Cooling method and cooling structure for gas turbine vanes

Abstract

The present disclosure provides a method for cooling static vanes of a turbine. The turbine comprises airfoils and a shroud arranged at radial end portions which are end portions of the airfoils along the radial direction of the turbine. The shroud includes a shroud body and a shroud end portion arranged around the outer periphery of the shroud body so as to surround the shroud body. The shroud end portion internally includes a shroud-end-portion flow passage. After cooling air is caused to flow inside the shroud-end-portion flow passage to cool the shroud end portion, the cooling air which has flowed inside the shroud-end-portion flow passage is used to cool the shroud body.

Description

[Technical Field]

[0001] The present disclosure relates to a cooling method for a stator vane of a gas turbine, and also to a cooling structure for a stator vane of a gas turbine. [Background technology]

[0002] Gas turbine stator vanes and rotor blades are exposed to high-temperature combustion gases. Therefore, the stator vanes and rotor blades need to be cooled with cooling air. FIG. 1 schematically illustrates a conventional structure for cooling the airfoils of gas turbine stator vanes. For example, in a conventional structure for cooling the airfoils of gas turbine stator vanes, as shown in FIG. 1, cooling air supplied to the airfoil insert of a first-stage stator vane is injected toward the inner surface of the airfoil through impingement cooling holes provided in the insert, impingement cooling the inner surface of the airfoil, and then discharged through film cooling holes in the airfoil to generate film cooling air that flows along the outer surface of the airfoil. In other words, the cooling air used for impingement cooling is discharged into the hot gas flow path through the film cooling holes (the discharge of the cooling air is illustrated by the arrows in FIG. 1).

[0003] FIG. 2 schematically illustrates a conventional structure for cooling a shroud of a gas turbine. The stationary vane includes a shroud structure that is cooled by cooling air. As shown in FIG. 2, the cooling air is taken in from the shroud body at the side edges of the shroud (the shroud ends) and flows along the side edges of the shroud toward the trailing edge of the shroud. The cooling air used to cool the shroud ends is discharged from the trailing edge of the shroud into the hot gas flow path (the discharge of the cooling air is illustrated by the arrows in FIG. 2). [Prior art documents] [Patent documents]

[0004] [Patent Document 1] Japanese Patent No. 6799702 [Patent Document 2] International Publication No. WO / 003590 Summary of the Invention [Problem to be solved by the invention]

[0005] In recent years, gas turbine inlet temperatures have increased, and therefore, there is a demand for further cooling of first-stage stator vanes. One approach to addressing this issue is to supply cooling air at higher pressure and lower temperature (compared to conventional techniques) to the first-stage stator vanes. The inventors have found that if higher-pressure, lower-temperature cooling air is used to cool the first-stage stator vanes, it may be possible to reuse the cooling air to cool other components of the first-stage stator vanes, even after it has been used to cool the airfoil or shroud end. However, in conventional techniques, the cooling air in the first-stage stator vanes is discharged into the hot gas flow path. This limits the efficiency of cooling air utilization.

[0006] It is desirable to provide a cooling method or structure for gas turbine stator vanes that can increase the efficiency of cooling air usage. [Means for solving the problem]

[0007] According to a first aspect of the present disclosure, there is provided a method for cooling a turbine vane, the vane including an airfoil and a shroud disposed at a radial end of the airfoil along a radial direction of the turbine, the shroud including a shroud body and a shroud end portion disposed around the shroud body to surround the shroud body and including a shroud end flow passage therein, the method including the following steps: (i) cooling the shroud end by flowing cooling air through the shroud end passage; (ii) After cooling the shroud end portion, the shroud body is cooled using the cooling air that has flowed through the inside of the shroud end flow passage.

[0008] The above-described features allow the cooling air used to cool the shroud end to be used to cool other components of the vane, such as the shroud body, without being discharged into the high-temperature gas flow path. This makes it possible to improve the efficiency of cooling air use. Furthermore, the shroud end, which is more severely exposed to high-temperature gas, can be first cooled with low-temperature cooling air, and the shroud body can then be cooled using that cooling air. This makes it possible to improve the efficiency of cooling air use.

[0009] In the first aspect, the method may further include a step of flowing cooling air inside the airfoil, and step (i) may further include, after cooling the airfoil, using the cooling air that has flowed inside the airfoil to flow the cooling air into the shroud end flow passage to cool the shroud end. According to the above-mentioned feature, the cooling air used to cool the airfoil can be used to cool other components of the vane, such as the shroud end, without being discharged into the hot gas flow passage. This makes it possible to improve the efficiency of use of the cooling air.

[0012] According to a third aspect of the present disclosure, there is provided a turbine vane including an airfoil and a shroud disposed at an end of the airfoil, the end being located at a radial end along a radial direction of the turbine. The shroud includes a shroud body having a first wall facing a hot gas flow path of the turbine and a second wall arranged on the opposite side of the first wall from the hot gas flow path, and a shroud end portion arranged around the shroud body to surround the shroud body and including a shroud end flow path therein. The shroud end includes a cooling air inlet for introducing cooling air into the shroud end flow passage and a cooling air outlet for exiting the cooling air from the shroud end flow passage. The shroud body includes a hollow space between the first wall and the second wall, and the hollow space is connected to the shroud end flow passage through the cooling air outlet.

[0013] With the above-described features, the cooling air introduced into the shroud end flow passage through the cooling air inlet and used to cool the shroud end flows through the cooling air outlet into the hollow space of the shroud body, and can be used to cool the shroud body without being discharged into the high-temperature gas flow passage. This makes it possible to improve the efficiency of cooling air use. Furthermore, the shroud end, which is more severely exposed to high-temperature gas, can be first cooled with low-temperature cooling air, and then the shroud body can be cooled using that cooling air. This makes it possible to improve the efficiency of cooling air use. [Brief explanation of the drawings]

[0014] [Figure 1] FIG. 1 is a diagram illustrating a conventional structure for cooling the airfoil of a gas turbine vane. [Figure 2] FIG. 2 is a diagram that schematically illustrates a conventional structure for cooling the shroud of a gas turbine stator blade. [Figure 3] FIG. 3 is a schematic cross-sectional view of a gas turbine according to an embodiment of the present disclosure. [Figure 4] FIG. 4 is a perspective view of the stator blade in the first embodiment. [Figure 5] FIG. 5 is a cross-sectional view taken along line VV in FIG. [Figure 6] FIG. 6 is a partially enlarged view of the stationary blade. [Figure 7] FIG. 7 is a partial perspective view of the stator blade in the first embodiment. [Figure 8] FIG. 8 is a flowchart illustrating the cooling method for the stationary blade according to the first embodiment. [Figure 9] FIG. 9 is a flowchart illustrating the cooling method for the stationary blade according to the second embodiment. [Figure 10] FIG. 10 is a diagram for explaining the cooling step of the second embodiment. [Figure 11] FIG. 11 is a flowchart illustrating a method for cooling a stationary blade according to the third embodiment. [Figure 12]FIG. 12 is a schematic cross-sectional view of a stator vane according to the fourth embodiment. DETAILED DESCRIPTION OF THE INVENTION

[0015] Embodiments of the present disclosure will be described in detail below with reference to the drawings. Fig. 3 is a schematic cross-sectional view of a gas turbine according to an embodiment of the present disclosure. As shown in Fig. 3, a gas turbine 10 according to this embodiment includes a turbine 20 driven by combustion gas generated by a combustor 30. The turbine 20 includes a rotor shaft 24, a turbine rotor 26 that rotates about an axis Ar, a turbine casing 22 that covers the turbine rotor 26, and multiple stages of stator vanes 28.

[0016] FIG. 4 schematically illustrates a stator vane of a gas turbine according to an embodiment of the present disclosure. FIG. 4 is a perspective view of the stator vane according to a first embodiment. FIG. 5 is a cross-sectional view taken along line VV in FIG. 4. FIG. 6 is a partially enlarged view of the stator vane. As shown in FIG. 4, the stator vane 50 includes a stator vane body (airfoil) 51 extending in the radial direction of the gas turbine, an inner shroud 60 disposed radially inside the stator vane body 51, and an outer shroud 70 disposed radially outside the stator vane body 51. The stator vane body 51 is disposed in a combustion gas flow path (high-temperature gas flow path) through which combustion gas passes. In general, the annular combustion gas flow path is defined radially inside by the inner shroud 60 and radially outside by the outer shroud 70. The inner shroud 60 and the outer shroud 70 are plate-shaped members that define part of the combustion gas flow path.

[0017] 4, the upstream end of the stator vane body 51 has a leading edge 52, and the downstream end of the stator vane body 51 has a trailing edge 53. Of the surfaces of the stator vane body 51, the convex surface is the suction side surface 54 (suction surface), and the concave surface is the pressure side surface 55 (pressure surface). For convenience, in the following description, the pressure side (pressure surface side) of the stator vane body 51 and the pressure side (suction surface side) of the stator vane body 51 will be referred to as the pressure side and the pressure side, respectively.

[0018] The inner shroud 60 and the outer shroud 70 have basically the same structure, and therefore, the outer shroud 70 will be mainly described below.

[0019] 4 and 5, the outer shroud 70 is a plate-shaped shroud member and includes a shroud body 72, a shroud end portion 74 disposed on the outer periphery of the shroud body 72, and a peripheral wall 76 extending along the shroud end portion 74. The peripheral wall 76 protrudes from the shroud body 72 toward the radially outer side of the gas turbine.

[0020] The outer shroud 70 has a forward end face that is an upstream end face, an aft end face that is a downstream end face, a ventral end face that is a downstream end face, and a ventral end face that is a downstream end face. The outer shroud 70 has a gas path surface 78 that faces radially inward and faces the hot gas flow path. The forward end face and the aft end face are substantially parallel to each other, and the ventral end face and the suction end face are substantially parallel to each other. Therefore, when viewed radially, the outer shroud 70 has a substantially parallelogram shape, as shown in FIG. 5 .

[0021] The shroud end 74 is a flange-like or edge-like structure that protrudes from the shroud body 72. The shroud end 74 is a forward shroud end 74 located upstream of the outer shroud 70. L and an aft shroud end 74 located downstream of the outer shroud 70. T and a rear shroud end 74 disposed on the rear side of the outer shroud 70. N and a ventral shroud end 74 disposed on the ventral side of the outer shroud 70. P For example, as shown in FIG. 5, the front shroud end 74 L , rear shroud end 74 T , dorsal shroud end 74 N , and the ventral shroud end 74 P is disposed on the outer periphery of the shroud body 72 and surrounds the entire shroud body 72.

[0022] Forward shroud end 74 L The front shroud end flow passage 75 is formed therein.L aft shroud end 74 T The rear shroud end flow passage 75 is disposed therein. T back shroud end 74 N The rear shroud end flow passage 75 is formed therein. N ventral shroud end 74 P The inner portion of the flow passage 75 is a ventral shroud end flow passage. P Includes.

[0023] In this embodiment, the forward shroud end flow passage 75 L At one end thereof, the rear shroud end flow passage 75 N and the other end of the flow passage 75 P The rear shroud end flow passage 75 is connected to the rear shroud end flow passage 75. T At one end thereof, the rear shroud end flow passage 75 N and the other end of the flow passage 75 P As shown in FIGS. 4, 5, and 6, the front shroud end passage 75 L has a shroud end flow passage inlet 171. T The forward shroud end flow passage 75 has a shroud end flow passage outlet 172. L A portion of the cooling air flowing into the suction side shroud end passage 75 N and ventral shroud end flow passage 75 P and then through the aft shroud end flow passage 75 T and exits through the shroud end flow passage outlet 172. As shown in FIG. L , 75 T , 75 P , 75 NThe shroud end passage includes turbulators 175. The turbulators 175 may be ribs disposed on the inner surface of the shroud end passage. To enhance cooling of the shroud end, the turbulators 175 may be disposed on a bottom surface of the passage that defines the radially inner surface of the passage. Here, the bottom surface of the passage may extend substantially parallel to the radially inner wall 81. The turbulators 175 may also be disposed on a side surface of the passage that defines the circumferential or axial side wall of the passage.

[0024] In this embodiment, the shroud end channel inlet 171 is connected to the forward shroud end channel 75 L The shroud end flow passage outlet 172 is provided at the aft shroud end flow passage 75 T However, the structure of the stator vane is not limited to this embodiment. The shroud end flow passage inlet 171 is provided in the suction side shroud end flow passage 75. N , ventral shroud end flow passage 75 P , or aft shroud end flow passage 75 T The shroud end flow passage outlet 172 may be provided in other shroud end flow passages such as the suction side shroud end flow passage 75. N , ventral shroud end flow passage 75 P , or forward shroud end flow passage 75 L Alternatively, multiple shroud end channel inlets 171 may be provided in one or more shroud end channels 75. L , 75 T , 75 N , 75 P Also, multiple shroud end flow passage outlets 172 may be provided in one or more shroud end flow passages 75. L , 75 T , 75 N , 75 P It may also be provided in.

[0025] The shroud body 72 includes a radially inner wall 81 and a radially outer wall 82 located on the opposite side thereof. The shroud body 72 includes a hollow space S between the radially inner wall 81 and the radially outer wall 82. The radially inner surface of the inner wall 81 forms the gas path surface 78 of the outer shroud 70. This radially inner wall 81 forms a part of the shroud body 72. This radially inner wall 81 may be continuously extended in the circumferential direction or the axial direction of the gas turbine so as to form a part of the shroud end 74. FIG. 4 shows, as an example, a configuration in which the radially inner wall 81 continuously extends in the axial direction of the gas turbine to form the aft shroud end 74. T The shroud body 72 includes an impingement plate 73 that divides the space S of the outer shroud 70 into an outer region on the radially outer side and an inner region (cavity) on the radially inner side. In the outer region, a part of the cooling air flows through the aft shroud end flow passage 75. T The inner region is defined between the radially inner wall 81 of the outer shroud 70 and the impingement plate 73.

[0026] In the impingement plate 73, a plurality of impingement cooling holes 79 are provided so as to penetrate the impingement plate 73 in the radial direction. A portion of the cooling air present in the outer region flows into the inner region through the impingement cooling holes 79 of the impingement plate 73. This cooling air is injected toward the radially outer surface of the radially inner wall 81, impingement cools the radially outer surface of the radially inner wall 81, and then passes through the outer wall 82 to be discharged to the outside. For example, the cooling air injected from the impingement cooling holes 79 toward the radially outer surface of the radially inner wall 81 in order to impingement cool the radially outer surface of the radially inner wall 81 is discharged through a passage that connects the inner region of the space S to an outer space located on the opposite side (outside) of the outer wall 82 from the space S. Such a passage may be isolated from the outer region of the space S. More specifically, in the present embodiment, the cooling air is discharged through a hole in the discharge pipe 83. The discharge pipe 83 is provided to pass through the radially outer wall 82 and the impingement plate 73 in a manner connecting the inner region with the exterior space.

[0027] The inside of the stator blade body 51 is formed with a partition wall 51 extending in the radial direction. P The outer shroud 70 is partitioned into a plurality of dividing regions 141, 142, 143 by the stator vane 51. A plurality of inserts 151, 152, 153 are inserted into the respective dividing regions 141, 142, 143. The plurality of inserts 151, 152, 153 include radially extending air channels 161, 162, 163, respectively, and extend radially from the outer shroud 70 through the stator vane body 51 toward the inner shroud 60. Each of the inserts 151, 152, 153 is formed continuously from the outer shroud 70 through the stator vane body 51 to the inner shroud 60. Each of the air channels 161, 162, 163 has an air intake port 58 that opens to the inside of the intake manifold 56.

[0028] Each insert 151, 152, 153 has a plurality of holes (through-holes) 59 communicating with the air channels 161, 162, 163, respectively. A portion of the cooling air supplied to the air channels 161, 162, 163 of the inserts 151, 152, 153 is injected from the plurality of holes 59 toward the inner surface of the vane body 51 to impingement-cool the inner surface of the airfoil 51. Each of the plurality of dividing regions 141, 142, 143 has an outer air channel defined between the insert 151, 152, 153 and the inner surface of the vane body 51. A portion of the cooling air injected through the holes 59 is guided by the outer air channel and flows through the outer air channel radially outward, radially inward, or radially outward and inward. As an example, FIG. 5 shows the outer air channel 57 provided between the side of the insert 151 and the inner surface of the front end of the vane body 51.

[0029] The intake manifold 56 and the discharge pipe 83 are connected to a forced air cooling system in which cooling air drawn from the inside of the combustor casing is cooled by an external cooler (not shown) and then compressed by an external compressor (not shown). The compressed air is used for cooling and then returned to the inside of the combustor casing. In the above description, an example in which an air-cooling system is applied to this embodiment has been described. However, the present vane is not limited to this embodiment. The present disclosure may also be applied to other types of cooling systems. For example, the intake manifold 56 and the discharge pipe 83 may be connected to a closed-loop steam cooling system or a closed-loop air-cooling system.

[0030] For example, in the insert 151, which is a leading end insert, a portion of the cooling air supplied to the air channel 161 through the air intake 58 is injected toward the inner surface of the leading end of the airfoil 51 and then flows radially outward through the outer air channel 57. The outer air channel 57, which is the space between the inner surface of the leading end of the vane body 51 and the insert 151, is connected to the leading shroud end flow passage 75. L A portion of the cooling air injected toward the inner surface of the forward end of the airfoil 51 passes through the outer air channels 57 and into the forward shroud end flow passage 75. L The flow passes through the shroud end flow passage inlet 171.

[0031] In this embodiment, at the insert 151 which is the leading end insert, a portion of the cooling air injected toward the inner surface of the leading end of the airfoil 51 flows radially outward through the outer air channel 57. However, the structure of the stator vane is not limited to this embodiment. At the insert 151 which is the leading end insert, a portion of the cooling air injected toward the inner surface of the leading end of the airfoil 51 may flow radially inward through the outer air channel 57, or may flow both radially inward and radially outward.

[0032] 7 is a partial perspective view of the stator vane in the first embodiment. For example, in insert 152, which is an intermediate insert, a portion of the cooling air supplied to air channel 162 through air intake 58 is injected toward the inner surface of the center portion of airfoil 51, then flows radially inward through the outer air channel toward inner shroud 60, and then flows into shroud end flow passage inlet 181 (located on the aft shroud end) of inner shroud 60, as shown in FIG. 7. Next, the cooling air passes through shroud end flow passage 65 of inner shroud 60, cools shroud end 64 of inner shroud 60, and then flows into shroud body 62 of shroud 60 through shroud end flow passage outlet 182 (located in the forward shroud end flow passage) of inner shroud 60. Similar to the outer shroud 70, cooling air is injected from air holes in the impingement plate 63 to cool the radially outer wall of the inner shroud 60, which has a gas path surface facing radially outward and facing the hot gas flow path.

[0033] 4, the airfoil 51 includes a second airfoil cooling structure 154 that includes passages with a plurality of pin fins 164 disposed therein. In the second airfoil cooling structure 154, a portion of the cooling air flows downstream through the passages with the pin fins 164 before being discharged into the hot gas path at the trailing edge 53 of the airfoil 51.

[0034] Next, a method for cooling the stator vane according to the first embodiment will be described. Fig. 8 is a flowchart illustrating the method for cooling the stator vane according to the first embodiment. As shown in Fig. 8, in step S102, part of the cooling air flows into the shroud end flow passage 75 through the shroud end flow passage inlet 171. The cooling air flows along the shroud end flow passage 75 and cools the shroud end 74.

[0035] In step S104, cooling air flows into the outer region of the shroud body 72 and is ejected through the impingement cooling holes 79 toward the radially outer surface of the radially inner wall 81, thereby impingement cooling the radially outer surface of the radially inner wall 81 and cooling the shroud body 72.

[0036] Next, a cooling method for a stator vane according to a second embodiment will be described. FIG. 9 is a flowchart illustrating the cooling method for a stator vane according to the second embodiment. FIG. 10 schematically illustrates the cooling process of the second embodiment. As shown in FIGS. 9 and 10(a), in step S202, a portion of the cooling air from the forced-air cooling system flows through the air intake 58 into the air channels 161 of the insert 151. The cooling air is then injected through the holes 59 toward the inner surface of the leading end of the airfoil 51 to cool the airfoil 51, and then flows radially outward through the outer air channels 57.

[0037] 10(b), in step S204, cooling air flows into the shroud end passage 75 through the shroud end passage inlet 171. The cooling air flows along the shroud end passage 75 and cools the shroud end 74.

[0038] As shown in FIG. 10(c), in step S206, cooling air flows into the outer region of the shroud body 72 and is injected through the impingement cooling holes 79 toward the radially outer surface of the radially inner wall 81, thereby impingement cooling the radially outer surface of the radially inner wall 81 and cooling the shroud body 72.

[0039] Next, a cooling method for a stator vane according to a third embodiment will be described. Fig. 11 is a flowchart illustrating the cooling method for a stator vane according to the third embodiment. As shown in Fig. 11, in step S302, a portion of the cooling air from the forced-air cooling system flows into the air channels of the insert through the air intake. The cooling air is then injected through the holes toward the inner surface of the leading end of the airfoil to cool the airfoil, and then flows radially outward through the outer air channel.

[0040] In step S304, cooling air enters the outer region of the shroud body and is injected through the impingement cooling holes toward the radially outer surface of the radially inner wall to cool the radially outer surface of the radially inner wall and thereby cool the shroud body.

[0041] In step S306, cooling air flows into the shroud end flow passages through the shroud end flow passage inlets, flows along the shroud ends to cool the shroud ends, and is returned to the forced air cooling system through the shroud end flow passage outlets.

[0042] Next, a fourth embodiment of the present invention will be described below. Fig. 12 is a schematic cross-sectional view of a stator vane according to the fourth embodiment. As shown in Fig. 12, in the fourth embodiment, a plurality of airfoils 51 (two in this embodiment) are arranged in a shroud end flow passage 75. L , 75 T , 75 N , 75 P Unlike the first embodiment (FIG. 5), two shroud end channel inlets 171 are provided in the forward shroud end channel 75. L It is set up in.

[0043] The outer air channels, which are the spaces between the inner surfaces of the forward ends of the two airfoils 51 and the inserts 151, are connected to the forward shroud end flow passage 75 via air passages provided at the outer ends of the outer air channels of the respective airfoils 51. L The cooling air is passed through each shroud end channel inlet 171 to the forward shroud end channel 75. L and flows into the dorsal shroud end flow passage 75 N , or ventral shroud end flow passage 75 P and flows through shroud end flow passage outlets 172 into the outer region of shroud body 72 .

[0044] The present disclosure is not limited to the above-described embodiments and can be implemented in various embodiments. For better understanding, specific embodiments have been described with reference to the drawings. However, the above description is provided by way of example only and does not limit the scope of the invention as defined by the appended claims. The scope of the present invention should be determined by the appended claims. Those skilled in the art can make various modifications without departing from the scope of the invention, and the appended claims are intended to cover such modifications. [Explanation of symbols]

[0045] 10. Gas turbine 20 Turbine 22 Turbine casing 24 rotor shaft 26 Turbine rotor Ar axis 30 Combustor 50 Stator blade 51 Stator wing body (airfoil) 51 P bulkhead 141,142,143 divided area 52 leading edge 53 Trailing edge 54 Dorsal aspect 55 Ventral aspect 56 Intake manifold 57 outer air channel 58 Air intake 59 Hole 151, 152, 153 Inserts 161, 162, 163 Air channels 154 Second blade cooling structure 164 Pinfin 60 Inner shroud 70 Outer shroud 72 Shroud body 73 Impingement Plate 74 Shroud end 75 Shroud end flow passage S space 171 Shroud end flow passage inlet 172 Shroud end flow passage outlet 175 Turbulator 76 Peripheral wall 78 Gas Pass Surface 79 Impingement cooling holes 81 radial inner wall 82 radial outer wall 83 Discharge pipe 181 Shroud end flow passage inlet 182 Shroud end flow passage outlet

Claims

1. A turbine vane, Airfoil and a shroud disposed at a radial end of the airfoil of the turbine; The shroud includes: a shroud body including a first wall facing a hot gas path of the turbine and a second wall disposed on an opposite side of the hot gas path from the first wall; a shroud end portion disposed around the shroud body so as to surround the shroud body, the shroud end portion including a shroud end flow passage therein; the shroud end includes a cooling air inlet for introducing cooling air into the shroud end flow passage and a cooling air outlet for discharging the cooling air from the shroud end flow passage; the shroud body includes a hollow space between the first wall and the second wall, the hollow space being connected to the shroud end flow passage through the cooling air outlet.

2. The shroud body includes: an impingement plate disposed between the first wall and the second wall and dividing the hollow space into a first region on the first wall side and a second region on the second wall side; a flow path that is isolated from the second region of the hollow space and connects the first region of the hollow space to an outer space on the opposite side of the second wall from the hollow space; the impingement plate includes a plurality of impingement cooling holes extending radially therethrough; The vane of claim 1 , wherein the second region of the hollow space is connected to the shroud end flow passage through the cooling air outlet.

3. The vane of claim 2 , wherein the flow passage is configured to penetrate the second wall and the impingement plate.

4. The vane of claim 2 , wherein the flow passage is configured to bypass the second wall.

5. The shroud body includes: a discharge pipe extending in the radial direction; The vane of claim 2 , wherein the exhaust pipe includes the flow passage therein.

6. The shroud end portion a forward shroud end including a forward shroud end flow passage therein; an aft shroud end including an aft shroud end flow passage therein; a suction shroud end portion having a suction shroud end flow passage therein; and a ventral shroud end portion having a ventral shroud end flow passage therein; 2. The vane of claim 1, wherein the cooling air inlet is disposed at one of the forward shroud end and the aft shroud end, and the cooling air outlet is disposed at the other of the suction shroud end, the ventral shroud end, or the forward shroud end and the aft shroud end.

7. The vane of claim 6 , wherein the cooling air outlet is located at the other of the forward shroud end and the aft shroud end.

8. The vane of claim 1 , wherein the shroud end flow passage comprises turbulators disposed on an inner surface of the shroud end flow passage.

9. The vane of claim 8 , wherein the turbulators are disposed on a bottom surface that is the surface closest to a hot gas path of the turbine among a plurality of inner surfaces that define the shroud end flowpath.

10. The vane of claim 1 , wherein the cooling air inlet is in communication with an interior of the airfoil and is configured to introduce cooling air from the airfoil into the shroud end flowpath.

11. The airfoil is an insert extending in said radial direction and having an air channel extending therein; an outer air channel disposed between the inner surface of the airfoil and the outer surface of the insert; the insert includes a hole extending through a sidewall of the insert from an inner surface to an outer surface of the insert; 11. The vane of claim 10, wherein the outer air channel is connected to the cooling air inlet and directs cooling air introduced into the air channel and injected from the air channel through the holes onto the inner surface of the airfoil to the cooling air inlet.

12. 3. The vane of claim 2, wherein the cooling air inlet at the shroud end is configured to receive cooling air extracted from inside a combustor casing and compressed by an external compressor, and the flow passage is configured to discharge the cooling air into the combustor casing interior.

13. the shroud end portion surrounds the entire periphery of the shroud body; The vane of claim 1 , wherein the cooling air flows along the entire shroud end.

14. A method for cooling a turbine vane, the vane comprising: an airfoil; and a shroud disposed at an end of the airfoil in a radial direction of the turbine, the shroud including a shroud body; and a shroud end portion disposed around the shroud body so as to surround the shroud body, the shroud end portion including a shroud end flow passage therein; The cooling method for the stator blade includes: (i) cooling the shroud end by flowing cooling air through the shroud end passage; (ii) A method for cooling a stator vane, characterized in that after cooling the shroud end, the shroud body is cooled using the cooling air that has flowed inside the shroud end flow passage.

15. further comprising the step of flowing cooling air inside the airfoil; 15. The method for cooling a stator vane according to claim 14, wherein step (i) further comprises: cooling the shroud end by using the cooling air that has flowed through the interior of the airfoil to flow the cooling air into the shroud end flow passage, after cooling the airfoil.

16. the airfoil includes an insert extending in the radial direction and having an air channel therein extending in the radial direction; The method of cooling a stator vane of claim 15 , wherein cooling air is introduced into the air channel to cool the airfoil, and then the cooling air is guided by the airfoil toward the shroud end flowpath.

17. The airfoil is an insert extending in said radial direction and having an air channel extending therein; an outer air channel disposed between the inner surface of the airfoil and the outer surface of the insert; the insert includes a hole extending through a sidewall of the insert from an inner surface to an outer surface of the insert; 16. The method of cooling a stator vane as recited in claim 15, wherein cooling air introduced into the air channel and ejected from the air channel through the holes onto the inner surface of the airfoil is directed by the outer air channel to the shroud end flow passage.

18. the radially extending air channels of the insert are configured to receive cooling air extracted from an interior of a combustor casing and compressed by an external compressor; The method of cooling a stator vane according to claim 16 , further comprising discharging the cooling air into the interior of the combustor casing after cooling the shroud body.

19. The shroud end portion a forward shroud end including a forward shroud end flow passage therein; an aft shroud end including an aft shroud end flow passage therein; a suction shroud end including a suction shroud end flow passage therein; a ventral shroud end including a ventral shroud end flow passage therein; the step (i) uses cooling air introduced from either the forward shroud end or the aft shroud end to flow through the shroud end flow passage to cool the shroud end; 15. The vane cooling method according to claim 14, wherein the step (ii) cools the shroud body using cooling air flowing inside the shroud end flow passage and discharged from the other of the suction side shroud end, the pressure side shroud end, or the forward side shroud end and the aft side shroud end.

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