JP7733247B2 - Gas turbine vane and gas turbine - Google Patents

Description

The present disclosure relates to a gas turbine vane and a gas turbine.
This application claims priority based on Japanese Patent Application No. 2022-138919, filed with the Japan Patent Office on September 1, 2022, the contents of which are incorporated herein by reference.

The leading edge of a gas turbine stator vane is exposed to the flow of high-temperature combustion gas. Patent Document 1 discloses that, in order to maintain the metal temperature of the gas turbine stator vane below the allowable temperature of the blade material, multiple film cooling holes are provided in the leading edge of the gas turbine stator vane, and the cooling air flowing out from the multiple film cooling holes performs film cooling on the surface (blade surface) of the gas turbine stator vane.

Japanese Patent Application Publication No. 10-220203

The gas turbine vane described in Patent Document 1 has room for improvement in terms of effectively film cooling the surface of the gas turbine vane with a smaller amount of cooling air.

In consideration of the above circumstances, at least one embodiment of the present disclosure aims to provide a gas turbine vane and a gas turbine equipped therewith that can suppress a decrease in film efficiency of film cooling on the blade surface while reducing the amount of cooling air.

In order to achieve the above object, a gas turbine stator vane according to at least one embodiment of the present disclosure comprises:
A gas turbine stator vane,
an airfoil portion having a cavity formed therein;
The leading edge of the airfoil has:
a first film cooling hole row including a plurality of first film cooling holes arranged along the blade height direction;
a second film cooling hole row including a plurality of second film cooling holes arranged along the blade height direction;
a third film cooling hole row including a plurality of third film cooling holes arranged along the blade height direction;
is formed,
the second row of film cooling holes is disposed between the first row of film cooling holes and the third row of film cooling holes;
the first film cooling hole extends in a direction inclined with respect to a thickness direction of the leading edge portion, and when a plane including the thickness direction of the leading edge portion and the blade height direction at the center of the inlet of the first film cooling hole is defined as a first plane, the center of the outlet of the first film cooling hole is located downstream of a position where the first plane intersects with an outer surface of the leading edge portion in a flow of combustion gas along the leading edge portion,
the second film cooling hole extends in a direction inclined with respect to a thickness direction of the leading edge portion, and when a plane including the thickness direction of the leading edge portion and the blade height direction at the center of an inlet of the second film cooling hole is defined as a second plane, an outlet of the second film cooling hole is located at a position where the second plane intersects with an outer surface of the leading edge portion,
The third film cooling hole extends in a direction inclined with respect to the thickness direction of the leading edge portion, and if a plane including the thickness direction of the leading edge portion and the blade height direction at the center of the inlet of the third film cooling hole is defined as a third plane, the center of the outlet of the third film cooling hole is located downstream of the position where the third plane intersects with the outer surface of the leading edge portion in the flow of combustion gas along the leading edge portion.

At least one embodiment of the present disclosure provides a gas turbine vane and a gas turbine equipped therewith that can suppress a decrease in film efficiency of film cooling on the blade surface while reducing the amount of cooling air.

FIG. 1 is a diagram illustrating a schematic configuration of a gas turbine 2 according to an embodiment of the present disclosure. FIG. 2 is a diagram illustrating an example of a cross section of a turbine stator blade 12 along the blade height direction according to an embodiment of the present disclosure. 3 is a diagram showing an example of a cross section perpendicular to the blade height direction of the turbine stator blade 12 shown in FIG. 2. FIG. FIG. 4 is an enlarged cross-sectional view of a front edge portion 21 shown in FIG. 3 . 2 is a view of the outer surface 25 of the leading edge 21 of the airfoil 20 viewed from a direction perpendicular to the outer surface 25. FIG. 1 is a diagram showing an example of a cross section along the blade height direction at the position of a first row of film cooling holes R71 in the leading edge portion 21 of the airfoil portion 20. FIG. FIG. 7 is an enlarged cross-sectional view of an outer first film cooling hole 711, which is a first film cooling hole 71 near the outer shroud 22 in FIG. 6. FIG. 7 is an enlarged cross-sectional view of an inner first film cooling hole 712, which is the first film cooling hole 71 near the inner shroud 24 in FIG. 6. FIG. 7B is a diagram showing the QG-QG cross section in FIG. 7A. FIG. 7C is a diagram showing the QH-QH cross section in FIG. 7B. 1 is a diagram showing an example of a cross section along the blade height direction at the position of a second row of film cooling holes R72 in the leading edge portion 21 of the airfoil portion 20. FIG. FIG. 10 is an enlarged cross-sectional view of an outer second film cooling hole 721, which is the second film cooling hole 72 near the outer shroud 22 in FIG. 9. FIG. 10 is an enlarged cross-sectional view of an inner second film cooling hole 722, which is the second film cooling hole 72 near the inner shroud 24 in FIG. 9. FIG. 10B is a diagram showing the cross section QJ-QJ in FIG. 10A. FIG. 10C is a diagram showing the QK-QK cross section in FIG. 10B. 1 is a diagram showing an example of a cross section along the blade height direction at the position of a third row of film cooling holes R73 in the leading edge portion 21 of the airfoil portion 20. FIG. FIG. 13 is an enlarged cross-sectional view of an outer third film cooling hole 731, which is the third film cooling hole 73 near the outer shroud 22 in FIG. 12. FIG. 13 is an enlarged cross-sectional view of an inner third film cooling hole 732, which is the third film cooling hole 73 near the inner shroud 24 in FIG. 12. FIG. 13B is a diagram showing the QL-QL cross section in FIG. 13A. FIG. 13C shows the QM-QM cross section in FIG. 13B. FIG. 10 is a diagram showing a cross section perpendicular to the blade height direction of the airfoil portion 020 according to the first comparative embodiment. 10A and 10B are diagrams showing the relationship between the distance from the second film cooling hole row R72 and the average film efficiency in the blade span direction in a turbine vane 12 according to one embodiment, and the relationship between the distance from the third film cooling hole row 07 (the film cooling hole row at a position corresponding to the second film cooling hole row R72) in a turbine vane 012 according to a first comparative embodiment and the average film efficiency in the blade span direction.

Hereinafter, several embodiments of the present disclosure will be described with reference to the accompanying drawings. However, the dimensions, materials, shapes, relative arrangements, etc. of components described as embodiments or shown in the drawings are merely illustrative examples and are not intended to limit the scope of the invention.
For example, expressions expressing relative or absolute arrangement such as "in a certain direction,""along a certain direction,""parallel,""orthogonal,""center,""concentric," or "coaxial" not only express such an arrangement exactly, but also express a state in which there is a relative displacement with a tolerance or an angle or distance to the extent that the same function is obtained.
For example, expressions such as "identical,""equal," and "homogeneous" that indicate that something is in an equal state not only indicate a state of strict equality, but also indicate a state in which there is a tolerance or a difference to the extent that the same function is obtained.
For example, expressions representing shapes such as a square shape or a cylindrical shape not only represent shapes such as a square shape or a cylindrical shape in the strict geometric sense, but also represent shapes including uneven portions, chamfered portions, etc., to the extent that the same effect can be obtained.
On the other hand, the expressions "comprise,""include,""have,""includes," or "have" of one element are not exclusive expressions that exclude the presence of other elements.

FIG. 1 is a diagram showing a schematic configuration of a gas turbine 2 according to one embodiment.
As shown in FIG. 1 , the gas turbine 2 includes a compressor 4, a combustor 6 for mixing compressed air generated by the compressor 4 with fuel and burning the mixture, and a turbine 8 for obtaining power from the combustion gas generated by the combustor 6.

1, the turbine 8 includes a rotor 9 (turbine rotor), a turbine casing 10 that houses the rotor 9, a plurality of turbine stator vanes 12 (gas turbine stator vanes) fixed to the inner surface of the turbine casing 10, and a plurality of turbine blades 16 that are implanted in the rotor 9 so as to be arranged alternately in the axial direction relative to the turbine stator vanes 12. Hereinafter, unless otherwise specified, "circumferential direction" means the circumferential direction of the gas turbine 2, i.e., the circumferential direction of the rotor 9, "axial direction" means the axial direction of the gas turbine 2, i.e., the axial direction of the rotor 9, unless otherwise specified, and "radial direction" means the radial direction of the gas turbine 2, i.e., the radial direction of the rotor 9, unless otherwise specified.

FIG. 2 is a diagram showing an example of a cross section of the turbine stator blade 12 along the blade height direction.
2 , the turbine stator vane 12 includes an airfoil portion 20, an outer shroud 22, and an inner shroud 24. Hereinafter, the term "blade height direction" refers to the blade height direction of the turbine stator vane 12, i.e., the blade height direction of the airfoil portion 20, and corresponds to the radial direction of the gas turbine 2, i.e., the radial direction of the rotor 9. Furthermore, the term "outside in the blade height direction" refers to the outside in the radial direction of the gas turbine 2, i.e., the outside in the radial direction of the rotor 9, and the term "inside in the blade height direction" refers to the inside in the radial direction of the gas turbine 2, i.e., the inside in the radial direction of the rotor 9.

The airfoil section 20 has an airfoil-shaped cross-sectional shape defined by a pressure surface and a suction surface. Inside the airfoil section 20, an intra-blade cavity (described below) is formed, which connects an outer cavity 32 formed outside the outer shroud 22 in the blade height direction with an inner cavity 34 formed inside the inner shroud 24 in the blade height direction. The outer cavity 32 is the space formed between the outer shroud 22 and the turbine casing 10 (see Figure 1), and the inner cavity 34 is the space formed between the inner shroud 24 and an inner diaphragm (not shown).

As shown in Figure 2, the leading edge 21 of the airfoil 20 is formed with a first row of film cooling holes R71, a second row of film cooling holes R72, and a third row of film cooling holes R73. In this specification, the term "leading edge" refers to the portion of the airfoil 20 that includes the most upstream position E2 of the airfoil 20 in the axial direction, as shown in Figure 3, for example, the portion opposite the trailing edge TE of the airfoil 20 across the leading edge bulkhead 48.

The first film cooling hole row R71 includes a plurality of first film cooling holes 71 arranged along the blade height direction. The second film cooling hole row R72 includes a plurality of second film cooling holes 72 arranged along the blade height direction. The third film cooling hole row R73 includes a plurality of third film cooling holes 73 arranged along the blade height direction. The second film cooling hole row R72 is arranged between the first film cooling hole row R71 and the third film cooling hole row R73.

The outer shroud 22 is connected to the outer end 20a of the airfoil portion 20 in the blade height direction and is formed in a generally plate-like shape along a plane intersecting the blade height direction. The outer shroud 22 forms the outer peripheral wall 33 of the combustion gas flow path 31 (the flow path of the main combustion gas in the turbine 8) in the turbine 8. In this specification, "upstream side in the axial direction" means the upstream side of the main combustion gas flow in the turbine 8 in the axial direction (the flow of combustion gas flowing through the flow path 31), and "downstream side in the axial direction" means the downstream side of the main combustion gas flow in the turbine 8 in the axial direction.

The inner shroud 24 connects to the inner end 20b of the airfoil section 20 in the blade height direction and is formed in a generally plate-like shape along a plane intersecting the blade height direction. The inner shroud 24 forms the inner circumferential wall 35 of the combustion gas flow path 31 in the turbine 8 (the flow path of the main combustion gas in the turbine 8).

Figure 3 is a diagram showing an example of a cross section perpendicular to the blade height direction of the turbine stator vane 12 shown in Figure 2. The turbine stator vane 12 described below may be, for example, a first-stage turbine stator vane 12 used under the highest temperature conditions in the turbine 8.

As shown in FIG. 3, the airfoil portion 20 of the turbine vane 12 includes a suction surface forming wall 38 that forms the suction surface 36, and a pressure surface forming wall 44 that forms the pressure surface 40 and forms an intra-blade cavity 42 between the suction surface forming wall 38.

Each of the suction surface forming wall 38 and the pressure surface forming wall 44 may have a curved plate shape with a substantially constant thickness. The intra-blade cavity 42 is formed inside the airfoil portion 20 along the blade height direction from one end of the airfoil portion 20 to the other end.

3, the intra-blade cavity 42 of the airfoil 20 is provided with a leading edge partition wall 48. The leading edge partition wall 48 is integrally formed with the airfoil 20, for example by casting, and is configured to extend from the inner surface 39 of the suction surface forming wall 38 to the inner surface 45 of the pressure surface forming wall 44, dividing the intra-blade cavity 42 into a leading edge cavity 46 and a trailing edge cavity 49. The leading edge partition wall 48 may have a plate shape with a substantially constant thickness.

In the exemplary embodiment shown in FIG. 3 , an insert 50 is provided in the leading edge cavity 46 of the airfoil 20. The insert 50 is formed, for example, from sheet metal in a tubular shape so as to extend from one end of the airfoil 20 to the other end along the blade height direction, and is inserted into the leading edge cavity 46. The internal space 51 of the insert 50 is connected to the outer cavity 32 (see FIG. 2 ), and compressed air supplied from the compressor 4 to the outer cavity 32 is supplied from the outer cavity 32 to the internal space 51 of the insert 50 as cooling air. The insert 50 is formed with a plurality of impingement cooling holes 52 for impingement cooling the wall surfaces (the inner surface 39 of the suction surface forming wall 38 and the inner surface 45 of the pressure surface forming wall 44) that form the leading edge cavity 46 of the airfoil 20. In the illustrated example, a plurality of impingement cooling holes 52 are formed in a portion of the insert 50 facing the suction surface forming wall 38 and a portion of the insert 50 facing the pressure surface forming wall 44 .

As shown in FIG. 3 , the leading edge 21 of the airfoil 20 is formed with the first film cooling hole row R71, second film cooling hole row R72, and third film cooling hole row R73. As described above, the first film cooling hole row R71 includes a plurality of first film cooling holes 71 arranged along the blade height direction, the second film cooling hole row R72 includes a plurality of second film cooling holes 72 arranged along the blade height direction, and the third film cooling hole row R73 includes a plurality of third film cooling holes 73 arranged along the blade height direction. Furthermore, the second film cooling hole row R72 is disposed between the first film cooling hole row R71 and the third film cooling hole row R73.

As shown in FIG. 3 , each of the first film cooling hole 71, the second film cooling hole 72, and the third film cooling hole 73 is located closer to the pressure surface 40 than the camber line CL in a cross section perpendicular to the blade height direction of the airfoil section 20. That is, each of the first film cooling hole row R71, the second film cooling hole row R72, and the third film cooling hole row R73 is formed in the pressure surface forming wall 44.

In the above-mentioned turbine vane 12, at least a portion of the cooling air supplied to the blade cavity 42 of the turbine vane 12 from outside the turbine vane 12 (e.g., the compressor 4) flows out to the outside of the airfoil portion 20 through either the first film cooling hole 71, the second film cooling hole 72, or the third film cooling hole 73, and flows along the surface of the airfoil portion 20 to perform film cooling on the surface of the airfoil portion 20.

3, a portion of the cooling air supplied to the internal space 51 of the insert 50 from outside the turbine vane 12 (e.g., the compressor 4) is ejected from the multiple impingement cooling holes 52 of the insert 50 to impingement cool the wall surface forming the leading edge cavity 46 of the airfoil 20. The cooling air that has been ejected from the multiple impingement cooling holes 52 of the insert 50 and performed impingement cooling of the pressure surface forming wall 44 passes through the first film cooling hole 71, the second film cooling hole 72, or the third film cooling hole 73 and flows outside the leading edge 21 of the airfoil 20, and flows along the surface of the airfoil 20 to perform film cooling of the surface of the airfoil 20.

FIG. 4 is an enlarged cross-sectional view of the leading edge portion 21 shown in FIG.
As shown in Figure 4, each of the first film cooling holes 71 is formed as a through hole that penetrates the leading edge portion 21 and extends from the inlet 71a to the outlet 71b of the first film cooling hole 71 along a direction inclined with respect to the thickness direction of the leading edge portion 21 (a direction perpendicular to each of the outer surface 25 and inner surface 26 of the leading edge portion 21).

In the exemplary configuration shown in Figure 4, each of the first film cooling holes 71 extends from the inlet 71a to the outlet 71b of the first film cooling hole 71 along a direction inclined with respect to a plane perpendicular to the blade height direction (the cross section of the airfoil portion 20 shown in Figure 4).

4, if a plane including the thickness direction and blade height direction of the leading edge portion 21 at the center C1a of the inlet 71a of each first film cooling hole 71 is defined as a first plane H1, each first film cooling hole 71 extends from the inlet 71a to the outlet 71b of the first film cooling hole 71 in a direction inclined with respect to the first plane H1. Furthermore, the center C1b of the outlet 71b of each first film cooling hole 71 is located downstream of the flow F of combustion gas along the leading edge portion 21 (the main stream of combustion gas flowing through the flow passage 31) relative to the position P1 where the first plane H1 intersects with the outer surface 25 of the leading edge portion 21.

As shown in Figure 4, each of the second film cooling holes 72 is formed as a through hole that penetrates the leading edge portion 21 and extends in a direction inclined relative to the thickness direction of the leading edge portion 21 from the inlet 72a to the outlet 72b of the second film cooling hole 72.

In the exemplary configuration shown in Figure 4, each of the second film cooling holes 72 extends from the inlet 72a to the outlet 72b of the second film cooling hole 72 along a direction inclined with respect to a plane perpendicular to the blade height direction.

Furthermore, as shown in Figure 4, if the plane including the thickness direction and blade height direction of the leading edge portion 21 at the center C2a of the inlet 72a of the second film cooling hole 72 is defined as the second plane H2, the outlet 72b of the second film cooling hole 72 is located at position P2 where the second plane H2 intersects with the outer surface 25 of the leading edge portion 21.

As shown in Figure 4, each of the third film cooling holes 73 is formed as a through hole that penetrates the leading edge portion 21 and extends in a direction inclined relative to the thickness direction of the leading edge portion 21 from the inlet 73a to the outlet 73b of the third film cooling hole 73.

In the exemplary configuration shown in Figure 4, each of the third film cooling holes 73 extends from the inlet 73a to the outlet 73b of the third film cooling hole 73 along a direction inclined with respect to a plane perpendicular to the blade height direction.

Also, as shown in Figure 4, if the plane including the thickness direction and blade height direction of the leading edge portion 21 at the center C3a of the inlet 73a of the third film cooling hole 73 is defined as the third plane H3, the center C3b of the outlet 73b of the third film cooling hole 73 is located downstream of the flow F of combustion gas along the leading edge portion 21 from the position P3 where the third plane H3 intersects with the outer surface 25 of the leading edge portion 21.

As shown in Figure 4, in a cross section perpendicular to the blade height direction of the airfoil section 20, if the intersection point between the camber line CL and the outer surface 25 of the leading edge 21 is E1 and the most upstream position of the airfoil section 20 in the axial direction of the gas turbine 2 is E2, the center C1b of the outlet 71b of the first film cooling hole 71 is located between the intersection point E1 and the most upstream position E2 on the outer surface 25 of the airfoil section 20, the center C2b of the outlet 72b of the second film cooling hole 72 is located on the opposite side of the intersection point E1 (towards the leading edge bulkhead 48) across the most upstream position E2 on the outer surface 25 of the airfoil section 20, and the center C3b of the outlet 73b of the third film cooling hole 73 is located on the opposite side of the intersection point E1 across the most upstream position E2 on the outer surface 25 of the airfoil section 20.

As shown in Figure 4, in a cross section perpendicular to the blade height direction of the airfoil section 20, each of the first film cooling hole 71, the second film cooling hole 72, and the third film cooling hole 73 is located upstream of the position G in the axial direction where the pressure surface forming wall 44 and the leading edge partition wall 48 connect.

Figure 5 is a view of the outer surface 25 of the leading edge 21 of the airfoil 20 as viewed from a direction perpendicular to the outer surface 25. Figure 6 is a diagram showing an example of a cross section along the blade height direction at the position of the first film cooling hole row R71 in the leading edge 21 of the airfoil 20. Figure 7A is an enlarged cross section of an outer first film cooling hole 711, which is a first film cooling hole 71 near the outer shroud 22 in Figure 6. Figure 7B is an enlarged cross section of an inner first film cooling hole 712, which is a first film cooling hole 71 near the inner shroud 24 in Figure 6. Figure 8A is a diagram showing the QG-QG cross section in Figure 7A. Figure 8B is a diagram showing the QH-QH cross section in Figure 7B.

Figure 9 is a diagram showing an example of a cross section along the blade height direction at the position of the second film cooling hole row R72 in the leading edge portion 21 of the airfoil 20. Figure 10A is an enlarged cross section of an outer second film cooling hole 721, which is a second film cooling hole 72 near the outer shroud 22 in Figure 9. Figure 10B is an enlarged cross section of an inner second film cooling hole 722, which is a second film cooling hole 72 near the inner shroud 24 in Figure 9. Figure 11A is a diagram showing the QJ-QJ cross section in Figure 10A. Figure 11B is a diagram showing the QK-QK cross section in Figure 10B.

Figure 12 is a diagram showing an example of a cross section along the blade height direction at the position of the third film cooling hole row R73 in the leading edge portion 21 of the airfoil 20. Figure 13A is an enlarged cross-sectional view of an outer third film cooling hole 731, which is the third film cooling hole 73 near the outer shroud 22 in Figure 12. Figure 13B is an enlarged cross-sectional view of an inner third film cooling hole 732, which is the third film cooling hole 73 near the inner shroud 24 in Figure 12. Figure 14A is a diagram showing the QL-QL cross section in Figure 13A. Figure 14B is a diagram showing the QM-QM cross section in Figure 13B.

5 and 6, the first film cooling holes 71 belonging to the first film cooling hole row R71 include a plurality of outer first film cooling holes 711 located outside a first position K1 in the blade height direction, and a plurality of inner first film cooling holes 712 located inside the first position K1 in the blade height direction. In the illustrated example, the first position K1 is located closer to the inner shroud 24 than the center of the airfoil 20 in the blade height direction.

For example, as shown in FIG. 7A, the outer first film cooling hole 711 extends inward in the blade height direction from the inlet 71a of the outer first film cooling hole 711 to the outlet 71b of the outer first film cooling hole 711.

For example, as shown in FIG. 7B, the inner first film cooling hole 712 extends outward in the blade height direction from the inlet 71a of the inner first film cooling hole 712 to the outlet 71b of the inner first film cooling hole 712.

For example, as shown in Figures 5 and 9, the multiple second film cooling holes 72 belonging to the second film cooling hole row R72 include multiple outer second film cooling holes 721 located outside the first position K1 in the blade height direction, and multiple inner second film cooling holes 722 located inside the first position K1 in the blade height direction.

For example, as shown in FIG. 10A, the outer second film cooling hole 721 extends inward in the blade height direction from the inlet 72a of the outer second film cooling hole 721 to the outlet 72b of the outer second film cooling hole 721.

For example, as shown in FIG. 10B, the inner second film cooling hole 722 extends outward in the blade height direction from the inlet 72a of the inner second film cooling hole 722 to the outlet 72b of the inner second film cooling hole 722.

For example, as shown in Figures 5 and 12, the multiple third film cooling holes 73 belonging to the third film cooling hole row R73 include multiple outer third film cooling holes 731 located outside the first position K1 in the blade height direction, and multiple inner third film cooling holes 732 located inside the first position K1 in the blade height direction.

For example, as shown in FIG. 13A, the outer third film cooling hole 731 extends inward in the blade height direction from the inlet 73a of the outer third film cooling hole 731 to the outlet 73b of the outer third film cooling hole 731.

For example, as shown in FIG. 13B, the inner third film cooling hole 732 extends outward in the blade height direction from the inlet 73a of the inner third film cooling hole 732 to the outlet 73b of the inner third film cooling hole 732.

As shown in Figures 5, 7A, and 8A, for example, the outer first film cooling hole 711 includes a circular hole portion 74 and a shaped hole portion 75. The circular hole portion 74 is the portion on the inlet 71a side of the outer first film cooling hole 711 and has a circular cross-sectional shape. The cross-sectional area of the circular hole portion 74 is constant regardless of the position in the direction of the axis L74 of the circular hole portion 74. The shaped hole portion 75 is the portion on the outlet 71b side of the outer first film cooling hole 711 and connects to the circular hole portion 74. The cross-sectional area of the shaped hole portion 75 increases toward the outlet 71b of the outer first film cooling hole 711 (the outlet of the shaped hole portion 75). When viewed in the direction of the axis L74 of the circular hole portion 74, the cross-sectional area of the shaped hole portion 75 expands only inward in the blade height direction relative to the circular hole portion 74, and does not expand outward in the blade height direction relative to the circular hole portion 74. Furthermore, when viewed from the direction of the axis L74 of the circular hole 74, the cross-sectional area of the shaped hole 75 is not expanded in comparison with the circular hole 74 in a direction perpendicular to the blade height direction.

7A, for example, the center C1b of the outlet 71b of the outer first film cooling hole 711 is located inward in the blade height direction from an extension of the axis L74 of the circular hole portion 74 of the outer first film cooling hole 711. More specifically, the center C1b of the outlet 71b of the outer first film cooling hole 711 is located inward in the blade height direction from an intersection S11 between the extension of the axis L74 of the circular hole portion 74 of the outer first film cooling hole 711 and the outlet 71b of the outer first film cooling hole 711.

For example, as shown in Figures 5, 7B, and 8B, the inner first film cooling hole 712 includes a circular hole portion 76 and a shaped hole portion 77. The circular hole portion 76 is the portion on the inlet 71a side of the inner first film cooling hole 712 and has a circular cross-sectional shape. The cross-sectional area of the circular hole portion 76 is constant regardless of the position in the direction of the axis L76 of the circular hole portion 76. The shaped hole portion 77 is the portion on the outlet 71b side of the inner first film cooling hole 712 and connects to the circular hole portion 76. The cross-sectional area of the shaped hole portion 77 increases toward the outlet 71b of the inner first film cooling hole 712 (the outlet of the shaped hole portion 77). When viewed in the direction of the axis L76 of the circular hole portion 76, the cross-sectional area of the shaped hole portion 77 expands only outward in the blade height direction relative to the circular hole portion 76, and does not expand inward in the blade height direction relative to the circular hole portion 76. Furthermore, when viewed from the direction of the axis L76 of the circular hole 76, the cross-sectional area of the shaped hole 77 is not expanded in comparison with the circular hole 76 in a direction perpendicular to the blade height direction.

7B, for example, the center C1b of the outlet 71b of the inner first film cooling hole 712 is located outward in the blade height direction from an extension of the axis L76 of the circular hole portion 76 of the inner first film cooling hole 712. More specifically, the center C1b of the outlet 71b of the inner first film cooling hole 712 is located outward in the blade height direction from an intersection S12 between the extension of the axis L76 of the circular hole portion 76 of the inner first film cooling hole 712 and the outlet 71b of the inner first film cooling hole 712.

5, 10A, and 11A, the outer second film cooling hole 721 includes a circular hole portion 78 and a shaped hole portion 79. The circular hole portion 78 is the portion on the inlet 72a side of the outer second film cooling hole 721 and has a circular cross-sectional shape. The cross-sectional area of the circular hole portion 78 is constant regardless of the position in the direction of the axis L78 of the circular hole portion 78. The shaped hole portion 79 is the portion on the outlet 72b side of the outer second film cooling hole 721 and connects to the circular hole portion 78. The cross-sectional area of the shaped hole portion 79 increases toward the outlet 72b of the outer second film cooling hole 721 (the outlet of the shaped hole portion 79). When viewed in the direction of the axis L78 of the circular hole portion 78, the cross-sectional area of the shaped hole portion 79 expands only inward in the blade height direction relative to the circular hole portion 78, and does not expand outward in the blade height direction relative to the circular hole portion 78. Furthermore, when viewed from the direction of the axis L78 of the circular hole 78, the cross-sectional area of the shaped hole 79 is not expanded in comparison with the circular hole 78 in a direction perpendicular to the blade height direction.

10A, for example, the center C2b of the outlet 72b of the outer second film cooling hole 721 is located inward in the blade height direction from an extension of the axis L78 of the circular hole portion 78 of the outer second film cooling hole 721. More specifically, the center C2b of the outlet 72b of the outer second film cooling hole 721 is located inward in the blade height direction from an intersection S21 between the extension of the axis L78 of the circular hole portion 78 of the outer second film cooling hole 721 and the outlet 72b of the outer second film cooling hole 721.

5, 10B, and 11B, the inner second film cooling hole 722 includes a circular hole portion 80 and a shaped hole portion 81. The circular hole portion 80 is the portion on the inlet 72a side of the inner second film cooling hole 722 and has a circular cross-sectional shape. The cross-sectional area of the circular hole portion 80 is constant regardless of the position in the direction of the axis L80 of the circular hole portion 80. The shaped hole portion 81 is the portion on the outlet 72b side of the inner second film cooling hole 722 and connects to the circular hole portion 80. The cross-sectional area of the shaped hole portion 81 increases toward the outlet 72b of the inner second film cooling hole 722 (the outlet of the shaped hole portion 81). When viewed in the direction of the axis L80 of the circular hole portion 80, the cross-sectional area of the shaped hole portion 81 expands only outward in the blade height direction relative to the circular hole portion 80, and does not expand inward in the blade height direction relative to the circular hole portion 80. Furthermore, when viewed from the direction of the axis L80 of the circular hole 80, the cross-sectional area of the shaped hole 81 is not expanded relative to the circular hole 80 in a direction perpendicular to the blade height direction.

10B, for example, the center C2b of the outlet 72b of the inner second film cooling hole 722 is located outward in the blade height direction from an extension of the axis L80 of the circular hole portion 80 of the inner second film cooling hole 722. More specifically, the center C2b of the outlet 72b of the inner second film cooling hole 722 is located outward in the blade height direction from an intersection S22 between the extension of the axis L80 of the circular hole portion 80 of the inner second film cooling hole 722 and the outlet 72b of the inner second film cooling hole 722.

5, 13A, and 14A, the outer third film cooling hole 731 includes a circular hole portion 82 and a shaped hole portion 83. The circular hole portion 82 is the portion on the inlet 73a side of the outer third film cooling hole 731 and has a circular cross-sectional shape. The cross-sectional area of the circular hole portion 82 is constant regardless of the position of the circular hole portion 82 in the direction of the axis L82. The shaped hole portion 83 is the portion on the outlet 73b side of the outer third film cooling hole 731 and connects to the circular hole portion 82. The cross-sectional area of the shaped hole portion 83 increases toward the outlet 73b of the outer third film cooling hole 731 (the outlet of the shaped hole portion 83). When viewed in the direction of the axis L82 of the circular hole portion 82, the cross-sectional area of the shaped hole portion 83 expands only inward in the blade height direction relative to the circular hole portion 82, and does not expand outward in the blade height direction relative to the circular hole portion 82. Furthermore, when viewed from the direction of the axis L82 of the circular hole 82, the cross-sectional area of the shaped hole 83 is not expanded in comparison with the circular hole 82 in a direction perpendicular to the blade height direction.

13A, for example, the center C3b of the outlet 73b of the outer third film cooling hole 731 is located inward in the blade height direction from an extension of the axis L82 of the circular hole portion 82 of the outer third film cooling hole 731. More specifically, the center C3b of the outlet 73b of the outer third film cooling hole 731 is located inward in the blade height direction from an intersection S31 between the extension of the axis L82 of the circular hole portion 82 of the outer third film cooling hole 731 and the outlet 73b of the outer third film cooling hole 731.

5, 13B, and 14B, the inner third film cooling hole 732 includes a circular hole portion 84 and a shaped hole portion 85. The circular hole portion 84 is the portion on the inlet 73a side of the inner third film cooling hole 732 and has a circular cross-sectional shape. The cross-sectional area of the circular hole portion 84 is constant regardless of the position in the direction of the axis L84 of the circular hole portion 84. The shaped hole portion 85 is the portion on the outlet 73b side of the inner third film cooling hole 732 and connects to the circular hole portion 84. The cross-sectional area of the shaped hole portion 85 increases toward the outlet 73b of the inner third film cooling hole 732 (the outlet of the shaped hole portion 83). When viewed in the direction of the axis L84 of the circular hole portion 84, the cross-sectional area of the shaped hole portion 85 is expanded only outward in the blade height direction relative to the circular hole portion 84, and is not expanded inward in the blade height direction relative to the circular hole portion 84. Furthermore, when viewed from the direction of the axis L84 of the circular hole 84, the cross-sectional area of the shaped hole 85 is not expanded in comparison with the circular hole 84 in the direction perpendicular to the blade height direction.

13B, for example, the center C3b of the outlet 73b of the inner third film cooling hole 732 is located outward in the blade height direction from an extension of the axis L84 of the circular hole portion 84 of the inner third film cooling hole 732. More specifically, the center C3b of the outlet 73b of the inner third film cooling hole 732 is located outward in the blade height direction from an intersection S32 between the extension of the axis L84 of the circular hole portion 84 of the inner third film cooling hole 732 and the outlet 73b of the inner third film cooling hole 732.

For example, as shown in Figure 12, in a cross section including the axis of the third film cooling hole 73 (axis L82 of the circular hole portion 82 of the outer third film cooling hole 731 in Figure 13, axis L84 of the circular hole portion 84 of the inner third film cooling hole 732 in Figure 13B and their extensions) and the blade height direction, if the outer shroud 22 or the inner shroud 24 which protrudes the greater amount from the leading edge portion 21 in a direction perpendicular to the blade height direction is defined as the large shroud 90, and the outer shroud 22 or the inner shroud 24 which is not the large shroud 90 is defined as the small shroud 91, the distance V90 between the first position K1 and the large shroud 90 in the blade height direction is greater than the distance V91 between the first position K1 and the small shroud 91 in the blade height direction. In the example shown in Figure 12, in a cross section including the axis of the third film cooling hole 73 and the blade height direction, if the amount of protrusion of the outer shroud 22 from the leading edge portion 21 in a direction perpendicular to the blade height direction is A3 and the amount of protrusion of the inner shroud 24 from the leading edge portion 21 in a direction perpendicular to the blade height direction is B3, A3 > B3, so the outer shroud 22 is the large shroud 90 and the inner shroud 24 is the small shroud 91.

For example, as shown in Figure 6, in a cross section including the axis of the first film cooling hole 71 (axis L74 of the circular hole portion 74 of the outer first film cooling hole 711 in Figure 7A, axis L76 of the circular hole portion 76 of the inner first film cooling hole 712 in Figure 7B and their extensions) and the blade height direction, the amount of protrusion of the outer shroud 22 from the leading edge portion 21 in a direction perpendicular to the blade height direction is A1, and the amount of protrusion of the inner shroud 24 from the leading edge portion 21 in a direction perpendicular to the blade height direction is B1. 9, for example, in a cross section including the axis of the second film cooling hole 72 (the axis L78 of the circular hole portion 78 of the outer second film cooling hole 721 in FIG. 10A, the axis L80 of the circular hole portion 80 of the inner second film cooling hole 722 in FIG. 10B and their extensions) and the blade height direction, the amount of protrusion of the outer shroud 22 from the leading edge portion 21 in a direction perpendicular to the blade height direction is denoted as A2, and the amount of protrusion of the inner shroud 24 from the leading edge portion 21 in a direction perpendicular to the blade height direction is denoted as B2. 12, for example, in a cross section along the blade height direction that includes the axis of the third film cooling hole 73 (the axis L82 of the circular hole portion 82 of the outer third film cooling hole 731 in FIG. 13A, the axis L84 of the circular hole portion 84 of the inner third film cooling hole 732 in FIG. 13B and their extensions), the amount of protrusion of the outer shroud 22 from the leading edge portion 21 in a direction perpendicular to the blade height direction is denoted as A3, and the amount of protrusion of the inner shroud 24 from the leading edge portion 21 in a direction perpendicular to the blade height direction is denoted as B3.

Here, the outer shroud 22 or the inner shroud 24 having the largest protrusion amount among protrusion amounts A1, A2, A3, B1, B2, and B3 is defined as the large shroud 92, and the outer shroud 22 or the inner shroud 24 that is not the large shroud 92 is defined as the small shroud 93. The distance V92 between the first position K1 in the blade height direction and the large shroud 92 is greater than the distance V93 between the first position K1 in the blade height direction and the small shroud 93. In the example shown in Figures 6, 9, and 12, the largest protrusion amount among protrusion amounts A1, A2, A3, B1, B2, and B3 is A3, so the outer shroud 22 is the large shroud 92 and the inner shroud 24 is the small shroud 93.

The effects achieved by the turbine vane 12 will be described below.
For example, as shown in FIG. 4 , each of the first film cooling hole 71, the second film cooling hole 72, and the third film cooling hole 73 extends in a direction inclined with respect to the thickness direction of the leading edge portion 21, and the center C1b of the outlet 71b of the first film cooling hole 71 is located downstream of the flow F of combustion gas along the leading edge portion 21 from the position P1 where the first plane H1 intersects with the outer surface 25 of the leading edge portion 21, the outlet 72b of the second film cooling hole 72 is located at the position P2 where the second plane H2 intersects with the outer surface 25 of the leading edge portion 21, and the center C3b of the outlet 73b of the third film cooling hole 73 is located downstream of the flow F of combustion gas along the leading edge portion 21 from the position P3 where the third plane H3 intersects with the outer surface 25 of the leading edge portion 21.

Therefore, by arranging the second film cooling hole row R72 so that the second film cooling holes 72 are located at or near the stagnation point P0 (see FIG. 4 ) of the flow of combustion gas on the outer surface 25 of the leading edge portion 21, cooling air for film cooling can be discharged from the first film cooling hole 71 and the third film cooling hole 73 along the outer surface 25 of the leading edge portion 21 toward the downstream side of the flow of combustion gas on both sides of the boundary of the stagnation point P0 of the flow of combustion gas on the outer surface 25 of the leading edge portion 21. Therefore, film cooling of the outer surface 25 of the leading edge portion 21 can be effectively performed with a small amount of cooling air, and a decrease in the film efficiency of film cooling of the blade surface can be suppressed while reducing the amount of cooling air.

Furthermore, as shown in Figures 5, 6 and 7A, the outer first film cooling hole 711 located on the outer shroud 22 side in the blade height direction extends inward in the blade height direction as it moves from the inlet 71a to the outlet 71b of the outer first film cooling hole 711. Therefore, when the outer first film cooling hole 711 is formed from the outer surface 25 side of the airfoil section 20 using a jig, for example by laser processing, interference between the jig and the outer shroud 22 can be suppressed, compared to a case where the outer first film cooling hole 711 extends outward in the blade height direction as it moves from the inlet 71a to the outlet 71b of the outer first film cooling hole 711.

Furthermore, as shown in Figures 5, 6 and 7B, the inner first film cooling hole 712 located on the inner shroud 24 side in the blade height direction extends outward in the blade height direction as it moves from the inlet 71a to the outlet 71b of the inner first film cooling hole 712. Therefore, when forming the inner first film cooling hole 712 using a jig from the outer surface 25 side of the airfoil portion 20, interference between the jig and the inner shroud 24 can be suppressed, compared to when the inner first film cooling hole 712 extends inward in the blade height direction as it moves from the inlet 71a to the outlet 71b of the inner first film cooling hole 712.

Furthermore, as shown in Figures 5, 9 and 10A, the outer second film cooling hole 721 located on the outer shroud 22 side in the blade height direction extends inward in the blade height direction from the inlet 72a to the outlet 72b of the outer second film cooling hole 721. Therefore, when the outer second film cooling hole 721 is formed from the outer surface 25 side of the airfoil portion 20 using a jig, for example by laser processing, interference between the jig and the outer shroud 22 can be suppressed, compared to a case where the outer second film cooling hole 721 extends outward in the blade height direction from the inlet 72a to the outlet 72b of the outer second film cooling hole 721.

Furthermore, as shown in Figures 5, 9 and 10B, the inner second film cooling hole 722 located on the inner shroud 24 side in the blade height direction extends outward in the blade height direction from the inlet 72a to the outlet 72b of the inner second film cooling hole 722. Therefore, when forming the inner second film cooling hole 722 using a jig from the outer surface 25 side of the airfoil 20, interference between the jig and the inner shroud 24 can be suppressed, compared to when the inner second film cooling hole 722 extends inward in the blade height direction from the inlet 72a to the outlet 72b of the inner second film cooling hole 722.

Furthermore, as shown in Figures 5, 12, and 13A, the outer third film cooling hole 731 located on the outer shroud 22 side in the blade height direction extends inward in the blade height direction from the inlet 73a to the outlet 73b of the outer third film cooling hole 731. Therefore, when the outer third film cooling hole 731 is formed from the outer surface 25 side of the airfoil section 20 using a jig, for example, by laser processing, interference between the jig and the outer shroud 22 can be suppressed, compared to a case where the outer third film cooling hole 731 extends outward in the blade height direction from the inlet 73a to the outlet 73b of the outer third film cooling hole 731.

Furthermore, as shown in Figures 5, 12, and 13B, the inner third film cooling hole 732 located on the inner shroud 24 side in the blade height direction extends outward in the blade height direction from the inlet 73a to the outlet 73b of the inner third film cooling hole 732. Therefore, when forming the inner third film cooling hole 732 using a jig from the outer surface 25 side of the airfoil 20, interference between the jig and the inner shroud 24 can be suppressed, compared to when the inner third film cooling hole 732 extends inward in the blade height direction from the inlet 73a to the outlet 73b of the inner third film cooling hole 732.

12 , when forming the outer third film cooling holes 731 and the inner third film cooling holes 732 from the outer surface 25 of the airfoil 20 using a jig by, for example, laser machining, the outer shroud 22, which protrudes more from the leading edge 21 in the direction perpendicular to the blade height direction, is more likely to interfere with the jig than the inner shroud 24, which protrudes less from the leading edge 21 in the direction perpendicular to the blade height direction. Therefore, as shown in FIG. 12 , the distance V90 between the outer shroud 22 and the first position K1, which is the boundary between the range in the blade height direction where the outer third film cooling holes 731 and the range in the blade height direction where the inner third film cooling holes 732 are formed, is set to be greater than the distance V91 between the first position K1 and the inner shroud 24 in the blade height direction. This makes it possible to form the outer third film cooling holes 731 and the inner third film cooling holes 732 while suppressing interference between the jig and the outer shroud 22 and the inner shroud 24.

Furthermore, when forming each of the first through third film cooling holes 71 through 73 from the outer surface 25 of the airfoil 20 using a jig, for example, by laser processing, the large shroud 92, which has the greatest protrusion amount A3 from the leading edge 21 in a direction perpendicular to the blade height direction, is more likely to cause interference with the jig than the small shroud 93. Therefore, by making the distance V92 between the first position K1 in the blade height direction and the large shroud 92 greater than the distance V93 between the first position K1 in the blade height direction and the small shroud 93, each of the first through third film cooling holes 71 through 73 can be formed while suppressing interference between the jig and each of the outer shroud 22 and the inner shroud 24.

Furthermore, as explained using Figures 5, 6, 9, 12, etc., the first film cooling hole 71, the second film cooling hole 72, and the third film cooling hole 73 each extend in a direction inclined with respect to a plane perpendicular to the blade height direction. Therefore, compared to when the first film cooling hole 71, the second film cooling hole 72, and the third film cooling hole 73 each extend in the thickness direction of the leading edge portion 21, film cooling of the outer surface 25 of the leading edge portion 21 can be effectively performed with a smaller amount of cooling air, and a decrease in the film efficiency of film cooling of the blade surface can be suppressed while reducing the amount of cooling air.

7A , 10A , 13A , etc., by expanding the cross-sectional area of the shaped hole portions 75, 79, 83 of each of the outer first film cooling hole 711, the outer second film cooling hole 721, and the outer third film cooling hole 731 only inward in the blade height direction relative to the circular hole portions 74, 78, 82, respectively, each of the outer first film cooling hole 711, the outer second film cooling hole 721, and the outer third film cooling hole 731 can be easily formed by laser processing, etc. Furthermore, by expanding the cross-sectional area of the shaped hole portions 77, 81, 85 of each of the inner first film cooling hole 712, the inner second film cooling hole 722, and the inner third film cooling hole 732 only outward in the blade height direction relative to the circular hole portions 76, 80, 84, respectively, each of the inner first film cooling hole 712, the inner second film cooling hole 722, and the inner third film cooling hole 732 can be easily formed by laser processing, etc.

Figure 15 is a diagram showing a cross section perpendicular to the blade height direction of a turbine stator vane 012 according to the first comparative embodiment. Figure 16 is a diagram showing the relationship between the distance from the second film cooling hole row R72 and the average film efficiency in the blade span direction in a turbine stator vane 12 according to the above embodiment, and the relationship between the distance from the third film cooling hole row 07 (the film cooling hole row at a position corresponding to the second film cooling hole row R72) in a turbine stator vane 012 according to the first comparative embodiment and the average film efficiency in the blade span direction.

15, five film cooling hole rows 07 are formed in the leading edge portion 021 of the turbine stator vane 012, and each of the film cooling hole rows 07 includes a plurality of film cooling holes 007 arranged along the blade height direction. Furthermore, in each of the film cooling hole rows 07, the film cooling holes 007 have a circular cross-sectional shape. If a plane including the thickness direction of the leading edge portion 021 and the blade height direction at the center of the inlet of the film cooling hole 007 is defined as a vertical plane, the outlet of the film cooling hole 007 is located at the position where the vertical plane intersects with the outer surface 025 of the leading edge portion 021.

In Figure 16, the solid line shows the relationship between the distance from the second film cooling hole row R72 in the turbine vane 12 according to the above embodiment and the average film efficiency in the blade span direction, the dashed line shows the relationship between the distance from the third film cooling hole row 07 in the turbine vane 012 according to the first comparative embodiment and the average film efficiency in the blade span direction, and the dotted line shows the relationship between the distance from the central film cooling hole row 07 in the turbine vane 012 according to the second comparative embodiment, which has two fewer film cooling hole rows than the configuration of the first comparative embodiment, for a total of three rows.

As shown in FIG. 16 , in the second comparative embodiment, the number of film cooling hole rows 07 is reduced from five to three compared to the first comparative embodiment, resulting in a decrease in film efficiency. In contrast, in the above embodiment, despite the number of film cooling hole rows being reduced from five to three compared to the first comparative embodiment, the third film cooling hole row R73 includes the shaped hole portions 83, 85, thereby improving film efficiency downstream of the third film cooling hole row R73 in the flow direction of the combustion gas along the leading edge portion 21. In this way, in the above turbine vane 12, film cooling of the outer surface 25 of the leading edge portion 21 can be effectively performed with a small amount of cooling air, and a decrease in film efficiency of film cooling on the blade surface can be suppressed while reducing the amount of cooling air.

The present disclosure is not limited to the above-described embodiments, but also includes variations on the above-described embodiments and appropriate combinations of these embodiments.

For example, when viewed in the direction of axis L74 of circular hole 74, shaped hole 75 may have a cross-sectional area expanded not only inward in the blade height direction but also outward relative to circular hole 74, and may also have a cross-sectional area expanded in a direction perpendicular to the blade height direction. Furthermore, when viewed in the direction of axis L76 of circular hole 76, shaped hole 77 may have a cross-sectional area expanded not only outward in the blade height direction but also inward relative to circular hole 76, and may also have a cross-sectional area expanded in a direction perpendicular to the blade height direction.

Furthermore, when viewed in the direction of axis L78 of circular hole 78, shaped hole 79 may have a cross-sectional area expanded not only inward in the blade height direction but also outward relative to circular hole 78, and may also have a cross-sectional area expanded in a direction perpendicular to the blade height direction. Furthermore, when viewed in the direction of axis L80 of circular hole 80, shaped hole 81 may have a cross-sectional area expanded not only outward in the blade height direction but also inward relative to circular hole 80, and may also have a cross-sectional area expanded in a direction perpendicular to the blade height direction.

Furthermore, when viewed in the direction of axis L82 of circular hole 82, shaped hole 83 may have a cross-sectional area expanded not only inward in the wing height direction but also outward relative to circular hole 82, and may also have a cross-sectional area expanded in a direction perpendicular to the wing height direction. Furthermore, when viewed in the direction of axis L84 of circular hole 84, shaped hole 85 may have a cross-sectional area expanded not only outward in the wing height direction but also inward relative to circular hole 84, and may also have a cross-sectional area expanded in a direction perpendicular to the wing height direction.

The contents described in each of the above embodiments can be understood, for example, as follows.

(1) A gas turbine vane according to at least one embodiment of the present disclosure includes:
1. A gas turbine vane (e.g., turbine vane 12 described above), comprising:
an airfoil (e.g., the airfoil 20 described above) having a cavity (e.g., the intra-airfoil cavity 42 described above) formed therein;
The leading edge of the airfoil (e.g., leading edge 21 described above) has:
a first film cooling hole row (e.g., the above-described first film cooling hole row R71) including a plurality of first film cooling holes (e.g., the above-described plurality of first film cooling holes 71) arranged along the blade height direction;
a second film cooling hole row (e.g., the above-described second film cooling hole row R72) including a plurality of second film cooling holes (e.g., the above-described plurality of second film cooling holes 72) arranged along the blade height direction;
a third film cooling hole row (e.g., the above-described third film cooling hole row R73) including a plurality of third film cooling holes (e.g., the above-described third film cooling holes 73) arranged along the blade height direction;
is formed,
the second row of film cooling holes is disposed between the first row of film cooling holes and the third row of film cooling holes;
the first film cooling hole extends in a direction inclined with respect to the thickness direction of the leading edge portion, and when a plane including the thickness direction of the leading edge portion and the blade height direction at a center (e.g., the above-mentioned center C1a) of an inlet (e.g., the above-mentioned inlet 71a) of the first film cooling hole is defined as a first plane (e.g., the above-mentioned first plane H1), the center (e.g., the above-mentioned center C1b) of an outlet (e.g., the above-mentioned outlet 71b) of the first film cooling hole is located downstream of a position (e.g., the above-mentioned position P1) where the first plane intersects with an outer surface of the leading edge portion (e.g., the above-mentioned outer surface 25), in a flow of combustion gas (e.g., the above-mentioned flow F), along the leading edge portion,
the second film cooling hole extends in a direction inclined with respect to the thickness direction of the leading edge portion, and a plane including the thickness direction of the leading edge portion and the blade height direction at the center (e.g., the above-mentioned center C2a) of the inlet (e.g., the above-mentioned inlet 72a) of the second film cooling hole is defined as a second plane (e.g., the above-mentioned second plane H2), the outlet (e.g., the above-mentioned outlet 72b) of the second film cooling hole is located at a position (e.g., the above-mentioned position P2) where the second plane intersects with the outer surface of the leading edge portion,
The third film cooling hole extends in a direction inclined with respect to the thickness direction of the leading edge portion, and if a plane including the thickness direction of the leading edge portion and the blade height direction at the center (e.g., the above-mentioned center C3a) of the inlet (e.g., the above-mentioned inlet 73a) of the third film cooling hole is defined as a third plane (e.g., the above-mentioned third plane H3), the center (e.g., the above-mentioned center C3b) of the outlet (e.g., the above-mentioned outlet 73b) of the third film cooling hole is located downstream of the flow of combustion gas (e.g., the above-mentioned flow F) along the leading edge portion relative to the position where the third plane intersects with the outer surface of the leading edge portion (e.g., the above-mentioned position P3).

According to the gas turbine vane described in (1) above, a first row of film cooling holes, a second row of film cooling holes, and a third row of film cooling holes are formed in the leading edge portion exposed to the flow of high-temperature combustion gas, and each row is arranged along the blade height direction. Here, each of the first, second, and third film cooling holes 71, 72, and 73 extends in a direction oblique to the thickness direction of the leading edge portion 21, the center of the outlet of the first film cooling hole is located downstream of the intersection of the first plane and the outer surface of the leading edge portion in the flow of combustion gas along the leading edge portion, the center of the outlet of the second film cooling hole is located at the intersection of the second plane and the outer surface of the leading edge portion, and the center of the outlet of the third film cooling hole is located downstream of the intersection of the third plane and the outer surface of the leading edge portion in the flow of combustion gas along the leading edge portion. Therefore, by arranging the second film cooling hole row so that the second film cooling holes are located at or near the stagnation point of the flow of combustion gas on the outer surface of the leading edge of the airfoil, cooling air for film cooling can be discharged from the first and third film cooling holes toward the downstream side of the flow of combustion gas on both sides of the boundary of the stagnation point of the flow of combustion gas on the outer surface of the leading edge of the airfoil. Therefore, it is possible to provide a gas turbine vane that can effectively perform film cooling on the outer surface of the leading edge with a small amount of cooling air and that can suppress a decrease in film efficiency of film cooling on the blade surface while reducing the amount of cooling air.

(2) In some embodiments, in the gas turbine stator vane described in (1),
Each of the first film cooling hole, the second film cooling hole, and the third film cooling hole is located closer to the pressure surface (e.g., the pressure surface 40) than the camber line (e.g., the camber line CL mentioned above) in a cross section of the airfoil section perpendicular to the blade height direction.

Since the stagnation point of the combustion gas flow on the outer surface of the leading edge of the airfoil occurs closer to the pressure surface than the camber line, according to the gas turbine vane described in (2) above, by arranging the second film cooling hole row so that the second film cooling hole is located at or near the stagnation point of the combustion gas flow on the outer surface of the leading edge of the airfoil, cooling air for film cooling can be discharged from the first film cooling hole and the third film cooling hole toward the downstream side of the combustion gas flow on both sides of the boundary of the stagnation point of the combustion gas flow on the outer surface of the leading edge of the airfoil. This allows film cooling of the outer surface of the leading edge to be effectively performed with a small amount of cooling air, making it possible to provide a gas turbine vane that can suppress a decrease in film cooling efficiency on the blade surface while reducing the amount of cooling air.

(3) In some embodiments, in the gas turbine vane described in (1) or (2),
In a cross section of the airfoil portion perpendicular to the blade height direction, if the intersection point between the camber line and the outer surface of the leading edge portion is E1 and the most upstream position of the airfoil portion in the axial direction of the gas turbine is E2, the outlet of the first film cooling hole is located between the intersection point E1 and the most upstream position E2 on the outer surface of the airfoil portion, and the outlet of the third film cooling hole is located on the opposite side of the intersection point E1 across the most upstream position E2 on the outer surface of the airfoil portion.

According to the gas turbine vane described in (3) above, cooling air for film cooling can be discharged from the first film cooling hole and the third film cooling hole on both sides of the most upstream position E2 toward the downstream side of the combustion gas flow, so that film cooling of the outer surface of the leading edge can be effectively performed with a small amount of cooling air, and a gas turbine vane can be provided that can suppress a decrease in the film efficiency of film cooling on the blade surface while reducing the amount of cooling air.

(4) In some embodiments, in the gas turbine stator vane described in (3),
In a cross section of the airfoil portion perpendicular to the blade height direction, the outlet of the second film cooling hole is located on the opposite side of the intersection point E1 across the most upstream position E2 on the outer surface of the airfoil portion.

The gas turbine vane described in (4) above can provide a gas turbine vane that can reduce the amount of cooling air while suppressing a decrease in the film efficiency of film cooling on the blade surface.

(5) In some embodiments, in the gas turbine stator vane according to any one of (1) to (4),
The airfoil portion is
a pressure surface forming wall (e.g., the above-mentioned pressure surface forming wall 44) that forms a pressure surface (e.g., the above-mentioned pressure surface 40);
a negative pressure surface forming wall (e.g., the above-mentioned negative pressure surface forming wall 38) that forms a negative pressure surface (e.g., the above-mentioned negative pressure surface 36);
a leading edge partition (e.g., the leading edge partition 48 described above) extending from the inner surface of the suction surface defining wall to the inner surface of the pressure surface defining wall so as to divide the cavity within the airfoil;
In a cross section of the airfoil portion perpendicular to the blade height direction, each of the first film cooling hole, the second film cooling hole, and the third film cooling hole is located upstream of the position where the pressure surface forming wall and the leading edge partition connect (for example, the above-mentioned position G) in the axial direction of the gas turbine.

The stagnation point of the combustion gas flow on the outer surface of the leading edge of the airfoil occurs upstream of the position where the pressure surface forming wall and the leading edge bulkhead connect. Therefore, according to the gas turbine vane described in (5) above, by arranging the second film cooling hole row so that the second film cooling hole is located at or near the position of the stagnation point of the combustion gas flow on the outer surface of the leading edge of the airfoil, film cooling of the outer surface of the leading edge can be effectively performed with a small amount of cooling air, and a gas turbine vane can be provided that can suppress a decrease in the film efficiency of film cooling on the blade surface while reducing the amount of cooling air.

(6) In some embodiments, in the gas turbine vane according to any one of (1) to (5),
Each of the first, second and third film cooling holes extends along a direction inclined with respect to a plane perpendicular to the blade height direction.

According to the gas turbine vane described in (6) above, film cooling of the outer surface of the leading edge portion can be effectively performed with a smaller amount of cooling air compared to when each of the first film cooling hole, the second film cooling hole, and the third film cooling hole extends along the thickness direction of the leading edge portion, and a gas turbine vane can be provided that can suppress a decrease in the film efficiency of film cooling on the blade surface while reducing the amount of cooling air.

(7) In some embodiments, in the gas turbine stator vane according to any one of (1) to (6),
the plurality of third film cooling holes include a plurality of outer third film cooling holes (e.g., the above-mentioned plurality of outer third film cooling holes 731) located outside a first position (e.g., the above-mentioned first position K1) in the blade height direction, and a plurality of inner third film cooling holes (e.g., the above-mentioned plurality of inner third film cooling holes 732) located inside the first position in the blade height direction,
the outer third film cooling hole extends inward in the blade height direction from an inlet (e.g., the inlet 73 a) to an outlet (e.g., the outlet 73 b) of the outer third film cooling hole,
The inner third film cooling hole extends outward in the blade height direction from the inlet (e.g., the inlet 73a mentioned above) to the outlet (e.g., the outlet 73b mentioned above) of the inner third film cooling hole.

The gas turbine vane described in (7) above allows for more effective film cooling of the outer surface of the leading edge portion with a smaller amount of cooling air, compared to when the outer third film cooling holes and the inner third film cooling holes each extend along the thickness direction of the leading edge portion. Furthermore, because the outer third film cooling holes extend inward in the blade height direction from their inlets to their outlets, when the outer third film cooling holes are formed from the outer surface of the airfoil portion using a jig by, for example, laser processing, interference between the jig and the outer shroud and the inner shroud can be reduced, compared to when the outer third film cooling holes extend outward in the blade height direction from their inlets to their outlets. Furthermore, since the inner third film cooling hole extends outward in the blade height direction as it moves from the inlet to the outlet of the inner third film cooling hole, when the inner third film cooling hole is formed from the outer surface side of the airfoil portion by laser processing or the like using a jig, interference between the jig and each of the outer shroud and the inner shroud can be suppressed, compared to when the inner third film cooling hole extends inward in the blade height direction as it moves from the inlet to the outlet of the inner third film cooling hole.

(8) In some embodiments, in the gas turbine stator vane described in (7),
The gas turbine stator blade comprises:
an outer shroud (e.g., the outer shroud 22 described above) connected to an outer end of the airfoil in the blade height direction;
an inner shroud (e.g., the inner shroud 24 described above) connected to an inner end of the airfoil portion in the blade height direction;
Including,
In a cross section including the axis of the third film cooling hole and the blade height direction (for example, the cross section shown in Figure 12), if the outer shroud or the inner shroud that protrudes more from the leading edge in a direction perpendicular to the blade height direction is defined as a large shroud (for example, the above-mentioned large shroud 90), and the outer shroud or the inner shroud that is not the large shroud is defined as a small shroud (for example, the above-mentioned small shroud 91), the distance between the first position in the blade height direction and the large shroud (for example, the above-mentioned distance V90) is greater than the distance between the first position in the blade height direction and the small shroud (for example, the above-mentioned distance V91).

When forming the outer third film cooling holes and the inner third film cooling holes from the outer surface of the airfoil using a jig, for example, by laser processing, the large shroud, which protrudes more from the leading edge in a direction perpendicular to the blade height direction, is more likely to interfere with the jig than the small shroud. Therefore, as described in (8) above, by making the distance between the first position in the blade height direction and the large shroud greater than the distance between the first position in the blade height direction and the small shroud, the outer third film cooling holes and the inner third film cooling holes can be formed while suppressing interference between the jig and the outer shroud and the inner shroud.

(9) In some embodiments, in the gas turbine vane according to (7) or (8),
the plurality of first film cooling holes include a plurality of outer first film cooling holes (e.g., the above-described plurality of outer first film cooling holes 711) located outside the first position in the blade height direction, and a plurality of inner first film cooling holes (e.g., the above-described plurality of inner first film cooling holes 712) located inside the first position in the blade height direction,
the outer first film cooling hole extends inward in the blade height direction from an inlet (e.g., the inlet 71 a) to an outlet (e.g., the outlet 71 b) of the outer first film cooling hole,
the inner first film cooling hole extends outward in the blade height direction from an inlet (e.g., the inlet 71 a) to an outlet (e.g., the outlet 71 b) of the inner first film cooling hole,
the plurality of second film cooling holes include a plurality of outer second film cooling holes (e.g., the above-described plurality of outer second film cooling holes 721) located outside the first position in the blade height direction, and a plurality of inner second film cooling holes (e.g., the above-described plurality of inner second film cooling holes 722) located inside the first position in the blade height direction,
the outer second film cooling hole extends inward in the blade height direction from an inlet (e.g., the inlet 72 a) to an outlet (e.g., the outlet 72 b) of the outer second film cooling hole,
The inner second film cooling hole extends outward in the blade height direction from the inlet (e.g., the inlet 72a described above) to the outlet (e.g., the outlet 72b described above) of the inner second film cooling hole.

The gas turbine vane described in (9) above allows for effective film cooling of the outer surface of the leading edge portion with a smaller amount of cooling air, compared to when the outer first film cooling holes and the inner first film cooling holes each extend along the thickness direction of the leading edge portion. Furthermore, because the outer first film cooling holes extend inward in the blade height direction from their inlets to their outlets, when the outer first film cooling holes are formed from the outer surface of the airfoil portion using a jig by, for example, laser processing, interference between the jig and the outer shroud and the inner shroud can be reduced, compared to when the outer first film cooling holes extend outward in the blade height direction from their inlets to their outlets. In addition, since the inner first film cooling holes extend outward in the blade height direction from their inlets to their outlets, when the inner first film cooling holes are formed from the outer surface of the airfoil by laser processing or the like using a jig, interference between the jig and each of the outer shroud and the inner shroud can be suppressed, compared to when the inner first film cooling holes extend inward in the blade height direction from their inlets to their outlets. Furthermore, compared to when the outer second film cooling holes and the inner second film cooling holes extend along the thickness direction of the leading edge portion, film cooling of the outer surface of the leading edge portion can be effectively performed with a smaller amount of cooling air. [0013] Furthermore, because the outer second film cooling holes extend inward in the blade height direction from their inlets to their outlets, when the outer second film cooling holes are formed using a jig from the outer surface side of the airfoil by, for example, laser machining, interference between the jig and each of the outer shroud and the inner shroud can be reduced, compared to when the outer second film cooling holes extend outward in the blade height direction from their inlets to their outlets. Furthermore, because the inner second film cooling holes extend outward in the blade height direction from their inlets to their outlets, when the inner second film cooling holes are formed using a jig from the outer surface side of the airfoil, interference between the jig and each of the outer shroud and the inner shroud can be reduced, compared to when the inner second film cooling holes extend inward in the blade height direction from their inlets to their outlets.

(10) In some embodiments, in the gas turbine vane according to any one of (9) above,
The gas turbine stator blade comprises:
an outer shroud (e.g., the outer shroud 22 described above) connected to an outer end of the airfoil in the blade height direction;
an inner shroud (e.g., the inner shroud 24 described above) connected to an inner end of the airfoil portion in the blade height direction;
Including,
In a cross section (e.g., the cross section shown in FIG. 6 ) including the axis of the first film cooling hole and the blade height direction, a protrusion amount of the outer shroud from the leading edge portion in a direction perpendicular to the blade height direction is defined as A1, and a protrusion amount of the inner shroud from the leading edge portion in the direction perpendicular to the blade height direction is defined as B1,
In a cross section including the axis of the second film cooling hole and the blade height direction (for example, the cross section shown in FIG. 9 ), a protrusion amount of the outer shroud from the leading edge portion in a direction perpendicular to the blade height direction is defined as A2, and a protrusion amount of the inner shroud from the leading edge portion in the direction perpendicular to the blade height direction is defined as B2,
In a cross section including the axis of the third film cooling hole and the blade height direction (for example, the cross section shown in FIG. 12 ), when the amount of protrusion of the outer shroud from the leading edge portion in the direction perpendicular to the blade height direction is A3 and the amount of protrusion of the inner shroud from the leading edge portion in the direction perpendicular to the blade height direction is B3,
If one of the outer shroud and the inner shroud that has the largest protrusion amount among the protrusion amounts A1, A2, A3, B1, B2, and B3 is defined as a large shroud (for example, the above-mentioned large shroud 90), and the other of the outer shroud and the inner shroud that is not the large shroud is defined as a small shroud (for example, the above-mentioned small shroud 91), the distance between the first position in the blade height direction and the large shroud (for example, the above-mentioned distance V90) is greater than the distance between the first position in the blade height direction and the small shroud (for example, the above-mentioned distance V91).

When forming each of the first through third film cooling holes from the outer surface of the airfoil using a jig, for example, by laser processing, the large shroud, which protrudes the most from the leading edge in a direction perpendicular to the blade height direction, is more likely to interfere with the jig than the small shroud. Therefore, as described in (10) above, by making the distance between the first position in the blade height direction and the large shroud greater than the distance between the first position in the blade height direction and the small shroud, each of the first through third film cooling holes can be formed while suppressing interference between the jig and the outer shroud and the inner shroud.

(11) In some embodiments, in the gas turbine vane according to any one of (7) to (10),
The outer third film cooling hole is
a circular hole portion (e.g., the circular hole portion 82 described above) having a circular cross-sectional shape in the inlet side portion of the outer third film cooling hole;
a shaped hole portion (e.g., the above-described shaped hole portion 83) that is connected to the circular hole portion and that is located on the outlet side of the outer third film cooling hole and has a cross-sectional area that increases toward the outlet of the outer third film cooling hole;
Including,
The inner third film cooling hole is
a circular hole portion (e.g., the circular hole portion 84 described above) having a circular cross-sectional shape in the inlet side portion of the inner third film cooling hole;
a shaped hole portion (e.g., the above-described shaped hole portion 85) that is connected to the circular hole portion and that is located on the outlet side of the inner third film cooling hole and has a cross-sectional area that increases toward the outlet of the inner third film cooling hole;
Including,
a center of the outlet of the outer third film cooling hole (e.g., center C3b in FIG. 13A ) is located inside, in the blade height direction, a point of intersection (e.g., the above-mentioned point of intersection S31) between an extension line of an axis of the circular hole portion of the outer third film cooling hole (e.g., the above-mentioned axis L82) and the outlet of the outer third film cooling hole,
The center of the outlet of the inner third film cooling hole (e.g., center C3b in Figure 13B) is located outside in the blade height direction of the intersection (e.g., the above-mentioned intersection S32) between an extension of the axis of the circular hole portion of the inner third film cooling hole (e.g., the above-mentioned axis L84) and the outlet of the inner third film cooling hole.

According to the gas turbine vane described in (11) above, film cooling of the outer surface of the leading edge portion can be effectively performed with a smaller amount of cooling air compared to when the center of the outlet of the outer third film cooling hole and the center of the outlet of the inner third film cooling hole are located on the extension of the axis of the circular hole portion of the outer third film cooling hole and the extension of the axis of the circular hole portion of the inner third film cooling hole, respectively, and a gas turbine vane can be provided that can suppress a decrease in the film efficiency of film cooling on the blade surface while reducing the amount of cooling air.

(12) In some embodiments, in the gas turbine stator vane according to any one of (9) to (11),
The outer first film cooling hole is
a circular hole portion (e.g., the circular hole portion 74 described above) having a circular cross-sectional shape in the inlet side portion of the outer first film cooling hole;
a shaped hole portion (e.g., the above-described shaped hole portion 75) that is connected to the circular hole portion and that is located on the outlet side of the outer first film cooling hole and has a cross-sectional area that increases toward the outlet of the outer first film cooling hole;
Including,
The inner first film cooling hole is
a circular hole portion (e.g., the circular hole portion 76 described above) having a circular cross-sectional shape in the inlet side portion of the inner first film cooling hole;
a shaped hole portion (e.g., the above-described shaped hole portion 77) that is located at the outlet side of the inner first film cooling hole and connects to the circular hole portion of the inner first film cooling hole and has a cross-sectional area that increases toward the outlet of the inner first film cooling hole;
Including,
a center of the outlet of the outer first film cooling hole (e.g., center C1b in FIG. 7A ) is located inside, in the blade height direction, of an intersection (e.g., the above-mentioned intersection S11) between an extension line of an axis of the circular hole portion of the outer first film cooling hole (e.g., the above-mentioned axis L74) and the outlet of the outer first film cooling hole,
a center of the outlet of the inner first film cooling hole (e.g., a center C1b in FIG. 7B ) is located outward in the blade height direction from an intersection (e.g., the above-mentioned intersection S12) between an extension line of an axis of the circular hole portion of the inner first film cooling hole (e.g., the above-mentioned axis L76) and the outlet of the inner first film cooling hole,
The outer second film cooling hole is
a circular hole portion (e.g., the circular hole portion 78 described above) having a circular cross-sectional shape in the inlet side portion of the outer second film cooling hole;
a shaped hole portion (e.g., the above-described shaped hole portion 79) that is located at the outlet side of the outer second film cooling hole and that is connected to the circular hole portion of the outer second film cooling hole and has a cross-sectional area that increases toward the outlet of the outer second film cooling hole;
Including,
The inner second film cooling hole is
a circular hole portion (e.g., the circular hole portion 80 described above) having a circular cross-sectional shape in the inlet side portion of the inner second film cooling hole;
a shaped hole portion (e.g., the above-described shaped hole portion 81) that is located at the outlet side of the inner second film cooling hole and is connected to the circular hole portion of the inner second film cooling hole and whose cross-sectional area increases toward the outlet of the inner second film cooling hole;
Including,
a center of the outlet of the outer second film cooling hole (e.g., center C2b in FIG. 10A ) is located inside, in the blade height direction, of an intersection (e.g., the above-mentioned intersection S21) between an extension line of an axis of the circular hole portion of the outer second film cooling hole (e.g., the above-mentioned axis L78) and the outlet of the outer second film cooling hole,
The center of the outlet of the inner second film cooling hole (e.g., center C2b in Figure 10B) is located outward in the blade height direction from the intersection (e.g., the above-mentioned intersection S22) between an extension of the axis of the circular hole portion of the inner second film cooling hole (e.g., the above-mentioned axis L80) and the outlet of the inner second film cooling hole.

According to the gas turbine stator vane described in (12) above, compared to when the outlet centers of the outer first film cooling holes and the inner first film cooling holes are located on the extensions of the axes of the circular holes of the outer first film cooling holes and the inner first film cooling holes, respectively, a gas turbine stator vane can be provided that can effectively perform film cooling of the outer surface of the leading edge portion with a smaller amount of cooling air, thereby suppressing a decrease in film efficiency of film cooling of the blade surface while reducing the amount of cooling air. Furthermore, compared to when the outlet centers of the outer second film cooling holes and the inner second film cooling holes are located on the extensions of the axes of the circular holes of the outer second film cooling holes and the inner second film cooling holes, respectively, a gas turbine stator vane can be provided that can effectively perform film cooling of the outer surface of the leading edge portion with a smaller amount of cooling air, thereby suppressing a decrease in film efficiency of film cooling of the blade surface while reducing the amount of cooling air.

(13) In some embodiments, in the gas turbine stator vane described in (12),
a first film cooling hole having a cross-sectional area enlarged only inward in a blade height direction relative to the circular hole portion of the first film cooling hole when viewed in an axial direction of the circular hole portion of the first film cooling hole,
a first film cooling hole having a cross-sectional area enlarged only outward in a blade height direction relative to the circular hole portion of the first film cooling hole when viewed in an axial direction of the circular hole portion of the first film cooling hole,
a cross-sectional area of the shaped hole portion of the outer second film cooling hole is expanded only inward in the blade height direction relative to the circular hole portion of the outer second film cooling hole when viewed in the axial direction of the circular hole portion of the outer second film cooling hole,
a second cooling film hole having a cross-sectional area expanded only outward in the blade height direction relative to the circular hole portion of the second cooling film hole when viewed in the axial direction of the circular hole portion of the second cooling film hole,
a cross-sectional area of the shaped hole portion of the outer third film cooling hole is expanded only inward in the blade height direction relative to the circular hole portion of the outer third film cooling hole when viewed in the axial direction of the circular hole portion of the outer third film cooling hole,
The shaped hole portion of the inner third film cooling hole has a cross-sectional area expanded only outward in the blade height direction relative to the circular hole portion of the inner third film cooling hole when viewed in the axial direction of the circular hole portion of the inner third film cooling hole.

According to the gas turbine vane described in (13) above, by expanding the cross-sectional area of the shaped hole portion of each of the outer first film cooling hole, outer second film cooling hole, and outer third film cooling hole only inward in the blade height direction relative to the circular hole portion, each of the outer first film cooling hole, outer second film cooling hole, and outer third film cooling hole can be easily formed by laser processing, etc. Furthermore, by expanding the cross-sectional area of the shaped hole portion of each of the inner first film cooling hole, inner second film cooling hole, and inner third film cooling hole only outward in the blade height direction relative to the circular hole portion, each of the inner first film cooling hole, inner second film cooling hole, and inner third film cooling hole can be easily formed by laser processing, etc. This facilitates the manufacture of the gas turbine vane.

(14) A gas turbine according to at least one embodiment of the present disclosure includes:
a compressor (e.g., the compressor 4 described above), a combustor (e.g., the combustor 6 described above), and a turbine (e.g., the turbine 8 described above);
The turbine includes the gas turbine vane according to any one of (1) to (13) above.

According to the gas turbine stator vane described in (14) above, since it is equipped with a gas turbine stator vane described in any of (1) to (13) above, film cooling of the outer surface of the leading edge can be effectively performed with a small amount of cooling air, and a gas turbine stator vane can be provided that can suppress a decrease in the film efficiency of film cooling on the blade surface while reducing the amount of cooling air.

2 Gas turbine 4 Compressor 6 Combustor 8 Turbine 9 Rotor 10 Turbine casing 12 Turbine stator vane 16 Turbine rotor blade 20 Airfoil section 20a Outer end 20b Inner end 21 Leading edge section 22 Outer shroud 24 Inner shroud 25 Outer surface 26, 39, 45 Inner surface 31 Flow passage 32 Outer cavity 33 Outer peripheral wall 34 Inner cavity 35 Inner peripheral wall 36 Suction surface 38 Suction surface forming wall 40 Pressure surface 42 Inner blade cavity 44 Pressure surface forming wall 46 Leading edge side cavity 48 Leading edge partition wall 49 Trailing edge side cavity 50 Insert 51 Internal space 52 Impingement cooling hole 71 First film cooling hole 711 Outer first film cooling hole 712 Inner first film cooling hole 72 Second film cooling hole 721 Outer second film cooling hole 722 Inner second film cooling holes 71a, 72a, 73a Inlets 71b, 72b, 73b Outlets 73 Third film cooling hole 731 Outer third film cooling hole 732 Inner third film cooling holes 74, 76, 78, 80, 82, 84 Circular hole sections 75, 77, 79, 81, 83, 85 Shaped hole sections 90, 92 Large shrouds 91, 93 Small shrouds

Claims (13)

A gas turbine stator vane,
an airfoil portion having a cavity formed therein;
The leading edge of the airfoil has:
a first film cooling hole row including a plurality of first film cooling holes arranged along the blade height direction;
a second film cooling hole row including a plurality of second film cooling holes arranged along the blade height direction;
a third film cooling hole row including a plurality of third film cooling holes arranged along the blade height direction;
is formed,
the second row of film cooling holes is disposed between the first row of film cooling holes and the third row of film cooling holes;
the first film cooling hole extends in a direction inclined with respect to a thickness direction of the leading edge portion, and when a plane including the thickness direction of the leading edge portion and the blade height direction at the center of the inlet of the first film cooling hole is defined as a first plane, the center of the outlet of the first film cooling hole is located downstream of a position where the first plane intersects with an outer surface of the leading edge portion in a flow of combustion gas along the leading edge portion,
the second film cooling hole extends in a direction inclined with respect to a thickness direction of the leading edge portion, and when a plane including the thickness direction of the leading edge portion and the blade height direction at the center of an inlet of the second film cooling hole is defined as a second plane, an outlet of the second film cooling hole is located at a position where the second plane intersects with an outer surface of the leading edge portion,
the third film cooling hole extends in a direction inclined with respect to a thickness direction of the leading edge portion, and when a plane including the thickness direction of the leading edge portion and the blade height direction at the center of the inlet of the third film cooling hole is defined as a third plane, the center of the outlet of the third film cooling hole is located downstream of a position where the third plane intersects with an outer surface of the leading edge portion in a flow of combustion gas along the leading edge portion,
the plurality of third film cooling holes include a plurality of outer third film cooling holes located outward from a first position in the blade height direction, and a plurality of inner third film cooling holes located inward from the first position in the blade height direction,
the outer third film cooling hole extends inward in the blade height direction from an inlet to an outlet of the outer third film cooling hole,
the inner third film cooling hole extends outward in the blade height direction from an inlet to an outlet of the inner third film cooling hole,
The gas turbine stator blade comprises:
an outer shroud connected to an outer end of the airfoil portion in the blade height direction;
an inner shroud connected to an inner end of the airfoil portion in the blade height direction;
Including,
a gas turbine stator vane, wherein, in a cross section including an axis of the third film cooling hole and the blade height direction, one of the outer shroud and the inner shroud that protrudes a greater amount from a portion of the airfoil where the third film cooling hole row is formed in a direction perpendicular to the blade height direction is defined as a large shroud, and the other of the outer shroud and the inner shroud that is not the large shroud is defined as a small shroud, the distance between the first position in the blade height direction and the large shroud is greater than the distance between the first position in the blade height direction and the small shroud . A gas turbine vane according to claim 1, wherein each of the first film cooling hole, the second film cooling hole, and the third film cooling hole is located closer to the pressure surface than the camber line in a cross section of the airfoil portion perpendicular to the blade height direction. A gas turbine vane according to claim 2, wherein, in a cross section of the airfoil perpendicular to the blade height direction, the intersection of the camber line and the outer surface of the leading edge is defined as E1, and the most upstream position of the airfoil in the axial direction of the gas turbine is defined as E2. The outlet of the first film cooling hole is located between the intersection E1 and the most upstream position E2 on the outer surface of the airfoil, and the outlet of the third film cooling hole is located on the opposite side of the most upstream position E2 on the outer surface of the airfoil from the intersection E1. A gas turbine vane according to claim 3, wherein, in a cross section of the airfoil perpendicular to the blade height direction, the outlet of the second film cooling hole is located on the opposite side of the most upstream position E2 on the outer surface of the airfoil from the intersection point E1. The airfoil portion is
a pressure surface forming wall that forms a pressure surface;
a negative pressure surface forming wall that forms a negative pressure surface;
a leading edge partition extending from the inner surface of the suction surface wall to the inner surface of the pressure surface wall so as to divide the cavity within the airfoil,
2. The gas turbine vane according to claim 1, wherein, in a cross section of the airfoil portion perpendicular to the blade height direction, each of the first film cooling hole, the second film cooling hole, and the third film cooling hole is located upstream of a position where the pressure surface forming wall and the leading edge partition are connected in an axial direction of the gas turbine. A gas turbine vane according to claim 1, wherein each of the first film cooling hole, the second film cooling hole, and the third film cooling hole extends along a direction inclined with respect to a plane perpendicular to the blade height direction. the plurality of first film cooling holes include a plurality of outer first film cooling holes located outward from the first position in the blade height direction, and a plurality of inner first film cooling holes located inward from the first position in the blade height direction,
the outer first film cooling hole extends inward in the blade height direction from an inlet to an outlet of the outer first film cooling hole,
the inner first film cooling hole extends outward in the blade height direction from an inlet to an outlet of the inner first film cooling hole,
the plurality of second film cooling holes include a plurality of outer second film cooling holes located outward from the first position in the blade height direction, and a plurality of inner second film cooling holes located inward from the first position in the blade height direction,
the outer second film cooling hole extends inward in the blade height direction from an inlet to an outlet of the outer second film cooling hole,
The gas turbine vane according to claim 1 , wherein the inner second film cooling hole extends outward in the blade height direction from an inlet to an outlet of the inner second film cooling hole. The gas turbine stator blade comprises:
an outer shroud connected to an outer end of the airfoil portion in the blade height direction;
an inner shroud connected to an inner end of the airfoil portion in the blade height direction;
Including,
In a cross section including the axis of the first film cooling hole and the blade height direction, a protrusion amount of the outer shroud from a portion of the airfoil where the first film cooling hole row is formed in a direction perpendicular to the blade height direction is defined as A1, and a protrusion amount of the inner shroud from a portion of the airfoil where the first film cooling hole row is formed in a direction perpendicular to the blade height direction is defined as B1,
In a cross section including the axis of the second film cooling hole and the blade height direction, a protrusion amount of the outer shroud from a portion of the airfoil where the second film cooling hole row is formed in a direction perpendicular to the blade height direction is defined as A2, and a protrusion amount of the inner shroud from a portion of the airfoil where the second film cooling hole row is formed in a direction perpendicular to the blade height direction is defined as B2,
In a cross section including the axis of the third film cooling hole and the blade height direction, when the amount of protrusion of the outer shroud from the portion of the airfoil where the third film cooling hole row is formed in the direction perpendicular to the blade height direction is defined as A3 and the amount of protrusion of the inner shroud from the portion of the airfoil where the third film cooling hole row is formed in the direction perpendicular to the blade height direction is defined as B3,
8. The gas turbine vane according to claim 7, wherein, when one of the outer shroud and the inner shroud having the largest protrusion amount among the protrusion amounts A1, A2, A3, B1, B2, and B3 is defined as a large shroud, and the other of the outer shroud and the inner shroud that is not the large shroud is defined as a small shroud, a distance between the first position in the blade height direction and the large shroud is larger than a distance between the first position in the blade height direction and the small shroud. The outer third film cooling hole is
a circular hole portion having a circular cross-sectional shape, the circular hole portion being an inlet side portion of the outer third film cooling hole;
a shaped hole portion, the shaped hole portion being connected to the circular hole portion and having a cross-sectional area increasing toward the outlet of the outer third film cooling hole;
Including,
The inner third film cooling hole is
a circular hole portion having a circular cross-sectional shape, the circular hole portion being an inlet side portion of the inner third film cooling hole;
a shaped hole portion, which is an outlet side portion of the inner third film cooling hole and is connected to the circular hole portion, and whose cross-sectional area increases toward the outlet of the inner third film cooling hole;
Including,
a center of the outlet of the outer third film cooling hole is located inward in the blade height direction with respect to an intersection between an extension of an axis of the circular hole portion of the outer third film cooling hole and the outlet of the outer third film cooling hole,
2. The gas turbine vane according to claim 1, wherein a center of the outlet of the inner third film cooling hole is located outward in the blade height direction from an intersection between an extension of an axis of the circular hole portion of the inner third film cooling hole and the outlet of the inner third film cooling hole. The outer first film cooling hole is
a circular hole portion having a circular cross-sectional shape, the circular hole portion being an inlet side portion of the outer first film cooling hole;
a shaped hole portion, the shaped hole portion being connected to the circular hole portion and having a cross-sectional area that increases toward the outlet of the outer first film cooling hole;
Including,
The inner first film cooling hole is
a circular hole portion having a circular cross-sectional shape, the circular hole portion being an inlet side portion of the inner first film cooling hole;
a first cooling film hole having a cross-sectional area that increases toward the outlet of the first cooling film hole;
Including,
a center of the outlet of the outer first film cooling hole is located inward in the blade height direction with respect to an intersection between an extension of an axis of the circular hole portion of the outer first film cooling hole and the outlet of the outer first film cooling hole,
a center of the outlet of the inner first film cooling hole is located outside, in the blade height direction, of an intersection between an extension of an axis of the circular hole portion of the inner first film cooling hole and the outlet of the inner first film cooling hole,
The outer second film cooling hole is
a circular hole portion having a circular cross-sectional shape, the circular hole portion being an inlet side portion of the outer second film cooling hole;
a shaped hole portion, the shaped hole portion being connected to the circular hole portion of the outer second film cooling hole at the outlet side of the outer second film cooling hole and having a cross-sectional area that increases toward the outlet of the outer second film cooling hole;
Including,
The inner second film cooling hole is
a circular hole portion having a circular cross-sectional shape, the circular hole portion being an inlet side portion of the inner second film cooling hole;
a shaped hole portion, the shaped hole portion being connected to the circular hole portion of the inner second film cooling hole at the outlet side of the inner second film cooling hole and having a cross-sectional area that increases toward the outlet of the inner second film cooling hole;
Including,
a center of the outlet of the outer second film cooling hole is located inward in the blade height direction with respect to an intersection between an extension of an axis of the circular hole portion of the outer second film cooling hole and the outlet of the outer second film cooling hole,
8. The gas turbine vane according to claim 7, wherein a center of the outlet of the inner second film cooling hole is located outward in the blade height direction from an intersection between an extension of an axis of the circular hole portion of the inner second film cooling hole and the outlet of the inner second film cooling hole. a first film cooling hole having a cross-sectional area enlarged only inward in a blade height direction relative to the circular hole portion of the first film cooling hole when viewed in an axial direction of the circular hole portion of the first film cooling hole,
a first film cooling hole having a cross-sectional area enlarged only outward in a blade height direction relative to the circular hole portion of the first film cooling hole when viewed in an axial direction of the circular hole portion of the first film cooling hole,
a cross-sectional area of the shaped hole portion of the outer second film cooling hole is expanded only inward in the blade height direction relative to the circular hole portion of the outer second film cooling hole when viewed in the axial direction of the circular hole portion of the outer second film cooling hole,
a second cooling film hole having a cross-sectional area expanded only outward in the blade height direction relative to the circular hole portion of the second cooling film hole when viewed in the axial direction of the circular hole portion of the second cooling film hole,
a cross-sectional area of the shaped hole portion of the outer third film cooling hole is expanded only inward in the blade height direction relative to the circular hole portion of the outer third film cooling hole when viewed in the axial direction of the circular hole portion of the outer third film cooling hole,
11. The gas turbine vane according to claim 10, wherein a cross-sectional area of the shaped hole portion of the inner third film cooling hole is expanded only outward in the blade height direction relative to the circular hole portion of the inner third film cooling hole when viewed in the axial direction of the circular hole portion of the inner third film cooling hole. A gas turbine stator vane,
an airfoil portion having a cavity formed therein;
The leading edge of the airfoil has:
a first film cooling hole row including a plurality of first film cooling holes arranged along the blade height direction;
a second film cooling hole row including a plurality of second film cooling holes arranged along the blade height direction;
a third film cooling hole row including a plurality of third film cooling holes arranged along the blade height direction;
is formed,
the second row of film cooling holes is disposed between the first row of film cooling holes and the third row of film cooling holes;
the first film cooling hole extends in a direction inclined with respect to a thickness direction of the leading edge portion, and when a plane including the thickness direction of the leading edge portion and the blade height direction at the center of the inlet of the first film cooling hole is defined as a first plane, the center of the outlet of the first film cooling hole is located downstream of a position where the first plane intersects with an outer surface of the leading edge portion in a flow of combustion gas along the leading edge portion,
the second film cooling hole extends in a direction inclined with respect to a thickness direction of the leading edge portion, and when a plane including the thickness direction of the leading edge portion and the blade height direction at the center of an inlet of the second film cooling hole is defined as a second plane, an outlet of the second film cooling hole is located at a position where the second plane intersects with an outer surface of the leading edge portion,
the third film cooling hole extends in a direction inclined with respect to a thickness direction of the leading edge portion, and when a plane including the thickness direction of the leading edge portion and the blade height direction at the center of the inlet of the third film cooling hole is defined as a third plane, the center of the outlet of the third film cooling hole is located downstream of a position where the third plane intersects with an outer surface of the leading edge portion in a flow of combustion gas along the leading edge portion,
the plurality of third film cooling holes include a plurality of outer third film cooling holes located outward from a first position in the blade height direction, and a plurality of inner third film cooling holes located inward from the first position in the blade height direction,
the outer third film cooling hole extends inward in the blade height direction from an inlet to an outlet of the outer third film cooling hole,
the inner third film cooling hole extends outward in the blade height direction from an inlet to an outlet of the inner third film cooling hole,
the plurality of first film cooling holes include a plurality of outer first film cooling holes located outward from the first position in the blade height direction, and a plurality of inner first film cooling holes located inward from the first position in the blade height direction,
the outer first film cooling hole extends inward in the blade height direction from an inlet to an outlet of the outer first film cooling hole,
the inner first film cooling hole extends outward in the blade height direction from an inlet to an outlet of the inner first film cooling hole,
the plurality of second film cooling holes include a plurality of outer second film cooling holes located outward from the first position in the blade height direction, and a plurality of inner second film cooling holes located inward from the first position in the blade height direction,
the outer second film cooling hole extends inward in the blade height direction from an inlet to an outlet of the outer second film cooling hole,
the inner second film cooling hole extends outward in the blade height direction from an inlet to an outlet of the inner second film cooling hole,
The gas turbine stator blade comprises:
an outer shroud connected to an outer end of the airfoil portion in the blade height direction;
an inner shroud connected to an inner end of the airfoil portion in the blade height direction;
Including,
In a cross section including the axis of the first film cooling hole and the blade height direction, a protrusion amount of the outer shroud from a portion of the airfoil where the first film cooling hole row is formed in a direction perpendicular to the blade height direction is defined as A1, and a protrusion amount of the inner shroud from a portion of the airfoil where the first film cooling hole row is formed in a direction perpendicular to the blade height direction is defined as B1,
In a cross section including the axis of the second film cooling hole and the blade height direction, a protrusion amount of the outer shroud from a portion of the airfoil where the second film cooling hole row is formed in a direction perpendicular to the blade height direction is defined as A2, and a protrusion amount of the inner shroud from a portion of the airfoil where the second film cooling hole row is formed in a direction perpendicular to the blade height direction is defined as B2,
In a cross section including the axis of the third film cooling hole and the blade height direction, when the amount of protrusion of the outer shroud from the portion of the airfoil where the third film cooling hole row is formed in the direction perpendicular to the blade height direction is defined as A3 and the amount of protrusion of the inner shroud from the portion of the airfoil where the third film cooling hole row is formed in the direction perpendicular to the blade height direction is defined as B3,
a gas turbine stator vane, wherein, of the outer shroud and the inner shroud, one having the largest protrusion amount among the protrusion amounts A1, A2, A3, B1, B2, B3 is defined as a large shroud, and the other of the outer shroud and the inner shroud that is not the large shroud is defined as a small shroud, a distance between the first position in the blade height direction and the large shroud is greater than a distance between the first position in the blade height direction and the small shroud. a compressor, a combustor, and a turbine;
A gas turbine, the turbine comprising a gas turbine vane according to any one of claims 1 to 12 .