Gas turbine stationary blade, grouped blade unit, and manufacturing method of gas turbine stationary blade

Abstract

To provide a gas turbine stationary blade capable of efficiently reducing an amount of antioxidation coating reduced in damage of the gas turbine stationary blade caused by oxidation.SOLUTION: In a gas turbine stationary blade including an airfoil portion, a surface of the airfoil portion includes a front edge, a rear edge, a pressure surface connecting the front edge and the rear edge and a negative pressure surface connecting the front edge and the rear edge, and an antioxidation coating layer is formed within at least a part of a range at the pressure surface side of the surface of the airfoil portion. In a cross section orthogonal to a blade height direction in the airfoil portion, an average value of a thickness of the antioxidation coating layer at the pressure surface side within a range of 0.90≤X / Xmax≤1.0 is smaller than an average value of a thickness of the antioxidation coating layer at the pressure surface side within a range of 0.0≤X / Xmax≤0.50, when a distance from the front edge to the rear edge in an axial direction of a gas turbine is Xmax and an axial position when the front edge is made an origin, is X.SELECTED DRAWING: Figure 2

Description

[Technical field]

[0001] The present disclosure relates to a gas turbine vane, a superstrate vane unit, and a method for manufacturing a gas turbine vane. [Background technology]

[0002] Patent Document 1 discloses that a foreign object that collides with a gas turbine rotor blade and is reflected off the gas turbine blade collide with the trailing edge side of the gas turbine stator blade, causing damage to the gas turbine stator blade, and that in order to solve this problem, a protective coating is partially formed only on the trailing edge side of the gas turbine stator blade. [Prior art documents] [Patent documents]

[0003] [Patent Document 1] Japanese Patent Application Publication No. 10-148101 Summary of the Invention [Problem to be solved by the invention]

[0004] However, when a gas turbine stator vane is exposed to a high-temperature ambient gas, the gas turbine stator vane may be damaged due to the progress of thinning caused by oxidation. For this reason, it is conceivable to apply an oxidation-resistant coating to the surface of the gas turbine stator vane, but applying the oxidation-resistant coating to the entire surface of the gas turbine stator vane requires a great deal of effort and cost for the coating. In this regard, Patent Document 1 does not disclose any knowledge for efficiently suppressing damage to gas turbine vanes caused by oxidation with a small amount of oxidation-resistant coating.

[0005] In consideration of the above circumstances, at least one embodiment of the present disclosure aims to provide a gas turbine stator vane, a vane assembly, and a method for manufacturing a gas turbine stator vane, which are capable of efficiently suppressing damage to the gas turbine stator vane caused by oxidation with a small amount of oxidation-resistant coating. [Means for solving the problem]

[0006] In order to achieve the above object, a gas turbine stator vane according to at least one embodiment of the present disclosure comprises: 1. A gas turbine vane comprising an airfoil, a surface of the airfoil including a leading edge, a trailing edge, a pressure side connecting the leading edge and the trailing edge, and a suction side connecting the leading edge and the trailing edge; an oxidation-resistant coating layer is formed on at least a portion of the pressure surface side of the surface of the airfoil; In a cross section perpendicular to the blade height direction of the airfoil portion, a distance from the leading edge to the trailing edge in the axial direction of the gas turbine is defined as Xmax, and a position in the axial direction when the leading edge is the origin is defined as X, The average thickness of the oxidation-resistant coating layer on the pressure surface side in the range of 0.90≦X / Xmax≦1.0 is smaller than the average thickness of the oxidation-resistant coating layer on the pressure surface side in the range of 0.0≦X / Xmax≦0.50.

[0007] In order to achieve the above object, the binding wing unit according to at least one embodiment of the present disclosure comprises: A splice blade unit including a plurality of gas turbine stator vanes, the plurality of gas turbine stator vanes being integrally molded, each of the plurality of gas turbine stator vanes being the gas turbine stator vane.

[0008] In order to achieve the above object, a method for manufacturing a gas turbine vane according to at least one embodiment of the present disclosure includes: Applying an oxidation-resistant coating layer to at least a portion of a pressure surface of a surface of an airfoil of a gas turbine vane, In a cross section of the airfoil portion perpendicular to the blade height direction, a position in the axial direction of the gas turbine is defined as a distance X from a leading edge of the airfoil portion to the position in the axial direction, and a position of a trailing edge of the airfoil portion in the axial direction is defined as Xmax. In the step, the oxidation-resistant coating layer is formed on at least a portion of the pressure surface side of the surface of the airfoil section such that an average value of a thickness of the oxidation-resistant coating layer on the pressure surface side in a range of 0.90≦X / Xmax≦1.0 is smaller than an average value of a thickness of the oxidation-resistant coating layer on the pressure surface in a range of 0.0≦X / Xmax≦0.50. Effect of the Invention

[0009] According to at least one embodiment of the present disclosure, there are provided a gas turbine stator vane, a vane assembly, and a method for manufacturing a gas turbine stator vane, which are capable of efficiently suppressing damage to the gas turbine stator vane caused by oxidation with a small amount of oxidation-resistant coating. [Brief description of the drawings]

[0010] [Figure 1] 1 is a diagram showing a cross-sectional shape perpendicular to the blade height direction in a blade unit 2 for a gas turbine according to one embodiment. FIG. [Diagram 2] 2 is an enlarged view showing a schematic cross-sectional shape perpendicular to the blade height direction of an airfoil portion 6 of the gas turbine stator blade 4 shown in FIG. [Diagram 3] FIG. 2 is a diagram for explaining an example of a method for applying an oxidation-resistant coating to a gas turbine vane 4. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0011] Hereinafter, some embodiments of the present disclosure will be described with reference to the accompanying drawings. However, the dimensions, materials, shapes, relative arrangements, etc. of components described as the embodiments or shown in the drawings are merely illustrative examples and are not intended to limit the scope of the invention. For example, expressions expressing relative or absolute configuration, such as "in a certain direction," "along a certain direction," "parallel," "orthogonal," "center," "concentric," or "coaxial," not only strictly express such a configuration, but also express a state in which there is a relative displacement with a tolerance or an angle or distance to the extent that the same function is obtained. For example, expressions indicating that things are in an equal state, such as "identical," "equal," and "homogeneous," not only indicate a state of strict equality, but also indicate a state in which there is a tolerance or a difference to the extent that the same function is obtained. For example, expressions describing shapes such as a rectangular shape or a cylindrical shape do not only refer to rectangular shapes, cylindrical shapes, etc. in the strict geometric sense, but also refer to shapes that include uneven portions, chamfered portions, etc., to the extent that the same effect is obtained. On the other hand, the expressions "comprise," "include," "have," "includes," or "have" of one element are not exclusive expressions excluding the presence of other elements.

[0012] FIG. 1 is a diagram illustrating a cross-sectional shape perpendicular to the blade height direction in a splice blade unit 2 for a gas turbine according to one embodiment.

[0013] 1, the splice vane unit 2 includes a plurality of gas turbine vanes 4, and is integrally configured by integrally molding the plurality of gas turbine vanes 4. Each of the plurality of gas turbine vanes 4 includes an airfoil portion 6, an outer shroud 8, and an inner shroud 10. The plurality of airfoil portions 6 of the splice vane unit 2 are arranged at intervals in the circumferential direction of the gas turbine, and the outer shrouds 8 of adjacent gas turbine vanes 4 in the plurality of gas turbine vanes 4 are connected to each other, and the inner shrouds 10 of adjacent gas turbine vanes 4 in the plurality of gas turbine vanes 4 are connected to each other.

[0014] 2 is an enlarged view showing a schematic cross-sectional shape perpendicular to the blade height direction in the airfoil portion 6 of the gas turbine stator vane 4 shown in FIG. 1. In this specification, the blade height direction may coincide with the radial direction of the gas turbine. Since the airfoil portions 6 of the gas turbine stator vanes 4 included in the splice vane unit 2 have the same structure, the configurations of the airfoil portions 6 of the gas turbine stator vanes 4 will be described below without any particular distinction. In addition, in the following, "axial direction" means the axial direction of the gas turbine unless otherwise specified, "circumferential direction" means the circumferential direction of the gas turbine unless otherwise specified, and "radial direction" means the radial direction of the gas turbine unless otherwise specified.

[0015] 2, a surface 7 of an airfoil 6 of a gas turbine stator vane 4 includes a leading edge 12, a trailing edge 14, a pressure side 16 connecting the leading edge 12 and the trailing edge 14, and a suction side 18 connecting the leading edge 12 and the trailing edge 14. In the illustrated exemplary embodiment, a cavity 9 is formed inside the airfoil 6, extending along the blade height direction.

[0016] 2, an oxidation-resistant coating layer 20 is formed on each of the pressure surface 16 and the suction surface 18 of the airfoil 6 of the gas turbine stator vane 4. In the illustrated exemplary embodiment, the oxidation-resistant coating layer 20 is formed on a partial range of the surface 7 of the airfoil 6 on the pressure surface 16 side and on the entire surface 7 of the airfoil 6 on the suction surface 18 side. Here, the oxidation-resistant coating layer 20 may be a coating layer made of, for example, CoNiCrAlY, NiCrAlY, CoCrAlY, or MCrAlX (wherein M is Ni, Co, or an alloy of Ni and Co, i.e., NiCo, and X is Hf, Si, or Ta).

[0017] 2, the oxidation-resistant coating layer 20 on the pressure surface 16 side is formed over a range from the position of the leading edge 12 on the pressure surface 16 to a position P1 in the vicinity of the trailing edge 14, but is not formed over a range from the position P1 in the vicinity of the trailing edge 14 on the pressure surface 16 to the trailing edge 14. Here, the position P1 is a position between the trailing edge 14 and a position P2 on the pressure surface 16 where the blade thickness t of the airfoil portion 6 of the gas turbine stator blade 4 is maximum.

[0018] The oxidation-resistant coating layer 20 formed on the pressure surface 16 side of the surface 7 of the airfoil 6 includes an oxidation-resistant coating layer leading edge side portion 20h1 formed over a range from the leading edge 12 to a position P3 on the trailing edge 14 side on the pressure surface 16, and an oxidation-resistant coating layer trailing edge side portion 20h2 formed over a range from the position P3 to a position P1 on the pressure surface 16. Here, the thickness dh1 of the oxidation-resistant coating layer leading edge side portion 20h1 is greater than the thickness dh2 of the oxidation-resistant coating layer trailing edge side portion 20h2. The thickness dh1 of the oxidation-resistant coating layer leading edge side portion 20h1 may be constant from the leading edge 12 to the position P3, and the thickness dh2 of the oxidation-resistant coating layer trailing edge side portion 20h2 may decrease as it approaches the position P1 from the position P3.

[0019] In the embodiment shown in Fig. 2, if the position on the suction surface 18 where the blade thickness t of the airfoil 6 is maximum is P4 in a cross section perpendicular to the blade height direction of the airfoil 6, the oxidation-resistant coating layer 20 formed on the suction surface 18 side of the surface 7 of the airfoil 6 includes an oxidation-resistant coating layer intermediate portion 20s2 formed in a range including the position P4 on the suction surface 18, an oxidation-resistant coating layer leading edge side portion 20s1 formed in a range between the oxidation-resistant coating layer intermediate portion 20s2 on the suction surface 18 and the leading edge 12, and an oxidation-resistant coating layer trailing edge side portion 20s3 formed in a range between the oxidation-resistant coating layer intermediate portion 20s2 on the suction surface 18 and the trailing edge 14. The thickness ds2 of the oxidation-resistant coating layer intermediate portion 20s2 is smaller than the thickness ds1 of the oxidation-resistant coating layer leading edge side portion 20s1 and smaller than the thickness ds3 of the oxidation-resistant coating layer trailing edge side portion 20s3.

[0020] In the embodiment shown in Figure 2, the oxidation-resistant coating layer leading edge side portion 20s1 is an oxidation-resistant coating layer formed over a range from position P5 between the leading edge 12 and position P4 on the surface 7 of the airfoil 6 to the leading edge 12, and the thickness ds1 of the oxidation-resistant coating layer leading edge side portion 20s1 may be constant from the leading edge 12 to position P5.

[0021] In the exemplary embodiment shown in FIG. 2, the oxidation-resistant coating layer trailing edge side portion 20s3 is an oxidation-resistant coating layer formed over a range from position P6 between the trailing edge 14 and position P4 on the surface 7 of the airfoil 6 to the trailing edge 14, and the thickness ds3 of the oxidation-resistant coating layer trailing edge side portion 20s3 may be constant from the trailing edge 14 to position P6.

[0022] In the embodiment shown in FIG. 2, in a cross section perpendicular to the blade height direction in the airfoil portion 6, if a position in the axial direction is defined as the distance X from the leading edge 12 to that position in the axial direction and the position of the trailing edge 14 in the axial direction (i.e., the distance from the leading edge 12 to the trailing edge 14 in the axial direction) is defined as Xmax, the average value of the thickness of the oxidation-resistant coating layer 20 on the pressure surface 16 side in the range of 0.90≦X / Xmax≦1.0 is smaller than the average value of the thickness of the oxidation-resistant coating layer 20 on the pressure surface 16 side in the range of 0.0≦X / Xmax≦0.50. That is, in an XY coordinate system (orthogonal coordinate system) with the position of the leading edge 12 as the origin in a cross section perpendicular to the blade height direction of the airfoil portion 6, if the coordinate axis parallel to the axial direction of the gas turbine is taken as the X-axis, the downstream side of the gas flow in the axial direction (X-axis direction) is taken as the positive direction of the X-axis, and the X-coordinate of the trailing edge 14 is taken as Xmax, the average thickness of the oxidation-resistant coating layer 20 on the pressure surface 16 side in the range where the X-coordinate satisfies 0.90≦X / Xmax≦1.0 is smaller than the average thickness of the oxidation-resistant coating layer 20 on the pressure surface 16 side in the range where the X-coordinate satisfies 0.0≦X / Xmax≦0.50. Here, X / Xmax is the value obtained by dividing X by Xmax.

[0023] In general, for the airfoil of a gas turbine vane, the bending strength against the gas flow (gas bending strength) is important as a physical strength from the viewpoint of damage prevention. However, when the airfoil is exposed to the flow of high-temperature gas and thinning of the airfoil due to oxidation progresses, the bending strength of the airfoil decreases.

[0024] In contrast, according to the gas turbine vane 4, the average thickness of the oxidation-resistant coating layer 20 on the pressure surface 16 side in the range of 0.90≦X / Xmax≦1.0, where the effect on the bending strength against the gas flow is small, is smaller than the average thickness of the oxidation-resistant coating layer 20 on the pressure surface 16 side in the range of 0.0≦X / Xmax≦0.50, where the effect on the bending strength against the gas flow is large, so that damage to the gas turbine vane 4 caused by oxidation (e.g. bending or chipping toward the suction surface 18 side) can be efficiently suppressed with a small amount of oxidation-resistant coating, and the cost required for the oxidation-resistant coating can be reduced. Also, the improvement in durability of the gas turbine vane 4 makes it possible to extend the maintenance cycle.

[0025] In the embodiment shown in FIG. 2 , the thickness of the oxidation-resistant coating layer 20 formed on the suction surface 18 decreases from a position on the leading edge 12 side of the suction surface 18 toward a position P4 where the blade thickness t of the gas turbine stator vane 4 is maximum, and decreases from a position on the trailing edge 14 side of the suction surface 18 toward the position P4 where the blade thickness t of the gas turbine stator vane 4 is maximum.

[0026] According to this configuration, the thickness of the oxidation-resistant coating layer 20 on the suction surface 18 decreases from the position on the leading edge 12 side on the suction surface 18 toward the position P4 where the blade thickness t of the gas turbine stator vane 4 is maximum, so that damage to the gas turbine stator vane 4 caused by oxidation can be efficiently suppressed with a small amount of oxidation-resistant coating. In addition, since the blade thickness on the trailing edge 14 side of the gas turbine stator vane 4 is generally thinner than that on the leading edge 12 side, there is a concern that the blade shape and the chord length of the blade will change significantly when the thinning caused by oxidation progresses, and the aerodynamic performance will be significantly reduced. However, according to the above-mentioned gas turbine stator vane 4, the thickness of the oxidation-resistant coating layer 20 decreases from the position on the trailing edge 14 side on the suction surface 18 toward the position P4 where the blade thickness of the gas turbine stator vane 4 is maximum, so that the thickness of the oxidation-resistant coating layer 20 is relatively larger on the trailing edge 14 side than the position P4 on the suction surface 18, so that the reduction in the aerodynamic performance of the gas turbine stator vane 4 caused by oxidation can be efficiently suppressed with a small amount of oxidation-resistant coating.

[0027] In the embodiment shown in FIG. 2 , in a cross section perpendicular to the blade height direction of the airfoil portion 6, if the axial distance from position P7 on the suction surface 18 at which the thickness of the oxidation-resistant coating layer 20 on the suction surface 18 side is smallest within the range from the leading edge 12 to the trailing edge 14 is defined as X1, the thickness of the oxidation-resistant coating layer 20 on the suction surface 18 side changes between the leading edge 12 and the trailing edge 14 so as to satisfy 0.25≦X1 / Xmax≦0.70.

[0028] According to this configuration, the position P7 at which the thickness of the oxidation-resistant coating layer 20 on the suction surface 18 side is minimum is included within the range of 0.25≦X / Xmax≦0.70, where the impact on damage to the gas turbine stator vane 4 and degradation of aerodynamic performance is small, so that damage to the gas turbine stator vane 4 and degradation of aerodynamic performance can be efficiently suppressed with a small amount of the oxidation-resistant coating.

[0029] In the embodiment shown in FIG. 2 , in a cross section perpendicular to the blade height direction in the airfoil portion 6, the average value of the thickness of the oxidation-resistant coating layer 20 on the suction surface 18 side in the range of 0.25≦X / Xmax≦0.70 is smaller than the average value of the thickness of the oxidation-resistant coating layer 20 on the suction surface 18 side in the range of 0.0≦X / Xmax≦0.25 and is smaller than the average value of the thickness of the oxidation-resistant coating layer 20 on the suction surface 18 side in the range of 0.70≦X / Xmax≦1.0.

[0030] According to this configuration, by making the thickness of the oxidation-resistant coating layer 20 on the suction surface 18 side relatively smaller within the range of 0.25≦X / Xmax≦0.70, where the impact on damage to the gas turbine stator vane 4 and degradation of aerodynamic performance is small, it is possible to efficiently suppress damage to the gas turbine stator vane 4 and degradation of aerodynamic performance with a small amount of oxidation-resistant coating.

[0031] In addition, the above-mentioned gas turbine vane 4 may be formed on the surface 7 of the airfoil 6 of the gas turbine vane 4 by spraying a material of the oxidation-resistant coating layer 20 (e.g., CoNiCrAlY, NiCrAlY, CoCrAlY or MCrAlX (wherein M is Ni, Co or an alloy of Ni and Co, and X is Hf, Si or Ta) using a thermal spray gun 30, as shown in FIG. 3, to form the above-mentioned oxidation-resistant coating layer 20 (20s1, 20s2, 20s3, 20h1, 20h2) on the surface 7 of the airfoil 6 of the gas turbine vane 4. In this case, for example, as shown in FIG. The oxidation-resistant coating layer leading edge side portion 20h1 on the pressure surface 16 side may be formed by spraying the material of the oxidation-resistant coating layer 20 from the axial upstream side of the airfoil portion 6 of the gas turbine vane 4 by a thermal spray gun 30 onto the range from the leading edge 12 to position P3 on the pressure surface 16 of the airfoil portion 6. Also, the oxidation-resistant coating layer trailing edge side portion 20h2 on the pressure surface 16 side may be formed by spraying (blowing) the material of the oxidation-resistant coating layer 20 from the axial upstream side of the airfoil portion 6 of the gas turbine vane 4 onto the range from position P3 to position P1 on the pressure surface 16 of the airfoil portion 6.

[0032] Alternatively, the thermal spray gun 30 may spray the thermal oxidation-resistant coating material from a position upstream of the airfoil 6 of the gas turbine vane 4 in the axial direction of the gas turbine to a range from the leading edge 12 to a position P5 on the suction surface 18 of the airfoil 6 to form the leading edge side portion 20s1 of the oxidation-resistant coating layer on the suction surface 18 side. Alternatively, the thermal spray gun 30 may spray the thermal oxidation-resistant coating material from a position upstream of the airfoil 6 of the gas turbine vane 4 in the axial direction of the gas turbine to a range from the position P5 to a position between the positions P5 and P6 on the suction surface 18 of the airfoil 6, and the thermal spray gun 30 may spray the thermal oxidation-resistant coating material from a position downstream of the airfoil 6 of the gas turbine vane 4 in the axial direction of the gas turbine to a range from the position P6 to a position between the positions P6 and P5 on the suction surface 18 of the airfoil 6 to form the oxidation-resistant coating layer intermediate portion 20s2 on the suction surface 18 side. In addition, the material of the oxidation-resistant coating layer 20 may be sprayed from the axially downstream side of the airfoil portion 6 of the gas turbine vane 4 using a thermal spray gun 30 onto the range from the trailing edge 14 on the negative pressure surface 18 of the airfoil portion 6 to position P6, thereby forming the trailing edge side portion 20s3 of the oxidation-resistant coating layer on the negative pressure surface 18 side.

[0033] 1 and 2, the splice vane unit 2 and the gas turbine vane 4 can be manufactured while avoiding interference between the thermal spray gun and the gas turbine vane 4. Furthermore, by applying an oxidation-resistant coating to the splice vane unit 2 in which the multiple gas turbine vanes 4 are integrally molded, it is not necessary to assemble the multiple gas turbine vanes, and therefore the labor and cost required for assembly can be reduced, compared to the case where an oxidation-resistant coating is applied to multiple gas turbine vanes molded separately and then the gas turbine vanes are assembled.

[0034] The present disclosure is not limited to the above-described embodiments, and includes modifications to the above-described embodiments and appropriate combinations of these modifications.

[0035] For example, in the embodiment shown in Fig. 2, the oxidation-resistant coating layer 20 is not formed in the range from the position P1 to the trailing edge 14 on the pressure surface 16 of the airfoil 6, but the oxidation-resistant coating layer 20 may also be formed in the range from the position P1 to the trailing edge 14 on the pressure surface 16 of the airfoil 6, or the oxidation-resistant coating layer 20 may be formed on the entire pressure surface 16 of the airfoil 6. In the gas turbine vane 4 of the present disclosure, the oxidation-resistant coating layer 20 is formed in at least a partial range on the pressure surface 16 side of the surface 7 of the airfoil 6, and the average thickness of the oxidation-resistant coating layer 20 on the pressure surface 16 side in the range of 0.90 <= X / Xmax <= 1.0 is smaller than the average thickness of the oxidation-resistant coating layer 20 on the pressure surface 16 side in the range of 0.0 <= X / Xmax <= 0.50, so that damage to the gas turbine vane 4 caused by oxidation can be efficiently suppressed with a small amount of oxidation-resistant coating.

[0036] The contents described in each of the above embodiments can be understood, for example, as follows.

[0037] (1) A gas turbine vane according to at least one embodiment of the present disclosure (e.g., the above-described gas turbine vane 4) has A gas turbine vane having an airfoil (e.g., the airfoil 6 described above), the surfaces of the airfoil include a leading edge (e.g., leading edge 12 described above), a trailing edge (e.g., trailing edge 14 described above), a pressure side connecting the leading edge and the trailing edge (e.g., pressure side 16 described above), and a suction side connecting the leading edge and the trailing edge (e.g., suction side 18 described above); an oxidation-resistant coating layer (e.g., the oxidation-resistant coating layer 20 of the pressure surface 16 described above) is formed on at least a portion of the pressure surface side of the surface of the airfoil portion; In a cross section (e.g., the cross section shown in FIG. 2) perpendicular to the blade height direction of the airfoil portion, a position in the axial direction of the gas turbine is defined as a distance X from the leading edge to the position in the axial direction, and a position of the trailing edge in the axial direction is defined as Xmax. The average thickness of the oxidation-resistant coating layer on the pressure surface side in the range of 0.90≦X / Xmax≦1.0 is smaller than the average thickness of the oxidation-resistant coating layer on the pressure surface side in the range of 0.0≦X / Xmax≦0.50.

[0038] For turbine vanes, the bending strength against gas flow (gas bending strength) is important as a physical strength from the viewpoint of damage prevention. However, when the turbine vane is exposed to the flow of high-temperature gas and thinning of the turbine vane due to oxidation progresses, the bending strength of the turbine vane decreases.

[0039] In contrast, according to the gas turbine vane described in (1) above, the average thickness of the oxidation-resistant coating layer on the pressure surface side in the range of 0.90≦X / Xmax≦1.0, where the effect on the bending strength against the gas flow is small, is smaller than the average thickness of the oxidation-resistant coating layer on the pressure surface side in the range of 0.0≦X / Xmax≦0.50, where the effect on the bending strength against the gas flow is large, so that damage to the gas turbine vane caused by oxidation (e.g. bending or chipping toward the suction surface side) can be efficiently suppressed with a small amount of oxidation-resistant coating, and the cost required for the oxidation-resistant coating can be reduced. Also, the maintenance cycle of the gas turbine vane can be extended by improving the durability of the gas turbine vane.

[0040] (2) In some embodiments, in the gas turbine stator vane described in (1) above, an oxidation-resistant coating layer (e.g., the oxidation-resistant coating layer 20 of the suction surface 18 described above) is formed on at least a portion of the surface of the airfoil portion on the suction surface side; The thickness of the oxidation-resistant coating layer formed on the suction surface decreases from the leading edge position on the suction surface toward the position where the blade thickness of the gas turbine stator vane is maximum, and decreases from the trailing edge position on the suction surface toward the position where the blade thickness of the gas turbine stator vane is maximum.

[0041] According to the gas turbine vane described in (2) above, the thickness of the oxidation-resistant coating layer on the suction surface decreases from the leading edge to the position where the blade thickness of the gas turbine vane is maximum, so that damage to the gas turbine vane caused by oxidation can be efficiently suppressed with a small amount of oxidation-resistant coating. In addition, since the trailing edge side of a gas turbine vane generally has a thinner blade thickness than the leading edge side, there is a concern that the blade profile and the chord length of the blade will change significantly when the thinning caused by oxidation progresses, and the aerodynamic performance will be significantly reduced. However, according to the gas turbine vane described in (2) above, the thickness of the oxidation-resistant coating layer decreases from the trailing edge side position on the suction surface to the position where the blade thickness of the gas turbine vane is maximum, so that the thickness of the oxidation-resistant coating layer is relatively large on the trailing edge side compared to the position where the blade thickness is maximum on the suction surface, so that the deterioration of the aerodynamic performance of the gas turbine vane caused by oxidation can be efficiently suppressed with a small amount of oxidation-resistant coating.

[0042] (3) In some embodiments, in the gas turbine vane described in (1) or (2), an oxidation-resistant coating layer (e.g., the oxidation-resistant coating layer 20 of the suction surface 18 described above) is formed on at least a portion of the surface of the airfoil portion on the suction surface side; When a distance in the axial direction of the gas turbine from a position on the suction surface where the thickness of the oxidation-resistant coating layer on the suction surface side is minimum to the leading edge is defined as X1 in the cross section perpendicular to the blade height direction of the airfoil portion, the thickness of the oxidation-resistant coating layer on the suction surface side varies between the leading edge and the trailing edge so as to satisfy 0.25≦X1 / Xmax≦0.70.

[0043] According to the gas turbine vane described in (3) above, the thickness of the oxidation-resistant coating layer 20 on the suction surface 18 side is minimized in the range of 0.25≦X / Xmax≦0.70, where the influence on damage to the gas turbine vane 4 and deterioration of its aerodynamic performance is small. Therefore, like the gas turbine vane described in (2) above, damage to the gas turbine vane and deterioration of its aerodynamic performance caused by oxidation can be efficiently suppressed with a small amount of oxidation-resistant coating.

[0044] (4) In some embodiments, in the gas turbine stator vane according to any one of (1) to (3), an oxidation-resistant coating layer (e.g., the above-mentioned oxidation-resistant coating layer 20) is formed on at least a portion of the surface of the airfoil portion on the suction side; In the cross section of the airfoil portion perpendicular to the blade height direction, an average value of a thickness of the oxidation-resistant coating layer on the suction surface side in a range of 0.25≦X / Xmax≦0.70 is smaller than an average value of a thickness of the oxidation-resistant coating layer on the suction surface side in a range of 0.0≦X / Xmax≦0.25 and is smaller than an average value of a thickness of the oxidation-resistant coating layer on the suction surface side in a range of 0.70≦X / Xmax≦1.0.

[0045] According to the gas turbine vane described in (4) above, by making the thickness of the oxidation-resistant coating layer 20 on the suction surface 18 side relatively smaller than that in other ranges within the range of 0.25≦X / Xmax≦0.70, where the influence on damage to the gas turbine vane 4 and degradation of its aerodynamic performance is small, damage to the gas turbine vane and degradation of its aerodynamic performance caused by oxidation can be efficiently suppressed with a small amount of oxidation-resistant coating, in the same way as in the gas turbine vane described in (2) above.

[0046] (5) In some embodiments, in the gas turbine stator vane according to any one of (1) to (4), The oxidation-resistant coating is composed of CoNiCrAlY, NiCrAlY, CoCrAlY, or MCrAlX (wherein M is Ni, Co, or an alloy of Ni and Co, and X is Hf, Si, or Ta).

[0047] According to the gas turbine vane described in (5) above, in the case where the oxidation-resistant coating is made of CoNiCrAlY, NiCrAlY, CoCrAlY, or MCrAlX (wherein M is Ni, Co, or an alloy of Ni and Co, and X is Hf, Si, or Ta), damage to the gas turbine vane caused by oxidation can be efficiently suppressed with a small amount of the oxidation-resistant coating.

[0048] (6) At least one embodiment of the binding unit according to the present disclosure (for example, the binding unit 2 described above) A splice blade unit comprising a plurality of gas turbine vanes (e.g., the plurality of gas turbine vanes 4 described above), the plurality of gas turbine vanes being integrally molded, each of the plurality of gas turbine vanes being the gas turbine vane described in (1) above.

[0049] In a splice vane unit in which a plurality of gas turbine vanes are integrally molded, when the gas turbine vanes are miniaturized, for example, when an oxidation-resistant coating layer is formed on the vane surface using a thermal spray gun, interference between the thermal spray gun and the gas turbine vane is likely to occur, but in the splice vane unit described in (5) above, the average value of the thickness of the oxidation-resistant coating layer on the pressure surface side in the range of 0.90≦X / Xmax≦1.0 is smaller than the average value of the thickness of the oxidation-resistant coating layer on the pressure surface side in the range of 0.0≦X / Xmax≦0.50, so that the splice vane unit can be manufactured while avoiding interference between the thermal spray gun and the gas turbine vane. In addition, compared to the case where an oxidation-resistant coating is applied to a plurality of gas turbine vanes molded separately and then the gas turbine vanes are assembled, by applying an oxidation-resistant coating to a splice vane unit in which a plurality of gas turbine vanes 4 are integrally molded, it is not necessary to assemble a plurality of gas turbine vanes, so that the labor and cost required for assembly can be reduced.

[0050] (7) A method for manufacturing a gas turbine vane according to at least one embodiment of the present disclosure includes: forming an oxidation-resistant coating layer (e.g., the above-mentioned oxidation-resistant coating layer 20) on at least a partial area of ​​a pressure surface (e.g., the above-mentioned pressure surface 16) side of a surface of an airfoil (e.g., the above-mentioned airfoil 6) of a gas turbine stator vane (e.g., the above-mentioned gas turbine stator vane 4); In a cross section (e.g., the cross section shown in FIG. 2) perpendicular to the blade height direction of the airfoil, a position in the axial direction of the gas turbine is defined as a distance X from a leading edge of the airfoil in the axial direction to the position, and a position of a trailing edge of the airfoil in the axial direction is defined as Xmax. In the step, the oxidation-resistant coating layer is formed on at least a portion of the pressure surface side of the surface of the airfoil section such that an average value of a thickness of the oxidation-resistant coating layer on the pressure surface side in a range of 0.90≦X / Xmax≦1.0 is smaller than an average value of a thickness of the oxidation-resistant coating layer on the pressure surface in a range of 0.0≦X / Xmax≦0.50.

[0051] For turbine vanes, the bending strength against gas flow (gas bending strength) is important as a physical strength from the viewpoint of damage prevention. However, when the turbine vane is exposed to the flow of high-temperature gas and thinning of the turbine vane due to oxidation progresses, the bending strength of the turbine vane decreases.

[0052] In contrast, according to the manufacturing method of a gas turbine vane described in (7) above, the oxidation-resistant coating layer is formed on at least a portion of the pressure surface side of the surface of the airfoil so that the average thickness of the oxidation-resistant coating layer on the pressure surface side in the range of 0.90≦X / Xmax≦1.0, where the effect on the bending strength to the gas flow is small, is smaller than the average thickness of the oxidation-resistant coating layer on the pressure surface side in the range of 0.0≦X / Xmax≦0.50, where the effect on the bending strength to the gas flow is large. Therefore, it is possible to realize a manufacturing method of a gas turbine vane that can efficiently suppress damage to the gas turbine vane due to oxidation with a small amount of oxidation-resistant coating. [Explanation of symbols]

[0053] 2 Spelling Wing Unit 4 Gas turbine vanes 6 Airfoil 7 surface 8 Outer Shroud 9 Cavity 10 Inner shroud 12 Leading Edge 14 Trailing edge 16 Pressure Surface 18 Suction surface 20 Oxidation-resistant coating layer 20h1, 20s1 Oxidation-resistant coating layer leading edge side 20h2, 20s3 Oxidation-resistant coating layer trailing edge side 20s2 Oxidation-resistant coating layer middle part 30 Thermal spray gun

Claims

1. 1. A gas turbine vane comprising an airfoil, a surface of the airfoil including a leading edge, a trailing edge, a pressure side connecting the leading edge and the trailing edge, and a suction side connecting the leading edge and the trailing edge; an oxidation-resistant coating layer is formed on at least a portion of the pressure surface side of the surface of the airfoil; In a cross section of the airfoil portion perpendicular to the blade height direction, a position in the axial direction of the gas turbine is defined as a distance X from the leading edge to the position in the axial direction, and a position of the trailing edge in the axial direction is defined as Xmax. a mean value of a thickness of the oxidation-resistant coating layer on the pressure surface side in a range of 0.90≦X / Xmax≦1.0 is smaller than a mean value of a thickness of the oxidation-resistant coating layer on the pressure surface side in a range of 0.0≦X / Xmax≦0.

50.

2. an oxidation-resistant coating layer is formed on at least a portion of the surface of the airfoil on the suction side; 2. The gas turbine vane according to claim 1, wherein a thickness of the oxidation-resistant coating layer formed on the suction surface side decreases from a position on the leading edge side of the suction surface toward a position where a blade thickness of the gas turbine vane is maximum, and decreases from a position on the trailing edge side of the suction surface toward a position where a blade thickness of the gas turbine vane is maximum.

3. an oxidation-resistant coating layer is formed on at least a portion of the surface of the airfoil on the suction side; 2. The gas turbine vane according to claim 1, wherein, in the cross section of the airfoil portion perpendicular to the blade height direction, a distance in an axial direction of the gas turbine from a position on the suction surface where a thickness of the oxidation-resistant coating layer on the suction surface side is minimum to the leading edge is defined as X1, and a thickness of the oxidation-resistant coating layer on the suction surface side varies between the leading edge and the trailing edge so as to satisfy 0.25≦X1 / Xmax≦0.

70.

4. an oxidation-resistant coating layer is formed on at least a portion of the surface of the airfoil on the suction side; 2. The gas turbine vane according to claim 1, wherein, in the cross section orthogonal to the blade height direction of the airfoil portion, an average value of a thickness of the oxidation-resistant coating layer on the suction surface side in a range of 0.25≦X / Xmax≦0.70 is smaller than an average value of a thickness of the oxidation-resistant coating layer on the suction surface side in a range of 0.0≦X / Xmax≦0.25 and is smaller than an average value of a thickness of the oxidation-resistant coating layer on the suction surface side in a range of 0.70≦X / Xmax≦1.

0.

5. 2. The gas turbine vane according to claim 1, wherein the oxidation-resistant coating layer is made of CoNiCrAlY, NiCrAlY, CoCrAlY, or MCrAlX (wherein M is Ni, Co, or an alloy of Ni and Co, and X is Hf, Si, or Ta).

6. 2. A splice vane unit comprising a plurality of gas turbine vanes, the plurality of gas turbine vanes being integrally formed, each of the plurality of gas turbine vanes being the gas turbine vane according to claim 1.

7. forming an oxidation-resistant coating layer on at least a portion of a pressure surface of a surface of an airfoil of a gas turbine vane; In a cross section of the airfoil portion perpendicular to the blade height direction, a position in the axial direction of the gas turbine is defined as a distance X from a leading edge of the airfoil portion to the position in the axial direction, and a position of a trailing edge of the airfoil portion in the axial direction is defined as Xmax. the step of forming the oxidation-resistant coating layer on at least a portion of the pressure surface side of the surface of the airfoil portion such that an average value of a thickness of the oxidation-resistant coating layer on the pressure surface side in a range of 0.90≦X / Xmax≦1.0 is smaller than an average value of a thickness of the oxidation-resistant coating layer on the pressure surface in a range of 0.0≦X / Xmax≦0.50.

Images