Gas turbine blade

Abstract

A gas turbine blade includes a platform portion in which a cooling passage is formed, and a blade main body portion that protrudes from an upper surface of the platform portion, in which a plurality of first ejection ports that eject gas supplied from the cooling passage are formed in an in-chord inner region of the upper surface on a side of the blade main body portion than a blade chord connecting a leading edge and a trailing edge of the blade main body portion, a plurality of second ejection ports that eject gas supplied from the cooling passage are formed in an out-chord outer region of the upper surface on a side opposite to the in-chord inner region with the blade chord as a reference, the out-chord outer region has an out-chord leading edge region on the leading edge side and an out-chord trailing edge region on the trailing edge side, and the plurality of second ejection ports are formed more in the out-chord leading edge region than in the out-chord trailing edge region.

Description

Technical Field

[0001] This invention relates to a gas turbine blade.

[0002] This application claims priority based on Japanese Patent Application No. 2020-174743, filed on October 16, 2020, the contents of which are incorporated herein by reference. Background Technology

[0003] The gas turbine blade has a plate-shaped platform section and a blade body section extending from the upper surface of the platform section. A cooling channel is formed inside the platform section for the flow of compressed air as a cooling medium. The ends of the cooling channel open onto the upper surface and side end faces of the platform section, as well as the surface of the blade body section.

[0004] As a specific example of such a cooling channel, the cooling channel described in Patent Document 1 is known. In the gas turbine blade involved in Patent Document 1, the opening end of the cooling channel is formed on the positive pressure surface (ventral surface) side of the blade body in the same manner.

[0005] Previous technical documents

[0006] Patent documents

[0007] Patent Document 1: Japanese Patent Application Publication No. 2012-102726 Summary of the Invention

[0008] The technical problem to be solved by the invention

[0009] Here, because a portion of the platform's surface away from the blade body (specifically, a portion of the area outside the blade chord connecting the leading and trailing edges of the blade body) is thinner than the area near the blade body, it is more susceptible to thermal stress from high-temperature combustion gases. If numerous cooling channel openings are formed at the ends of such areas, the fatigue strength of the platform may be reduced. In other words, the technology described in Patent Document 1 above still has room for improvement.

[0010] The present invention was made to solve the above-mentioned problems, and its purpose is to provide a gas turbine blade with further improved fatigue strength.

[0011] means for solving technical problems

[0012] To address the aforementioned issues, the gas turbine blade of the present invention comprises: a platform portion having a cooling channel formed therein; and a blade body portion protruding from the upper surface of the platform portion. In the upper surface, a plurality of first nozzles are formed in an in-chord region closer to the blade body portion than the blade chord connecting the leading and trailing edges of the blade body portion, for ejecting gas supplied from the cooling channel. In the upper surface, a plurality of second nozzles are formed in an out-of-chord region on the opposite side of the in-chord region, with reference to the blade chord. The out-of-chord region has an out-of-chord leading edge region on the leading edge side and an out-of-chord trailing edge region on the trailing edge side. The plurality of second nozzles are formed more extensively in the out-of-chord leading edge region than in the out-of-chord trailing edge region.

[0013] Invention Effects

[0014] According to the present invention, a gas turbine blade with further improved fatigue strength can be provided. Attached Figure Description

[0015] Figure 1 This is a schematic diagram illustrating the structure of a gas turbine according to an embodiment of the present invention.

[0016] Figure 2 This is an unfolded view of the moving blade layer according to an embodiment of the present invention.

[0017] Figure 3 This is a perspective view showing the structure of the moving blade (gas turbine blade) according to an embodiment of the present invention.

[0018] Figure 4 yes Figure 3 A cross-sectional view below IV-IV line.

[0019] Figure 5 This is a top view of the moving blade according to an embodiment of the present invention.

[0020] Figure 6 This is a cross-sectional view of the platform section according to an embodiment of the present invention. Detailed Implementation

[0021] (Structure of a gas turbine)

[0022] The following is for reference. Figures 1 to 6 The gas turbine 10 and gas turbine blades (moving blades 50) according to embodiments of the present invention will be described.

[0023] The gas turbine 10 of this embodiment includes: a compressor 20 for compressing air; a combustor 30 for burning fuel F in the air compressed by the compressor 20 to generate combustion gas G; and a turbine 40 for being driven by the combustion gas G.

[0024] The compressor 20 has: a compressor rotor 21 that rotates about an axis Ar; a compressor housing 25 that covers the compressor rotor 21 to enable rotation; and a plurality of stationary blade layers 26. The turbine 40 has a turbine rotor 41 that rotates about an axis Ar; a turbine housing 45 that covers the turbine rotor 41 to enable rotation; and a plurality of stationary blade layers 46.

[0025] The compressor rotor 21 and the turbine rotor 41 are located on the same axis Ar and are connected to each other to form the gas turbine rotor 11. For example, a generator rotor is connected to this gas turbine rotor 11. Furthermore, the compressor housing 25 and the turbine housing 45, sandwiching an intermediate housing 16, are connected to each other to form the gas turbine housing 15. The burner 30 is disposed in the intermediate housing 16. Hereinafter, the direction in which the axis Ar extends is referred to as the axial direction Da, the circumferential direction centered on the axis Ar is simply referred to as the circumferential direction Dc, and the direction perpendicular to the axis Ar is referred to as the radial direction Dr. Furthermore, along the axial direction Da, with the turbine 40 as a reference, the side facing the compressor 20 is referred to as the upstream side, and the side opposite to it is referred to as the downstream side.

[0026] The compressor rotor 21 has: a rotor shaft 22 extending along axis Ar in the axial direction Da; and a plurality of moving blade layers 23 mounted on the rotor shaft 22. The plurality of moving blade layers 23 are arranged along the axial direction Da. Each moving blade layer 23 is composed of a plurality of moving blades 24 arranged along the circumferential direction Dc. A stationary blade layer 26 is disposed on the downstream side of each of the plurality of moving blade layers 23. Each stationary blade layer 26 is disposed inside the compressor housing 25. Each stationary blade layer 26 is composed of a plurality of stationary blades 27 arranged along the circumferential direction Dc.

[0027] The turbine rotor 41 has: a rotor shaft 42 extending along axis Ar in the axial direction Da; and a plurality of moving blade layers 43 mounted on the rotor shaft 42. The plurality of moving blade layers 43 are arranged along the axial direction Da. Each moving blade layer 43 is composed of a plurality of moving blades 50 arranged along the circumferential direction Dc. A stationary blade layer 46 is disposed on the upstream side of each of the plurality of moving blade layers 43. Each stationary blade layer 46 is disposed inside the turbine housing 45. Each stationary blade layer 46 is composed of a plurality of stationary blades 47 arranged along the circumferential direction Dc. The annular space between the outer circumference of the rotor shaft 42 and the inner circumference of the turbine housing 45, in the region where the stationary blades 47 and moving blades 50 are disposed in the axial direction Da, forms a combustion gas flow path through which the combustion gas G from the combustor 30 flows.

[0028] (Structure of the moving blade)

[0029] like Figure 2 and Figure 3As shown, the moving blade 50 has: a blade body 51 (blade main body) extending radially Dr; a platform 61 (platform portion) formed radially inside the blade body 51; a shank 58 formed radially inside the platform 61; and a blade root 59 disposed radially inside the shank 58. A combustion gas flow path for combustion gas G from the burner 30 is formed radially outside the platform 61 (i.e., the region where the blade body 51 exists).

[0030] The blade body 51 has a leading edge 52 at its upstream end and a trailing edge 53 at its downstream end. The side of the blade body 51 facing circumferentially Dc is smoothly convex. On the surface of the blade body 51, the convex surface facing circumferentially Dc forms a dorsal surface (= negative pressure surface) 55, and the concave surface forms a ventral surface (= positive pressure surface) 54. For ease of explanation, the ventral surface (= positive pressure surface side) of the blade body 51 in circumferentially Dc is referred to as the circumferential ventral surface, and the dorsal surface (= negative pressure surface side) of the blade body 51 is referred to as the circumferential dorsal surface. Furthermore, the upstream side in axial direction Da is sometimes referred to as the leading side, and the downstream side in axial direction Da is referred to as the trailing side.

[0031] The cross-sectional shape of the blade root 59, perpendicular to the blade chord of the blade body 51, is a Christmas tree shape with alternating widening and narrowing sections facing radially inward. A blade root groove for embedding the blade root 59 is formed on the rotor shaft 42.

[0032] Platform 61 has a front end face 62, which is an axial upstream end face; a rear end face 63, which is an axial downstream end face; a ventral end face 64, which is a circumferential ventral end face; and a dorsal end face 65, which is a circumferential dorsal end face. When viewed radially from Dr, as... Figure 2 As shown, platform 61 is parallelogram-shaped. In the platforms 61 of adjacent moving blades 50 along the circumferential direction Dc, the ventral end face 64 of one platform 61 faces the dorsal end face 65 of the other platform 61. Furthermore, a gas flow surface 66 (upper surface) is formed on platform 61 as a radially outer surface, and an inner surface 67 as a radially inner surface. The gas flow surface 66 forms a portion of the radially inner side of the surface defining the combustion gas flow path and is in contact with the high-temperature combustion gas.

[0033] (Structure of the air passage in the blade)

[0034] like Figure 4As shown, a plurality of blade air channels 71 (cooling channels) extending radially Dr are formed on the moving blade 50. As an example, the moving blade 50 of this embodiment has seven blade air channels 71. Furthermore, an example of seven blade air channels 71 is shown, but the invention is not limited to this. Each blade air channel 71 is continuously formed from the blade body 51, platform 61, shank 58, and blade root 59 to the platform 61. The plurality of blade air channels 71 are arranged along the blade chord of the blade body 51. A portion of adjacent blade air channels 71 is connected to each other in the radially outer portion within the blade body 51 or the radially inner portion within the platform 61. Furthermore, any one of the plurality of blade air channels 71 is continuously formed across the blade body 51, platform 61, shank 58, and blade root 59, and opens at the radially inner end of the blade root 59. Compressed air generated by the compressor 20 flows into this blade air channel 71 as a cooling medium.

[0035] The blade body 51 has a blade tip passage 56 that extends upstream from the first blade air passage 71a, the most upstream of the plurality of blade air passages 71, and opens at the leading edge 52 of the blade body 51. A portion of the cooling air flowing into the first blade air passage 71a exits through the plurality of blade tip passages 56 of the blade body 51 to the combustion gas flow path. Another portion of this cooling air exits through the seventh blade air passage 71b, the most downstream of the blade air passages 71, to the combustion gas flow path. The leading edge 52 and trailing edge 53 of the blade body 51 are cooled by this cooling air.

[0036] (Structure of the ventral side passage and the first ejection outlet)

[0037] like Figure 5 As shown, a plurality of ventral side channels 75 are formed on the platform 61, extending from the blade air passage 71 toward the gas flow surface 66. These ventral side channels 75 form part of the cooling channel. The ventral side channels 75 extend parallel to each other toward the circumferential ventral side of the blade body 51. Furthermore, the term "parallel" here refers to substantial parallelism, allowing for design tolerances and manufacturing errors. One end of each ventral side channel 75 is a first nozzle 75a opening onto the gas flow surface 66. These first nozzles 75a are arranged in a curved shape along the ventral surface of the blade body 51. In this embodiment, as an example, seven first nozzles 75a are formed. Furthermore, the number of first nozzles 75a can be appropriately varied according to design and specifications.

[0038] Here, taking the line (blade chord Ch) connecting the leading edge 52 and trailing edge 53 of the blade body 51 on the gas flow surface 66 as a reference, the region closer to the blade body 51 than the blade chord Ch is called the inner chord region A. Furthermore, taking the blade chord Ch on the gas flow surface 66 as a reference, the region located on the opposite side to the inner chord region A is called the outer chord region B. At this time, all the aforementioned first nozzles 75a are located within the inner chord region A. That is, all first nozzles 75a are arranged closer to the blade body 51 than the blade chord Ch.

[0039] Furthermore, the outer chord region B is divided by a dividing line D orthogonal to the blade chord Ch into an outer leading edge region B1 located further towards the leading edge 52 than the dividing line D, and an outer trailing edge region B2 located further towards the trailing edge 53 than the dividing line D. When the length of the blade chord Ch is set to 100%, the position of the dividing line D is appropriately set within a range of less than 50% from the leading edge 52. The opening direction of the aforementioned first nozzle 75a is set to eject gas from the inner chord region A toward the outer trailing edge region B2.

[0040] (Structure of the leading edge side channel, branch channel, and second nozzle)

[0041] A leading-edge side passage 76 is connected to the first blade air passage 71a located on the side closest to the leading edge 52 in the blade air passage 71. The leading-edge side passage 76 extends from the first blade air passage 71a toward the aforementioned outer leading edge region B1. Furthermore, multiple (for example, three) branch passages 77 are connected along the middle of this leading-edge side passage 76. These leading-edge side passages 76 and branch passages 77 are part of the cooling passage.

[0042] Each branch channel 77 extends from the leading-edge side channel 76 along the blade chord Ch. One end of each branch channel 77 is a second nozzle 77a opening in the outer leading-edge region B1. As an example, these second nozzles 77a are positioned identically along the blade chord Ch, and are located within 25% of the length of the blade chord Ch from the leading edge 52. The opening direction of each second nozzle 77a is set to eject gas toward the outer trailing-edge region B2. In other words, the second nozzles 77a open further along the blade chord Ch than the first nozzle 75a.

[0043] Alternatively, a structure can be adopted in which the second nozzle 77a is also formed in the outer trailing edge region B2. However, in this case, the number of second nozzles 77a formed in the outer leading edge region B1 is greater than the number of second nozzles 77a formed in the outer trailing edge region B2. That is, the second nozzles 77a are preferably locally disposed on the outer leading edge region B1 side. Furthermore, the number of second nozzles 77a is less than the number of first nozzles 75a.

[0044] (Structure of the side passage and the third nozzle)

[0045] like Figure 6 As shown, the platform 61 has multiple side channels 72 extending from a portion of the blade air passage 71. More specifically, the side channels 72 are formed in a portion of the blade air passage 71 located on the leading edge 52 side and the blade air passage 71 located on the trailing edge 53 side. Furthermore, the side channels 72 are formed more on the trailing edge region B2 side than on the outer leading edge region B1 side. Alternatively, the side channels 72 may be formed only in the outer trailing edge region B2. In this embodiment, as an example, two side channels 72 are connected to each blade air passage 71. The side channels 72 extend from the blade air passage 71 toward the ventral end face 64. The end of each side channel 72 is a third nozzle 72a.

[0046] (Effects)

[0047] Here, because a portion of the platform 61 in the gas flow surface 66 that is far from the blade body 51 (specifically, a portion of the outer side of the blade chord connecting the leading and trailing edges of the blade body) is thinner than the region closer to the blade body 51, it is prone to thermal stress due to the high-temperature combustion gases. If more cooling channel openings are formed at the ends of such regions, the fatigue strength of the platform 61 may be reduced.

[0048] In addition, in order to improve the sealing performance between platforms 61 near the ventral end face 64 that contacts the back end face 65 of the adjacent platform 61, the area near the end face is sometimes locally thickened. However, since part of the outer region of the blade chord Ch is thinner than the area near the blade body 51, if more cooling channels are formed at the opening end of the outer region, the fatigue strength of the platform 61 will decrease.

[0049] In this embodiment, by forming the first nozzle 75a in the region near the blade body 51, the impact on strength can be minimized. Furthermore, in the outer chord region B, the number of second nozzles 77a in the outer leading edge region B1 is greater than that in the outer trailing edge region B2. In other words, in the outer chord region B, the second nozzles 77a are locally distributed on the leading edge 52 side compared to the first nozzles 75a in the inner chord region A. Thus, for example, compared to the case where the second nozzles 77a are uniformly formed in the outer chord region B, the impact on structural strength can be minimized.

[0050] Furthermore, according to the above structure, the number of second nozzles 77a in the outer region B is less than the number of first nozzles 75a in the inner region A. Since the outer region B is thinner than the region near the blade body 51, reducing the number of second nozzles 77a can further reduce the impact on structural strength.

[0051] Furthermore, according to the above structure, a second nozzle 77a formed in the outer leading edge region B1 ejects gas toward the outer trailing edge region B2. This effectively cools the surface of the outer trailing edge region B2, which has a relatively small number of second nozzles 77a.

[0052] Furthermore, according to the above structure, the first nozzle 75a ejects gas toward the trailing edge region B2. This effectively cools the surface of the relatively few trailing edge regions B2 caused by the second nozzle 77a.

[0053] Furthermore, according to the above structure, the second nozzle 77a opens further along the blade chord Ch than the first nozzle 75a. Therefore, the gas supplied from the second nozzle 77a can be used to actively perform film cooling on the outer trailing edge region B2.

[0054] Furthermore, according to the above structure, a plurality of first nozzles 75a are arranged along the shape of the ventral side 54 of the blade body. That is, these first nozzles 75a are arranged along the shape of the portion where the platform 61 contacts the blade body 51. Since the strength of this portion is higher than that of the region near the blade chord, the impact on strength caused by the formation of the first nozzles 75a can be further reduced.

[0055] Furthermore, according to the above structure, the side channels 72 communicating with the third nozzle 72a are formed more extensively on the outer trailing edge region B2 side than on the outer leading edge region B1 side. Therefore, the side channels 72 passing through the interior can provide convective cooling to the outer trailing edge region B2, where the number of second nozzles 77a is small and cooling is difficult. As a result, the outer trailing edge region B2 can be cooled more effectively.

[0056] The embodiments of the present invention have been described above. Furthermore, various changes and modifications can be made to the above structure without departing from the spirit of the invention. For example, the number and shape of the blade air passages 71 described in the above embodiments are merely examples and can be appropriately changed according to design and specifications. Additionally, other cooling passages communicating with the blade air passages 71 may also be formed.

[0057] [Postscript]

[0058] The gas turbine blades (moving blades 50) described in each embodiment are as follows.

[0059] (1) The gas turbine blade (moving blade 50) according to the first embodiment includes: a platform portion (platform 61) having a cooling channel (blade air channel 71) formed therein; and a blade body portion (blade body 51) protruding from the upper surface (gas flow surface 66) of the platform portion. In the upper surface, a plurality of first nozzles 75a are formed in an inner region A that is closer to the blade body portion than the blade chord Ch connecting the leading edge 52 and the trailing edge 53 of the blade body portion. In the upper surface, a plurality of second nozzles 77a are formed in an outer region B that is located on the opposite side of the inner region A with reference to the blade chord Ch. The outer region B has an outer leading edge region B1 located on the leading edge 52 side and an outer trailing edge region B2 located on the trailing edge 53 side. The plurality of second nozzles 77a are formed in the outer leading edge region B1 more than in the outer trailing edge region B2.

[0060] Here, since the in-chord region A, located closer to the blade body than the blade chord Ch, is thicker than the out-of-chord region B, its strength is relatively high, and the impact of thermal stress is smaller. By forming the first nozzle 75a in this region, the impact on strength can be minimized. Furthermore, in the out-of-chord region B, the number of second nozzles 77a in the out-of-chord leading edge region B1 is greater than that in the out-of-chord trailing edge region B2. In other words, in the out-of-chord region B, the second nozzles 77a are locally distributed on the leading edge 52 side, compared to the first nozzles 75a in the in-chord region A. Thus, for example, compared to the case where the second nozzles 77a are uniformly formed in the out-of-chord region B, the impact on structural strength can be minimized.

[0061] (2) In the gas turbine blades involved in the second method, the number of the second nozzles 77a may be less than the number of the first nozzles 75a.

[0062] According to the above structure, the number of second nozzles 77a in the outer region B is less than the number of first nozzles 75a in the inner region A. Since the outer region B is thinner than the region near the blade body 51, reducing the number of second nozzles 77a can further reduce the impact on structural strength.

[0063] (3) In the gas turbine blade involved in the third method, the position of the second nozzle 77a can be set in the outer leading edge region B1 within a range of less than 25% of the length of the blade chord Ch, with the leading edge 52 along the direction of the blade chord Ch as a reference.

[0064] Based on the above structure, the position of the second nozzle 77a can be set within a range of less than 25% of the length of the blade chord in the outer leading edge region B1, with the leading edge 52 in the direction of the blade chord Ch as a reference. This further reduces the impact on structural strength.

[0065] (4) In the gas turbine blade involved in the fourth method, the second nozzle 77a formed in the outer leading edge region B1 of the plurality of second nozzles 77a can be configured to spray the gas toward the outer trailing edge region B2.

[0066] According to the above structure, a second nozzle 77a formed in the outer leading edge region B1 ejects gas toward the outer trailing edge region B2. This effectively cools the surface of the outer trailing edge region B2, which has a relatively small number of second nozzles 77a.

[0067] (5) In the gas turbine blade involved in the fifth method, the plurality of first nozzles 75a may be configured to eject the gas toward the outer trailing edge region B2.

[0068] According to the above structure, the first nozzle 75a ejects gas toward the outer trailing edge region B2. This effectively cools the surface of the relatively few outer trailing edge regions B2 vented by the second nozzle 77a.

[0069] (6) In the gas turbine blade involved in the sixth method, the second nozzle 77a may open further along the direction of the blade chord Ch than the first nozzle 75a.

[0070] According to the above structure, the second nozzle 77a opens further along the blade chord Ch than the first nozzle 75a. Therefore, the gas supplied from the second nozzle 77a can be used to actively perform film cooling on the outer trailing edge region B2.

[0071] (7) In the gas turbine blade involved in the seventh method, the plurality of first nozzles 75a may be arranged along the shape of the positive pressure surface (ventral side 54) of the blade body.

[0072] According to the above structure, a plurality of first nozzles 75a are arranged along the shape of the positive pressure surface of the blade body. That is, these first nozzles 75a are arranged along the shape of the portion where the platform portion and the blade body portion contact. Since this portion has high strength, the impact of forming the first nozzles 75a on strength can be further reduced.

[0073] (8) In the gas turbine blade of the eighth embodiment, a plurality of third nozzles 72a may be formed on the side end face (ventral end face 64) of the platform portion to eject the gas supplied from the cooling channel, and the cooling channel (side channel 72) communicating with the plurality of third nozzles 72a may be formed on the side of the outer trailing edge region B2 more than on the side of the outer leading edge region B1.

[0074] According to the above structure, the cooling channels communicating with the third nozzle 72a are formed more extensively on the outer trailing edge region B2 side than on the outer leading edge region B1 side. Therefore, the cooling channels (side channels 72) passing through the interior can provide convective cooling to the outer trailing edge region B2, where the number of second nozzles 77a is small and cooling is difficult to achieve. As a result, the outer trailing edge region B2 can be cooled more effectively.

[0075] Industrial availability

[0076] According to the present invention, a gas turbine blade with further improved fatigue strength can be provided.

[0077] Symbol Explanation

[0078] 10-Gas turbine, 11-Gas turbine rotor, 15-Gas turbine housing, 16-Intermediate housing, 20-Compressor, 21-Compressor rotor, 23-Moving blade layer, 24-Moving blade, 25-Compressor housing, 26-Stationary blade layer, 27-Stationary blade, 30-Burner, 40-Turbine, 41-Turbine rotor, 42-Rotor shaft, 43-Moving blade layer, 45-Turbine housing, 46-Stationary blade layer, 47-Stationary blade, 50-Moving blade, 51-Blade body, 52-Leading edge, 53-Leading edge, 54-Ventral side, 55-Back side, 58-Stalk, 59-Blade Root, 61-Platform, 62-Front end face, 63-Rear end face, 64-Ventral end face, 65-Dorsal end face, 66-Gas flow surface, 67-Inner surface, 71-Blade air passage, 71a-First blade air passage, 71b-Seventh blade air passage, 72-Side passage, 72a-Third nozzle, 75-Ventral side passage, 75a-First nozzle, 76-Leading edge side passage, 77-Branch passage, 77a-Second nozzle, A-Inner chord region, B-Outer chord region, B1-Outer chord leading edge region, B2-Outer chord trailing edge region, Ch-Blade chord, D-Dividing line.

Claims

1. A gas turbine blade, comprising: The platform section has internal cooling channels; and The main body of the blade protrudes from the upper surface of the platform section. The upper surface has an in-chord region located closer to the blade body than the blade chord connecting the leading and trailing edges of the blade body, and an out-of-chord region located on the opposite side of the in-chord region, with the blade chord as a reference. The cooling channel includes a ventral side channel extending toward the upper surface, a leading edge side channel extending toward the outer chord region, and a branch channel extending from the leading edge side channel in a direction along the blade chord. Multiple first nozzles are formed in the inner region of the string to eject gas supplied from the ventral side channel. Multiple second nozzles are formed in the outer region to eject gas supplied from the leading edge side channel and the branch channel. The outer-chord region includes an outer-chord leading-edge region located on the leading-edge side and an outer-chord trailing-edge region located on the trailing-edge side. The plurality of second nozzles are formed more extensively in the leading edge region than in the trailing edge region.

2. The gas turbine blade according to claim 1, wherein, The number of the second spray outlets is less than the number of the first spray outlets.

3. The gas turbine blade according to claim 1 or 2, wherein, The position of the second nozzle is set in the outer leading edge region of the blade chord, within a range of less than 25% of the length of the blade chord, with the leading edge along the direction of the blade chord as a reference.

4. The gas turbine blade according to claim 1 or 2, wherein, The second nozzle formed in the outer leading edge region of the plurality of second nozzles is configured to eject the gas toward the outer trailing edge region.

5. The gas turbine blade according to claim 1 or 2, wherein, The plurality of first nozzles are configured to eject the gas toward the outer trailing edge region.

6. The gas turbine blade according to claim 1 or 2, wherein, Compared to the first nozzle, the second nozzle opens further in the direction along the blade chord.

7. The gas turbine blade according to claim 1 or 2, wherein, The plurality of first nozzles are arranged along the shape of the positive pressure surface of the blade body.

8. The gas turbine blade according to claim 1 or 2, wherein, The cooling channel also has a side channel extending toward the side end face of the platform portion. A plurality of third nozzles are formed on the side end face of the platform section to eject the gas supplied from the side channel. The side passages that communicate with the plurality of third nozzles are formed more extensively on the trailing edge side than on the leading edge side.

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