Turbine rotor blade and gas turbine

Abstract

In a turbine rotor blade of a gas turbine, a plurality of first film cooling holes each including a cooling fluid inlet formed on a flow passage wall surface of a first cooling flow passage part, and a plurality of second film cooling holes each including a cooling fluid inlet formed on a flow passage wall surface of a second cooling flow passage part adjacent to the first cooling flow passage part are formed in a suction surface forming wall part. When a range on a tip side of an airfoil part is defined as a blade tip side range with a center position of the airfoil part in a blade height direction as a reference, the number of second film cooling holes belonging to the blade tip side range is smaller than the number of first film cooling holes belonging to the blade tip side range.

Description

Turbine Rotor Blade and Gas Turbine

[0001] This disclosure relates to a turbine rotor blade and a gas turbine. This application claims priority based on Japanese Patent Application No. 2024-172985 filed with the Japan Patent Office on October 2, 2024, the content of which is incorporated herein by reference.

[0002] Patent Document 1 discloses a structure for a turbine rotor blade of a gas turbine, aiming to improve the durability of the turbine rotor blade while suppressing the amount of cooling air used. In this turbine rotor blade, a plurality of cooling flow path portions extending along the blade height direction are arranged inside the airfoil portion in the direction from the leading edge side to the trailing edge side of the airfoil portion. On the suction surface of the airfoil portion, there are a row of ejection holes for ejecting the cooling fluid flowing through the cooling flow path portion closest to the leading edge of the airfoil portion among the plurality of cooling flow path portions from the leading edge portion of the airfoil portion, and a row of film cooling holes for allowing the cooling fluid flowing through the cooling flow path portion second closest to the leading edge of the airfoil portion among the plurality of cooling flow path portions to flow out from the base end side of the airfoil portion.

[0003] Japanese Unexamined Patent Application Publication No. 2023-183113

[0004] By the way, as a result of the inventors' intensive studies, it has been clarified that in a turbine rotor blade with a thermal barrier coating (TBC) applied to its surface, the thermal barrier coating peels off due to the impact of flying objects in the region on the tip side near the leading edge of the airfoil portion, and the thinning of the airfoil portion tends to progress. Regarding this point, Patent Document 1 does not disclose any findings for achieving both suppressing the progress of the thinning of the airfoil portion caused by the impact of such flying objects and suppressing the increase in the flow rate of the cooling fluid.

[0005] In view of the above circumstances, at least one embodiment of this disclosure aims to provide a turbine rotor blade of a gas turbine capable of suppressing the progress of the thinning of the airfoil portion caused by the impact of flying objects while suppressing the increase in the flow rate of the cooling fluid, and a gas turbine including the same.

[0006] To achieve the above objective, a turbine blade according to at least one embodiment of the present disclosure is a turbine blade for a gas turbine, comprising: an airfoil portion; a platform portion connected to the base end of the airfoil portion; and a blade root portion provided on the opposite side of the platform portion from the airfoil portion, wherein the airfoil portion includes a pressure surface forming wall portion that forms a pressure surface and a negative pressure surface forming wall portion that forms a negative pressure surface, and a plurality of cooling passage portions extending along the blade height direction of the airfoil portion are formed between the pressure surface forming wall portion and the negative pressure surface forming wall portion in the airfoil portion, the plurality of cooling passage portions are arranged along the direction from the leading edge to the trailing edge of the airfoil portion, and the cooling passage portion closest to the leading edge among the plurality of cooling passage portions is defined as the first cooling passage portion, and the cooling passage portion adjacent to the first cooling passage portion among the plurality of cooling passage portions is defined as the second cooling passage portion, the negative pressure surface forming wall portion of the airfoil portion includes, A plurality of first film cooling holes are formed, each including a cooling fluid inlet formed on the wall surface of the first cooling channel and a cooling fluid outlet formed on the negative pressure surface of the airfoil; and a plurality of second film cooling holes are formed, each including a cooling fluid inlet formed on the wall surface of the second cooling channel and a cooling fluid outlet formed on the negative pressure surface of the airfoil. If the intersection point of the line indicating the thickness direction of the wall portion forming the negative pressure surface at the center position of the cooling fluid inlet of the first film cooling hole and the negative pressure surface is defined as the first intersection point, then the center of the cooling fluid outlet of the first film cooling hole is located downstream of the first intersection point in the combustion gas flow direction along the negative pressure surface. If the range on the tip side of the airfoil is defined as the tip side range with respect to a position at 2 / 3 of the airfoil height, then the number of second film cooling holes belonging to the tip side range is less than the number of first film cooling holes belonging to the tip side range.

[0007] To achieve the above objective, a gas turbine according to at least one embodiment of the present disclosure is a gas turbine comprising a compressor, a combustor, and a turbine, wherein the turbine includes the turbine blades.

[0008] According to at least one embodiment of the present disclosure, a turbine blade for a gas turbine and a gas turbine equipped therewith are provided that can suppress the progression of thinning of the airfoil section caused by the collision of flying objects while suppressing an increase in the flow rate of the cooling fluid.

[0009] This figure shows the schematic configuration of a gas turbine 2 according to one embodiment. This is a schematic perspective view showing an example of a turbine blade 16. This figure shows an example of a cross-section perpendicular to the blade height direction in the airfoil portion 20. This is a schematic cross-sectional view showing an enlarged view of the vicinity of the leading edge 21 of the airfoil portion 20 shown in Figure 3. This is a schematic cross-sectional view for explaining the extending directions of the first ejection hole 60, the first film cooling hole 62, and the second film cooling hole 64 in the cross-section shown in Figure 4. This is a schematic perspective view showing a modified example of the turbine blade 16. This is a schematic perspective view showing yet another

[0010] Hereinafter, several embodiments of this disclosure will be described with reference to the attached drawings. However, the dimensions, materials, shapes, relative arrangements, etc., of the components described as embodiments or shown in the drawings are not intended to limit the scope of the invention, but are merely illustrative examples. For example, expressions describing relative or absolute arrangements such as "in a certain direction," "along a certain direction," "parallel," "orthogonal," "center," "concentric," or "coaxial" should not only strictly represent such arrangements, but also represent states where there is a tolerance, or a relative displacement of an angle or distance sufficient to achieve the same function. For example, expressions describing things as being in an equal state such as "identical," "equal," and "homogeneous" should not only strictly represent states of equality, but also represent states where there is a tolerance, or a difference sufficient to achieve the same function. For example, expressions describing shapes such as a square shape or a cylindrical shape should not only represent geometrically precise shapes such as square shapes or cylindrical shapes, but also represent shapes including concave and concave parts, chamfered parts, etc., to the extent that the same effect can be achieved. On the other hand, expressions such as "to possess," "to be equipped with," "to have," "to include," or "to have" a single component are not exclusive expressions that exclude the existence of other components.

[0011] Figure 1 is a diagram showing the schematic configuration of a gas turbine 2 according to one embodiment. As shown in Figure 1, the gas turbine 2 comprises a compressor 4, a combustor 6 that mixes the compressed air generated by the compressor 4 with fuel and burns the fuel, and a turbine 8 that is driven by the combustion gas generated by the combustor 6.

[0012] As shown in Figure 1, the turbine 8 includes a rotor 9 (turbine rotor), a turbine casing 10 that houses the rotor 9, a plurality of turbine stator blades 12 (turbine stator blades) fixed to the inner surface of the turbine casing 10, and a plurality of turbine rotor blades 16 that are mounted on the rotor 9 so as to be alternately arranged in the axial direction with respect to the turbine stator blades 12. Hereinafter, unless otherwise specified, "circumferential direction" means the circumferential direction of the gas turbine 2, i.e., the circumferential direction of the rotor 9; unless otherwise specified, "axial direction" means the axial direction of the gas turbine 2, i.e., the axial direction of the rotor 9; and unless otherwise specified, "radial direction" means the radial direction of the gas turbine 2, i.e., the radial direction of the rotor 9.

[0013] Figure 2 is a schematic perspective view showing an example of a turbine blade 16. As shown in Figure 2, the turbine blade 16 includes an airfoil section 20, a platform section 22, and a blade root section 24.

[0014] The airfoil section 20 includes a leading edge 30, a trailing edge 31, a pressure surface 32 connecting the leading edge 30 and the trailing edge 31, and a negative pressure surface 33 connecting the leading edge 30 and the trailing edge 31. In the following description, "airfoil height direction" means the airfoil height direction of the airfoil section 20 (the direction along the airfoil height H, i.e., the distance from the base end 20h to the tip end 20t of the airfoil section 20 is H), "base end 20h side of the airfoil section 20" means the base end 20h side of the airfoil section 20 in the airfoil height direction (platform section 22 side, i.e., hub side), and "tip end 20t side of the airfoil section 20" means the tip end 20t side of the airfoil section 20 in the airfoil height direction (opposite side of the platform section 22, i.e., tip side). The airfoil height direction of the airfoil section 20 may also be the radial direction.

[0015] The platform portion 22 has a plate shape and includes an outward-facing surface 22a that faces outward in the radial direction and an inward-facing surface 22b that faces inward in the radial direction. The outward-facing surface 22a constitutes the flow path wall of the combustion gas flow path. The outward-facing surface 22a is connected to the base end 20h of the airfoil portion 20.

[0016] The blade root portion 24 is provided on the opposite side of the airfoil portion 20 from the platform portion 22, and the radial outer end of the blade root portion 24 is connected to the inward-facing surface 22b of the platform portion 22. The blade root portion 24 is mounted on the rotor 9 (see Figure 1). The airfoil portion 20 is made of a heat-resistant alloy such as a nickel-based alloy, and a heat-shielding coating layer 23 (such as a ceramic layer) is formed on its surface via a metal bond layer.

[0017] Figure 3 shows an example of a cross-section of the airfoil portion 20 perpendicular to the airfoil height direction. As shown in Figure 3, the airfoil portion 20 includes a pressure surface forming wall portion 40 that forms a pressure surface 32 and a negative pressure surface forming wall portion 42 that forms a negative pressure surface 33.

[0018] The pressure surface forming wall portion 40 is a curved plate-like portion extending from the front edge 30 to the rear edge 31, and the pressure surface 32 includes a concave curved surface and a convex curved surface. The negative pressure surface forming wall portion 42 is a curved plate-like portion extending from the front edge 30 to the rear edge 31, and the negative pressure surface 33 includes a convex curved surface.

[0019] Multiple cooling channel sections 44 are formed between the pressure surface forming wall section 40 and the negative pressure surface forming wall section 42 in the airfoil section 20. The multiple cooling channel sections 44 are arranged inside the airfoil section 20 along the direction from the leading edge 30 to the trailing edge 31 (i.e., along the camber line CL of the airfoil section 20), and each of the multiple cooling channel sections 44 extends along the airfoil height direction. The ends of adjacent cooling channel sections 44 in the multiple cooling channel sections 44 (the ends on the tip 20t side of the airfoil section 20 or the ends on the base side of the airfoil section 20) may be connected by a U-turn channel (not shown), and two or more cooling channel sections 44 in the multiple cooling channel sections 44 may constitute a serpentine channel connected by the U-turn channel.

[0020] In the illustrated example, the multiple cooling channel sections 44 include a first cooling channel section 44A, a second cooling channel section 44B, a third cooling channel section 44C, a fourth cooling channel section 44D, and a fifth cooling channel section 44E. The first cooling channel section 44A, the second cooling channel section 44B, the third cooling channel section 44C, the fourth cooling channel section 44D, and the fifth cooling channel section 44E are arranged in this order within the airfoil section 20, along the direction from the leading edge 30 to the trailing edge 31.

[0021] In the illustrated example, the first cooling channel section 44A is the cooling channel section 44 closest to the leading edge 30 among the multiple cooling channel sections 44, and the second cooling channel section 44B is the cooling channel section adjacent to the first cooling channel section 44A among the multiple cooling channel sections 44. The third cooling channel section 44C is located between the second cooling channel section 44B and the fourth cooling channel section 44D, and the fourth cooling channel section 44D is located between the third cooling channel section 44C and the fifth cooling channel section 44E. The fifth cooling channel section 44E is the cooling channel section 44 closest to the trailing edge 31 among the multiple cooling channel sections 44.

[0022] As shown in Figure 3, the airfoil section 20 includes a plurality of partition walls 50 connecting the inner surface 46 of the pressure surface forming wall section 40 and the inner surface 48 of the negative pressure surface forming wall section 42. The plurality of partition walls 50 are arranged inside the airfoil section 20 along the direction from the leading edge 30 to the trailing edge 31 (i.e., along the camber line CL of the airfoil section 20), and each of the plurality of partition walls 50 extends along the wing height direction. In the illustrated example, the plurality of partition walls 50 include a first partition wall section 50A, a second partition wall section 50B, a third partition wall section 50C, and a fourth partition wall section 50D. The first partition wall section 50A, the second partition wall section 50B, the third partition wall section 50C, and the fourth partition wall section 50D are arranged in this order inside the airfoil section 20 along the direction from the leading edge 30 to the trailing edge 31.

[0023] In the illustrated example, the first partition wall 50A is the partition wall 50 closest to the front edge 30 among the multiple partition wall 50, and the second partition wall 50B is the partition wall 50 adjacent to the first partition wall 50A among the multiple partition wall 50. Furthermore, the third partition wall 50C is located between the second partition wall 50B and the fourth partition wall 50D, and the fourth partition wall 50D is the partition wall 50 closest to the rear edge 31 among the multiple partition wall 50.

[0024] In the illustrated example, the first cooling channel section 44A is formed by a pressure surface forming wall section 40, a negative pressure surface forming wall section 42, and a first partition wall section 50A; the second cooling channel section 44B is formed by a pressure surface forming wall section 40, a negative pressure surface forming wall section 42, a first partition wall section 50A, and a second partition wall section 50B; the third cooling channel section 44C is formed by a pressure surface forming wall section 40, a negative pressure surface forming wall section 42, a second partition wall section 50B, and a third partition wall section 50C; the fourth cooling channel section 44D is formed by a pressure surface forming wall section 40, a negative pressure surface forming wall section 42, a third partition wall section 50C, and a fourth partition wall section 50D; and the fifth cooling channel section 44E is formed by a pressure surface forming wall section 40, a negative pressure surface forming wall section 42, and a fourth partition wall section 50D.

[0025] Figure 4 is a schematic cross-sectional view showing an enlarged view of the vicinity of the leading edge 21 of the airfoil portion 20 shown in Figure 3. For example, as shown in Figures 2 and 4, the negative pressure surface forming wall portion 42 of the airfoil portion 20 has a plurality of first ejection holes 60, a plurality of first film cooling holes 62, and a plurality of second film cooling holes 64. The pressure surface forming wall portion 40 of the airfoil portion 20 has a plurality of second ejection holes 68.

[0026] As shown in Figure 2, the multiple first ejection holes 60 are arranged in a row along the blade height direction, forming a first ejection hole row 61. The multiple first film cooling holes 62 are arranged in a row along the blade height direction, forming a first film cooling hole row 63. The multiple second film cooling holes 64 are arranged in a row along the blade height direction, forming a second film cooling hole row 65. In the following description, "multiple first ejection holes 60" means all first ejection holes 60 provided by the airfoil section 20 (all first ejection holes 60 constituting the first ejection hole row 61) unless otherwise specified, "multiple first film cooling holes 62" means all first film cooling holes 62 provided by the airfoil section 20 (all first film cooling holes 62 constituting the first film cooling hole row 63) unless otherwise specified, and "multiple second film cooling holes 64" means all second film cooling holes 64 provided by the airfoil section 20 (all second film cooling holes 64 constituting the second film cooling hole row 65) unless otherwise specified. Note that the multiple first ejection holes 60 may or may not be arranged at equal intervals. The multiple first film cooling holes 62 may or may not be arranged at equal intervals. The multiple second film cooling holes 64 may or may not be arranged at equal intervals.

[0027] As shown in Figure 4, each of the multiple first ejection holes 60 includes a cooling fluid inlet 60a formed in the flow path wall surface 45 of the first cooling flow path section 44A and a cooling fluid outlet 60b formed in the negative pressure surface 33 of the leading edge 21 of the airfoil section 20. The leading edge 21 of the airfoil section 20 is the portion of the airfoil section 20 that includes the leading edge 30, and is, for example, the portion of the pressure surface forming wall section 40 and the negative pressure surface forming wall section 42 that is on the leading edge 30 side of the multiple first film cooling holes 62 in the direction along the camber line of the airfoil section 20.

[0028] Each of the multiple first film cooling holes 62 includes a cooling fluid inlet 62a formed in the flow path wall surface 45 of the first cooling flow path section 44A and a cooling fluid outlet 62b formed in the negative pressure surface 33.

[0029] Each of the multiple second film cooling holes 64 includes a cooling fluid inlet 64a formed in the flow path wall surface 47 of the second cooling flow path section 44B and a cooling fluid outlet 64b formed in the negative pressure surface 33.

[0030] As shown in Figure 4, the cooling fluid outlet 62b of the first film cooling hole 62 is located downstream of the cooling fluid outlet 60b of the first ejection hole 60 in the combustion gas flow direction F along the negative pressure surface 33, and the cooling fluid outlet 64b of the second film cooling hole 64 is located downstream of the cooling fluid outlet 62b of the first film cooling hole 62 in the combustion gas flow direction F along the negative pressure surface 33.

[0031] Figure 5 is a schematic cross-sectional view illustrating the respective directions of extension of the first ejection hole 60, the first film cooling hole 62, and the second film cooling hole 64 in the cross-section shown in Figure 4.

[0032] In the exemplary embodiment shown in Figure 5, if we define the straight line L0 as the line indicating the direction of the thickness of the wall portion 42 forming the negative pressure surface at the center position Ps of the cooling fluid inlet 60a of the first ejection hole 60, and define the intersection point P0 as the intersection point of the straight line L0 and the negative pressure surface 33, then the center Pt of the cooling fluid outlet 60b of the first ejection hole 60 is located upstream (towards the leading edge 30) in the combustion gas flow direction F along the negative pressure surface 33 from the intersection point P0. The cooling fluid (for example, cooling air) that flows from the first cooling flow channel 44A through the cooling fluid inlet 60a to the first ejection hole 60 is ejected from the cooling fluid outlet 60b toward the upstream side in the combustion gas flow direction, thereby cooling the leading edge portion 21 of the airfoil portion 20 (so-called showerhead cooling).

[0033] Furthermore, if we define a straight line L1 as the line indicating the direction of the thickness of the wall portion 42 forming the negative pressure surface at the center position Pa of the cooling fluid inlet 62a of the first film cooling hole 62, and define the intersection point P1 as the intersection point of the straight line L1 and the negative pressure surface 33, then the center Pb of the cooling fluid outlet 62b of the first film cooling hole 62 is located downstream (towards the trailing edge 31) in the combustion gas flow direction F along the negative pressure surface 33 from the intersection point P1. The cooling fluid (for example, cooling air) that flows from the first cooling channel portion 44A to the first film cooling hole 62 via the cooling fluid inlet 62a flows out from the cooling fluid outlet 62b toward the downstream side in the combustion gas flow direction along the negative pressure surface 33, thereby cooling the film of the negative pressure surface 33.

[0034] Furthermore, if we define the straight line L2 as the line indicating the thickness direction of the negative pressure surface forming wall portion 42 at the center position Pd of the cooling fluid inlet 64a of the second film cooling hole 64, and define the intersection point P2 of the straight line L2 and the negative pressure surface 33, then the center Pe of the cooling fluid outlet 64b of the second film cooling hole 64 is located downstream (towards the trailing edge 31) in the combustion gas flow direction F along the negative pressure surface 33 from the intersection point P2. The cooling fluid (for example, cooling air) that flows from the second cooling channel portion 44B through the cooling fluid inlet 64a to the second film cooling hole 64 flows out from the cooling fluid outlet 64b toward the downstream side in the combustion gas flow direction along the negative pressure surface 33, thereby cooling the film of the negative pressure surface 33.

[0035] In some embodiments, as shown in Figure 2, for example, if the range on the tip side 20t of the airfoil 20 is defined as the tip-side range W1 with respect to a position Pm at 2 / 3 of the airfoil height H of the airfoil 20, then the number of second film cooling holes 64 belonging to the tip-side range W1 is less than the number of first film cooling holes 62 belonging to the tip-side range W1. Here, the position Pm at 2 / 3 of the airfoil height H of the airfoil 20 means the position where the distance from the base end 20h of the airfoil 20 in the airfoil height direction is 2H / 3, and the range on the tip side 20t of the airfoil 20 with respect to position Pm means the range from position Pm to the tip 20t of the airfoil 20 in the airfoil height direction. In the exemplary embodiment shown in Figure 2, the number of second film cooling holes 64 belonging to the tip-side range W1 is 0, and the number of first film cooling holes 62 belonging to the tip-side range W1 is 6, but these numbers are not particularly limited.

[0036] Furthermore, in the exemplary embodiment shown in Figure 2, the range Wb in the wing height direction where multiple second film cooling holes 64 are formed is located closer to the base end in the wing height direction than the range Wa in the wing height direction where multiple first film cooling holes 62 are formed. Therefore, no second film cooling holes 64 are formed in the range Wa in the wing height direction where multiple first film cooling holes 62 are formed, and no first film cooling holes 62 are formed in the range Wb in the wing height direction where multiple second film cooling holes 64 are formed.

[0037] Furthermore, in the exemplary embodiment shown in Figure 2, if the range on the base end 20h side of the airfoil 20 is defined as the base end range W2, with reference to position Pc, which is half the wing height H of the airfoil 20, then the number of first film cooling holes 62 belonging to the base end range W2 is less than the number of second film cooling holes 64 belonging to the base end range W2. Here, position Pc, which is half the wing height H of the airfoil 20, means the position where the distance from the base end 20h of the airfoil 20 in the wing height direction is H / 2, that is, the central position of the airfoil 20 in the wing height direction, and the range on the base end 20h side of the airfoil 20 with reference to position Pc means the range from the base end 20h of the airfoil 20 to position Pc in the wing height direction. In the exemplary configuration shown in Figure 2, the number of first film cooling holes 62 belonging to the wing base end area W2 is 0, and the number of second film cooling holes 64 belonging to the wing base end area W2 is 11, but these numbers are not particularly limited.

[0038] Here, the effects of the turbine blade 16 described above will be explained. Even if the heat-shielding coating layer 23 peels off due to an object colliding with the blade tip side range W1 of the negative pressure surface 33 near the leading edge 21 of the airfoil portion 20 of the turbine blade 16, the film cooling effect of the cooling fluid flowing out from the first film cooling holes 62, which are formed in the blade tip side range W1 of the negative pressure surface 33 downstream of the multiple first ejection holes 60 for cooling the leading edge 21 in the direction of combustion gas flow F, can suppress the progression of thinning of the airfoil portion 20. Furthermore, since the combustion gas pressure near the negative pressure surface 33 is lower than the combustion gas pressure near the pressure surface 32, the flow rate of the cooling fluid tends to increase when the first film cooling holes 62 are provided. However, by reducing the number of second film cooling holes 64 belonging to the wingtip side range W1 to less than the number of first film cooling holes 62 belonging to the wingtip side range W1, it is possible to suppress the increase in the flow rate of the cooling fluid while suppressing the progression of thinning of the airfoil section 20 caused by collisions with flying objects.

[0039] Furthermore, in the wing base side range W2 of the negative pressure surface 33 near the leading edge 21 of the airfoil portion 20, peeling of the heat-shielding coating layer 23 due to impacts from flying objects is less likely to occur compared to the wing tip side range W1. Therefore, there is little benefit in providing the first film cooling holes 62 communicating with the first cooling channel 44A in the wing base side range W2. For this reason, as described above, by reducing the number of first film cooling holes 62 belonging to the wing base side range W2 to less than the number of second film cooling holes 64 belonging to the wing base side range W2, it is possible to suppress the progression of thinning of the airfoil portion 20 due to impacts from flying objects while suppressing an increase in the flow rate of the cooling fluid.

[0040] Figure 6 is a schematic perspective view showing a modified example of the turbine blade 16. Figure 7 is a schematic perspective view showing another modified example of the turbine blade 16. Note that the above explanation relating to Figures 3 to 5 is also common to several embodiments shown in Figures 6 and 7, so redundant explanations will be omitted.

[0041] In some embodiments shown in Figures 6 and 7, the number of second film cooling holes 64 belonging to the wingtip side range W1 is less than the number of first film cooling holes 62 belonging to the wingtip side range W1, and the number of first film cooling holes 62 belonging to the wing base side range W2 is less than the number of second film cooling holes 64 belonging to the wing base side range W2. Therefore, it is possible to suppress the progression of thinning of the airfoil portion 20 caused by collisions with flying objects while suppressing an increase in the flow rate of the cooling fluid.

[0042] In some embodiments shown in Figures 6 and 7, at least a portion of the range Wa in which a plurality of first film cooling holes 62 are formed in the wing height direction overlaps with a portion of the range Wb in which a plurality of second film cooling holes 64 are formed in the wing height direction. Therefore, compared to the embodiment shown in Figure 2, the flow rate of the cooling fluid used for film cooling of the negative pressure surface 33 increases, but the effect of film cooling of the negative pressure surface 33 can be enhanced.

[0043] In the exemplary embodiment shown in FIG. 6, among the plurality of first film cooling holes 62, the position of the first film cooling hole 62 closest to the tip 20t of the airfoil portion 20 is defined as the first position A1, and the position of the first film cooling hole 62 farthest from the tip 20t of the airfoil portion 20 among the plurality of first film cooling holes 62 is defined as the second position A2. Among the plurality of second film cooling holes 64, the position of the second film cooling hole 64 closest to the tip 20t of the airfoil portion 20 is defined as the third position A3, and the position of the second film cooling hole 64 farthest from the tip 20t of the airfoil portion 20 among the plurality of second film cooling holes 64 is defined as the fourth position A4. In the airfoil height direction, the third position A3 is located between the first position A1 and the second position A2, and the second position A2 is located between the third position A3 and the fourth position A4. Further, in the exemplary embodiment shown in FIG. 6, the number of the first film cooling holes 62 provided in the airfoil portion 20 is less than the number of the second film cooling holes 64 provided in the airfoil portion 20.

[0044] When the plurality of first film cooling holes 62 are arranged at equal intervals, on the downstream side of the first film cooling hole row 63 in the flow direction F of the combustion gas along the negative pressure surface 33, since there is more margin in the metal temperature on the tip 20t side of the airfoil portion 20 than on the base end 20h side, as in the configuration shown in FIG. 6, in the airfoil height direction, the third position A3 is located between the first position A1 and the second position A2, and the second position A2 is located between the third position A3 and the fourth position A4, so that while suppressing an increase in the flow rate of the cooling fluid, it is possible to suppress the progress of the thinning of the airfoil portion 20 caused by the collision of flying objects.

[0045] In the exemplary embodiment shown in FIG. 7, the pitch E2 of the cooling fluid outlets 64b of the second film cooling holes 64 in the range Wa where the plurality of first film cooling holes 62 are formed in the airfoil height direction is larger than the pitch E1 of the cooling fluid outlets 62b of the first film cooling holes 62. Further, the pitch E2 of the cooling fluid outlets 64b of the second film cooling holes 64 in the range Wa is larger than the pitch E3 of the cooling fluid outlets 64b of the second film cooling holes 64 in the range Wc on the base end 20h side in the airfoil height direction than the range Wa. Even with such a configuration, it is possible to suppress the progress of the thinning of the airfoil portion 20 caused by the collision of flying objects while suppressing an increase in the flow rate of the cooling fluid.

[0046] FIG. 8 is a schematic perspective view showing still another modification of the turbine rotor blade 16. Since the above description related to FIGS. 3 to 5 is common to some of the embodiments shown in FIG. 8, duplicate descriptions are omitted.

[0047] In some embodiments, as shown in FIG. 8, the total area of the cooling fluid outlets 64b of the second film cooling holes 64 belonging to the blade tip side range W1 is smaller than the total area of the cooling fluid outlets 62b of the first film cooling holes 62 belonging to the blade tip side range W1. In the exemplary embodiment shown in FIG. 8, the total area of the cooling fluid outlets 62b of the first film cooling holes 62 belonging to the blade root side range W2 is smaller than the total area of the cooling fluid outlets 64b of the second film cooling holes 64 belonging to the blade root side range W2. Also, the total area of the cooling fluid outlets 64b of the second film cooling holes 64 belonging to the blade tip side range W1 is smaller than the total area of the cooling fluid outlets 64b of the second film cooling holes 64 belonging to the blade root side range W2. In the example shown in FIG. 8, since the first film cooling holes 62 are not formed in the blade root side range W2, the total area of the cooling fluid outlets 62b of the first film cooling holes 62 belonging to the blade root side range W2 is 0.

[0048] Also, in the exemplary embodiment shown in FIG. 8, the hole diameter D2a of the second film cooling holes 64 belonging to the range Wa in which a plurality of the first film cooling holes 62 are formed in the blade height direction is smaller than the hole diameter D1 of the first film cooling holes 62. Also, the hole diameter D2b of the second film cooling holes 64 existing on the root 20h side of the above range Wa in the blade height direction among the plurality of the second film cooling holes 64 is larger than the hole diameter D2a of the second film cooling holes 64 belonging to the range Wa.

[0049] Even if the heat-shielding coating layer 23 peels off due to an object colliding with the blade tip side range W1 of the negative pressure surface 33 near the leading edge 21 of the airfoil portion 20 of the turbine blade 16, the film cooling effect of the cooling fluid flowing out from the first film cooling holes 62, which are formed in the blade tip side range W1 of the negative pressure surface 33 downstream of the multiple first ejection holes 60 for cooling the leading edge 21 in the combustion gas flow direction F, can suppress the progression of thinning of the airfoil portion 20. Furthermore, since the combustion gas pressure near the negative pressure surface 33 is lower than the combustion gas pressure near the pressure surface 32, the flow rate of the cooling fluid tends to increase when the first film cooling holes 62 are provided. However, by making the total area of ​​the cooling fluid outlets 64b of the second film cooling holes 64 belonging to the wingtip side range W1 smaller than the total area of ​​the cooling fluid outlets 64b of the first film cooling holes 62 belonging to the wingtip side range W1, it is possible to suppress the increase in the flow rate of the cooling fluid while suppressing the progression of thinning of the airfoil section 20 caused by collisions with flying objects.

[0050] Furthermore, in the wing base side range W2 of the negative pressure surface 33 near the leading edge 21 of the airfoil portion 20, peeling of the heat-shielding coating layer 23 due to impacts from flying objects is less likely to occur compared to the wing tip side range W1. Therefore, there is little benefit in providing the first film cooling hole 62 communicating with the first cooling channel 44A in the wing base side range W2. For this reason, as described above, by making the total area of ​​the cooling fluid outlets 62b of the first film cooling holes 62 belonging to the wing base side range W2 smaller than the total area of ​​the cooling fluid outlets 64b of the second film cooling holes 64 belonging to the wing base side range W2, it is possible to suppress the progression of thinning of the airfoil portion 20 due to impacts from flying objects while suppressing an increase in the flow rate of the cooling fluid.

[0051] This disclosure is not limited to the embodiments described above, but also includes modified forms of the embodiments described above, as well as forms that combine these forms as appropriate.

[0052] For example, in the embodiment shown in Figure 3, etc., an airfoil section 20 in which the first cooling channel section 44A to the fifth cooling channel section 44E are formed is illustrated, but the number of cooling channel sections 44 provided in the airfoil section 20 is not particularly limited and can be two or more.

[0053] The contents described in each of the above embodiments can be understood, for example, as follows:

[0054] [1] A turbine blade according to at least one embodiment of the present disclosure is a turbine blade (e.g., a turbine blade 16) of a gas turbine (e.g., the gas turbine 2 described above), comprising: an airfoil portion (e.g., the airfoil portion 20 described above); a platform portion (e.g., the platform portion 22 described above) connected to the base end of the airfoil portion; and a blade root portion (e.g., the blade root portion 24 described above) provided on the opposite side of the airfoil portion from the platform portion, wherein the airfoil portion includes a pressure surface forming wall portion (e.g., the pressure surface forming wall portion 40 described above) that forms a pressure surface (e.g., the pressure surface 32 described above) and a negative pressure surface forming wall portion (e.g., the negative pressure surface forming wall portion 42 described above) that forms a negative pressure surface (e.g., the negative pressure surface 33 described above), wherein a plurality of cooling channel portions (e.g., the plurality of cooling channel portions 44 described above) extending along the blade height direction of the airfoil portion are formed between the pressure surface forming wall portion and the negative pressure surface forming wall portion of the airfoil portion, and the plurality of cooling channel portions are arranged along the direction from the leading edge to the trailing edge of the airfoil portion. If we define the cooling channel section closest to the leading edge among the plurality of cooling channel sections as the first cooling channel section (for example, the first cooling channel section 44A described above), and the cooling channel section adjacent to the first cooling channel section among the plurality of cooling channel sections as the second cooling channel section (for example, the second cooling channel section 44B described above), then the negative pressure surface forming wall section of the airfoil section is formed with: a plurality of first film cooling holes (for example, the plurality of first film cooling holes 62 described above), each containing a cooling fluid inlet (for example, the cooling fluid inlet 62a described above) formed in the channel wall surface of the first cooling channel section (for example, the channel wall surface 45 described above) and a cooling fluid outlet (for example, the cooling fluid outlet 62b described above) formed in the negative pressure surface of the airfoil section; and a plurality of second film cooling holes (for example, the plurality of second film cooling holes 64 described above), each containing a cooling fluid inlet (for example, the cooling fluid inlet 62a described above) formed in the channel wall surface of the second cooling channel section (for example, the channel wall surface 47 described above) and a cooling fluid outlet formed in the negative pressure surface of the airfoil section.If we define the intersection point of the line indicating the thickness direction of the wall portion forming the negative pressure surface (for example, the line L1 described above) and the negative pressure surface at the center position of the cooling fluid inlet of the first film cooling hole (for example, position Pa described above) as the first intersection point (for example, intersection point P1 described above), then the center of the cooling fluid outlet of the first film cooling hole (for example, center Pb described above) is located downstream of the first intersection point in the flow direction of the combustion gas along the negative pressure surface (for example, flow direction F described above). If we define the range on the tip side of the airfoil portion with respect to the position at 2 / 3 of the airfoil height (for example, position Pm described above) as the tip side range (for example, tip side range W1 described above), then the number of second film cooling holes belonging to the tip side range is less than the number of first film cooling holes belonging to the tip side range.

[0055] With the turbine blade described in [1] above, even if the heat-shielding coating layer peels off due to an impact of flying debris on the wingtip side of the negative pressure surface near the leading edge of the airfoil, the film cooling effect of the cooling fluid flowing out from the first film cooling holes formed in the wingtip side area can suppress the progression of thinning of the airfoil. Furthermore, since the pressure of the combustion gas near the negative pressure surface is lower than the pressure of the combustion gas near the pressure surface, the flow rate of the cooling fluid tends to increase when the first film cooling holes are provided. However, by making the number of second film cooling holes belonging to the wingtip side area less than the number of first film cooling holes belonging to the wingtip side area, it is possible to suppress the progression of thinning of the airfoil caused by the impact of flying debris while suppressing the increase in the flow rate of the cooling fluid.

[0056] [2] In some embodiments, in the turbine blade described in [1] above, if the range on the base end side of the airfoil is defined as the base end range (for example, the base end range W2) with respect to a position half the height of the airfoil (for example, the position Pc described above), then the number of second film cooling holes belonging to the base end range is greater than the number of first film cooling holes belonging to the base end range.

[0057] According to the turbine blade described in [2] above, in the base-end range near the leading edge of the airfoil, peeling of the heat-shielding coating layer due to impact of flying objects is less likely to occur compared with the tip-end range. Therefore, there is little benefit in providing the first film cooling holes communicating with the first cooling channel in the base-end range. For this reason, as described in [2] above, by making the number of first film cooling holes belonging to the base-end range less than the number of second film cooling holes belonging to the base-end range, it is possible to suppress the progression of thinning of the airfoil due to impact of flying objects while suppressing an increase in the flow rate of the cooling fluid.

[0058] [3] In some embodiments, in the turbine blade described in [1] or [2] above, the range in the blade height direction in which the plurality of second film cooling holes are formed (for example, the range Wb described above) is closer to the base end of the airfoil in the blade height direction than the range in the blade height direction in which the plurality of first film cooling holes are formed (for example, the range Wa described above).

[0059] According to the turbine blade described in [3] above, it is possible to suppress the progression of thinning of the airfoil section caused by collisions with flying objects while suppressing an increase in the flow rate of the cooling fluid.

[0060] [4] In some embodiments, in the turbine blade described in [3] above, the second film cooling holes are not formed in the range in the blade height direction in which the plurality of first film cooling holes are formed (for example, the range Wa described above).

[0061] According to the turbine blade described in [4] above, it is possible to suppress the progression of thinning of the airfoil section caused by collisions with flying objects while suppressing an increase in the flow rate of the cooling fluid.

[0062] [5] In some embodiments, in the turbine blade described in [1] above, if the position of the first film cooling hole closest to the tip of the airfoil is defined as the first position (for example, the first position A1 described above), the position of the first film cooling hole furthest from the tip of the airfoil is defined as the second position (for example, the second position A2 described above), the position of the second film cooling hole closest to the tip of the airfoil is defined as the third position (for example, the third position A3 described above), and the position of the second film cooling hole furthest from the tip of the airfoil is defined as the fourth position (for example, the fourth position A4 described above), then in the blade height direction, the third position is located between the first position and the second position, and the second position is located between the third position and the fourth position.

[0063] With the turbine blade described in [5] above, compared to the turbine blade described in [4] above, the flow rate of the cooling fluid used for film cooling on the negative pressure surface increases, but the effect of negative pressure film cooling can be enhanced.

[0064] [6] In some embodiments, in the turbine blade described in [5] above, the number of first film cooling holes is less than the number of second film cooling holes.

[0065] In the wing root end region near the leading edge of the airfoil, peeling of the heat-shielding coating layer due to impacts from flying objects is less likely to occur compared to the wing tip region. Therefore, there is little benefit in providing the first film cooling holes communicating with the first cooling channel in the wing root end region. For this reason, as described in [6] above, by reducing the number of first film cooling holes to fewer than the number of second film cooling holes, it is possible to suppress the progression of thinning of the airfoil due to impacts from flying objects while suppressing an increase in the flow rate of the cooling fluid.

[0066] [7] A turbine blade according to at least one embodiment of the present disclosure is a turbine blade (e.g., a turbine blade 16) of a gas turbine (e.g., the gas turbine 2 described above), comprising: an airfoil portion (e.g., the airfoil portion 20 described above); a platform portion (e.g., the platform portion 22 described above) connected to the base end of the airfoil portion; and a blade root portion (e.g., the blade root portion 24 described above) provided on the opposite side of the platform portion from the airfoil portion, wherein the airfoil portion includes a pressure surface forming wall portion (e.g., the pressure surface forming wall portion 40 described above) that forms a pressure surface (e.g., the pressure surface 32 described above) and a negative pressure surface forming wall portion (e.g., the negative pressure surface forming wall portion 42 described above) that forms a negative pressure surface (e.g., the negative pressure surface 33 described above), wherein a plurality of cooling channel portions (e.g., the plurality of cooling channel portions 44 described above) extending along the blade height direction of the airfoil portion are formed between the pressure surface forming wall portion and the negative pressure surface forming wall portion of the airfoil portion, and the plurality of cooling channel portions are arranged along the direction from the leading edge to the trailing edge of the airfoil portion. If we define the cooling channel section closest to the leading edge among the plurality of cooling channel sections as the first cooling channel section (for example, the first cooling channel section 44A described above), and the cooling channel section adjacent to the first cooling channel section among the plurality of cooling channel sections as the second cooling channel section (for example, the second cooling channel section 44B described above), then the negative pressure surface forming wall section of the airfoil section is formed with: a plurality of first film cooling holes (for example, the plurality of first film cooling holes 62 described above), each containing a cooling fluid inlet (for example, the cooling fluid inlet 62a described above) formed in the channel wall surface of the first cooling channel section (for example, the channel wall surface 45 described above) and a cooling fluid outlet (for example, the cooling fluid outlet 62b described above) formed in the negative pressure surface of the airfoil section; and a plurality of second film cooling holes (for example, the plurality of second film cooling holes 64 described above), each containing a cooling fluid inlet (for example, the cooling fluid inlet 62a described above) formed in the channel wall surface of the second cooling channel section (for example, the channel wall surface 47 described above) and a cooling fluid outlet formed in the negative pressure surface of the airfoil section.If we define the intersection point of the line indicating the thickness direction of the wall portion forming the negative pressure surface (e.g., the line L1 described above) and the negative pressure surface at the center position of the cooling fluid inlet of the first film cooling hole (e.g., position Pa described above) as the first intersection point (e.g., intersection point P1 described above), then the center of the cooling fluid outlet of the first film cooling hole (e.g., center Pb described above) is located downstream of the first intersection point in the combustion gas flow direction along the negative pressure surface (e.g., flow direction F described above). If we define the range on the tip side of the airfoil portion as the tip side range (e.g., tip side range W1 described above) with respect to the position 2 / 3 of the airfoil height (e.g., position Pm described above), then the total area of ​​the cooling fluid outlets of the second film cooling holes belonging to the tip side range is smaller than the total area of ​​the cooling fluid outlets of the first film cooling holes belonging to the tip side range.

[0067] With the turbine blade described in [7] above, even if the heat-shielding coating layer peels off due to an impact of flying debris on the wingtip side of the negative pressure surface near the leading edge of the airfoil, the film cooling effect of the cooling fluid flowing out from the first film cooling holes formed in the wingtip side area can suppress the progression of thinning of the airfoil. Furthermore, since the pressure of the combustion gas near the negative pressure surface is lower than the pressure of the combustion gas near the pressure surface, the flow rate of the cooling fluid tends to increase compared to the case where the first film cooling holes are not provided. However, by making the total area of ​​the cooling fluid outlets of the second film cooling holes belonging to the wingtip side area smaller than the total area of ​​the cooling fluid outlets of the first film cooling holes belonging to the wingtip side area, it is possible to suppress the increase in the flow rate of the cooling fluid while suppressing the progression of thinning of the airfoil caused by the impact of flying debris.

[0068] [8] In some embodiments, in the turbine blade described in [7] above, if the range of the airfoil on the base end side of the airfoil is defined as the base end range (for example, the base end range W2 above) with respect to a position half the height of the airfoil (for example, the position Pc above), then the total area of ​​the cooling fluid outlets of the first film cooling holes belonging to the base end range is smaller than the total area of ​​the cooling fluid outlets of the second film cooling holes belonging to the base end range.

[0069] In the wing root end region near the leading edge of the airfoil, peeling of the heat-shielding coating layer due to impacts from flying objects is less likely to occur compared to the wing tip region. Therefore, there is little benefit in providing the first film cooling holes communicating with the first cooling channel in the wing root end region. For this reason, as described in [8] above, by making the total area of ​​the cooling fluid outlets of the first film cooling holes belonging to the wing root end region smaller than the total area of ​​the cooling fluid outlets of the second film cooling holes belonging to the wing root end region, it is possible to suppress the progression of thinning of the airfoil due to impacts from flying objects while suppressing an increase in the flow rate of the cooling fluid.

[0070] [9] A gas turbine according to at least one embodiment of the present disclosure is a gas turbine comprising a compressor, a combustor, and a turbine, wherein the turbine includes turbine blades as described in any of [1] to [8] above.

[0071] According to the gas turbine described in [9] above, it is possible to suppress the progression of thinning of the airfoil section caused by collisions with flying objects while suppressing an increase in the flow rate of the cooling fluid.

[0072] 2 Gas turbine 4 Compressor 6 Combustor 8 Turbine 9 Rotor 10 Turbine casing 12 Turbine stator blade 16 Turbine rotor blade 20 Airfoil section 20h Base 20t Tip 21 Leading edge 22 Platform section 22a Outward surface 22b Inward surface 23 Heat-shielding coating layer 24 Blade root 30 Leading edge 31 Trailing edge 32 Pressure surface 33 Negative pressure surface 40 Pressure surface forming wall section 42 Negative pressure surface forming wall section 44 Cooling passage section 44A First cooling passage section 44B Second cooling passage section 44C Third cooling passage section 44D Fourth cooling passage section 44E Fifth cooling passage section 45, 47 Passage wall surface 46, 48 Inner surface 50 Partition wall section 50A First partition wall section 50B Second partition wall section 50C Third partition wall section 50D Fourth partition wall section 60 First ejection holes 60a, 62a, 64a Cooling fluid inlet 60b, 62b, 64b Cooling fluid outlet 61 First ejection hole row 62 First film cooling hole 63 First film cooling hole row 64 Second film cooling hole 65 Second film cooling hole row 68 Second ejection hole

Claims

1. A turbine blade for a gas turbine, comprising: an airfoil portion; a platform portion connected to the base end of the airfoil portion; and a blade root portion provided on the opposite side of the platform portion from the airfoil portion, wherein the airfoil portion includes a pressure surface forming wall portion that forms a pressure surface and a negative pressure surface forming wall portion that forms a negative pressure surface, and a plurality of cooling channel portions extending along the blade height direction of the airfoil portion are formed between the pressure surface forming wall portion and the negative pressure surface forming wall portion of the airfoil portion, the plurality of cooling channel portions are arranged along the direction from the leading edge to the trailing edge of the airfoil portion, and the cooling channel portion closest to the leading edge is defined as the first cooling channel portion, and the cooling channel portion adjacent to the first cooling channel portion is defined as the second cooling channel portion, wherein the negative pressure surface forming wall portion of the airfoil portion includes a plurality of first film cooling holes, each including a cooling fluid inlet formed in the channel wall surface of the first cooling channel portion and a cooling fluid outlet formed in the negative pressure surface of the airfoil portion, A turbine blade is provided, wherein a plurality of second film cooling holes are formed, each including a cooling fluid inlet formed in the flow channel wall of the second cooling channel and a cooling fluid outlet formed in the negative pressure surface of the airfoil, and the intersection point of the line indicating the thickness direction of the wall portion forming the negative pressure surface at the center position of the cooling fluid inlet of the first film cooling hole and the negative pressure surface is defined as the first intersection point, the center of the cooling fluid outlet of the first film cooling hole is located downstream of the first intersection point in the combustion gas flow direction along the negative pressure surface, and the range on the tip side of the airfoil is defined as the tip side range with respect to a position at 2 / 3 of the blade height of the airfoil, the number of second film cooling holes belonging to the tip side range is less than the number of first film cooling holes belonging to the tip side range.

2. The turbine blade according to claim 1, wherein, with reference to a position half the blade height of the airfoil, the range on the base end side of the airfoil is defined as the base end range, and the number of first film cooling holes belonging to the base end range is less than the number of second film cooling holes belonging to the base end range.

3. The turbine blade according to claim 1, wherein the range in which the plurality of second film cooling holes are formed in the blade height direction is closer to the base end of the airfoil in the blade height direction than the range in which the plurality of first film cooling holes are formed in the blade height direction.

4. The turbine blade according to claim 3, wherein the second film cooling holes are not formed in the range in the blade height direction in which the plurality of first film cooling holes are formed.

5. Defining the position of the first film cooling hole closest to the tip of the airfoil among the plurality of first film cooling holes as the first position, the position of the first film cooling hole furthest from the tip of the airfoil among the plurality of first film cooling holes as the second position, the position of the second film cooling hole closest to the tip of the airfoil among the plurality of second film cooling holes as the third position, and the position of the second film cooling hole furthest from the tip of the airfoil among the plurality of second film cooling holes as the fourth position, the turbine blade according to claim 1, wherein in the blade height direction, the third position is located between the first position and the second position, and the second position is located between the third position and the fourth position.

6. The turbine blade according to claim 5, wherein the number of first film cooling holes is less than the number of second film cooling holes.

7. A turbine blade for a gas turbine, comprising: an airfoil portion; a platform portion connected to the base end of the airfoil portion; and a blade root portion provided on the opposite side of the platform portion from the airfoil portion, wherein the airfoil portion includes a pressure surface forming wall portion that forms a pressure surface and a negative pressure surface forming wall portion that forms a negative pressure surface, and a plurality of cooling passage portions extending along the blade height direction of the airfoil portion are formed between the pressure surface forming wall portion and the negative pressure surface forming wall portion of the airfoil portion, the plurality of cooling passage portions are arranged along the direction from the leading edge to the trailing edge of the airfoil portion, and the cooling passage portion closest to the leading edge is defined as the first cooling passage portion, and the cooling passage portion adjacent to the first cooling passage portion is defined as the second cooling passage portion, the negative pressure surface forming wall portion of the airfoil portion includes a plurality of first film cooling holes, each including a cooling fluid inlet formed in the passage wall surface of the first cooling passage portion and a cooling fluid outlet formed in the negative pressure surface of the airfoil portion, A turbine blade is provided, wherein a plurality of second film cooling holes are formed, each including a cooling fluid inlet formed in the flow channel wall of the second cooling channel and a cooling fluid outlet formed in the negative pressure surface of the airfoil, and the intersection point of the line indicating the thickness direction of the wall portion forming the negative pressure surface at the center position of the cooling fluid inlet of the first film cooling hole and the negative pressure surface is defined as the first intersection point, the center of the cooling fluid outlet of the first film cooling hole is located downstream of the first intersection point in the combustion gas flow direction along the negative pressure surface, and the range on the tip side of the airfoil is defined as the tip side range with respect to a position at 2 / 3 of the blade height of the airfoil, the sum of the areas of the cooling fluid outlets of the second film cooling holes belonging to the tip side range is smaller than the sum of the areas of the cooling fluid outlets of the first film cooling holes belonging to the tip side range.

8. The turbine blade according to claim 7, wherein, with reference to a position half the blade height of the airfoil, the range on the base end side of the airfoil is defined as the base end range, and the total area of ​​the cooling fluid outlets of the first film cooling holes belonging to the base end range is smaller than the total area of ​​the cooling fluid outlets of the second film cooling holes belonging to the base end range.

9. A gas turbine comprising a compressor, a combustor, and a turbine, wherein the turbine includes turbine blades as described in any one of claims 1 to 8.

Images