Turbine blade and gas turbine

Abstract

A turbine blade according to at least one embodiment of the present disclosure comprises a plurality of pressure-side turbulators and a plurality of suction-side turbulators aligned at intervals in the blade height direction. Film cooling holes open at a first inner wall surface, which is either a pressure-side inner wall surface or a suction-side inner wall surface. Among the pressure-side turbulators or suction-side turbulators, first turbulators provided to the first inner wall surface and second turbulators provided to a second inner wall surface are configured such that a value (P1 / e1) obtained by dividing a first arrangement pitch P1 of the first turbulators in the blade height direction by a first height e1 of the first turbulators from the first inner wall surface is greater than a value (P2 / e2) obtained by dividing a second arrangement pitch P2 of the second turbulators in the blade height direction by a second height e2 of the second turbulators from the second inner wall surface.

Description

Turbine blade and gas turbine

[0001] The present disclosure relates to a turbine blade and a gas turbine. The present application claims priority based on Japanese Patent Application No. 2025-047830 filed with the Japan Patent Office on March 24, 2025, the content of which is incorporated herein by reference.

[0002] In turbine blades for gas turbines and the like, it is known to cool a turbine blade exposed to high-temperature gas flow by causing a cooling fluid to flow through a cooling passage formed inside the turbine blade. A rib-shaped turbulator may be provided on the inner wall surface of such a cooling passage in order to promote turbulence of the flow of the cooling fluid in the cooling passage and improve the heat transfer coefficient between the cooling fluid and the turbine blade.

[0003] For example, Patent Document 1 discloses a turbine blade in which a plurality of turbulators are provided along the flow direction of a cooling fluid on the inner wall surface of a cooling passage extending along the blade height direction.

[0004] Japanese Unexamined Patent Publication No. 2004-225690

[0005] For example, in the turbine blade disclosed in Patent Document 1, it is conceivable to further cool the airfoil portion with film air from film cooling holes provided in the blade wall. In this case, when the film cooling holes are provided only in either one of the ventral blade wall portion and the dorsal blade wall portion that define a certain cooling passage, cooling is promoted in the blade wall portion provided with the film cooling holes. Therefore, if the temperature of the blade wall portion provided with the film cooling holes becomes lower than the temperature of the blade wall portion not provided with the film cooling holes, it is conceivable that tensile stress acts on the blade wall portion provided with the film cooling holes, and compressive stress acts on the blade wall portion not provided with the film cooling holes. In such a case, the tensile stress may become excessive, which may shorten the service life of the turbine blade.

[0006] In view of the above circumstances, at least one embodiment of the present disclosure aims to reduce thermal stress acting on a turbine blade.

[0007] (1) A turbine blade according to at least one embodiment of the present disclosure is a turbine blade for a gas turbine, comprising: an airfoil having at least one cooling passage extending in the blade height direction and a film cooling hole communicating with the at least one cooling passage; a plurality of ventral turbulators formed on the ventral inner wall surface of the airfoil defining the at least one cooling passage and arranged at intervals in the blade height direction; and a plurality of dorsal turbulators formed on the dorsal inner wall surface and arranged at intervals in the blade height direction, wherein the film cooling hole opens to a first inner wall surface which is either the ventral inner wall surface or the dorsal inner wall surface, and of the ventral turbulators or dorsal turbulators, a first turbulator provided on the first inner wall surface and a second turbulator provided on the second inner wall surface which is the other of the ventral inner wall surface or the dorsal inner wall surface, As a first condition, the value obtained by dividing the first arrangement pitch P1 in the wing height direction of the first turbulator by the first height e1 from the first inner wall surface of the first turbulator (P1 / e1) is greater than the value obtained by dividing the second arrangement pitch P2 in the wing height direction of the second turbulator by the second height e2 from the second inner wall surface of the second turbulator (P2 / e2), or as a second condition, the first turbulator has a first region in which the first height e1 does not change as it moves from the leading edge to the trailing edge of the airfoil portion of the first turbulator, and a second region connected to the trailing edge end of the first region in which the first height e1 gradually decreases as it moves from the leading edge to the trailing edge. At least one of these conditions is satisfied.

[0008] (2) A gas turbine according to at least one embodiment of the present disclosure comprises a turbine blade having the configuration of (1) above, and a combustor for generating combustion gas flowing through a combustion gas passage on which the turbine blade is provided.

[0009] According to at least one embodiment of the present disclosure, the thermal stress acting on the turbine blade can be reduced.

[0010] These are schematic diagrams showing partial cross-sectional structures of gas turbines according to several embodiments. This is a cross-sectional view of the airfoil along the II-II section of Figure 3. This is a cross-sectional view of the turbine blade along the III-III section of Figure 2. This is a diagram showing an example of the trailing edge region of the airfoil in the III-III cross-sectional view of the turbine blade in Figure 2. This is a diagram showing another example of the trailing edge region of the airfoil in the III-III cross-sectional view of the turbine blade in Figure 2. This is a schematic diagram of an example of the V-V section in Figures 4A and 4B. This is a schematic diagram of another example of the V-V section in Figures 4A and 4B.

[0011] 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 this disclosure, 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 describe such arrangements strictly, but also represent states of relative displacement with tolerances, or angles or distances to the extent that the same function is achieved. For example, expressions describing things being in an equal state such as "identical," "equal," and "homogeneous" should not only describe states of being strictly equal, but also represent states where tolerances, or differences to the extent that the same function is achieved, exist. For example, expressions describing shapes such as square shapes or cylindrical shapes should not only describe geometrically precise square shapes or cylindrical shapes, but also represent shapes including concave and concave parts, chamfered parts, etc., to the extent that the same effect is 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.

[0012] <Overview of Gas Turbine> Figure 1 is a schematic diagram showing a partial cross-sectional structure of a gas turbine 6 according to several embodiments. This gas turbine 6 comprises a compressor 91 and a turbine 92 directly connected to each other. The compressor 91 is configured as, for example, an axial flow compressor, and draws in air or a predetermined gas as a working fluid from an intake port and pressurizes it. A combustor 8 is connected to the discharge port of the compressor 91, and the working fluid discharged from the compressor 91 is heated by the combustor 8 to a predetermined turbine inlet temperature. The working fluid (combustion gas) heated to the predetermined temperature is then supplied to the turbine 92. As shown in Figure 1, multiple stages of turbine stator blades 5 are provided inside the casing of the turbine 92, and turbine rotor blades 3 are attached to the rotor 64 so as to form a set of stages with each stator blade 5. One end of the rotor 64 is connected to the rotating shaft 65 of the compressor 91, and the other end is connected to the rotating shaft of a generator (not shown).

[0013] With this configuration, when high-temperature, high-pressure combustion gas is supplied from the combustor 8 into the casing of the turbine 92, the combustion gas expands within the casing, causing the rotor 64 to rotate and drive a generator (not shown) connected to the gas turbine 6. Specifically, the pressure is reduced by each stationary blade 5 fixed to the casing, and the resulting kinetic energy is converted into rotational torque via each rotor blade 3 attached to the rotor 64. This generated rotational torque is then transmitted to the rotor 64, driving the generator.

[0014] <Overview of Turbine Blades> Turbine blades according to several embodiments of the present disclosure are shown in Figures 2 to 4B. Figure 2 is a cross-sectional view of the airfoil along the section II-II in Figure 3, and Figure 3 is a cross-sectional view of the turbine blade of Figure 2 along the section III-III. Figure 4A is a diagram showing an example of the trailing edge region of the airfoil in the section III-III of the turbine blade of Figure 2. Figure 4B is a diagram showing another example of the trailing edge region of the airfoil in the section III-III of the turbine blade of Figure 2.

[0015] In some embodiments, the turbine blade 3 comprises an airfoil section 31, a platform 32, and a blade root 33. The blade root 33 is embedded in the rotor 64 of the gas turbine 6, and the turbine blade 3 rotates together with the rotor 64. The platform 32 is integrally formed with the blade root 33.

[0016] In some embodiments of the turbine blade 3, as shown in Figure 2, there is a meandering passage (leading edge meandering passage 21) that extends meanderingly from the central portion of the blade toward the leading edge 34, and a meandering passage (trailing edge meandering passage 22) that extends meanderingly from the central portion of the blade toward the trailing edge 35. In some embodiments of the turbine blade 3, the leading edge meandering passage 21 and the trailing edge meandering passage 22 are independent passages. In some embodiments of the turbine blade 3, there is a plurality of cooling passages 41 to 46 that extend in the blade height direction and constitute the leading edge meandering passage 21 and the trailing edge meandering passage 22. The plurality of cooling passages 41 to 46 include, for example, six cooling passages 41 to 46 that are provided in order from the leading edge 34 side.

[0017] In the following description, the six cooling passages 41 to 46, which are provided in order from the leading edge 34, will also be referred to as the first passage 41, the second passage 42, the third passage 43, the fourth passage 44, the fifth passage 45, and the sixth passage 46, in order from the leading edge 34.

[0018] A cooling passage 48, for example, which is provided with a number of pin fins 49, is connected to the sixth passage 46, which is formed on the trailing edge 35 side. In some embodiments of the turbine blade 3, there may be a cooling passage (not shown) which is a passage independent of the leading edge meandering passage 21 and the trailing edge meandering passage 22, extends in the blade height direction, and is formed on the leading edge 34 side of the first passage 41.

[0019] In some embodiments of the turbine blade 3, a plurality of ventral turbulators 53 (see Figures 4A and 4B) are formed on the ventral side inner wall surface 51p of the airfoil portion 31 defining each cooling passage 41 to 46, which is located on the ventral side 38, in order to improve the heat transfer coefficient between the cooling fluid flowing through each cooling passage 41 to 46 and the ventral side inner wall surface 51p, which is located on the ventral side 39, which is located on the ventral side inner wall surface 51s of the airfoil portion 31 defining each cooling passage 41 to 46, which is located on the ventral side 39, which is located on the ventral side inner wall surface 51s, which is located on the ventral side 39, in order to improve the heat transfer coefficient between the cooling fluid flowing through each cooling passage 41 to 46 and the ventral side inner wall surface 51s, which is located on the ventral side turbulators 55 (see Figures 4A and 4B) which is located on the ventral side 39, which is located on the ventral side inner wall surface 51s, which is located on the ventral side 39, which is located on the ventral side turbulators 55 (see Figures 4A and 4B) Note that the fifth passage 45 in Figures 4A and 4B, and the ventral turbulator 53 and dorsal turbulator 55 in Figures 2 and 3 are omitted from the description.

[0020] The turbine blade 3 shown in Figure 2 has a plurality of cooling holes 41b that open in the blade wall surface 31s near the leading edge 34, and a plurality of cooling holes 46b that open in the blade wall surface 31s on the belly side 38 on the trailing edge 35 side, as film cooling holes for blowing out film cooling air (see Figure 3). For example, the plurality of cooling holes 41b are connected to the first passage 41, and the plurality of cooling holes 46b are connected to the sixth passage 46.

[0021] In some embodiments of the turbine blade 3, the first passage 41, the second passage 42, and the third passage 43, which are provided in order from the leading edge 34 side, are sequentially connected to form a meandering passage (leading edge meandering passage 21) that extends meanderingly from the central part of the blade toward the leading edge 34. In addition, the fourth passage 44, the fifth passage 45, and the sixth passage 46 are sequentially connected toward the trailing edge 35 to form a meandering passage (trailing edge meandering passage 22).

[0022] The third passage 43, which constitutes the leading edge meandering channel 21, has an opening 43a at one end (inlet side) that is formed at the base of the wing, i.e., the bottom 33a of the wing root 33. Similarly, the fourth passage 44, which constitutes the trailing edge meandering channel 22, has an opening 44a at one end (inlet side) that is formed at the bottom 33a of the wing root 33.

[0023] The second passage 42 extends in the wing height direction and is formed closer to the leading edge 34 of the airfoil 31 than the third passage 43, and is connected to the third passage 43 at the folded portion on the tip 36 side of the airfoil 31. The first passage 41 extends in the wing height direction and is formed closest to the leading edge 34, and is connected to the second passage 42 at the folded portion on the base end 37 side of the airfoil 31.

[0024] The fifth passage 45 extends in the wing height direction and is formed closer to the trailing edge 35 of the airfoil 31 than the fourth passage 44, and is connected to the fourth passage 44 at the folded portion on the tip 36 side of the airfoil 31. The sixth passage 46 extends in the wing height direction and is formed closest to the trailing edge 35, and is connected to the fifth passage 45 at the folded portion on the base end 37 side of the airfoil 31.

[0025] In some embodiments of the turbine blade 3, the cooling air supplied from the opening 43a, which is the intake for cooling air as a cooling fluid, flows from the third passage 43 through the second passage 42 toward the first passage 41, that is, toward the leading edge 34, in the leading edge meandering passage 21. The cooling air as a cooling fluid is, for example, compressed air from the compressor 91. A portion of the cooling air that flows into the first passage 41 is blown out as film cooling air 11 from a plurality of cooling holes 41b to film cool the airfoil portion 31 from the outside. In addition, a portion of the cooling air that flows into the first passage 41 is blown out to the outside of the turbine blade 3 from the opening 41a formed at the tip 36 of the airfoil portion 31.

[0026] In some embodiments of the turbine blade 3, the cooling air supplied from the opening 44a, which is the intake for cooling air, flows from the fourth passage 44 through the fifth passage 45 toward the sixth passage 46, that is, toward the trailing edge 35, in the trailing edge meandering passage 22. A portion of the cooling air that flows into the sixth passage 46 is blown out from a plurality of cooling holes 46b as film cooling air 13 (see Figure 3) to film-cool the belly 38 side of the airfoil portion 31 from the outside. The remaining cooling air that flows into the sixth passage 46 is blown out from the cooling passage 48, which is provided with a large number of pin fins 49, as trailing edge blown air 12 (see Figure 2) to the outside of the turbine blade 3.

[0027] In some embodiments of the turbine blade 3, a plurality of cooling holes 46b are provided as film cooling holes, spaced apart in the blade height direction, only in the blade wall portion 50 on the belly side 38 in the sixth passage 46. In this way, when film cooling holes are provided only in either the blade wall portion 50 on the belly side 38 or the blade wall portion 50 on the back side 39 that defines a certain cooling passage, cooling is promoted in the blade wall portion 50 on the side with the film cooling holes. Therefore, if the temperature of the blade wall portion 50 on the side with the film cooling holes falls below the temperature of the blade wall portion 50 on the side without the film cooling holes, it is conceivable that tensile stress will act on the blade wall portion 50 on the side with the film cooling holes and compressive stress will act on the blade wall portion 50 on the side without the film cooling holes. In such a case, the tensile stress may become excessive, potentially shortening the lifespan of the turbine blade 3.

[0028] For example, in the case of turbine blades 3 according to some embodiments, cooling is more efficient on the blade wall portion 50 on the ventral side 38 (hereinafter also referred to as the first blade wall portion 501) within the range overlapping with the ventral inner wall surface 51p of the sixth passage 46 along the camber line Lc than on the blade wall portion 50 on the dorsal side 39 (hereinafter also referred to as the second blade wall portion 502) within the range overlapping with the dorsal inner wall surface 51s of the sixth passage 46 along the camber line Lc. Therefore, if the temperature of the first blade wall portion 501 falls below the temperature of the second blade wall portion 502, it is conceivable that tensile stress will act on the first blade wall portion 501 and compressive stress will act on the second blade wall portion 502. In such a case, the tensile stress in the first blade wall portion 501 may become excessive, potentially shortening the lifespan of the turbine blade 3.

[0029] Therefore, in some embodiments of the turbine blade 3, in order to reduce the temperature difference between the first blade wall portion 501 and the second blade wall portion 502, at least one of the first or second conditions described below is satisfied. In addition, in some embodiments of the turbine blade 3, either the first or second condition described below is satisfied.

[0030] <Regarding the First and Second Conditions> In some embodiments of the turbine blade 3, as described above, in the sixth passage 46, a plurality of cooling holes 46b are provided as film cooling holes, spaced apart in the blade height direction, only in the blade wall portion 50 on the ventral side 38. In the following description, the ventral turbulator 53 formed on the ventral inner wall surface 51p of the first blade wall portion 501, which is the blade wall portion 50 on which the plurality of cooling holes 46b are provided, will also be referred to as the first turbulator 71, and the dorsal turbulator 55 formed on the dorsal inner wall surface 51s of the second blade wall portion 502, which is the blade wall portion 50 on which the plurality of cooling holes 46b are not provided, will also be referred to as the second turbulator 72. The inner wall surface of the first blade wall portion 501 will also be referred to as the first inner wall surface 81, and the inner wall surface of the second blade wall portion 502 will also be referred to as the second inner wall surface 82.

[0031] If, for example, in the sixth passage 46, multiple cooling holes 46b are provided as film cooling holes only in the wing wall portion 50 on the back side 39, spaced apart in the wing height direction, then the wing wall portion 50 on the back side 39 where the multiple cooling holes 46b are provided will be referred to as the first wing wall portion 501, and the dorsal turbulator 55 formed on the dorsal inner wall surface 51s of the first wing wall portion 501 will also be referred to as the first turbulator 71, and the wing wall portion 50 on the ventral side 38 where the multiple cooling holes 46b are not provided will be referred to as the second wing wall portion 502, and the ventral turbulator 53 formed on the ventral inner wall surface 51p of the second wing wall portion 502 will also be referred to as the second turbulator 72.

[0032] Figure 5 is a schematic diagram of an example of a cross-section taken along the line V-V in Figures 4A and 4B. Figure 6 is a schematic diagram of another example of a cross-section taken along the line V-V in Figures 4A and 4B.

[0033] <First Condition> In some embodiments of the turbine blade 3, the first condition is that, as shown in Figures 5 and 6, the value obtained by dividing the first arrangement pitch P1 in the blade height direction of the first turbulator 71 by the first height e1 from the first inner wall surface 81 of the first turbulator 71 (P1 / e1) is greater than the value obtained by dividing the second arrangement pitch P2 in the blade height direction of the second turbulator 72 by the second height e2 from the second inner wall surface 82 of the second turbulator 72 (P2 / e2). By satisfying the first condition, the heat transfer coefficient between the cooling air and the first inner wall surface 81 can be made smaller than the heat transfer coefficient between the cooling air and the second inner wall surface 82. This reduces the temperature difference between the blade wall portion 50 (first blade wall portion 501) in the area overlapping with the first inner wall surface 81 along the camber line Lc and the blade wall portion 50 (second blade wall portion 502) in the area overlapping with the second inner wall surface 82 along the camber line Lc, thereby reducing the tensile stress acting on the first blade wall portion 501 and thus reducing the thermal stress acting on the turbine blade 3. Furthermore, according to several embodiments of the gas turbine 6, the thermal stress acting on the turbine blade 3 can be reduced, thereby improving the durability of the gas turbine 6.

[0034] To satisfy the first condition, for example, as shown in Figure 5, the first height e1 may be equal to the second height e2, and the first arrangement pitch P1 may be greater than the second arrangement pitch P2. This means that even when the first height e1 and the second height e2 are equal, the first condition can be satisfied by making the first arrangement pitch P1 greater than the second arrangement pitch P2. Also, for example, in a conventional turbine blade where the first height e1 and the second height e2 are equal, and the first arrangement pitch P1 and the second arrangement pitch P2 are equal, the first condition can be satisfied by redesigning the first arrangement pitch P1 to be greater than the second arrangement pitch P2.

[0035] To satisfy the first condition, for example, as shown in Figure 5, the first arrangement pitch P1 may be an integer multiple (for example, twice) of the second arrangement pitch P2. This means that even when the first height e1 and the second height e2 are equal, the first condition can be satisfied by making the first arrangement pitch P1 twice the second arrangement pitch P2. Also, for example, in a conventional turbine blade where the first height e1 and the second height e2 are equal and the first arrangement pitch P1 and the second arrangement pitch P2 are equal, the first condition can be satisfied by changing the design to thin out the turbulators 53X shown by the dashed line in Figure 5. Note that by thinning out the turbulators that existed before the design change, every other turbulator becomes twice the second arrangement pitch P2.

[0036] To satisfy the first condition, for example, as shown in Figure 6, the first height e1 may be smaller than the second height e2, and the first arrangement pitch P1 may be equal to the second arrangement pitch P2. This means that even when the first arrangement pitch P1 and the second arrangement pitch P2 are equal, the first condition can be satisfied by making the first height e1 smaller than the second height e2. Also, for example, in a conventional turbine blade where the first height e1 and the second height e2 are equal, and the first arrangement pitch P1 and the second arrangement pitch P2 are equal, the first condition can be satisfied by redesigning the first height e1 to be smaller than the second height e2.

[0037] To satisfy the first condition, the first height e1 may be smaller than the second height e2, and the first arrangement pitch P1 may be larger than the second arrangement pitch P2. This allows the first condition to be satisfied.

[0038] <Second Condition> In some embodiments of the turbine blade 3, the second condition is that, as shown in Figures 4A and 4B, the first turbulator 71 has a first region 711 in which the first height e1 does not change as it moves from the leading edge 34 to the trailing edge 35, and a second region 712 connected to the trailing edge 35 end 711e of the first region 711, in which the first height e1 gradually decreases from the same height as the trailing edge 35 end 711e of the first region 711 as it moves from the leading edge 34 to the trailing edge 35. By satisfying the second condition, the heat transfer coefficient between the cooling air and the first inner wall surface 81 can be reduced compared to the case in which the first turbulator 71 does not have the second region 712, that is, compared to the case in which the first height e1 does not change as the first turbulator 71 reaches the trailing edge 35 end. This reduces the temperature difference between the first blade wall 501 and the second blade wall 502, thereby reducing the tensile stress acting on the first blade wall 501. This reduces the thermal stress acting on the turbine blade 3, and also reduces stress concentration near the connection between the trailing edge 35 side end of the first turbulator 71 (the trailing edge 35 side end 712e of the second region 712) and the first blade wall 501. Furthermore, according to some embodiments of the gas turbine 6, the thermal stress acting on the turbine blade 3 can be reduced, thereby improving the durability of the gas turbine 6.

[0039] In the following explanation, for example, as shown in Figures 5 and 6, if the positions of the first turbulator 71 and the second turbulator 72 are different in the wing height direction, it is desirable that the conditions described below be satisfied when comparing any first turbulator 71 with the second turbulator 72 that is closest to the first turbulator 71 in the wing height direction, either on the base end or the tip end. Alternatively, if the positions of the first turbulator 71 and the second turbulator 72 are different in the wing height direction, it is desirable that the conditions described below be satisfied when comparing any second turbulator 72 with the first turbulator 71 that is closest to the second turbulator 72 in the wing height direction, either on the base end or the tip end.

[0040] In some embodiments of the turbine blade 3, the first height e1 is preferably zero at the end 712e on the trailing edge 35 side of the second region 712. The end 712e on the trailing edge 35 side of the second region 712 is preferably located away from the trailing edge side inner wall surface 51t, which is connected to the ventral inner wall surface 51p and the dorsal inner wall surface 51s on the trailing edge 35 side of the inner wall surface 51. As a result, there is no first turbulator 71 on the trailing edge 35 side of the end 712e on the trailing edge 35 side of the second region 712, which reduces the heat transfer coefficient between the cooling air and the first inner wall surface 81 and reduces the temperature difference between the first blade wall portion 501 and the second blade wall portion 502.

[0041] In some embodiments of the turbine blade 3, the second turbulator 72 may have a third region 723 in which the second height e2 does not change as it moves from the leading edge 34 to the trailing edge 35. The trailing edge 35 end 711e of the first region 711 may be located closer to the leading edge 34 than the trailing edge 35 end 723e of the third region 723. This makes it possible to relatively reduce the length of the first region 711 of the first turbulator 71 which has a relatively large first height e1, and to relatively increase the length of the third region 723 of the second turbulator 72 which has a relatively large second height e2. Therefore, the heat transfer coefficient between the cooling air and the first inner wall surface 81 can be relatively reduced, and the heat transfer coefficient between the cooling air and the second inner wall surface 82 can be relatively increased, thereby reducing the temperature difference between the first blade wall 501 and the second blade wall 502.

[0042] In some embodiments of the turbine blade 3, as shown in Figure 4A, the end portion 72e on the trailing edge 35 side of the second turbulator 72 may be the end portion 723e on the trailing edge 35 side of the third region 723. This means that the second turbulator 72 does not have a region like the second region 712 in the first turbulator 71 where the second height e2 gradually decreases from the leading edge 34 side to the trailing edge 35 side. Therefore, the length of the third region 723 of the second turbulator 72 can be secured, making it possible to relatively increase the heat transfer coefficient between the cooling air and the second inner wall surface 82, and reducing the temperature difference between the first blade wall portion 501 and the second blade wall portion 502.

[0043] In the turbine moving blade 3 according to some embodiments, as shown in FIG. 4B, the second turbulator 72 is connected to an end 723e of a third region 723 on the trailing edge 35 side, and may include a fourth region 724 in which a second height e2 gradually decreases from the same height as the end 723e of the third region 723 on the trailing edge 35 side as it goes from the leading edge 34 side toward the trailing edge 35 side. Accordingly, by providing the fourth region 724, the second height e2 gradually decreases as the shape of the second turbulator 72 on the trailing edge 35 side extends from the leading edge 34 side toward the trailing edge 35 side, so that the shape of the second turbulator 72 does not change abruptly on the trailing edge 35 side of the second turbulator 72. As a result, stress concentration in the vicinity of the connection between the end 72e of the second turbulator 72 on the trailing edge 35 side and the second blade wall portion 502 can be reduced.

[0044] In the turbine moving blade 3 according to some embodiments, the opening 46c of the cooling hole 46b on the first inner wall surface 81 is preferably positioned away from the leading edge side inner wall surface 51f connected to the pressure side inner wall surface 51p and the suction side inner wall surface 51s toward the trailing edge 35 side on the leading edge 34 side of the inner wall surface 51. For example, the cooling hole 46b is formed by machining the turbine moving blade 3 after casting. When the cooling hole 46b is formed such that the opening 46c on the first inner wall surface 81 of the cooling hole 46b is located closer to the leading edge 34 side than the opening 46d of the cooling hole 46b on the blade wall surface 31s, which is the outer surface of the airfoil portion 31, a tool is advanced from the outside of the airfoil portion 31 toward the inside of the airfoil portion 31 from the trailing edge 35 side toward the leading edge 34 side to drill the cooling hole 46b. Therefore, in order to prevent the tool from damaging the leading edge side inner wall surface 51f, the opening 46c on the first inner wall surface 81 of the cooling hole 46b needs to be set at a position away from the leading edge side inner wall surface 51f toward the trailing edge 35 side. According to the turbine moving blade 3 according to some embodiments, the possibility of damage to the leading edge side inner wall surface 51f during formation of the cooling hole 46b can be reduced.

[0045] In some embodiments of the turbine blade 3, the position of the opening 46d of the cooling hole 46b in the blade wall surface 31s, which is the outer surface of the airfoil portion 31, is preferably on the leading edge 34 side of the first region 711 than the trailing edge 35 side end 711e. As a result, the region cooled by the cooling air from the cooling hole 46b overlaps with the second region 712, where the first height e1 gradually decreases from the leading edge 34 side to the trailing edge 35 side, and there is a portion that does not overlap with the first region 711, where the first height e1 does not change from the leading edge 34 side to the trailing edge 35 side. Therefore, in the range where the first region 711 exists along the camber line Lc in the first blade wall portion 501, the first turbulator 71 having a relatively large first height e1 ensures a heat transfer coefficient between the cooling air and the first inner wall surface 81, thereby cooling the first blade wall portion 501. Furthermore, in the area of ​​the first wing wall 501 where the second region 712 exists along the camber line Lc, the first wing wall 501 can be cooled by the film cooling air 13 from the cooling holes 46b while ensuring a consistent heat transfer coefficient between the cooling air and the first inner wall surface 81 through the second region 712, where the first height e1 gradually decreases from the leading edge 34 to the trailing edge 35. Therefore, even if the position along the camber line Lc is different, the first wing wall 501 can be cooled relatively uniformly, thus reducing temperature differences in the first wing wall 501.

[0046] In some embodiments of the turbine blade 3, the sixth passage 46, located furthest to the trailing edge 35 among the plurality of cooling passages 41 to 46, is preferable to satisfy at least one of the first or second conditions. When the plurality of cooling passages 41 to 46 constitute a meandering flow path, as in some embodiments of the turbine blade 3, the temperature of the cooling air flowing into the sixth passage 46, located furthest to the trailing edge 35, is relatively high because it has already flowed through the fourth passage 44 and fifth passage 45, which are located upstream of the sixth passage 46 in the flow of the cooling air. Therefore, in addition to convective cooling in the sixth passage 46, it is desirable to cool the airfoil portion 31 with film cooling air 13 from the cooling holes 46b. According to some embodiments of the turbine blade 3, when cooling holes 46b are provided, the temperature difference between the first blade wall portion 501 and the second blade wall portion 502 can be reduced, and the tensile stress acting on the first blade wall portion 501 can be reduced, thereby reducing the thermal stress acting on the turbine blade 3.

[0047] In some embodiments of the turbine blade 3, the first inner wall surface 81 is the ventral inner wall surface 51p, and the second inner wall surface 82 is the dorsal inner wall surface 51s. In the turbine blade 3, the metal temperature of the blade wall portion 50 on the ventral side 38 is generally more likely to rise than that of the blade wall portion 50 on the dorsal side 39. In particular, in the region of the blade wall portion 50 on the ventral side 38 that is close to the trailing edge 35, it is desirable to cool the airfoil portion 31 with film cooling air from the film cooling holes in addition to convection cooling in the cooling passage, because it is close to the throat portion and the flow velocity of the combustion gas is high and the heat transfer coefficient is high. According to several embodiments of the turbine blade 3, for the reasons mentioned above, when a film cooling hole is provided in the blade wall portion 50 on the belly 38 side, the temperature difference between the blade wall portion 50 on the belly 38 side (first blade wall portion 501) and the blade wall portion 50 on the back 39 side (second blade wall portion 502) can be reduced, and the tensile stress acting on the blade wall portion 50 on the belly 38 side (first blade wall portion 501) can be reduced, thereby reducing the thermal stress acting on the turbine blade 3.

[0048] The present disclosure is not limited to the embodiments described above, and also includes forms obtained by modifying the above-described embodiments and forms obtained by appropriately combining these forms. For example, in the above description, the cooling holes 46b, which are film cooling holes, are formed so as to open only to the ventral inner wall surface 51p in the sixth passage 46, the blade wall portion 50 on the ventral 38 side is defined as a first blade wall portion 501, and the blade wall portion 50 on the dorsal 39 side is defined as a second blade wall portion 502. However, in the turbine rotor blade 3 according to some embodiments, the cooling holes 46b, which are film cooling holes, may be formed so as to open only to the dorsal inner wall surface 51s in the sixth passage 46, with the blade wall portion 50 on the dorsal 39 side serving as the first blade wall portion 501 and the blade wall portion 50 on the ventral 38 side serving as the second blade wall portion 502.

[0049] Note that even in a case where, in any of the cooling passages 41 to 45 other than the sixth passage 46, the film cooling holes open to only one of the ventral inner wall surface 51p and the dorsal inner wall surface 51s, it is preferable that the ventral turbulator 53 and the dorsal turbulator 55 satisfy at least one of the first condition or the second condition described above, as mentioned above.

[0050] In the above description, the case where film cooling holes are formed only in either one of the blade wall portion 50 on the ventral 38 side and the blade wall portion 50 on the dorsal 39 side has been described. However, the content of the present disclosure described above can also be applied to a case where, in any of the cooling passages 41 to 46, film cooling holes are provided in both the blade wall portion 50 on the ventral 38 side and the blade wall portion 50 on the dorsal 39 side, but the number of arranged film cooling holes differs between the blade wall portion 50 on the ventral 38 side and the blade wall portion 50 on the dorsal 39 side. In this case, among the blade wall portion 50 on the ventral 38 side and the blade wall portion 50 on the dorsal 39 side, the blade wall portion 50 having a larger number of arranged film cooling holes is defined as the first blade wall portion 501, and the other blade wall portion 50 is defined as the second blade wall portion 502. It is preferable that the ventral turbulator 53 and the dorsal turbulator 55 satisfy at least one of the first condition or the second condition described above.

[0051] The above description explains the structure of the turbine blade 3 for reducing the thermal stress acting on the turbine blade 3, but the turbine stator blade 5 may have a similar structure. Furthermore, the first turbulator and the second turbulator in the turbine stator blade 5 may satisfy at least one of the first or second conditions described above. This makes it possible to reduce the thermal stress acting on the turbine stator blade 5.

[0052] The contents described in each of the above embodiments can be understood, for example, as follows: (1) A turbine blade according to at least one embodiment of the present disclosure is a turbine blade (turbine rotor blade 3) of a gas turbine 6, comprising: an airfoil portion 31 having at least one cooling passage (cooling passages 41 to 46) extending in the blade height direction, and a film cooling hole (cooling hole 46b) communicating with the at least one cooling passage (cooling passages 41 to 46); a plurality of ventral turbulators 53 formed on the ventral inner wall surface 51p of the inner wall surface 51 of the airfoil portion 31 defining the at least one cooling passage (cooling passages 41 to 46) and arranged at intervals in the blade height direction; and a plurality of dorsal turbulators 55 formed on the dorsal inner wall surface 51s of the inner wall surface 51 and arranged at intervals in the blade height direction. The film cooling hole (cooling hole 46b) opens into a first inner wall surface 81 which is either the ventral inner wall surface 51p or the dorsal inner wall surface 51s. Of the ventral turbulator 53 or dorsal turbulator 55, the first turbulator 71 provided on the first inner wall surface 81, and the second turbulator 72 provided on the second inner wall surface 82, which is the other of the ventral inner wall surface 51p or the dorsal inner wall surface 51s, satisfy at least one of the first or second conditions. The first condition is that the value obtained by dividing the first arrangement pitch P1 in the wing height direction of the first turbulator 71 by the first height e1 of the first turbulator 71 from the first inner wall surface 81 (P1 / e1) is greater than the value obtained by dividing the second arrangement pitch P2 in the wing height direction of the second turbulator 72 by the second height e2 of the second turbulator 72 from the second inner wall surface 82 (P2 / e2). The second condition is that the first turbulator 71 has a first region 711 in which the first height e1 does not change as it moves from the leading edge 34 side to the trailing edge 35 side of the airfoil portion 31 of the first turbulator 71, and a second region 712 connected to the trailing edge 35 side end 711e of the first region 711 in which the first height e1 gradually decreases as it moves from the leading edge 34 side to the trailing edge 35 side.

[0053] According to the configuration of (1) above, by satisfying the first condition, the heat transfer coefficient between the cooling air and the first inner wall surface 81 can be made smaller than the heat transfer coefficient between the cooling air and the second inner wall surface 82. As a result, the temperature difference between the blade wall portion 50 (first blade wall portion 501) in the area overlapping with the first inner wall surface 81 along the camber line Lc and the blade wall portion 50 (second blade wall portion 502) in the area overlapping with the second inner wall surface 82 along the camber line Lc can be reduced, and the tensile stress acting on the first blade wall portion 501 can be reduced, thereby reducing the thermal stress acting on the turbine blade (turbine rotor blade 3). Furthermore, according to the configuration of (1) above, by satisfying the second condition, the heat transfer coefficient between the cooling air and the first inner wall surface 81 can be made smaller compared to the case where the first turbulator 71 does not have a second region 712. This reduces the temperature difference between the first wing wall 501 and the second wing wall 502, thereby reducing the tensile stress acting on the first wing wall 501, and thus reducing the thermal stress acting on the turbine blade (turbine rotor blade 3).

[0054] (2) In some embodiments, in the configuration of (1) above, the first height e1 may be equal to the second height e2, and the first arrangement pitch P1 may be greater than the second arrangement pitch P2.

[0055] According to the configuration in (2) above, even when the first height e1 and the second height e2 are equal, the first condition can be satisfied by making the first arrangement pitch P1 larger than the second arrangement pitch P2.

[0056] (3) In some embodiments, in the configuration of (2) above, the first arrangement pitch P1 may be twice the second arrangement pitch P2.

[0057] According to the configuration described in (3) above, even when the first height e1 and the second height e2 are equal, the first condition can be satisfied by making the first arrangement pitch P1 twice the second arrangement pitch P2.

[0058] (4) In some embodiments, in the configuration of (1) above, the first height e1 may be smaller than the second height e2, and the first arrangement pitch P1 may be equal to the second arrangement pitch P2.

[0059] According to the configuration described in (4) above, even when the first arrangement pitch P1 and the second arrangement pitch P2 are equal, the first condition can be satisfied by making the first height e1 smaller than the second height e2.

[0060] (5) In some embodiments, in the configuration of (1) above, the first height e1 may be smaller than the second height e2, and the first arrangement pitch P1 may be larger than the second arrangement pitch P2.

[0061] According to the configuration in (5) above, the first condition can be satisfied.

[0062] (6) In some embodiments, in any of the configurations (1) to (5) above, the position of the opening 46d of the film cooling hole (cooling hole 46b) on the outer surface (wing wall surface 31s) of the airfoil portion 31 is preferably on the leading edge 34 side of the end portion 711e on the trailing edge 35 side of the first region 711.

[0063] According to the configuration of (6) above, the region cooled by the cooling air (film cooling air 13) from the film cooling holes (cooling holes 46b) overlaps with the second region 712, where the first height e1 gradually decreases as you move from the leading edge 34 to the trailing edge 35, but does not overlap with the first region 711, where the first height e1 does not change as you move from the leading edge 34 to the trailing edge 35. As a result, in the region of the first wing wall 501 where the first region 711 exists along the camber line Lc, the first turbulator 71, which has a relatively large first height e1, ensures a high heat transfer coefficient between the cooling air and the first inner wall surface 81, thereby cooling the first wing wall 501. Furthermore, in the area of ​​the first wing wall 501 where the second region 712 exists along the camber line Lc, the first wing wall 501 can be cooled by the cooling air (film cooling air 13) from the film cooling holes (cooling holes 46b) while ensuring a consistent heat transfer coefficient between the cooling air and the first inner wall surface 81 through the second region 712, where the first height e1 gradually decreases from the leading edge 34 to the trailing edge 35. Therefore, even if the position along the camber line Lc is different, the first wing wall 501 can be cooled relatively uniformly, thus reducing temperature differences in the first wing wall 501.

[0064] (7) In some embodiments, in the configuration of (6) above, the first height e1 at the end 712e on the trailing edge 35 side of the second region 712 is preferably zero. The end 712e on the trailing edge 35 side of the second region 712 is preferably located away from the trailing edge side inner wall surface 51t, which is connected to the ventral inner wall surface 51p and the dorsal inner wall surface 51s on the trailing edge 35 side of the inner wall surface 51, toward the leading edge 34 side.

[0065] According to the configuration of (7) above, the first turbulator 71 is not present on the trailing edge 35 side of the end portion 712e on the trailing edge 35 side in the second region 712, so the heat transfer coefficient between the cooling air and the first inner wall surface 81 can be reduced, and the temperature difference between the first wing wall portion 501 and the second wing wall portion 502 can be reduced.

[0066] (8) In some embodiments, in the configuration of (6) or (7) above, the second turbulator 72 may have a third region 723 in which the second height e2 does not change as it moves from the front edge 34 side to the rear edge 35 side of the second turbulator 72. The end 711e on the rear edge 35 side of the first region 711 may be located closer to the front edge 34 than the end 723e on the rear edge 35 side of the third region 723.

[0067] According to the configuration of (8) above, the length of the first region 711 of the first turbulator 71 having a relatively large first height e1 can be made relatively small, and the length of the third region 723 of the second turbulator 72 having a relatively large second height e2 can be made relatively large. As a result, the heat transfer coefficient between the cooling air and the first inner wall surface 81 can be made relatively small, and the heat transfer coefficient between the cooling air and the second inner wall surface 82 can be made relatively large, thereby reducing the temperature difference between the first wing wall portion 501 and the second wing wall portion 502.

[0068] (9) In some embodiments, in the configuration of (8) above, the end portion 72e on the trailing edge 35 side of the second turbulator 72 may be the end portion 723e on the trailing edge 35 side of the third region 723.

[0069] According to the configuration of (9) above, the second turbulator 72 does not have a region in which the second height e2 gradually decreases from the leading edge 34 side to the trailing edge 35 side, as in the second region 712 of the first turbulator 71. As a result, the length of the third region 723 of the second turbulator 72 can be secured, so that the heat transfer coefficient between the cooling air and the second inner wall surface 82 can be made relatively large, and the temperature difference between the first wing wall portion 501 and the second wing wall portion 502 can be reduced.

[0070] (10) In some embodiments, in the configuration of (8) above, the second turbulator 72 may have a fourth region 724 connected to the end 723e on the trailing edge 35 side of the third region 723, the second height e2 gradually decreases from the leading edge 34 side toward the trailing edge 35 side.

[0071] According to the configuration of (10) above, by providing the fourth region 724, the shape of the trailing edge 35 side of the second turbulator 72 gradually decreases as it moves from the leading edge 34 side to the trailing edge 35 side, and the shape of the second turbulator 72 does not change abruptly at the trailing edge 35 side of the second turbulator 72. As a result, stress concentration near the connection between the second turbulator 72 and the second wing wall 502 can be reduced.

[0072] (11) In some embodiments, in any of the configurations (6) to (10) above, the opening 46c in the first inner wall surface 81 of the film cooling hole (cooling hole 46b) is preferably located away from the rear edge 35 side from the front edge side inner wall surface 51f, which is connected to the ventral inner wall surface 51p and the dorsal inner wall surface 51s on the front edge 34 side of the inner wall surface 51.

[0073] According to the configuration described in (11) above, the possibility of damage to the leading edge inner wall surface 51f when forming the cooling holes 46b can be reduced.

[0074] (12) In some embodiments, in any of the configurations (1) to (11) above, at least one cooling passage (cooling passages 41 to 46) may include a plurality of cooling passages 41 to 46. The cooling passage 46 located furthest to the trailing edge 35 among the plurality of cooling passages (cooling passages 41 to 46) may satisfy at least one of the first or second conditions.

[0075] If the multiple cooling passages 41 to 46 constitute a meandering flow path, the cooling air in the cooling passage 46 located closest to the trailing edge 35 will be at a relatively high temperature. Therefore, in addition to convective cooling in the cooling passage 46, it is conceivable to cool the airfoil section 31 with film cooling air 13 from the film cooling holes (cooling holes 46b).

[0076] According to the configuration of (12) above, when film cooling holes (cooling holes 46b) are provided, the temperature difference between the first blade wall portion 501 and the second blade wall portion 502 can be reduced, and the tensile stress acting on the first blade wall portion 501 can be reduced, thereby reducing the thermal stress acting on the turbine blade (turbine rotor blade 3).

[0077] (13) In some embodiments, in any of the configurations (1) to (12) above, the first inner wall surface 81 is the ventral inner wall surface 51p, and the second inner wall surface 82 is the dorsal inner wall surface 51s.

[0078] In turbine blades (turbine rotor blades 3), the metal temperature of the blade wall 50 on the ventral side 38 is generally more prone to rising than that of the blade wall 50 on the dorsal side 39. In particular, in the region of the blade wall 50 on the ventral side 38 that is close to the trailing edge 35, the combustion gas flow velocity is high and the heat transfer coefficient is high because it is close to the throat. Therefore, it is desirable to cool the airfoil 31 with film cooling air from the film cooling holes in addition to convection cooling in the cooling passage.

[0079] According to the configuration of (13) above, for the reasons stated above, when a film cooling hole (cooling hole 46b) is provided in the wing wall portion 50 on the ventral side 38, the temperature difference between the wing wall portion 50 on the ventral side 38 (first wing wall portion 501) and the wing wall portion 50 on the dorsal side 39 (second wing wall portion 502) can be reduced, and the tensile stress acting on the wing wall portion 50 on the ventral side 38 (first wing wall portion 501) can be reduced, thereby reducing the thermal stress acting on the turbine blade (turbine blade 3).

[0080] (14) In some embodiments, in any of the configurations (1) to (13) above, the turbine blade may be a turbine rotor blade 3.

[0081] According to the configuration described in (14) above, the thermal stress acting on the turbine blades 3 can be reduced.

[0082] (14) A gas turbine 6 according to at least one embodiment of the present disclosure comprises a turbine blade (turbine rotor blade 3) having any of the configurations of (1) to (13) above, and a combustor 8 for generating combustion gas that flows through a combustion gas passage on which the turbine blade (turbine rotor blade 3) is provided.

[0083] According to the configuration described in (14) above, the thermal stress acting on the turbine blades (turbine rotor blades 3) can be reduced, thereby improving the durability of the gas turbine 6.

[0084] 3 Turbine rotor blade 5 Turbine stator blade 6 Gas turbine 8 Combustor 21 Meandering flow path (leading edge side meandering flow path) 22 Meandering flow path (trailing edge side meandering flow path) 31 Airfoil section 31s Airfoil wall surface 34 Leading edge 35 Trailing edge 36 Tip 37 Base 38 Ventral side 39 Dorsal side 41 Cooling passage (first passage) 42 Cooling passage (second passage) 43 Cooling passage (third passage) 44 Cooling passage (fourth passage) 45 Cooling passage (fifth passage) 46 Cooling passage (sixth passage) 46b Cooling hole 46c Opening 46d Opening 48 Cooling passage 50 Airfoil wall section 51 Inner wall surface 51p Ventral inner wall surface 51s Dorsal inner wall surface 51f Leading edge side inner wall surface 51t Trailing edge side inner wall surface 53 Ventral turbulator 55 Dorsal turbulator 71 First turbulator 72 Second turbulator 72e End 81 First inner wall surface 82 Second inner wall surface 501 First wing wall 502 Second wing wall 711 First region 711e End 712 Second region 712e End 723 Third region 723e End 724 Fourth region

Claims

1. A turbine blade for a gas turbine, comprising: an airfoil having at least one cooling passage extending in the blade height direction and a film cooling hole communicating with the at least one cooling passage; a plurality of ventral turbulators formed on the ventral inner wall surface of the airfoil defining the at least one cooling passage and arranged at intervals in the blade height direction; and a plurality of dorsal turbulators formed on the dorsal inner wall surface and arranged at intervals in the blade height direction, wherein the film cooling hole opens to a first inner wall surface which is either the ventral inner wall surface or the dorsal inner wall surface; and of the ventral turbulators or dorsal turbulators, a first turbulator provided on the first inner wall surface and a second turbulator provided on the second inner wall surface which is the other of the ventral inner wall surface or the dorsal inner wall surface, A turbine blade that satisfies at least one of the following conditions: firstly, the value obtained by dividing the first arrangement pitch P1 in the blade height direction of the first turbulator by the first height e1 from the first inner wall surface of the first turbulator (P1 / e1) is greater than the value obtained by dividing the second arrangement pitch P2 in the blade height direction of the second turbulator by the second height e2 from the second inner wall surface of the second turbulator (P2 / e2); or secondly, the first turbulator has a first region in which the first height e1 does not change as it moves from the leading edge to the trailing edge of the airfoil portion of the first turbulator, and a second region connected to the trailing edge end of the first region in which the first height e1 gradually decreases as it moves from the leading edge to the trailing edge.

2. The turbine blade according to claim 1, wherein the first height e1 is equal to the second height e2, and the first arrangement pitch P1 is greater than the second arrangement pitch P2.

3. The turbine blade according to claim 2, wherein the first arrangement pitch P1 is twice the second arrangement pitch P2.

4. The turbine blade according to claim 1, wherein the first height e1 is smaller than the second height e2, and the first arrangement pitch P1 is equal to the second arrangement pitch P2.

5. The turbine blade according to claim 1, wherein the first height e1 is smaller than the second height e2, and the first arrangement pitch P1 is larger than the second arrangement pitch P2.

6. The turbine blade according to claim 1, wherein the position of the opening of the film cooling hole on the outer surface of the airfoil portion is closer to the leading edge than the end on the trailing edge side of the first region.

7. The turbine blade according to claim 6, wherein at the trailing edge end in the second region, the first height e1 is zero, and the trailing edge end in the second region is located away from the trailing edge side of the inner wall surface that connects the ventral inner wall surface and the dorsal inner wall surface on the trailing edge side of the inner wall surface.

8. The turbine blade according to claim 6 or 7, wherein the second turbulator has a third region in which the second height e2 does not change as it moves from the leading edge side to the trailing edge side of the second turbulator, and the end of the trailing edge side of the first region is located closer to the leading edge than the end of the trailing edge side of the third region.

9. The turbine blade according to claim 8, wherein the trailing edge end of the second turbulator is the trailing edge end of the third region.

10. The turbine blade according to claim 8, wherein the second turbulator is connected to the end of the trailing edge side of the third region, and the second height e2 gradually decreases from the leading edge side toward the trailing edge side.

11. The turbine blade according to claim 6 or 7, wherein the opening in the first inner wall surface of the film cooling hole is located away from the trailing edge side of the inner wall surface, which is connected to the ventral inner wall surface and the dorsal inner wall surface on the leading edge side of the inner wall surface.

12. The turbine blade according to any one of claims 1 to 7, wherein the at least one cooling passage includes a plurality of cooling passages, and the cooling passage located furthest to the trailing edge among the plurality of cooling passages satisfies at least one of the first or second conditions.

13. The turbine blade according to any one of claims 1 to 7, wherein the first inner wall surface is the ventral inner wall surface, and the second inner wall surface is the dorsal inner wall surface.

14. The turbine blade according to any one of claims 1 to 7, wherein the turbine blade is a turbine rotor blade.

15. A gas turbine comprising: a turbine blade according to any one of claims 1 to 7; and a combustor for generating combustion gas flowing through a combustion gas passage on which the turbine blade is provided.

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