Turbine blades and gas turbines

Abstract

The present invention provides a turbine blade capable of suppressing the progression of airfoil thinning caused by collisions with flying objects while suppressing an increase in the flow rate of the cooling fluid. [Solution] The turbine blade of a gas turbine has a negative pressure surface forming wall portion which has a plurality of first film cooling holes, each containing a cooling fluid inlet formed in the flow channel wall surface of a first cooling channel portion, and a plurality of second film cooling holes, each containing a cooling fluid inlet formed in the flow channel wall surface of a second cooling channel portion adjacent to the first cooling channel portion. If the range on the tip side of the airfoil portion is defined as the tip side range with respect to the central position of the airfoil portion in the blade height direction, 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.

Description

Technical Field

[0001] The present disclosure relates to a turbine rotor blade and a gas turbine.

Background Art

[0002] Patent Document 1 discloses a structure for a turbine rotor blade of a gas turbine, aiming to enhance the durability of the turbine rotor blade while suppressing the amount of cooling air used. In this turbine rotor blade, a plurality of cooling channel 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 channel portion closest to the leading edge of the airfoil portion among the plurality of cooling channel 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 channel portion second closest to the leading edge of the airfoil portion among the plurality of cooling channel portions to flow out from the base end side of the airfoil portion.

Prior Art Documents

Patent Documents

[0003]

Patent Document 1

Summary of the Invention

Problems to be Solved by the Invention

[0004] By the way, as a result of the intensive study by the inventor of the present application, 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 collision 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. In this regard, Patent Document 1 does not disclose any findings for achieving both suppressing the progress of the thinning of the airfoil portion caused by the collision of such flying objects and suppressing the increase in the flow rate of the cooling fluid.

[0005] In view of the circumstances described above, at least one embodiment of the present disclosure aims to provide a turbine blade for a gas turbine and a gas turbine equipped therewith that can suppress the progression of thinning of the airfoil portion caused by collisions with flying objects while suppressing an increase in the flow rate of the cooling fluid. [Means for solving the problem]

[0006] To achieve the above objective, the turbine blades according to at least one embodiment of this disclosure are: A turbine blade for a gas turbine, Wing section and, A platform portion connected to the base end of the airfoil portion, A wing root portion is provided on the opposite side of the airfoil portion from the platform portion, Equipped with, 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. Between the pressure surface forming wall and the negative pressure surface forming wall in the airfoil portion, a plurality of cooling channel portions are formed that extend along the airfoil height direction of the airfoil portion. The plurality of cooling channel sections are arranged along the direction from the leading edge to the trailing edge of the airfoil section, If we define the cooling channel portion closest to the leading edge among the plurality of cooling channel portions as the first cooling channel portion, and the cooling channel portion adjacent to the first cooling channel portion among the plurality of cooling channel portions as the second cooling channel portion, The negative pressure surface forming wall portion of the airfoil portion, A plurality of first film cooling holes, 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 portion, A plurality of second film cooling holes, each including a cooling fluid inlet formed on the wall surface of the flow path section of the second cooling flow path section and a cooling fluid outlet formed on the negative pressure surface of the airfoil section, A structure has been formed, If the intersection point of the line indicating the thickness direction of the wall portion forming the negative pressure surface at the center 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 direction of combustion gas flow along the negative pressure surface, If the area 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 this disclosure is: A gas turbine comprising a compressor, a combustor, and a turbine, The turbine includes the turbine blades. [Effects of the Invention]

[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. [Brief explanation of the drawing]

[0009] [Figure 1] This figure shows a schematic configuration of a gas turbine 2 according to one embodiment. [Figure 2] This is a schematic perspective view showing an example of a turbine blade 16. [Figure 3] This figure shows an example of a cross-section perpendicular to the wing height direction in the airfoil section 20. [Figure 4] Figure 3 is a schematic cross-sectional view showing an enlarged view of the vicinity of the leading edge 21 of the airfoil section 20. [Figure 5] Figure 4 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. [Figure 6]It is a schematic perspective view showing a modified example of the turbine rotor blade 16. [Figure 7] It is a schematic perspective view showing another modified example of the turbine rotor blade 16. [Figure 8] It is a schematic perspective view showing still another modified example of the turbine rotor blade 16.

Mode for Carrying Out the Invention

[0010] Hereinafter, some embodiments of the present disclosure will be described with reference to the accompanying drawings. However, the dimensions, materials, shapes, relative arrangements, etc. of the components described as embodiments or shown in the drawings are not intended to limit the scope of the invention thereto, but are merely illustrative examples. For example, expressions representing relative or absolute arrangements such as "in a certain direction", "along a certain direction", "parallel", "orthogonal", "center", "concentric", or "coaxial" not only strictly represent such arrangements, but also represent a state in which there are tolerances or relative displacements with an angle or distance that can obtain the same function. For example, expressions representing that things such as "identical", "equal", and "homogeneous" are in an equal state not only strictly represent an equal state, but also represent a state in which there are tolerances or differences that can obtain the same function. For example, expressions representing shapes such as a rectangular shape or a cylindrical shape not only represent shapes such as a rectangular shape or a cylindrical shape in a geometrically strict sense, but also represent shapes including concave and convex portions, chamfered portions, etc. within a range where the same effect can be obtained. On the other hand, the expressions "comprising", "having", "including", or "possessing" one component are not exclusive expressions excluding the existence of other components.

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

[0012] As shown in FIG. 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 implanted in the rotor 9 so as to be alternately arranged axially with respect to the turbine stator blades 12. Hereinafter, unless otherwise specified, the "circumferential direction" means the circumferential direction of the gas turbine 2, that is, the circumferential direction of the rotor 9, the "axial direction" means the axial direction of the gas turbine 2, that is, the axial direction of the rotor 9, and the "radial direction" means the radial direction of the gas turbine 2, that is, the radial direction of the rotor 9.

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

[0014] The airfoil portion 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 suction surface 33 connecting the leading edge 30 and the trailing edge 31. In the following description, the "blade height direction" means the blade height direction of the airfoil portion 20 (assuming that the blade height of the airfoil portion 20, that is, the distance from the base end 20h of the airfoil portion 20 to the tip end 20t of the airfoil portion 20 is H, the direction along the blade height H), the "base end 20h side of the airfoil portion 20" means the base end 20h side of the airfoil portion 20 in the blade height direction (the platform portion 22 side, that is, the hub side), and the "tip end 20t side of the airfoil portion 20" means the tip end 20t side of the airfoil portion 20 in the blade height direction (the side opposite to the platform portion 22, that is, the tip side). The blade height direction of the airfoil portion 20 may be the radial direction.

[0015] The platform portion 22 has a plate shape and includes an outward-facing surface 22a facing the outside in the radial direction and an inward-facing surface 22b facing the inside in the radial direction. The outward-facing surface 22a constitutes the flow path wall surface 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 24 is located on the opposite side of the airfoil 20 from the platform 22, and the radial outer end of the blade root 24 connects to the inward-facing surface 22b of the platform 22. The blade root 24 is mounted on the rotor 9 (see Figure 1). The airfoil 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 perpendicular to the wing height direction in the airfoil section 20. 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. Furthermore, 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. Furthermore, 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 section 50A is the partition wall section 50 closest to the front edge 30 among the multiple partition wall sections 50, and the second partition wall section 50B is the partition wall section 50 adjacent to the first partition wall section 50A among the multiple partition wall sections 50. Furthermore, the third partition wall section 50C is located between the second partition wall section 50B and the fourth partition wall section 50D, and the fourth partition wall section 50D is the partition wall section 50 closest to the rear edge 31 among the multiple partition wall sections 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 section 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 (all first ejection holes 60 constituting the first ejection hole row 61) provided by the airfoil section 20 unless otherwise specified, "multiple first film cooling holes 62" means all first film cooling holes 62 (all first film cooling holes 62 constituting the first film cooling hole row 63) provided by the airfoil section 20 unless otherwise specified, and "multiple second film cooling holes 64" means all second film cooling holes 64 (all second film cooling holes 64 constituting the second film cooling hole row 65) provided by the airfoil section 20 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 configuration 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 (e.g., 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 the 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 (e.g., 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 direction of the thickness of the wall portion 42 forming the negative pressure surface 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 (e.g., 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 position Pm, which is 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, position Pm, which is 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 configuration 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, we will explain the effects of the turbine blades 16 described above. 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 reducing the number of second film cooling holes 64 belonging to the blade tip side range W1 to fewer than the number of first film cooling holes 62 belonging to the blade tip side range W1, it is possible to suppress the progression of thinning of the airfoil portion 20 caused by the collision of an object while suppressing the increase in the flow rate of the cooling fluid.

[0039] Furthermore, in the wing base side range W2 of the negative pressure surface 33 near the leading edge 21 of the airfoil section 20, peeling of the heat-shielding coating layer 23 due to impacts from flying objects is less likely to occur compared with 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 section 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 are 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 section 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 configuration shown in Figure 6, if we define the position of the first film cooling hole 62 closest to the tip 20t of the airfoil portion 20 as the first position A1, the position of the first film cooling hole 62 furthest from the tip 20t of the airfoil portion 20 as the second position A2, the position of the second film cooling hole 64 closest to the tip 20t of the airfoil portion 20 as the third position A3, and the position of the second film cooling hole 64 furthest from the tip 20t of the airfoil portion 20 as the fourth position A4, then 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. Furthermore, in the exemplary embodiment shown in Figure 6, the number of first film cooling holes 62 provided in the airfoil portion 20 is less than the number of second film cooling holes 64 provided in the airfoil portion 20.

[0044] When multiple first film cooling holes 62 are arranged at equal intervals, downstream of the row of first film cooling holes 63 in the combustion gas flow direction F along the negative pressure surface 33, the tip 20t side of the airfoil 20 has more margin in metal temperature than the base 20h side. Therefore, as shown in the configuration in Figure 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. This suppresses the increase in the flow rate of the cooling fluid while suppressing the progression of thinning of the airfoil 20 caused by collisions with flying objects.

[0045] In the exemplary embodiment shown in Figure 7, the pitch E2 of the cooling fluid outlet 64b of the second film cooling hole 64 in the range Wa where multiple first film cooling holes 62 are formed in the wing height direction is greater than the pitch E1 of the cooling fluid outlet 62b of the first film cooling hole 62. Furthermore, the pitch E2 of the cooling fluid outlet 64b of the second film cooling hole 64 in range Wa is greater than the pitch E3 of the cooling fluid outlet 64b of the second film cooling hole 64 in range Wc on the base end 20h side in the wing height direction than range Wa. This configuration also suppresses the progression of thinning of the airfoil section 20 caused by collisions with flying objects while suppressing an increase in the flow rate of the cooling fluid.

[0046] Figure 8 is a schematic perspective view showing yet 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 Figure 8, so redundant explanations will be omitted.

[0047] In some embodiments, as shown in Figure 8, the total area of ​​the cooling fluid outlets 64b of the second film cooling holes 64 belonging to the wingtip 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 wingtip side range W1. In the exemplary embodiment shown in Figure 8, the total area of ​​the cooling fluid outlets 62b of the first film cooling holes 62 belonging to the wing base 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 wing base side range W2. Also, the total area of ​​the cooling fluid outlets 64b of the second film cooling holes 64 belonging to the wingtip 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 wing base side range W2. Note that in the example shown in Figure 8, since no first film cooling holes 62 are formed in the wing base side range W2, the total area of ​​the cooling fluid outlets 62b of the first film cooling holes 62 belonging to the wing base side range W2 is 0.

[0048] Furthermore, in the exemplary embodiment shown in Figure 8, the diameter D2a of the second film cooling holes 64 that belong to the range Wa in the wing height direction where multiple first film cooling holes 62 are formed is smaller than the diameter D1 of the first film cooling holes 62. Also, among the multiple second film cooling holes 64, the diameter D2b of the second film cooling holes 64 that are located on the base end 20h side of the range Wa in the wing height direction is larger than the diameter D2a of the second film cooling holes 64 that belong 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 section 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 section 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 (for example, the turbine blade 16) of a gas turbine (for example, the gas turbine 2 described above), The airfoil section (for example, the airfoil section 20 described above) and A platform portion (for example, the platform portion 22 described above) connected to the base end of the airfoil portion, A wing root portion (for example, the wing root portion 24 described above) is provided on the opposite side of the airfoil portion from the platform portion, Equipped with, The airfoil portion includes a pressure surface forming wall portion (e.g., the pressure surface forming wall portion 40) 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) that forms a negative pressure surface (e.g., the negative pressure surface 33 described above). Between the pressure surface forming wall and the negative pressure surface forming wall in the airfoil portion, a plurality of cooling channel portions (for example, the plurality of cooling channel portions 44 described above) are formed, extending along the height direction of the airfoil portion. The plurality of cooling channel sections are arranged along the direction from the leading edge to the trailing edge of the airfoil section, 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), The negative pressure surface forming wall portion of the airfoil portion, A plurality of first film cooling holes (for example, the plurality of first film cooling holes 62 described above) each include a cooling fluid inlet (for example, the cooling fluid inlet 62a described above) formed in the flow channel wall surface (for example, the flow channel wall surface 45 described above) of the first cooling channel section, and a cooling fluid outlet (for example, the cooling fluid outlet 62b described above) formed in the negative pressure surface of the airfoil section, A plurality of second film cooling holes (for example, a plurality of second film cooling holes 64) each include a cooling fluid inlet (for example, a cooling fluid inlet 62a) formed in the flow channel wall surface (for example, the flow channel wall surface 47) of the second cooling channel section, and a cooling fluid outlet formed on the negative pressure surface of the airfoil section, A structure has been formed, If we define the intersection point of the line indicating the thickness direction of the wall portion forming the negative pressure surface (e.g., line L1) and the negative pressure surface at the position of the center of the cooling fluid inlet of the first film cooling hole (e.g., position Pa) as the first intersection point (e.g., intersection point P1), then the center of the cooling fluid outlet of the first film cooling hole (e.g., center Pb) is located downstream of the first intersection point in the flow direction of the combustion gas along the negative pressure surface (e.g., flow direction F), If we define the area on the tip side of the airfoil as the tip side range (for example, the tip side range W1) with respect to a position at 2 / 3 of the airfoil height (for example, the position Pm 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] According to 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 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 less than the number of first film cooling holes belonging to the wingtip side, 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.

[0056] [2] In some embodiments, in the turbine blade described in [1] above, If we define the range on the base end side of the airfoil 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 region 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 region. Therefore, there is little benefit in providing the first film cooling holes communicating with the first cooling channel in the base-end region. For this reason, as described in [2] above, by reducing the number of first film cooling holes belonging to the base-end region to less than the number of second film cooling holes belonging to the base-end region, 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 wing height direction where 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 wing height direction than the range in the wing height direction where the plurality of first film cooling holes are formed (for example, the range Wa described above).

[0059] The turbine blade described in [3] above can suppress the increase in the flow rate of the cooling fluid while suppressing the progression of thinning of the airfoil section caused by collisions with flying objects.

[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 wing height direction where the plurality of first film cooling holes are formed (for example, the range Wa described above).

[0061] The turbine blade described in [4] above can suppress the increase in the flow rate of the cooling fluid while suppressing the progression of thinning of the airfoil section caused by collisions with flying objects.

[0062] [5] In some embodiments, the turbine blade described in [1] above, If we define the position of the first film cooling hole closest to the tip of the airfoil 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 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 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 as the fourth position (for example, the fourth position A4 described above), then, In the wing 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 is increased, but the effect of negative pressure film cooling can be enhanced.

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

[0065] In the wing root end region near the leading edge of the airfoil, delamination of the heat-shielding coating layer due to impacts from flying objects is less likely to occur compared with 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 (for example, the turbine blade 16) of a gas turbine (for example, the gas turbine 2 described above), The airfoil section (for example, the airfoil section 20 described above) and A platform portion (for example, the platform portion 22 described above) connected to the base end of the airfoil portion, A wing root portion (for example, the wing root portion 24 described above) is provided on the opposite side of the airfoil portion from the platform portion, Equipped with, The airfoil portion includes a pressure surface forming wall portion (e.g., the pressure surface forming wall portion 40) 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) that forms a negative pressure surface (e.g., the negative pressure surface 33 described above). Between the pressure surface forming wall and the negative pressure surface forming wall in the airfoil portion, a plurality of cooling channel portions (for example, the plurality of cooling channel portions 44 described above) are formed, extending along the height direction of the airfoil portion. The plurality of cooling channel sections are arranged along the direction from the leading edge to the trailing edge of the airfoil section, 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), The negative pressure surface forming wall portion of the airfoil portion, A plurality of first film cooling holes (for example, the plurality of first film cooling holes 62 described above) each include a cooling fluid inlet (for example, the cooling fluid inlet 62a described above) formed in the flow channel wall surface (for example, the flow channel wall surface 45 described above) of the first cooling channel section, and a cooling fluid outlet (for example, the cooling fluid outlet 62b described above) formed in the negative pressure surface of the airfoil section, A plurality of second film cooling holes (for example, a plurality of second film cooling holes 64) each include a cooling fluid inlet (for example, a cooling fluid inlet 62a) formed in the flow channel wall surface (for example, the flow channel wall surface 47) of the second cooling channel section, and a cooling fluid outlet formed on the negative pressure surface of the airfoil section, A structure has been formed, If we define the intersection point of the line indicating the thickness direction of the wall portion forming the negative pressure surface (e.g., line L1) and the negative pressure surface at the position of the center of the cooling fluid inlet of the first film cooling hole (e.g., position Pa) as the first intersection point (e.g., intersection point P1), then the center of the cooling fluid outlet of the first film cooling hole (e.g., center Pb) is located downstream of the first intersection point in the flow direction of the combustion gas along the negative pressure surface (e.g., flow direction F), If we define the area on the tip side of the airfoil as the tip side range (for example, the tip side range W1) with respect to a position at 2 / 3 of the airfoil height (for example, the 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 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 smaller than the total area of ​​the cooling fluid outlets of the first film cooling holes belonging to the wingtip side, 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, the turbine blade described in [7] above, If we define the range of the airfoil on the base end side of the airfoil as the base end range (for example, the base end range W2 described above) with respect to a position half the height of the airfoil (for example, the position Pc described 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, 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. [Explanation of symbols]

[0072] 2 Gas Turbines 4. Compressor 6 Combustor 8 turbines 9 rotors 10 Turbine casing 12 Turbine stator blades 16 Turbine blades 20 Airfoil 20h proximal end 20t tip 21 Front edge 22 Platform Section 22a outward facing surface 22b Inward-facing surface 23 Heat-shielding coating layer 24 Wing root 30 Anterior edge 31 Trailing edge 32 Pressure surface 33 Suction surface 40 Pressure surface forming wall 42 Negative pressure surface forming wall 44 Cooling channel section 44A First cooling channel section 44B Second Cooling Channel Section 44C Third Cooling Channel Section 44D Fourth Cooling Channel Section 44E Fifth Cooling Channel Section 45,47 Channel wall 46,48 Inner 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 1st vent 60a, 62a, 64a Cooling fluid inlet 60b,62b,64b Cooling fluid outlet 61 1st nozzle row 62 First film cooling hole 63 First film cooling hole row 64 Second film cooling holes 65 Second film cooling hole row 68 2nd vent

Claims

1. A turbine blade for a gas turbine, Wing section and, A platform portion connected to the base end of the airfoil portion, A wing root portion is provided on the opposite side of the airfoil portion from the platform portion, Equipped with, 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. Between the pressure surface forming wall and the negative pressure surface forming wall in the airfoil portion, a plurality of cooling channel portions are formed that extend along the height direction of the airfoil portion. The plurality of cooling channel sections are arranged along the direction from the leading edge to the trailing edge of the airfoil section, 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, 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, The negative pressure surface forming wall portion of the airfoil portion, A plurality of first film cooling holes, 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 portion, A plurality of second film cooling holes, each including a cooling fluid inlet formed on the wall surface of the flow path section of the second cooling flow path section and a cooling fluid outlet formed on the negative pressure surface of the airfoil section, A structure has been formed, If we define the intersection point of the line indicating the thickness direction of the negative pressure surface forming wall portion at the center of the cooling fluid inlet of the first film cooling hole and the negative pressure surface 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, A turbine blade in which, with respect to a position 2 / 3 of the blade height of the airfoil, the area on the tip side of the airfoil is defined as the tip side range, and 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 the blade height direction in which the plurality of second film cooling holes are formed 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.

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. If we define 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, Wing section and, A platform portion connected to the base end of the airfoil portion, A wing root portion is provided on the opposite side of the airfoil portion from the platform portion, Equipped with, 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. Between the pressure surface forming wall and the negative pressure surface forming wall in the airfoil portion, a plurality of cooling channel portions are formed that extend along the height direction of the airfoil portion. The plurality of cooling channel sections are arranged along the direction from the leading edge to the trailing edge of the airfoil section, 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, 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, The negative pressure surface forming wall portion of the airfoil portion, A plurality of first film cooling holes, 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 portion, A plurality of second film cooling holes, each including a cooling fluid inlet formed on the wall surface of the flow path section of the second cooling flow path section and a cooling fluid outlet formed on the negative pressure surface of the airfoil section, A structure has been formed, If we define the intersection point of the line indicating the thickness direction of the negative pressure surface forming wall portion at the center of the cooling fluid inlet of the first film cooling hole and the negative pressure surface 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, A turbine blade in which, if the area 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 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.

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, The turbine is a gas turbine comprising turbine blades according to any one of claims 1 to 8.

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