio (e.g., 0.1 for cBN compared to the current approximately 0.4), improving the clearance between the tip 138 (Figure 2) and the shroud 120 (Figure 2), resulting in a smaller tip gap and less leakage compared to conventional systems (e.g., an improvement of 380 micrometers (0.015 inches)). The heat-treated polishing layer 170 is also resistant to oxidation at high temperatures exceeding 1090°C (approximately 2000°F), such as those caused by the increasing use of high-temperature fuels like hydrogen.

[0061] As part of quality assurance testing, the turbine blade section 114 is subjected to a thermal cycling test, sometimes called a furnace cycle test (FCT). The FCT involves exposing the turbine blade section body 124, more specifically the polishing layer 170, to repeated high-temperature cycles. Figure 10 is a graphical representation of an exemplary FCT process cycle. As illustrated, in one non-limiting example of the FCT, the temperature of the turbine blade section body 124 and the polishing layer 170 can be raised to over 1180°C (approximately 2160°F) over 10 minutes, held at that temperature for 30 minutes, and then forcibly cooled to approximately 93°C (approximately 200°F) over 10 minutes (with a fan). The cycle is then repeated many times until a failure occurs in the tested turbine blade section body 124, such as delamination or cracking of at least a portion of the polishing layer 170 on it.

[0062] Vacuum heat treatment according to embodiments of this disclosure results in several performance improvements of the polishing layers 170 observed through thermal cycling tests. As described herein, depending on the material used, a significant increase in hardness of the outermost layers of the polishing layers 170, i.e., fifth layers 180E, 182E, is observed. For example, for the gradient oxide ceramic layers 174: a) when ZTA is used as the second ceramic, a 17.5% increase in hardness is observed; b) when 8YSZ is used as the second ceramic, a 12.3% increase in hardness is observed; and c) when alumina is used as the second ceramic, a 13.0% increase in hardness is observed. The layers between the bond coat 172 and the outermost layer of the polishing layers 170, i.e., fifth layers 180E, 182E, can be called adhesive layers because they act to bond the outermost layer to the bond coat 172. FCT is also used to test the lifespan of these adhesive layers, which is indicated, for example, by the number of thermal cycles remaining before adhesion fails due to peeling or separation. As described herein, depending on the material used, a significant increase in the lifespan of the adhesive layer is observed for the polished layer 170. For example, in the case of the gradient oxide ceramic layer 174 in which 8YSZ and alumina are used as the first and second ceramics (i.e., having layers of 100 wt% 8YSZ, 75 wt% 8YSZ + 25 wt% Al2O3, 50 wt% 8YSZ + 50 wt% Al2O3, 15 wt% 8YSZ + 85 wt% Al2O3, and 100 wt% Al2O3), a 141% increase in remaining cycles is observed.

[0063] Embodiments of this disclosure offer various technical and commercial advantages, examples of which are discussed herein. Vacuum heat treatment of turbine blades having the polished layer described herein to harden the layer increases the friction ratio (e.g., 0.1 for cBN compared to the current approximately 0.4), improves the clearance between blade tips, reduces the tip gap and reduces leakage compared to conventional systems (e.g., an improvement of 380 micrometers (0.015 inches)). The heat-treated polished layer is also resistant to oxidation at high temperatures, e.g., above 1090°C (approximately 2000°F), caused by high temperatures in the high-temperature gas path. Vacuum heat treatment also results in improved thermal cycling testing, as described herein.

[0064] Throughout this specification and the claims, the approximation language can be applied to modify any quantitative expression that can vary to a reasonable extent without altering the fundamental function of the expression. Thus, values ​​modified by terms such as “approximately,” “about,” and “substantially” are not limited to the exact value specified. In at least some examples, the approximation language may correspond to the precision of the instrument used to measure the value. Here, and throughout this specification and the claims, limitations on ranges are interchangeable and / or replaceable, and unless the context or wording specifically indicates otherwise, such ranges are identified and include all subranges encompassed therein. “Approximately” or “about” applied to a particular value within a range may indicate + / - 10% of the stated value, unless applied to the values ​​at both ends and particularly dependent on the precision of the instrument used to measure the value.

[0065] All means or step-plus-function elements in the following claims are intended to include any structures, materials, or actions for performing a function in combination with other claimed elements specifically claimed. The descriptions in this disclosure are presented for illustrative and explanatory purposes and are not intended to be exhaustive or to limit the disclosure to the forms disclosed herein. Many modifications and variations will be apparent to those skilled in the art without departing from the scope and spirit of this disclosure. The embodiments have been selected and described to best illustrate the principles of this disclosure and the practical applications of the art, and to enable those skilled in the art to understand this disclosure in order to consider various modifications to these embodiments that may be suitable for the particular use under consideration. [Explanation of symbols]

[0066] 100 Gas Turbine Systems 102 Compressor 104 Combustor 106 Turbine 108 Turbine Rotor 110 Fuel Nozzle 112 Air supply source / Fuel supply source 114 Turbine Blades 116 High-temperature combustion gases 117 Rotor Disc 120 Annular stationary blade shroud / Annular stationary turbine shroud 122 Base 124 Airfoil section 126 Integrated Platform 128 Positive pressure sidewall 130 Negative pressure sidewall 132 Leading edge 134 Trailing edge 138 Radial outer end / tip / blade tip 142 Tip cavity 146 Tip plate 148 Tip end 150 Front Rail 152 Positive pressure side wall rail 154 Negative pressure side wall rail 156 Internal cooling cavity 157 Inner rail surface 159 Outer rail surface 160 Radial outermost surface 162 Cooling passage 170 Outermost polishing layer / polishing layer 172 Bond Coat 172A First bond coat layer 172B Second Bond Coat Layer 174. Gradient Oxide Ceramic Layer 176 Gradient Metal-Ceramic Layer 180A First layer 180B Second Layer 180C Third Layer 180D, fourth layer 180E Fifth Layer 182A First layer 182B Second Layer 182C Third Layer 182D The fourth layer 182E Fifth Layer 184 Arrow T1 Total thickness T2 thickness T3 thickness T4 thickness T5 thickness T6 thickness TB1 Thickness TB2 thickness

Claims

1. A turbine blade (114), A blade-shaped section body (124) having the outermost radial surface (160) relative to the turbine rotor (108), The bond coat (172) on the radially outermost surface (160), The outermost polished layer (170) on the bond coat (172), wherein the outermost polished layer (170) is A first layer (180A, 182A) on the bond coat (172) comprising 100% by weight of a first ceramic and 0% by weight of a second ceramic different from the first ceramic, wherein the first ceramic comprises a first layer (180A, 182A) consisting of 8% by weight of yttria-stabilized zirconia (8YSZ), A second layer (180B, 182B) on the first layer (180A, 182A) comprising 65-85% by weight of the first ceramic and 15-35% by weight of the second ceramic, A third layer (180C, 182C) on the second layer (180B, 182B) comprising 40-60 wt% of the first ceramic and 40-60 wt% of the second ceramic, A fourth layer (180D, 182D) on the third layer (180C, 182C) comprising 5 to 25% by weight of the first ceramic and 75 to 95% by weight of the second ceramic, A fifth layer (180E, 182E) on the fourth layer (180D, 182D) comprising 0% by weight of the first ceramic and 100% by weight of the second ceramic, and The outermost polished layer (170) includes a gradient oxide ceramic layer (174) and A turbine blade (114) is provided.

2. The turbine blade (114) according to claim 1, wherein the bond coat (172) comprises at least one of cobalt-nickel-chromium-aluminum-yttrium (CoNiCrAlY), nickel-chromium-aluminum-yttrium (NiCrAlY), nickel-cobalt-chromium-aluminum-yttrium (NiCoCrAlY), and cobalt-chromium-aluminum-yttrium (CoCrAlY).

3. The second ceramic described above is pure alumina (Al ２ O ３ A turbine blade (114) according to claim 1, comprising one of the following: ) or zirconia-reinforced alumina (ZTA).

4. The turbine blade (114) according to claim 1, wherein the bond coat (172) has a thickness of 85 to 125 micrometers (μm) (T1, T4), each of the first to fourth layers (180A, 180B, 180C, 180D, 182A, 182B, 182C, 182D) has a thickness of 55 to 75 μm (T2, T5), and the fifth layer (180E, 182E) has a thickness of 100 to 125 μm (T3, T6).

5. The turbine blade (114) according to claim 1, wherein the outermost polished layer (170) is formed using thermal plasma spraying and post-formation vacuum heat treatment, which includes exposing the gradient oxide ceramic layer (174) to a temperature exceeding 1000°C for a period of at least one hour, and then cooling the gradient oxide ceramic layer (174).

6. A method for forming an outermost polished layer (170) for the radial outermost surface (160) of a turbine blade (114), Forming a bond coat (172) on the radially outermost surface (160) of the turbine blade (114), Forming one of a gradient oxide ceramic layer (174) and a gradient metal ceramic layer (176) on the bond coat (172) on the radially outermost surface (160) of the turbine blade (114), Heat treatment of one of the gradient oxide ceramic layer (174) and the gradient metal ceramic layer (176) on the radially outermost surface (160) of the turbine blade (114) Methods that include...

7. The method according to claim 6, wherein the heat treatment includes a vacuum heat treatment comprising exposing the outermost polished layer (170) to a temperature exceeding 1000°C for a period of at least one hour, and then cooling the outermost polished layer (170).

8. The method according to claim 7, wherein the vacuum heat treatment includes raising the temperature of the outermost polished layer (170) to a temperature exceeding 1200°C for a period of at least 1.5 hours, and then cooling the outermost polished layer (170).

9. The method according to claim 6, wherein the heat treatment comprises exposing the outermost polished layer (170) to a temperature of 750°C to 1000°C in air for a period of 30 minutes to 4 hours.

10. The method according to claim 6, wherein the bond coat (172) comprises at least one of cobalt-nickel-chromium-aluminum-yttrium (CoNiCrAlY), nickel-chromium-aluminum-yttrium (NiCrAlY), nickel-cobalt-chromium-aluminum-yttrium (NiCoCrAlY), and cobalt-chromium-aluminum-yttrium (CoCrAlY).

11. The outermost polished layer (170) is A first layer (180A, 182A) on the bond coat (172) comprising 100% by weight of a first ceramic and 0% by weight of a second ceramic different from the first ceramic, wherein the first ceramic comprises a first layer (180A, 182A) consisting of 8% by weight of yttria-stabilized zirconia (8YSZ), A second layer (180B, 182B) on the first layer (180A, 182A) comprising 65-85% by weight of the first ceramic and 15-35% by weight of the second ceramic, A third layer (180C, 182C) on the second layer (180B, 182B) comprising 40-60 wt% of the first ceramic and 40-60 wt% of the second ceramic, A fourth layer (180D, 182D) on the third layer (180C, 182C) comprising 5 to 25% by weight of the first ceramic and 75 to 95% by weight of the second ceramic, A fifth layer (180E, 182E) on the fourth layer (180D, 182D) comprising 0% by weight of the first ceramic and 100% by weight of the second ceramic, and The method according to claim 6, comprising a gradient oxide ceramic layer (174) containing the following.

12. The second ceramic described above is pure alumina (Al ２ O ３ The method according to claim 11, comprising one of the following: ) or zirconia-reinforced alumina (ZTA).

13. The method according to claim 11, wherein the bond coat (172) has a thickness of 85 to 125 micrometers (μm) (T1, T4), each of the first to fourth layers (180A, 180B, 180C, 180D, 182A, 182B, 182C, 182D) has a thickness of 55 to 75 μm (T2, T5), and the fifth layer (180E, 182E) has a thickness of 100 to 125 μm (T3, T6).

14. The outermost polished layer (170) is A first layer (180A, 182A) on the bond coat (172) is made of a first ceramic consisting of 8 wt% yttria-stabilized zirconia (8YSZ) having a porosity of 5-25%, A second layer (180B, 182B) on top of the first layer (180A, 182A) comprising 65-85% by weight of metal and 15-35% by weight of a second ceramic, A third layer (180C, 182C) on the second layer (180B, 182B) comprising 40-60% by weight of the metal and 40-60% by weight of the second ceramic, A fourth layer (180D, 182D) on the third layer (180C, 182C) comprising 5 to 25% by weight of the metal and 75 to 95% by weight of the second ceramic, and a fifth layer on the fourth layer (180D, 182D) comprising 0% by weight of the metal and 100% by weight of the second ceramic. The method according to claim 6, comprising a gradient metal ceramic layer (176) containing the following.

15. Unlike the first ceramic, the second ceramic is made of pure alumina (Al ２ O ３ The method according to claim 14, comprising one of the following: ) or zirconia-reinforced alumina (ZTA).

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