Turbine moving blade and gas turbine

Abstract

A turbine moving blade according to at least one embodiment of the present disclosure includes: a second cooling flow path; and a first cooling flow path that extends in the blade height direction, is formed on the leading edge side relative to the second cooling flow path, and is connected to the second cooling flow path at a first folded section on the base end side of an airfoil section. The first folded section includes an upstream-side region extending along the extension direction of the first folded section, from the inlet section at a position in the same blade height direction as the base end of the airfoil section to the central position in the wall thickness direction of the first partition wall when viewed from the blade height direction. The upstream-side region has, in at least a part of the blade height direction, a region in which an inner wall surface of a back-side blade wall and a partition wall surface on a leading edge side of a second partition wall is connected by a curved surface that protrudes outward from the inside of the airfoil section when viewed from the blade height direction, the region having a radius of curvature of the curved surface that increases from the tip end side toward the base end side of the airfoil section.

Description

Turbine blades and gas turbines

[0001] This application claims priority to Japanese Patent Application No. 2023-092056, filed on June 5, 2023, with the Japan Patent Office, the contents of which are incorporated herein by reference.

[0002] For example, turbine blades used in gas turbines and the like are used in high-temperature combustion gases, so they are equipped with cooling channels inside for cooling, and the temperature rise of the blade metal is suppressed by circulating cooling air through the cooling channels (see Patent Document 1).

[0003] Japanese Patent Application Laid-Open No. 2012-177377

[0004] In the turbine blade described in Patent Document 1, the cooling flow path is formed as a serpentine flow path in which multiple internal passages extending in the blade height direction are connected. In a turbine blade, the length along the blade surface from the leading edge to the trailing edge differs significantly between the suction-side wall surface and the pressure-side wall surface at the base end of the airfoil. Therefore, in the internal passages constituting the serpentine flow path, two adjacent wall surfaces along the inner periphery of the internal passage may be connected at a relatively small angle when viewed from the blade height direction. In this case, if the base end of the airfoil and the bottom of the turn-up portion of the internal passage are relatively close to each other in the blade height direction, the two adjacent wall surfaces along the inner periphery of the internal passage remain connected at a relatively small angle and are further bent along the extension direction of the turn-up portion. Therefore, the flow rate of cooling air decreases near the connection where two adjacent wall surfaces along the inner periphery of the turn-up portion are connected at a relatively small angle, which may result in a decrease in cooling performance and insufficient suppression of temperature rise in the blade metal.

[0005] In view of the above circumstances, at least one embodiment of the present disclosure has an object to suppress a temperature rise in a blade metal of a turbine blade.

[0006] (1) A turbine blade according to at least one embodiment of the present disclosure comprises: a blade body including an airfoil portion having a plurality of cooling channels therein, wherein the plurality of cooling channels include, in order from a trailing edge side of the airfoil portion: a third cooling channel extending in the blade height direction and opening at a root of the blade; a second cooling channel extending in the blade height direction and formed on a leading edge side of the airfoil portion relative to the third cooling channel and connected to the third cooling channel at a second turning portion on a tip side of the airfoil portion; and a first cooling channel extending in the blade height direction and formed on a leading edge side of the second cooling channel and connected to the second cooling channel at a first turning portion on a base end side of the airfoil portion; a first partition wall separating the first cooling channel and the second cooling channel; and a second partition wall separating the second cooling channel and the third cooling channel. The first folded portion includes an upstream region extending along the extension direction of the first folded portion from an inlet portion at the same position in the blade height direction as the base end of the airfoil portion to a central position in the wall thickness direction of the first partition wall when viewed from the blade height direction, and a part of the upstream region when viewed from the blade height direction is defined at the base end of the airfoil portion by an inner wall surface of a suction side blade wall of the airfoil portion and a partition wall surface on the leading edge side of the second partition wall, and the upstream region has a region in at least a part of the blade height direction where the inner wall surface and the partition wall surface are connected by a curved surface that is convex from the inside to the outside of the airfoil portion when viewed from the blade height direction, and the radius of curvature of the curved surface increases from the tip side to the base end side.

[0007] (2) A turbine blade according to at least one embodiment of the present disclosure comprises: a blade body including an airfoil portion having a plurality of cooling channels therein, wherein the plurality of cooling channels include, in order from a trailing edge side of the airfoil portion: a third cooling channel extending in the blade height direction and opening at a root of the blade; a second cooling channel extending in the blade height direction and formed on a leading edge side of the airfoil portion relative to the third cooling channel and connected to the third cooling channel at a second turning portion on a tip side of the airfoil portion; and a first cooling channel extending in the blade height direction and formed on a leading edge side of the second cooling channel and connected to the second cooling channel at a first turning portion on a base end side of the airfoil portion; a first partition wall separating the first cooling channel and the second cooling channel; and a second partition wall separating the second cooling channel and the third cooling channel. the first folded portion includes an upstream region extending along the extension direction of the first folded portion from an inlet portion at the same position in the blade height direction as the base end of the airfoil portion to a center position in the wall thickness direction of the first partition wall when viewed from the blade height direction, a part of the upstream region when viewed from the blade height direction is defined at the base end of the airfoil portion by an inner wall surface of a suction side blade wall of the airfoil portion and a partition wall surface on the leading edge side of the second partition wall, when viewed from the blade height direction, the partition wall surface includes a first partition wall surface extending in a first direction from the pressure side to the suction side of the airfoil portion, and a second partition wall surface extending in a second direction intersecting the first direction on the suction side of the first partition wall surface, the second direction being inclined with respect to the first direction so as to be more toward the leading edge side than the first direction as it goes from the pressure side to the suction side, The upstream region has a region in which the angle difference between the extending directions of the first direction and the second direction increases from the tip side toward the base side.

[0008] (3) A gas turbine according to at least one embodiment of the present disclosure includes a turbine rotor blade having the configuration described in (1) or (2) above.

[0009] According to at least one embodiment of the present disclosure, it is possible to suppress a temperature rise in the blade metal of the turbine blade.

[0010] 3A is a schematic configuration diagram of a gas turbine according to some embodiments. It is a schematic diagram of a turbine blade according to some embodiments, as seen from the circumferential direction. It is a schematic diagram showing a cross section along the camber line of an airfoil portion of a turbine blade according to some embodiments. It is an enlarged view of a main portion of FIG. 3A. It is a schematic cross section of an airfoil portion near the base end of a turbine blade according to some embodiments, showing a cross section as viewed along arrows IV-IV in FIG. 3A. It is a schematic cross section of a turbine blade according to one embodiment, taken along arrows A-A in FIG. 3A. It is a schematic cross section of a turbine blade according to one embodiment, taken along arrows B-B in FIG. 3A. It is a schematic cross section of a turbine blade according to one embodiment, taken along arrows C-C in FIG. 3A. It is a schematic cross section of a turbine blade according to another embodiment, taken along arrows A-A in FIG. 3A. It is a schematic cross section of a turbine blade according to another embodiment, taken along arrows B-B in FIG. 3A. It is a schematic cross section of a turbine blade according to another embodiment, taken along arrows C-C in FIG. 3A. FIG. 10 is a diagram schematically illustrating a cross section of a turbine rotor blade according to yet another embodiment, taken along a camber line of an airfoil portion.

[0011] Several embodiments of the present disclosure will be described below with reference to the accompanying drawings. However, the dimensions, materials, shapes, relative arrangements, etc. of components described as embodiments or shown in the drawings are not intended to limit the scope of the present disclosure and are merely illustrative examples. For example, expressions expressing relative or absolute arrangements, such as "in a certain direction," "along a certain direction," "parallel," "orthogonal," "center," "concentric," or "coaxial," not only strictly express such arrangements, but also express relative displacements with a tolerance or angle or distance to the extent that the same function is achieved. For example, expressions expressing the equality of things, such as "same," "equal," and "homogeneous," not only express strict equality, but also express tolerance or differences to the extent that the same function is achieved. For example, expressions expressing shapes such as a square or cylindrical shape not only express shapes such as a square or cylindrical shape in the strict geometric sense, but also express shapes including concave and convex portions, chamfered portions, etc., to the extent that the same effect is achieved. On the other hand, the expressions "comprise," "include," "have," "includes," or "have" of one element are not exclusive expressions that exclude the presence of other elements.

[0012] (Overall Configuration of Gas Turbine 1) First, the configuration of a gas turbine to which turbine rotor blades according to some embodiments are applied will be described with reference to Fig. 1. Fig. 1 is a schematic configuration diagram of a gas turbine 1 according to some embodiments.

[0013] 1 , a gas turbine 1 according to some embodiments includes a compressor 2 for generating compressed air, a combustor 4 for generating combustion gas using the compressed air and fuel, and a turbine 6 configured to be rotationally driven by the combustion gas. In the case of a gas turbine 1 for power generation, a generator (not shown) is connected to the turbine 6, and the gas turbine 1 is configured to generate electricity using the rotational energy of the turbine 6.

[0014] In the gas turbine 1 shown in FIG. 1 , the compressor 2 includes a rotor 30 that is rotatable about a central axis AX, and a stator 5 that is disposed around the rotor 30 .

[0015] The stator 5 has a compressor casing 10 and a plurality of compressor stator vanes 16 fixed to the compressor casing 10. The rotor 30 has a rotor shaft 8 that is rotatable about a central axis AX, a plurality of rotor disks 31 fixed to the rotor shaft 8, and a plurality of compressor rotor blades 18 attached to each of the plurality of rotor disks 31. The rotor shaft 8 is arranged to penetrate both the compressor casing 10 and a turbine casing 22 described below.

[0016] A plurality of compressor rotor blades 18 are arranged in the circumferential direction of the central axis AX on the outer peripheries of the plurality of rotor disks 31. The rotor disks 31 are also arranged in a plurality of stages at intervals in a direction parallel to the central axis AX. Therefore, the compressor rotor blades 18 are arranged in a plurality of stages at intervals in a direction parallel to the central axis AX.

[0017] A plurality of compressor stator vanes 16 are arranged in the circumferential direction of the central axis AX. The compressor stator vanes 16 are also arranged in a plurality of stages spaced apart in a direction parallel to the central axis AX. The compressor stator vanes 16 are arranged in a plurality of stages so as to be positioned between the compressor rotor blades 18 in the direction parallel to the central axis AX.

[0018] In the gas turbine 1 shown in Fig. 1 , the compressor 2 is provided on the inlet side of the compressor casing 10, and includes an air intake 12 for taking in air, and an inlet guide vane 14 provided on the air intake 12 side. The compressor 2 may also include other components such as an air bleed chamber (not shown). In this compressor 2, the air taken in through the air intake 12 passes through a plurality of compressor stator vanes 16 and a plurality of compressor rotor blades 18 and is compressed to generate compressed air. The compressed air is then sent from the compressor 2 to the combustor 4 located downstream.

[0019] In the gas turbine 1 shown in Fig. 1 , the combustor 4 is disposed in a casing (combustor casing) 20. As shown in Fig. 1 , a plurality of combustors 4 may be disposed in an annular shape around the rotor shaft 8 inside the casing 20. Fuel and compressed air generated by the compressor 2 are supplied to the combustor 4, and the combustor 4 burns the fuel to generate high-temperature, high-pressure combustion gas, which serves as a working fluid for the turbine 6. The combustion gas is then sent from the combustor 4 to the turbine 6 at a downstream stage.

[0020] 1 , the turbine 6 includes a rotor 33 that is rotatable about a central axis AX, and a stator 7 that is disposed around the rotor 33. The stator 7 includes a turbine casing 22 and a plurality of turbine stator blades 26 that are fixed to the turbine casing 22 side.

[0021] The rotor 33 has the rotor shaft 8 described above, a plurality of rotor disks 35 fixed to the rotor shaft 8, and a plurality of turbine blades 24 attached to each of the plurality of rotor disks 35.

[0022] A plurality of turbine rotor blades 24 are arranged in the circumferential direction of the central axis AX on the outer periphery of each of the plurality of rotor disks 35. The rotor disks 35 are also arranged in a plurality of stages at intervals in a direction parallel to the central axis AX. Therefore, the turbine rotor blades 24 are arranged in a plurality of stages at intervals in a direction parallel to the central axis AX.

[0023] A plurality of turbine stator blades 26 are arranged in the circumferential direction of the central axis AX. The turbine stator blades 26 are also arranged in a plurality of stages at intervals in a direction parallel to the central axis AX. The turbine stator blades 26 are arranged in a plurality of stages so as to be positioned between the turbine rotor blades 24 in the direction parallel to the central axis AX.

[0024] In the turbine 6, the rotor shaft 8 extends in the axial direction (left-right direction in FIG. 1 ), and the combustion gas flows from the combustor 4 side toward the exhaust casing 28 side (from left to right in FIG. 1 ). Therefore, in FIG. 1 , the left side is the axially upstream side, and the right side is the axially downstream side. In the following description, the term "axial direction" simply refers to a direction parallel to the central axis AX, and the term "radial direction" simply refers to a radial direction centered on the central axis AX. In the following description, the term "circumferential direction of the rotor" or simply the term "circumferential direction" refers to a circumferential direction centered on the central axis AX.

[0025] The turbine rotor blades 24, together with the turbine stator blades 26, are configured to generate a rotational driving force from the high-temperature, high-pressure combustion gas flowing inside the turbine casing 22. This rotational driving force is transmitted to the rotor shaft 8, thereby driving a generator (not shown) connected to the rotor shaft 8.

[0026] An exhaust chamber 29 is connected to the axial downstream side of the turbine casing 22 via an exhaust casing 28. After driving the turbine 6, the combustion gas passes through the exhaust casing 28 and the exhaust chamber 29 and is discharged to the outside.

[0027] (Configuration of turbine rotor blade 24) Next, the turbine rotor blade 24 according to some embodiments will be described. Fig. 2 is a schematic diagram of the turbine rotor blade 24 according to some embodiments as viewed from the circumferential direction. Fig. 3A is a schematic diagram showing a cross section along the camber line of the airfoil portion 44 of the turbine rotor blade 24 according to some embodiments. Fig. 3B is an enlarged view of a main part of Fig. 3A. Fig. 4 is a schematic cross section of the airfoil portion 44 near the base end 44b of the turbine rotor blade 24 according to some embodiments, and is a schematic cross section as seen from the arrows IV-IV in Fig. 3A.

[0028] As shown in Figures 2 to 4, the turbine blade 24 according to one embodiment includes a platform 42, an airfoil portion 44 and a blade root portion 47 located on opposite sides of the platform 42 in the blade height direction, and a shank 48 located between the platform 42 and the blade root portion 47. The airfoil portion 44, the platform 42, the blade root portion 47, and the shank 48 may be integrally formed by casting or the like. Note that, when the turbine blade 24 is attached to the rotor disk 35, the blade height direction of the turbine blade 24 coincides with the radial direction. In the following description, the tip 44a side in the blade height direction is the radially outer side when the turbine blade 24 is attached to the rotor disk 35, and the base end 44b side in the blade height direction is the radially inner side when the turbine blade 24 is attached to the rotor disk 35. In the following description, the tip 44a side in the blade height direction will also be simply referred to as the tip 44a side, and the base end 44b side in the blade height direction will also be simply referred to as the base end 44b side.

[0029] The airfoil portion 44 is provided to extend in the blade height direction relative to the rotor disk 35. The airfoil portion 44 has a leading edge 45 and a trailing edge 46 extending along the blade height direction, and also has a pressure side blade surface (pressure surface) 41 and a suction side blade surface (suction surface) 43 extending between the leading edge 45 and the trailing edge 46.

[0030] In the turbine 6, the blade root portion 47 is engaged with a blade groove portion 37 provided in the rotor disk 35. In this way, the turbine rotor blade 24 is implanted in the rotor disk 35 of the turbine 6 and is adapted to rotate together with the rotor disk 35 about the central axis AX.

[0031] 3A , the turbine rotor blade 24 according to some embodiments has a serpentine flow passage (leading-edge-side serpentine flow passage 51) that extends in a serpentine manner from the blade central portion toward the leading edge 45, and a serpentine flow passage (trailing-edge-side serpentine flow passage 52) that extends in a serpentine manner from the blade central portion toward the trailing edge 46. In the turbine rotor blade 24 according to some embodiments, the leading-edge-side serpentine flow passage 51 and the trailing-edge-side serpentine flow passage 52 are flow passages independent of each other. In the turbine rotor blade 24 according to some embodiments, six cooling flow passages 60, for example, that constitute the leading-edge-side serpentine flow passage 51 and the trailing-edge-side serpentine flow passage 52 are provided in order from the leading edge 45 side, and a cooling flow passage 67 provided with a large number of pin fins 68 is provided on the trailing edge side.

[0032] In the turbine rotor blade 24 according to some embodiments, three cooling channels 60 provided in order from the leading edge 45 side are sequentially connected to form a serpentine channel (leading-edge-side serpentine channel 51) that extends in a serpentine manner from the blade central portion toward the leading edge 45. In addition, in the turbine rotor blade 24 according to some embodiments, three cooling channels 60 provided closer to the trailing edge 46 than the leading-edge-side serpentine channel 51 form a serpentine channel (trailing-edge-side serpentine channel 52) that is sequentially connected toward the trailing edge 46.

[0033] The three cooling passages 60 that make up the leading-edge-side serpentine passage 51 are, in order from the trailing edge 46 side, a third cooling passage 63 that extends in the blade height direction; a second cooling passage 62 that extends in the blade height direction, is formed closer to the leading edge 45 than the third cooling passage 63, and is connected to the third cooling passage 63 at a second turning portion 72 on the tip 44a side of the airfoil section 44; and a first cooling passage 61 that extends in the blade height direction, is formed closer to the leading edge 45 than the second cooling passage 62, and is connected to the second cooling passage 62 at a first turning portion 71 on the base end 44b side of the airfoil section.

[0034] The three cooling passages 60 that make up the trailing-edge-side serpentine passage 52 are, in order from the leading edge 45 side, a fourth cooling passage 64 that extends in the blade height direction; a fifth cooling passage 65 that extends in the blade height direction, is formed closer to the trailing edge 46 than the fourth cooling passage 64, and is connected to the fourth cooling passage 64 at a third turning portion 73 on the tip 44a side of the airfoil section 44; and a sixth cooling passage 66 that extends in the blade height direction, is formed closer to the trailing edge 46 than the fifth cooling passage 65, and is connected to the fifth cooling passage 65 at a fourth turning portion 74 on the base end 44b side of the airfoil section 44.

[0035] The first cooling flow path 61 and the second cooling flow path 62 are separated by a first partition wall 91, and the second cooling flow path 62 and the third cooling flow path 63 are separated by a second partition wall 92. The third cooling flow path 63 and the fourth cooling flow path 64 are separated by a third partition wall 93, the fourth cooling flow path 64 and the fifth cooling flow path 65 are separated by a fourth partition wall 94, and the fifth cooling flow path 65 and the sixth cooling flow path 66 are separated by a fifth partition wall 95.

[0036] The third cooling passage 63 constituting the leading-edge-side serpentine passage 51 has an opening 63 a, which is an opening on one end side (inlet side), formed at the base of the blade, i.e., the bottom 47 a of the blade root 47. Similarly, the fourth cooling passage 64 constituting the trailing-edge-side serpentine passage 52 has an opening 64 a, which is an opening on one end side (inlet side), formed at the bottom 47 a of the blade root 47.

[0037] The turbine rotor blade 24 according to some embodiments has a plurality of cooling holes 69 that open near the leading edge 45 as film cooling holes that blow out film cooling air. For example, the plurality of cooling holes 69 are connected to the first cooling flow passage 61.

[0038] In some embodiments of the turbine blade 24, in the leading edge side serpentine flow passage 51, cooling air supplied from the opening 63a, which is the cooling air intake, flows from the third cooling flow passage 63 through the second cooling flow passage 62 toward the first cooling flow passage 61, i.e., toward the leading edge 45.

[0039] A portion of the cooling air that has flowed into the first cooling flow passage 61 is blown out as film cooling air 81 from the multiple cooling holes 69 to film cool the airfoil portion 44 from the outside. In addition, a portion of the cooling air that has flowed into the first cooling flow passage 61 is blown out to the outside of the turbine rotor blade 24 from openings 61 a formed in the tip 44 a of the airfoil portion 44.

[0040] In the turbine rotor blade 24 according to some embodiments, in the trailing-edge-side serpentine flow passage 52, cooling air supplied from the opening 64a, which is a cooling air intake, flows from the fourth cooling passage 64 through the fifth cooling passage 65 and the sixth cooling passage 66 in this order toward the cooling passage 67, i.e., toward the trailing edge 46. This cooling air is blown out as trailing-edge blown air 82 from the cooling passage 67, which is provided with a large number of pin fins 68. In this way, the turbine rotor blade 24 according to some embodiments includes a blade body 49 including an airfoil portion 44 having a plurality of cooling passages 60 therein.

[0041] (Regarding the reduction in cooling air flow velocity at the first turning portion 71) In a turbine blade having a serpentine flow passage such as the turbine blade 24 according to some embodiments, the length along the blade surface from the leading edge 45 to the trailing edge 46 differs significantly between the suction blade surface 43 and the pressure blade surface 41 on the base end 44b side of the airfoil portion 44. Therefore, in the cooling passage 60 constituting the serpentine flow passage, two adjacent wall surfaces along the inner circumference of the cooling passage 60 when viewed from the blade height direction may be connected at a relatively small angle. For example, in the case of the turbine blade 24 according to some embodiments, as shown in FIG. 4 , an inner wall surface 96i of the suction blade wall 96 that defines the second cooling passage 62 and a partition wall surface 92Li on the leading edge 45 side of the second partition wall 92 are connected at a relatively small angle.

[0042] In this way, when two adjacent wall surfaces along the inner circumference of the cooling flow channel 60 are connected at a relatively small angle when viewed from the blade height direction, if the base end 44 b of the airfoil section 44 and the bottom (radially inner surface 71 b of the first folded portion 71 (see FIGS. 3A and 3B )) of the folded portion (for example, the first folded portion 71) of the cooling flow channel 60 are relatively close to each other in the blade height direction, the two adjacent wall surfaces along the inner circumference of the cooling flow channel 60 remain connected at a relatively small angle and are further bent at the folded portion along the extension direction of the folded portion. Therefore, near the connection portion where two adjacent wall surfaces along the inner circumference of the folded portion are connected at a relatively small angle, the flow rate of the cooling air decreases, which reduces cooling performance and may result in insufficient suppression of temperature rise in the blade metal. For example, in the case of the first fold portion 71, there is a risk that the flow rate of the cooling air may decrease near a position downstream of the flow of cooling air along the extension direction of the first fold portion 71 from the connection portion 97 between the inner wall surface 96i of the suction side wing wall 96 shown in Figure 4 and the partition wall surface 92Li on the leading edge 45 side of the second partition wall 92.

[0043] Therefore, in the turbine rotor blade 24 according to some embodiments, the flow velocity of the cooling air at the first turn-up portion 71 is ensured as follows. Fig. 5A is a schematic diagram of the turbine rotor blade 24 according to one embodiment, showing the region surrounded by the two-dot chain line in Fig. 4 in a cross section taken along arrow A in Fig. 3A . Fig. 5B is a schematic diagram of the turbine rotor blade 24 according to one embodiment, showing the region surrounded by the two-dot chain line in Fig. 4 in a cross section taken along arrow B in Fig. 3A . Fig. 5C is a schematic diagram of the turbine rotor blade 24 according to one embodiment, showing the region surrounded by the two-dot chain line in Fig. 4 in a cross section taken along arrow C in Fig. 3A . Fig. 6A is a schematic diagram of the turbine rotor blade 24 according to another embodiment, showing the region surrounded by the two-dot chain line in Fig. 4 in a cross section taken along arrow A in Fig. 3A . Fig. 6B is a schematic diagram of the turbine rotor blade 24 according to another embodiment, showing the region surrounded by the two-dot chain line in Fig. 4 in a cross section taken along arrow B in Fig. 3A . FIG. 6C is a schematic diagram showing a region surrounded by a two-dot chain line in FIG. 4 in a cross section taken along the arrow C in FIG. 3A, of a turbine rotor blade 24 according to another embodiment.

[0044] In the turbine rotor blade 24 according to some embodiments shown in Figures 5A to 6C, the first turned-up portion 71 includes an upstream region 110 extending along the extension direction of the first turned-up portion 71 from an inlet portion 101 (see Figures 3A and 3B) at the same position in the blade height direction as the base end 44b of the airfoil portion 44 to a position C1 (see Figures 3A, 3B, and 5A to 6C) that is the center of the first partition wall 91 in the wall thickness direction when viewed from the blade height direction, and a downstream region 130 extending along the extension direction of the first turned-up portion 71 from the position C1 to an outlet portion 103 at the same position in the blade height direction as the base end 44b of the airfoil portion 44. A portion of the upstream region 110 when viewed from the blade height direction is defined at the base end 44b of the airfoil portion 44 by the inner wall surface 96i of the suction side blade wall 96 of the airfoil portion 44 and the partition wall surface 92Li on the leading edge 45 side of the second partition wall 92. Note that since the upstream region 110 is the region from the inlet portion 101 at the same position in the blade height direction as the base end 44b of the airfoil portion 44 toward the blade root side, strictly speaking, the upstream region 110 is defined by the inner wall surface on the blade root side continuing from the inner wall surface 96i of the suction side blade wall 96 of the airfoil portion 44 and the partition wall surface 92Li on the leading edge 45 side of the second partition wall 92. In the following description, for convenience, the inner wall surface on the blade root side continuing from the inner wall surface 96i of the suction side blade wall 96, i.e., the inner wall surface closer to the blade root than the inner wall surface 96i, will also be referred to as the inner wall surface 96i of the suction side blade wall 96. In the turbine rotor blade 24 according to one embodiment shown in Figures 5A to 5C , in at least a portion of the upstream region 110 in the blade height direction, the inner wall surface 96i and the partition wall surface 92Li are connected by a curved surface 121 that is convex from the inside to the outside of the airfoil section 44 when viewed in the blade height direction. The upstream region 110 has a region 112 (see Figure 3B ) in which the radius of curvature R of the curved surface 121 increases from the tip 44a side toward the base end 44b side. That is, in the turbine rotor blade 24 according to one embodiment shown in Figures 5A to 5C , the region 112 in which the radius of curvature R of the curved surface 121 increases from the tip 44a side toward the base end 44b side exists in at least a portion of the range along the extension direction of the first turned portion 71 from the inlet portion 101 to the position C1.

[0045] In the turbine rotor blade 24 according to one embodiment shown in Figures 5A to 5C, the radius of curvature R of the curved surface 121 shown in Figure 5B is larger than the radius of curvature R of the curved surface 121 shown in Figure 5A, and the radius of curvature R of the curved surface 121 shown in Figure 5C is larger than the radius of curvature R of the curved surface 121 shown in Figure 5B.

[0046] According to the turbine rotor blade 24 of one embodiment shown in Figures 5A to 5C, the upstream region 110 has a region 112 in which the radius of curvature R of the curved surface 121 increases from the tip end 44a toward the base end 44b. This makes it difficult for the flow velocity of the cooling air flowing through the region in the upstream region 110 between the inner wall surface 96i of the suction side blade wall 96 of the airfoil section 44 and the partition wall surface 92Li on the leading edge 45 side of the second partition wall 92 to decrease. This makes it possible to suppress a temperature rise of the blade metal. Furthermore, according to the turbine rotor blade 24 of one embodiment shown in Figures 5A to 5C, the upstream region 110 has the region 112, making it possible to reduce pressure loss of the cooling air in the upstream region 110. 5A to 5C according to one embodiment, the upstream region 110 includes the region 112, and therefore the position of the curved surface 121 moves from the suction side to the ventral side as it moves from the tip end 44a to the base end 44b in the region 112. This makes it easier to ensure a sufficient wall thickness on the suction side, specifically, the wall thickness of the suction side wall portion 48Ws of the shank 48.

[0047] 5A to 5C , the end 91 a on the base end 44 b side of the first partition wall 91 (see FIGS. 3A and 3B ) is preferably located farther away in the direction from the tip 44 a toward the base end 44 b than the base end 44 b of the airfoil portion 44. The position Pmax at which the radius of curvature R is maximum is preferably located farther away in the direction from the tip 44 a toward the base end 44 b than the end 91 a on the base end 44 b side of the first partition wall 91.

[0048] 5A to 5C , the region 112 in which the radius of curvature R of the curved surface 121 increases from the tip 44 a side toward the base end 44 b side can be secured relatively large in the blade height direction, making it more difficult for the flow velocity of the cooling air to decrease in the upstream region 110. Furthermore, according to the turbine rotor blade 24 according to an embodiment shown in Figures 5A to 5C , the region 112 in which the radius of curvature R of the curved surface 121 increases from the tip 44 a side toward the base end 44 b side can be secured relatively large in the blade height direction, making it possible to further reduce pressure loss of the cooling air in the upstream region 110. Furthermore, according to the turbine rotor blade 24 according to an embodiment shown in Figures 5A to 5C , the region 112 in which the radius of curvature R of the curved surface 121 increases from the tip 44 a side toward the base end 44 b side can be secured relatively large in the blade height direction, making it possible to secure a relatively large range in the blade height direction in which the wall thickness on the suction side can be easily secured.

[0049] 5A to 5C , an end 91 a on the base end 44 b side of the first partition wall 91 is located at a position farther away in a direction from the tip 44 a toward the base end 44 b than the base end 44 b of the airfoil portion 44. For example, as shown in FIG. 3B , a distance L1 in the blade height direction from the inlet portion 101 to the end 91 a on the base end 44 b side of the first partition wall 91 may be 50% or more of a distance L2 in the blade height direction from the inlet portion 101 to a position Pb in the first turn portion 71 that is farthest from the inlet portion 101.

[0050] This expands the range in which the upstream region 110 exists in the blade height direction, thereby increasing the distance in the blade height direction from the inlet 101 to the position Pmax where the radius of curvature R is maximum. This makes it easier to increase the radius of curvature R, making it more difficult for the flow velocity of the cooling air in the upstream region 110 to decrease. Furthermore, with the turbine rotor blade 24 according to one embodiment shown in Figures 5A to 5C , the distance in the blade height direction from the inlet 101 to the position Pmax where the radius of curvature R is maximum can be increased, thereby further reducing the pressure loss of the cooling air in the upstream region 110. Furthermore, with the turbine rotor blade 24 according to one embodiment shown in Figures 5A to 5C , the distance in the blade height direction from the inlet 101 to the position Pmax where the radius of curvature R is maximum can be increased, thereby ensuring a relatively large range in the blade height direction where the wall thickness on the suction side can be easily ensured.

[0051] 6A to 6C , when viewed from the blade height direction, the partition wall surface 92Li includes a first partition wall surface 921Li extending in a first direction Dr1 from the pressure side (pressure-side blade surface 41 side) of the airfoil portion 44 toward the suction side (suction-side blade surface 43 side) and a second partition wall surface 922Li extending in a second direction Dr2 that intersects with the first direction Dr1 on the suction side of the first partition wall surface 921Li. The second direction Dr2 is inclined with respect to the first direction Dr1 so as to be closer to the leading edge 45 than the first direction Dr1 as it extends from the pressure side to the suction side. The upstream region 110 has a region 114 (see FIG. 3B ) in which the angular difference Δθ between the extension directions of the first direction Dr1 and the second direction Dr2 increases as it extends from the tip end 44a toward the base end 44b.

[0052] In a region 114 where the angular difference Δθ between the extending directions of the first direction Dr1 and the second direction Dr2 increases from the tip 44a side to the base end 44b side, a connection portion 97 between the second partition wall surface 922Li and the inner wall surface 96i of the suction side blade wall 96 as viewed from the blade height direction moves toward the leading edge 45 as viewed from the tip 44a side to the base end 44b side. Therefore, in the turbine blade 24 according to another embodiment shown in Figures 6A to 6C, the angle θa between the second partition wall surface 922Li and the inner wall surface 96i of the suction side blade wall 96 as viewed from the blade height direction increases as viewed from the tip 44a side to the base end 44b side. Therefore, according to the turbine blade 24 according to another embodiment shown in Figures 6A to 6C, the flow velocity of the cooling air flowing through the region sandwiched between the second partition wall surface 922Li and the inner wall surface 96i of the suction side blade wall 96 is less likely to decrease. 6A to 6C , in the region 114 where the angle difference Δθ between the extending directions of the first direction Dr1 and the second direction Dr2 increases from the tip 44 a side toward the base end 44 b side, the angle θa formed between the second partition wall surface 922Li and the inner wall surface 96i of the suction-side blade wall 96 when viewed from the blade height direction increases from the tip 44 a side toward the base end 44 b side, so that the pressure loss of the cooling air in the upstream region 110 can be reduced. Furthermore, in the turbine blade 24 according to another embodiment shown in FIGS. 6A to 6C , in the region 114 where the angular difference Δθ between the extension directions of the first direction Dr1 and the second direction Dr2 increases from the tip 44 a side toward the base end 44 b side, the connection portion 97 between the second partition wall surface 922Li and the inner wall surface 96i of the suction side blade wall 96 when viewed from the blade height direction moves toward the leading edge 45 side from the tip 44 a side toward the base end 44 b side, so that it becomes easier to ensure the suction side wall thickness of the upstream region 110, specifically the wall thickness of the suction side wall portion 48Ws of the shank 48.

[0053] In the turbine rotor blade 24 according to another embodiment shown in Figures 6A to 6C, the upstream region 110 may have a region (region 114) in which the connection portion 97 moves toward the leading edge 45 as it moves from the tip 44a toward the base end 44b.

[0054] 6A to 6C , in the upstream region 110, in a region 114 where the connecting portion 97 moves toward the leading edge 45 from the tip 44 a side toward the base end 44 b side, the angle θa formed between the second partition wall surface 922Li and the inner wall surface 96i of the suction side blade wall 96 when viewed from the blade height direction increases from the tip 44 a side toward the base end 44 b side. Therefore, according to the turbine blade 24 according to the other embodiment shown in FIGS. 6A to 6C , the flow velocity of the cooling air flowing through the region sandwiched between the second partition wall surface 922Li and the inner wall surface 96i of the suction side blade wall 96 is less likely to decrease. This makes it possible to suppress a temperature rise in the blade metal. 6A to 6C , in a region (region 114) in the upstream region 110 where the connecting portion 97 moves toward the leading edge 45 from the tip 44 a side toward the base end 44 b side, the angle θa formed between the second partition wall surface 922Li and the inner wall surface 96i of the suction-side blade wall 96 when viewed from the blade height direction increases from the tip 44 a side toward the base end 44 b side, thereby reducing pressure loss of the cooling air in the upstream region 110. Furthermore, in the turbine blade 24 according to the other embodiment shown in FIG. 6A to 6C , the connecting portion 97 moves toward the leading edge 45 from the tip 44 a side toward the base end 44 b side, making it easier to ensure the wall thickness on the suction side of the upstream region 110.

[0055] 5A to 6C , in the turbine rotor blade 24 according to some embodiments, the airfoil portion 44 may have therein a trailing-edge-side serpentine passage 52 that is a passage independent of the leading-edge-side serpentine passage 51 and that extends in a serpentine manner toward the trailing edge 46 on the trailing edge 46 side of the leading-edge-side serpentine passage 51. This allows the region of the airfoil portion 44 on the leading edge 45 side to be cooled by circulating cooling air through the leading-edge-side serpentine passage 51, and the region of the airfoil portion 44 on the trailing edge 46 side to be cooled by circulating cooling air through the trailing-edge-side serpentine passage 52.

[0056] 5A to 6C , in the turbine rotor blade 24 according to some embodiments, the first cooling passage 61 may be adjacent to the leading edge 45 of the airfoil portion 44 via a blade wall 44W on the leading edge 45 side of the airfoil portion 44. This makes it possible to reduce the amount of cooling air used to cool the airfoil portion 44, compared to when the airfoil portion 44 has a cooling passage 60 that is independent of the first cooling passage 61, extends in the blade height direction, and is formed closer to the leading edge 45 than the first cooling passage 61.

[0057] Figure 7 is a schematic diagram showing a cross section along the camber line of the airfoil portion 44 of a turbine rotor blade 24A according to yet another embodiment. In the turbine rotor blade 24A shown in Figure 7, the multiple cooling channels 60 may include a cooling channel 85 that is independent of the first cooling channel 61, extends in the blade height direction, and is formed closer to the leading edge 45 than the first cooling channel 61. For example, in the turbine rotor blade 24A shown in Figure 7, the cooling channel 85 has an opening 85a, which is an opening on one end side (inlet side), formed at the base of the blade, i.e., the bottom 47a of the blade root portion 47. In the turbine rotor blade 24A shown in Figure 7, the multiple cooling holes 69 are connected to the cooling channel 85.

[0058] 7 , cooling air supplied from an opening 85a, which is a cooling air intake, flows from the base end 44b toward the tip 44a in the cooling flow passage 85. A portion of the cooling air that has flowed into the cooling flow passage 85 is blown out as film cooling air 81 from multiple cooling holes 69 to film-cool the airfoil portion 44 from the outside. A portion of the cooling air that has flowed into the cooling flow passage 85 is blown out of the turbine blade 24A from an opening 85b formed in the tip 44a of the airfoil portion 44.

[0059] According to the turbine rotor blade 24A shown in FIG. 7, the cooling air flowing into the cooling flow passage 85 can further suppress the temperature rise of the blade metal in the region on the leading edge 45 side of the airfoil portion 44.

[0060] The gas turbine 1 according to some embodiments includes the turbine rotor blades 24 according to some embodiments described above, and therefore can suppress a temperature rise in the blade metal and improve the life of the turbine rotor blades 24. This reduces the frequency of maintenance of the gas turbine 1 and reduces the maintenance costs of the gas turbine 1.

[0061] The present disclosure is not limited to the above-described embodiments, but also includes modifications to the above-described embodiments and appropriate combinations of these modifications.

[0062] The contents of the above embodiments can be understood as follows, for example. (1) A turbine rotor blade 24, 24A according to at least one embodiment of the present disclosure includes a blade body 49 including an airfoil portion 44 having a plurality of cooling channels 60 therein. The plurality of cooling channels 60 include, in order from the trailing edge 46 side of the airfoil portion 44, a third cooling channel 63 extending in the blade height direction and opening at the blade root (bottom portion 47 a), a second cooling channel 62 extending in the blade height direction and formed closer to the leading edge 45 of the airfoil portion 44 than the third cooling channel 63 and connected to the third cooling channel 63 at a second turning portion 72 on the tip 44 a side of the airfoil portion 44, and a first cooling channel 61 extending in the blade height direction and formed closer to the leading edge 45 than the second cooling channel 62 and connected to the second cooling channel 62 at a first turning portion 71 on the base end 44 b side of the airfoil portion 44. The turbine blade 24, 24A according to at least one embodiment of the present disclosure includes a first partition wall 91 that separates the first cooling flow passage 61 and the second cooling flow passage 62, and a second partition wall 92 that separates the second cooling flow passage 62 and the third cooling flow passage 63. The first folded portion 71 includes an upstream region 110 that extends along the extension direction of the first folded portion 71 from an inlet portion 101 at the same position in the blade height direction as the base end 44b of the airfoil portion 44 to a position C1 that is the center of the first partition wall 91 in the wall thickness direction when viewed from the blade height direction. A portion of the upstream region 110 when viewed from the blade height direction is defined at the base end 44b of the airfoil portion 44 by an inner wall surface 96i of the suction-side airfoil wall 96 of the airfoil portion 44 and a partition wall surface 92Li on the leading edge 45 side of the second partition wall 92. In at least a portion of the upstream region 110 in the blade height direction, the inner wall surface 96i and the partition wall surface 92Li are connected by a curved surface 121 that is convex from the inside to the outside of the airfoil portion 44 when viewed in the blade height direction. The upstream region 110 has a region 112 in which the radius of curvature R of the curved surface 121 increases from the tip end 44a toward the base end 44b.

[0063] According to the configuration (1) above, the upstream region 110 has a region 112 in which the radius of curvature R of the curved surface 121 increases from the tip 44a side toward the base 44b side. This makes it difficult for the flow velocity of the cooling air flowing through the region in the upstream region 110 between the inner wall surface 96i of the suction side blade wall 96 of the airfoil section 44 and the partition wall surface 92Li on the leading edge 45 side of the second partition wall 92 to decrease. This makes it possible to suppress a temperature rise of the blade metal. Furthermore, according to the configuration (1) above, the upstream region 110 has a region 112 in which the radius of curvature R of the curved surface 121 increases from the tip 44a side toward the base 44b side. This makes it possible to reduce pressure loss of the cooling air in the upstream region 110. Furthermore, according to the configuration (1) above, the upstream region 110 has a region 112 in which the radius of curvature R of the curved surface 121 increases from the tip end 44a toward the base end 44b, and therefore the position of the curved surface 121 moves from the dorsal side toward the ventral side from the tip end 44a toward the base end 44b in the region 112. This makes it easier to ensure the wall thickness on the dorsal side.

[0064] (2) In some embodiments, in the configuration of (1) above, the end 91 a on the base end 44 b side of the first partition wall 91 may be located at a position farther away in the direction from the tip end 44 a toward the base end 44 b than the base end 44 b of the airfoil portion 44. The position Pmax at which the radius of curvature R is maximum may be located at a position farther away in the direction from the tip end 44 a toward the base end 44 b than the end 91 a on the base end 44 b side of the first partition wall 91.

[0065] According to the above configuration (2), a relatively large region 112 in which the radius of curvature R of the curved surface 121 increases from the tip 44a side toward the base 44b side can be ensured in the blade height direction, making it more difficult for the flow rate of the cooling air to decrease in the upstream region 110. Furthermore, according to the above configuration (2), a relatively large region 112 in which the radius of curvature R of the curved surface 121 increases from the tip 44a side toward the base 44b side can be ensured in the blade height direction, making it possible to further reduce pressure loss of the cooling air in the upstream region 110. Furthermore, according to the above configuration (2), a relatively large region 112 in which the radius of curvature R of the curved surface 121 increases from the tip 44a side toward the base 44b side can be ensured in the blade height direction, making it possible to ensure a relatively large range in the blade height direction in which the wall thickness on the suction side can be easily ensured.

[0066] (3) In some embodiments, in the configuration described in (1) or (2) above, the end 91 a of the first partition wall 91 on the base end 44 b side is located at a position farther away in the direction from the tip 44 a side toward the base end 44 b side than the base end 44 b of the airfoil portion 44. The distance L1 in the blade height direction from the inlet portion 101 to the end 91 a of the first partition wall 91 on the base end 44 b side may be 50% or more of the distance L2 in the blade height direction from the inlet portion 101 to a position Pb in the first folded portion 71 that is farthest from the inlet portion 101.

[0067] According to the configuration (3), the range in which the upstream region 110 exists is expanded in the blade height direction, so that the distance in the blade height direction from the inlet portion 101 to the position Pmax where the radius of curvature R is maximum can be increased. This makes it easier to increase the radius of curvature R, and makes it more difficult for the flow velocity of the cooling air in the upstream region 110 to decrease. Furthermore, according to the configuration (3), the distance in the blade height direction from the inlet portion 101 to the position Pmax where the radius of curvature R is maximum can be increased, so that the pressure loss of the cooling air in the upstream region 110 can be further reduced. Furthermore, according to the configuration (3), the distance in the blade height direction from the inlet portion 101 to the position Pmax where the radius of curvature R is maximum can be increased, so that a relatively large range in the blade height direction can be secured where the wall thickness on the suction side can be easily secured.

[0068] (4) In some embodiments, in any of the configurations (1) to (3) above, the first cooling channel 61, the second cooling channel 62, and the third cooling channel 63 may form a leading-edge-side serpentine channel 51 that extends in a serpentine manner toward the leading edge 45. The airfoil portion 44 may have therein a trailing-edge-side serpentine channel 52 that is a channel independent of the leading-edge-side serpentine channel 51 and that extends in a serpentine manner toward the trailing edge 46 on the trailing edge 46 side of the leading-edge-side serpentine channel 51.

[0069] According to the configuration (4) above, the area on the leading edge 45 side of the airfoil portion 44 can be cooled by circulating cooling air through the leading edge side serpentine flow path 51, and the area on the trailing edge 46 side of the airfoil portion 44 can be cooled by circulating cooling air through the trailing edge side serpentine flow path 52.

[0070] (5) The turbine blade 24, 24A according to at least one embodiment of the present disclosure includes a blade body 49 including an airfoil portion 44 having a plurality of cooling channels 60 therein. The plurality of cooling channels 60 include, in order from the trailing edge 46 side of the airfoil portion 44, a third cooling channel 63 extending in the blade height direction and opening at the root (bottom portion 47 a) of the blade, a second cooling channel 62 extending in the blade height direction and formed closer to the leading edge 45 of the airfoil portion 44 than the third cooling channel 63 and connected to the third cooling channel 63 at a second turning portion 72 on the tip 44 a side of the airfoil portion 44, and a first cooling channel 61 extending in the blade height direction and formed closer to the leading edge 45 than the second cooling channel 62 and connected to the second cooling channel 62 at a first turning portion 71 on the base end 44 b side of the airfoil portion 44. The turbine blade 24, 24A according to at least one embodiment of the present disclosure includes a first partition wall 91 that separates the first cooling flow passage 61 and the second cooling flow passage 62, and a second partition wall 92 that separates the second cooling flow passage 62 and the third cooling flow passage 63. The first folded portion 71 includes an upstream region 110 that extends along the extension direction of the first folded portion 71 from an inlet portion 101 at the same position in the blade height direction as the base end 44b of the airfoil portion 44 to a position C1 that is the center of the first partition wall 91 in the wall thickness direction when viewed from the blade height direction. A portion of the upstream region 110 when viewed from the blade height direction is defined at the base end 44b of the airfoil portion 44 by an inner wall surface 96i of the suction-side airfoil wall 96 of the airfoil portion 44 and a partition wall surface 92Li on the leading edge 45 side of the second partition wall 92. When viewed from the blade height direction, the partition wall surface 92Li includes a first partition wall surface 921Li extending in a first direction Dr1 from the ventral side toward the suction side of the airfoil portion 44, and a second partition wall surface 922Li extending in a second direction Dr2 that intersects with the first direction Dr1 on the suction side of the first partition wall surface 921Li. The second direction Dr2 is inclined with respect to the first direction Dr1 so as to be closer to the leading edge 45 than the first direction Dr1 as it extends from the ventral side to the suction side. The upstream region 110 has a region 114 in which the angular difference Δθ between the extension directions of the first direction Dr1 and the second direction Dr2 increases as it extends from the tip end 44a toward the base end 44b.

[0071] In the configuration (5) above, in the region 114 where the angle difference Δθ between the extending directions of the first direction Dr1 and the second direction Dr2 increases from the tip 44a side toward the base end 44b side, the connection portion 97 between the second partition wall surface 922Li and the inner wall surface 96i of the suction side blade wall 96 as viewed from the blade height direction moves toward the leading edge 45 as viewed from the tip 44a side toward the base end 44b side. Therefore, in the configuration (5) above, the angle θa between the second partition wall surface 922Li and the inner wall surface 96i of the suction side blade wall 96 as viewed from the blade height direction increases as viewed from the tip 44a side toward the base end 44b side. Therefore, according to the configuration (5) above, the flow velocity of the cooling air flowing through the region sandwiched between the second partition wall surface 922Li and the inner wall surface 96i of the suction side blade wall 96 is less likely to decrease. This suppresses the temperature rise of the blade metal. Furthermore, according to the configuration (5) above, in the region 114 where the angle difference Δθ between the extending directions of the first direction Dr1 and the second direction Dr2 increases from the tip end 44a side toward the base end 44b side, the angle Δθ formed between the second partition wall surface 922Li and the inner wall surface 96i of the suction-side blade wall 96 when viewed from the blade height direction increases from the tip end 44a side toward the base end 44b side, so that the pressure loss of the cooling air in the upstream region 110 can be reduced. Furthermore, according to the configuration (5) above, in the region 114 where the angle difference Δθ between the extension directions of the first direction Dr1 and the second direction Dr2 increases from the tip end 44a side toward the base end 44b side, the connection portion 97 between the second partition wall surface 922Li and the inner wall surface 96i of the suction side blade wall 96 when viewed from the blade height direction moves toward the leading edge 45 side from the tip end 44a side toward the base end 44b side, making it easier to ensure the suction side wall thickness of the upstream region 110.

[0072] (6) In some embodiments, in the configuration of (5) above, the upstream region 110 may include a connection portion 97 between the inner wall surface 96i of the suction-side blade wall 96 and the second partition wall surface 922Li. The upstream region 110 may have a region (region 114) where the connection portion 97 moves toward the leading edge 45 as it moves from the tip end 44a toward the base end 44b.

[0073] In the configuration (6) above, in the upstream region 110, in a region (region 114) where the connecting portion 97 moves toward the leading edge 45 as it moves from the tip 44a side toward the base end 44b side, the angle θa formed between the second partition wall surface 922Li and the inner wall surface 96i of the suction side blade wall 96 when viewed from the blade height direction increases as it moves from the tip 44a side toward the base end 44b side. Therefore, according to the configuration (6) above, the flow velocity of the cooling air flowing through the region sandwiched between the second partition wall surface 922Li and the inner wall surface 96i of the suction side blade wall 96 is less likely to decrease. This makes it possible to suppress a temperature rise of the blade metal. According to the configuration (6) above, in a region (region 114) in the upstream region 110 where the connecting portion 97 moves toward the leading edge 45 as it moves from the tip 44 a side toward the base end 44 b side, the angle θa formed between the second partition wall surface 922Li and the inner wall surface 96i of the suction-side blade wall 96 when viewed from the blade height direction increases as it moves from the tip 44 a side toward the base end 44 b side, thereby reducing pressure loss of the cooling air in the upstream region 110. Furthermore, according to the configuration (6) above, the connecting portion 97 moves toward the leading edge 45 as it moves from the tip 44 a side toward the base end 44 b side, making it easier to ensure the wall thickness on the suction side of the upstream region 110.

[0074] (7) In some embodiments, in the configuration of (5) or (6) above, the first cooling channel 61, the second cooling channel 62, and the third cooling channel 63 may form a leading-edge-side serpentine channel 51 that extends in a serpentine manner toward the leading edge 45. The airfoil portion 44 may have therein a trailing-edge-side serpentine channel 52 that is a channel independent of the leading-edge-side serpentine channel 51 and that extends in a serpentine manner toward the trailing edge 46 on the trailing edge 46 side of the leading-edge-side serpentine channel 51.

[0075] According to the configuration (7) above, the area on the leading edge 45 side of the airfoil portion 44 can be cooled by circulating cooling air through the leading edge side serpentine flow path 51, and the area on the trailing edge 46 side of the airfoil portion 44 can be cooled by circulating cooling air through the trailing edge side serpentine flow path 52.

[0076] (8) In some embodiments, in any of the configurations (1) to (7) above, the first cooling flow path 61 may be adjacent to the leading edge 45 via the wing wall 44W on the leading edge 45 side of the airfoil portion 44.

[0077] According to the configuration (8) above, the amount of cooling air used to cool the airfoil portion 44 can be reduced compared to when the airfoil portion 44 has a cooling flow path 60 that is independent of the first cooling flow path 61, extends in the blade height direction, and is formed closer to the leading edge 45 than the first cooling flow path 61.

[0078] (9) In some embodiments, in any of the configurations (1) to (7) above, the multiple cooling passages 60 may include a cooling passage 85 that is a passage independent of the first cooling passage 61, extends in the blade height direction, and is formed closer to the leading edge 45 than the first cooling passage 61.

[0079] According to the above configuration (9), the temperature rise of the blade metal in the region on the leading edge 45 side of the airfoil portion 44 can be further suppressed.

[0080] (10) The gas turbine 1 according to at least one embodiment of the present disclosure includes turbine rotor blades 24, 24A having any of the configurations described in (1) to (9) above.

[0081] According to the configuration (10) above, the temperature rise of the blade metal can be suppressed, thereby improving the life of the turbine rotor blades 24, 24A. As a result, the frequency of maintenance of the gas turbine 1 can be reduced, and the maintenance costs of the gas turbine 1 can be reduced.

[0082] REFERENCE SIGNS LIST 1 Gas turbine 6 Turbine 24, 24A Turbine rotor blade 41 Pressure side blade surface (pressure surface) 42 Platform 43 Suction side blade surface (suction surface) 44 Airfoil section 44a Tip 44b Base end 44W Blade wall 45 Leading edge 46 Trailing edge 47 Blade root 47a Bottom 48 Shank 51 Serpentine flow path (leading edge side serpentine flow path) 52 Serpentine flow path (trailing edge side serpentine flow path) 60 Cooling flow path 61 First cooling flow path 62 Second cooling flow path 63 Third cooling flow path 63a Opening 64 Fourth cooling flow path 65 Fifth cooling flow path 66 Sixth cooling flow path 67 Cooling flow path 71 First turning portion 71b Surface 72 Second turning portion 73 Third turning portion 74 Fourth turning portion 85 Cooling flow path 91 First partition wall 91a End 92 Second partition wall 92Li Partition wall surface 93 Third partition wall 94 Fourth partition wall 95 Fifth partition wall 96 Suction side blade wall 96i Inner wall surface 97 Connection portion 101 Inlet portion 110 Upstream side region 112 Region 114 Region 121 Curved surface 921Li First partition wall surface 922Li Second partition wall surface

Claims

1. A blade body including an airfoil portion having a plurality of cooling passages therein, the plurality of cooling passages including, in order from the trailing edge side of the airfoil portion, a third cooling passage extending in the airfoil height direction and opening at a base of the blade; a second cooling passage extending in the airfoil height direction and formed on the leading edge side of the airfoil portion relative to the third cooling passage and connected to the third cooling passage at a second turn-back portion on the tip side of the airfoil portion; a first cooling passage extending in the airfoil height direction and formed on the leading edge side of the second cooling passage and connected to the second cooling passage at a first turn-back portion on the base side of the airfoil portion; a first partition wall separating the first cooling passage and the second cooling passage; and a second partition wall separating the second cooling passage and the third cooling passage. the first turn-back portion includes an upstream region extending along an extension direction of the first turn-back portion from an inlet portion at the same position in the blade height direction as a base end of the airfoil portion to a center position in the wall thickness direction of the first partition wall when viewed from the blade height direction, a part of the upstream region when viewed from the blade height direction is defined by an inner wall surface of a suction side airfoil wall of the airfoil portion and the leading edge side partition wall surface of the second partition wall at the base end of the airfoil portion, the upstream region having a region in which the inner wall surface and the partition wall surface are connected by a curved surface that is convex from the inside to the outside of the airfoil portion when viewed from the blade height direction, and a radius of curvature of the curved surface increases from the tip side to the base end side.

2. A turbine blade as described in claim 1, wherein the base end side end of the first partition wall is located at a position farther away in the direction from the tip side to the base end side than the base end of the airfoil portion, and the position at which the radius of curvature is maximum is located at a position farther away in the direction from the tip side to the base end side than the base end side end of the first partition wall.

3. A turbine blade as described in claim 1 or 2, wherein the base end side end of the first partition wall is located at a position farther away from the base end of the airfoil portion in the direction from the tip side to the base end side, and the distance in the blade height direction from the inlet portion to the base end side end of the first partition wall is 50% or more of the distance in the blade height direction from the inlet portion to a position in the first turn portion that is farthest from the inlet portion.

4. A turbine blade as described in claim 1 or 2, wherein the first cooling passage, the second cooling passage and the third cooling passage form a leading edge side serpentine passage extending in a serpentine manner toward the leading edge side, and the airfoil portion has therein a trailing edge side serpentine passage which is a passage independent of the leading edge side serpentine passage and which extends in a serpentine manner toward the trailing edge side on the trailing edge side of the leading edge side serpentine passage.

5. A blade body including an airfoil portion having a plurality of cooling passages therein, the plurality of cooling passages including, in order from the trailing edge side of the airfoil portion, a third cooling passage extending in the airfoil height direction and opening at the base of the blade, a second cooling passage extending in the airfoil height direction and formed on the leading edge side of the airfoil portion relative to the third cooling passage and connected to the third cooling passage at a second turn-back portion on the tip side of the airfoil portion, and a first cooling passage extending in the airfoil height direction and formed on the leading edge side of the second cooling passage and connected to the second cooling passage at a first turn-back portion on the base side of the airfoil portion; a first partition wall separating the first cooling passage and the second cooling passage, and a second partition wall separating the second cooling passage and the third cooling passage, the first folded portion includes an upstream region extending along an extending direction of the first folded portion from an inlet portion at the same position in the wing height direction as the base end of the airfoil portion to a center position in the wall thickness direction of the first partition wall when viewed from the wing height direction, a part of the upstream region when viewed from the wing height direction is defined by an inner wall surface of a suction side airfoil wall of the airfoil portion and a partition wall surface of the leading edge side of the second partition wall at the base end of the airfoil portion, when viewed from the wing height direction, the partition wall surface includes a first partition wall surface extending in a first direction from the ventral side to the suction side of the airfoil portion, and a second partition wall surface extending in a second direction intersecting the first direction on the suction side of the first partition wall surface, the second direction being inclined with respect to the first direction so as to be more toward the leading edge side than the first direction as it goes from the ventral side to the suction side, the upstream region has a region in which an angle difference between the extending directions of the first direction and the second direction increases from the tip side toward the base end side.

6. The turbine blade according to claim 5, wherein the upstream region includes a connection portion between the inner wall surface of the suction side blade wall and the second partition wall surface, and has a region in which the connection portion moves toward the leading edge side as it moves from the tip side to the base end side.

7. A turbine blade as described in claim 5 or 6, wherein the first cooling passage, the second cooling passage and the third cooling passage form a leading edge side serpentine passage extending in a serpentine manner toward the leading edge side, and the airfoil portion has therein a trailing edge side serpentine passage which is a passage independent of the leading edge side serpentine passage and which extends in a serpentine manner toward the trailing edge side on the trailing edge side of the leading edge side serpentine passage.

8. A turbine rotor blade according to claim 1 or 5, wherein the first cooling passage is adjacent to the leading edge via a blade wall on the leading edge side of the airfoil portion.

9. A turbine blade according to claim 1 or 5, wherein the plurality of cooling passages include a cooling passage that is independent of the first cooling passage, extends in the blade height direction, and is formed closer to the leading edge than the first cooling passage.

10. A gas turbine comprising a turbine blade according to claim 1 or 5.

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