turbine blades

Abstract

A turbine stator according to at least one embodiment of the present disclosure comprises an airfoil, an inner shroud provided on the inner circumference side of the airfoil, and an outer shroud provided on the outer circumference side of the airfoil. The inner shroud has a first recess formed on a surface opposite to the airfoil, with the gas pass surface of the inner shroud in between. The outer shroud has a second recess formed on a surface opposite to the airfoil, with the gas pass surface of the outer shroud in between, and at least one outer passage that communicates with the second recess and is not in communication with the space inside the airfoil. The number of outer passages is greater than the number of inner passages in the inner shroud that communicate with the first recess and are not in communication with the space inside the airfoil.

Description

Technology Field

[0001] The present disclosure relates to a turbine static blade.

[0002] The present application claims priority based on Patent Application No. 2021-193277 filed with the Japan Patent Office on November 29, 2021, and incorporates the contents thereof herein. Background Technology

[0003] For example, turbine stator blades used in gas turbines, etc., are exposed to high-temperature fluids such as combustion gases, so they have a structure for cooling. For example, in the turbine stator blade described in Patent Document 1, the stator blade body (airfoil section), the inner shroud, and the outer shroud are each cooled by cooling air (see Patent Document 1). Prior art literature

[0004] International Publication No. 2020 / 213381 The problem to be solved

[0005] As a result of measuring the combustion gas temperature in the actual machine, it was found that the combustion gas temperature on the hub side was lower than previously assumed, and the metal temperature of the inner shroud was also found to be lower than previously assumed. Since a lower metal temperature of the turbine stator blades than the previously assumed temperature leads to a decrease in the efficiency of the gas turbine, it is desired that the metal temperature of the turbine stator blades be close to the previously assumed temperature.

[0006] At least one embodiment of the present disclosure aims to improve turbine efficiency by appropriately cooling the turbine stator blades in light of the circumstances described above. means of solving the problem

[0007] 1) A turbine stator according to at least one embodiment of the present disclosure is,

[0008] Airfoil and,

[0009] An inner shroud provided on the inner circumference side of the airfoil portion, and

[0010] It is provided with an outer shroud provided on the outer circumference side of the airfoil portion, and

[0011] The inner shroud has a first concave portion formed on a surface opposite to the airfoil portion, with the gas pass surface of the inner shroud in between.

[0012] The outer shroud has a second concave portion formed on a surface opposite to the airfoil portion with the gas pass surface of the outer shroud in between, and at least one outer passage that communicates with the second concave portion and is not in communication with the internal space of the airfoil portion.

[0013] The number of the above outer passages is greater than the number of inner passages that communicate with the first concave portion in the inner shroud and are not in communication with the space inside the airfoil portion. Effects of the invention

[0014] According to at least one embodiment of the present disclosure, turbine efficiency can be improved by appropriately cooling the turbine stator blades. Brief explanation of the drawing

[0015] FIG. 1 is a schematic diagram showing a gas turbine (1) of one embodiment in which turbine stator blades according to several embodiments are used. FIG. 2 is a schematic internal cross-sectional view of a turbine stator of one embodiment. FIG. 3 is a schematic internal cross-sectional view of a turbine stator of another embodiment. FIG. 4 is a schematic internal cross-sectional view of a turbine stator of another embodiment. FIG. 5 is a cross-section of the trailing edge of the outer shroud in the turbine stator blades shown in FIG. 2 to 4, viewed from the downstream side. Figure 6 is a cross-section of the trailing edge of the inner shroud in the turbine stator shown in Figure 2, viewed from the downstream side. Specific details for implementing the invention

[0016] 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 the components described as embodiments or shown in the drawings are not intended to limit the scope of the present disclosure and are merely illustrative examples.

[0017] For example, expressions indicating relative or absolute arrangements, such as "in a certain direction," "along a certain direction," "parallel," "orthogonal," "center," "concentric," or "coaxial," are meant to indicate not only strictly such arrangements but also tolerances, or a state of relative displacement by an angle or distance sufficient to obtain the same function.

[0018] For example, expressions indicating that things are in an equivalent state, such as "identical," "equivalent," and "homogeneous," are meant to indicate not only a strictly equivalent state, but also a state in which tolerances or differences in the degree to which the same function is obtained exist.

[0019] For example, expressions representing shapes such as square or cylindrical shapes represent not only shapes such as square or cylindrical shapes in the strictly geometric sense, but also shapes including uneven surfaces or chamfers within the scope where the same effect is obtained.

[0020] Meanwhile, expressions such as "possesses," "equips," "includes," or "has" a component are not exclusive expressions that exclude the existence of other components.

[0021] FIG. 1 is a schematic diagram showing a gas turbine (1) of one embodiment in which turbine stator blades according to several embodiments are used.

[0022] As shown in FIG. 1, a gas turbine (1) according to one embodiment comprises a compressor (2) for generating compressed air, a combustor (4) for generating combustion gas using compressed air and fuel, and a turbine (6) configured to be driven by rotation by combustion gas. In the case of a gas turbine (1) for power generation, a generator not shown is connected to the turbine (6), and power generation is performed by the rotational energy of the turbine (6).

[0023] Specific configuration examples of each part in the gas turbine (1) will be explained using FIG. 1.

[0024] The compressor (2) comprises a compressor chamber (10), an air intake port (12) provided on the inlet side of the compressor chamber (10) for injecting air, a rotor shaft (8) provided to penetrate both the compressor chamber (10) and the turbine chamber (22) described later, and various blades arranged within the compressor chamber (10). The various blades include an inlet guide blade (14) provided on the side of the air intake port (12), a plurality of compressor static blades (16) fixed on the side of the compressor chamber (10), and a plurality of compressor dynamic blades (18) installed on the rotor shaft (8) so as to be arranged alternately in the axial direction with respect to the compressor static blades (16). Additionally, the compressor (2) may be provided with other components such as an exhaust chamber not shown. In such a compressor (2), air blown in from the air intake port (12) passes through a plurality of compressor stator vanes (16) and a plurality of compressor moving vanes (18) and is compressed to produce compressed air. Then, the compressed air is transferred from the compressor (2) to a downstream combustor (4).

[0025] The combustor (4) is placed inside a casing (combustor chamber) (20). As shown in FIG. 1, the combustor (4) may be arranged in a ring around the rotor shaft (8) inside the casing (20). Fuel and compressed air generated by the compressor (2) are supplied to the combustor (4), and by burning the fuel, high-temperature, high-pressure combustion gas, which is the working fluid of the turbine (6), is generated. Then, the combustion gas is transferred from the combustor (4) to the turbine (6) at the rear end.

[0026] The turbine (6) comprises a turbine casing (22) and various turbine blades arranged within the turbine casing (22). The various turbine blades include a plurality of turbine static blades (100) fixed to the side of the turbine casing (22) and a plurality of turbine moving blades (24) installed on a rotor shaft (8) so as to be arranged alternately in the axial direction with respect to the turbine static blades (100).

[0027] In addition, in the turbine (6), the rotor shaft (8) extends in the axial direction (left-right direction in FIG. 1), and the combustion gas flows from the combustor (4) side toward the exhaust chamber (28) side (from left to right in FIG. 1). Therefore, in FIG. 1, the left side shown is the axial upstream side, and the right side shown is the axial downstream side. Also, in the following description, when simply referred to as the radial direction, it indicates the same direction as the radial direction perpendicular to the rotor shaft (8).

[0028] In a gas turbine (1) according to one embodiment, the turbine stator (100) includes a first stage stator (31), a second stage stator (32), and a third stage stator (33) arranged in order from the axial upstream side.

[0029] The turbine driving blade (24) is configured to generate rotational driving force from high-temperature, high-pressure combustion gas flowing through the turbine cabin (22) together with the turbine stator blade (100). This rotational driving force is transmitted to the rotor shaft (8), thereby driving the generator connected to the rotor shaft (8).

[0030] On the axial downstream side of the turbine chamber (22), an exhaust chamber (29) is connected through an exhaust chamber (28). The combustion gas after driving the turbine (6) passes through the exhaust chamber (28) and the exhaust chamber (29) and is discharged to the outside.

[0031] FIG. 2 is a schematic internal cross-sectional view of a turbine stator (100) of one embodiment, showing a cross-section along the camber line of the airfoil.

[0032] FIG. 3 is a schematic internal cross-sectional view of a turbine stator (100) of another embodiment, showing a cross-section along the camber line of the airfoil.

[0033] FIG. 4 is a schematic internal cross-sectional view of a turbine stator (100) of another embodiment, showing a cross-section along the camber line of the airfoil.

[0034] Below, the structure of a turbine blade (100) according to several embodiments is described.

[0035] In the following description, the structure of the third stage stator (33) is described, but the first stage stator (31) or the second stage stator (32) may have the same structure as the third stage stator (33). Also, in the case where the gas turbine (1) according to one embodiment has a turbine stator (100) located axially downstream from the third stage stator (33), this turbine stator (100) may have the same structure as the third stage stator (33).

[0036] As shown in FIGS. 2 to 4, a turbine stator (100) according to several embodiments is equipped with a blade body (101), an impingement plate (130), and an air pipe (127).

[0037] A wing body (101) according to several embodiments comprises an airfoil (110) having a plurality of cooling channels (111) partitioned by a bulkhead (140), an outer shroud (121) provided on the leading edge (110c) side of the airfoil (110), i.e., on the outer side in the diameter direction, and an inner shroud (122) provided on the base edge (110d) side of the airfoil (110), i.e., on the inner side in the diameter direction. In addition, in the following description, the diameter direction is also referred to as the wing height direction of the airfoil (110), or simply the wing height direction. Also, for convenience of explanation, the plurality of cooling channels (111) are referred to as the first cooling channel (111a), the second cooling channel (111b), the third cooling channel (111c), the fourth cooling channel (111d), and the fifth cooling channel (111e), in order from the leading edge (110a) side of the airfoil (110) side to the trailing edge (110b) side. However, in the following description, if there is no need to distinguish each cooling channel (111a, 111b, 111c, 111d, 111e), the alphabet following the number in the symbol may be omitted and simply referred to as a cooling channel (111).

[0038] In a turbine stator (100) according to several embodiments, a first cooling channel (111a) and a second cooling channel (111b) are separated by a bulkhead (140), and a second cooling channel (111b) and a third cooling channel (111c) are separated by a bulkhead (140). A third cooling channel (111c) and a fourth cooling channel (111d) are separated by a bulkhead (140), and a fourth cooling channel (111d) and a fifth cooling channel (111e) are separated by a bulkhead (140).

[0039] In a turbine stator (100) according to some embodiments, the gas pass surface is the surface to which combustion gas contacts when the turbine stator (100) according to some embodiments is placed in a turbine, and corresponds to the outer surface (121a, 122a) of the outer shroud (121) and inner shroud (122) shown in FIGS. 2 to 4. In a turbine stator (100) according to some embodiments, the airfoil (110) and the shroud (121, 122) are manufactured, for example, by casting.

[0040] In the turbine stator (100) according to several embodiments shown in FIGS. 2 to 4, for example, four bending passages (112) are formed. Specifically, in order from the leading edge (110a) side, the first bending passage (112a) is connected to the first cooling passage (111a) and the second cooling passage (111b), and the second bending passage (112b) is connected to the second cooling passage (111b) and the third cooling passage (111c). The third bending passage (112c) is connected to the third cooling passage (111c) and the fourth cooling passage (111d), and the fourth bending passage (112d) is connected to the fourth cooling passage (111d) and the fifth cooling passage (111e).

[0041] In the turbine stator (100) according to several embodiments shown in FIGS. 2 to 4, a plurality of cooling holes (113) are formed near the trailing edge (110b) of the airfoil (110), which are connected to the fifth cooling channel (111e) on the upstream side in the direction of flow of the cooling medium, and whose downstream side is opened at the end of the trailing edge (110b).

[0042] In addition, in the turbine stator (100) according to several embodiments shown in FIGS. 2 and 3, at least one outlet passage (118) is formed in the inner shroud (122), which communicates with the fifth cooling channel (111e) on the upstream side in the direction of flow of the cooling medium, and on the downstream side, for example, opens to the cross section (122c) on the rear edge side of the inner shroud (122).

[0043] In the turbine stator (100) according to the embodiment shown in FIG. 4, at least one outlet passage (118) is formed in the inner shroud (122), which communicates with the fifth cooling channel (111e) on the upstream side in the direction of flow of the cooling medium, and on the downstream side, opens to the cross-section (161c) on the trailing edge side of the downstream rib (161b) described later, for example.

[0044] In the turbine stator (100) according to several embodiments shown in FIGS. 2 to 4, a serpentine channel (115) including a plurality of cooling channels (111) and a plurality of bending channels (112) is formed.

[0045] In the turbine stator (100) according to several embodiments shown in FIGS. 2 to 4, the outer shroud (121) has a bottom portion (124) forming a gas pass surface and an outer wall portion (123) extending from both ends of the bottom portion (124) in the axial and circumferential directions to the side opposite to the gas pass surface in the wing height direction. The outer shroud (121) has an impingement plate (130) fixed to the outer wall portion (123) and having a plurality of through holes (114) (outer impingement holes (114o)).

[0046] In the outer shroud (121), at least one outer passage (172) is formed that connects the space between the impingement plate (130) and the inner surface (121b) of the outer shroud (121) and the outside of the turbine stator (100) (the outside of the airfoil portion (110)) in the second recess (i.e., the inner space (116) of the outer shroud (121)) surrounded by the outer wall portion (123).

[0047] In the turbine stator (100) according to several embodiments shown in FIGS. 2 to 4, at least one outer passage (172) is formed at the trailing end (125) on the axially downstream side of the inner space (116).

[0048] In the turbine stator (100) according to several embodiments shown in FIGS. 2 to 4, the outlet end (172o) of the outer passage (172) is opened, for example, in the cross section (125c) on the rear side of the outer shroud (121).

[0049] In the turbine stator (100) according to several embodiments shown in FIGS. 2 to 4, the inner shroud (122) has an upstream rib (161a) positioned on the leading edge (110a) and a downstream rib (161b) positioned on the trailing edge (110b) that protrudes inward in the wing height direction from the inner surface (122b) of the inner shroud (122).

[0050] In the turbine stator (100) shown in FIG. 2, an impingement plate (130) having a plurality of through holes (114) (inner impingement holes (114i)) is disposed between the upstream rib (161a) and the downstream rib (161b), which partitions the internal space (117). In addition, in the turbine stator (100) shown in FIG. 3 and FIG. 4, the impingement plate (130) is not disposed.

[0051] In the turbine stator (100) shown in FIG. 2, at least one inner passage (171) is formed in the inner shroud (122) that communicates the space between the impingement plate (130) and the inner surface (122b) of the inner shroud (122) in the first recess (i.e., the inner space (117) of the inner shroud (122)) between the upstream rib (161a) and the downstream rib (161b), and the outside of the turbine stator (100) (the outside of the airfoil (110)).

[0052] In the turbine stator (100) shown in FIG. 2, at least one inner passage (171) is formed at the trailing end (126) on the axially downstream side of the inner space (117). In FIG. 2, for convenience of illustration, the inner passage (171) and the outlet passage (118) are shown offset in the wing height direction, but as shown in FIG. 6 which will be described later, the inner passage (171) and the outlet passage (118) may be formed at the same position in the wing height direction.

[0053] In the turbine stator (100) shown in FIG. 2, the inner passage (171) is not in communication with the space (110i) inside the airfoil (110). Here, the fact that the inner passage (171) is not in communication with the space (110i) inside the airfoil (110) means that the direct connection point of the inner passage (171) is not the space (110i) inside the airfoil (110). That is, the fact that the inner passage (171) is not in communication with the space (110i) inside the airfoil (110) means that all end openings (inlet end (171i) and outlet end (171o)) of the inner passage (171) do not face the space (110i) inside the airfoil (110).

[0054] Additionally, in some embodiments, the space (110i) inside the airfoil (110) is a void formed inside the airfoil (110) (the area between the outer surface (121a) of the outer shroud (121) and the outer surface (122a) of the inner shroud (122), and includes, for example, a serpentine channel (115).

[0055] In addition, in the turbine stator (100) shown in FIGS. 3 and 4, an inner passage (171) is not formed.

[0056] The turbine stator (100) according to several embodiments shown in FIGS. 2 to 4 is provided with an air pipe (127) penetrating the airfoil (110) in the wing height direction, for example, within the second cooling channel (111b). One end of the air pipe (127) is open into an internal space (117) formed in a support ring (162) supported by an inner shroud (122). The support ring (162) is supported by the inner shroud (122) through an upstream rib (161a) and a downstream rib (161b) of the inner shroud (122).

[0057] Also, the support ring (162) is provided with, for example, a flow hole (162a) on the bottom surface.

[0058] For example, compressed air extracted from a compressor (2) is used as the cooling medium supplied to the turbine stator (100).

[0059] In the turbine stator (100) according to several embodiments shown in FIGS. 2 to 4, the cooling medium supplied to the serpentine path (115) is supplied from the outside to the internal space (116) of the outer shroud (121), as indicated by arrow a. The cooling medium flows into the first cooling path (111a) through an opening (133) formed in the outer shroud (121) and flows within the first cooling path (111a) along the blade height direction from the leading edge (110c) side toward the leading edge (110d) side, as indicated by arrow b. Afterward, the cooling medium introduced into the first cooling path (111a) flows through each bending path (112) and each cooling path (111) in sequence, as indicated by arrows c to j. In this way, the cooling medium flows in the same direction as the main flow of combustion gas, from the leading edge (110a) side to the trailing edge (110b) side within the airfoil (110).

[0060] The cooling medium introduced into the fifth cooling channel (111e) is discharged into the combustion gas outside the airfoil (110) through a plurality of cooling holes (113) that open in the trailing edge (110b), as indicated by arrow k1.

[0061] Additionally, the cooling medium introduced into the fifth cooling channel (111e) is discharged into the combustion gas outside the airfoil (110) through the outlet passage (118), as indicated by arrow k2.

[0062] In addition, in the turbine stator (100) according to several embodiments shown in FIGS. 2 to 4, a cooling medium supplied from the outside is sprayed onto the inner surface (121b) of the bottom portion (124) of the outer shroud (121) through a plurality of through holes (114) of the impingement plate (130) that is radially outward (toward end (110c) side) of the impingement plate (130) (inner space (116)). The cooling medium impingement cools the inner surface (121b). By doing so, the bottom portion (124) of the outer shroud (121) can be cooled with the cooling medium.

[0063] The cooling medium that cools the bottom portion (124) of the outer shroud (121) is discharged into the combustion gas outside the airfoil portion (110) from the second concave portion (internal space (116)) through a plurality of outer passages (172).

[0064] The cooling medium (Ac) supplied from the internal space (116) of the outer shroud (121) is supplied through the air pipe (127) to the internal space (117) formed in the support ring (162) on the inner shroud (122) side.

[0065] In the turbine stator (100) shown in FIG. 2, a portion of the cooling medium (Ac) is applied as cooling air to impingement cool (impingement cool) the inner surface (122b) of the inner shroud (122) through a plurality of through holes (114) of the impingement plate (130) disposed in the inner shroud (122).

[0066] In the turbine stator (100) shown in FIG. 2, the remaining cooling medium (Ac) is supplied from the flow hole (162a) to the inter-stage cavity not shown, and as purge air, prevents the combustion gas from flowing back into the inter-stage cavity.

[0067] In the turbine stator (100) shown in FIG. 2, the cooling medium (Ac) that cools the inner surface (122b) of the inner shroud (122) is discharged into the combustion gas outside the airfoil (110) from the first concave portion (inner space (117)) through the inner passage (171).

[0068] In addition, in the turbine stator (100) shown in FIGS. 3 and 4, the cooling medium (Ac) supplied to the internal space (117) is supplied to an unillustrated inter-stage cavity through the flow hole (162a), thereby preventing the combustion gas from flowing back into the inter-stage cavity as purge air.

[0069] As described above, the turbine stator (100) according to several embodiments shown in FIGS. 2 to 4 comprises an airfoil (110), an inner shroud (122) provided on the inner side of the airfoil (110), and an outer shroud (121) provided on the outer side of the airfoil (110).

[0070] In the turbine stator (100) according to several embodiments shown in FIGS. 2 to 4, the inner shroud (122) has a first concave portion (internal space (117)) formed on the side opposite to the airfoil portion (110) with the gas pass surface (outer surface (122a)) of the inner shroud (122) in between.

[0071] In the turbine stator (100) according to several embodiments shown in FIGS. 2 to 4, the outer shroud (121) has a second recess (internal space (116)) formed on a side opposite to the airfoil (110) with the gas pass surface (outer surface (121a)) of the outer shroud (121) in between, and at least one outer passage (172) that communicates with the second recess (internal space (116)) and is not in communication with the internal space (110i) of the airfoil (110).

[0072] Here, the fact that the outer passage (172) is not connected to the internal space (110i) of the airfoil (110) means that the direct connection point of the outer passage (172) is not the internal space (110i) of the airfoil (110). That is, the fact that the outer passage (172) is not connected to the internal space (110i) of the airfoil (110) means that all end openings (inlet end (172i) and outlet end (172o)) of the outer passage (172) do not face the internal space (110i) of the airfoil (110).

[0073] As a result of measuring the temperature of the combustion gas in the actual operation of the gas turbine (1), it was found that the combustion gas temperature on the hub side of the third stage stator (33) was lower than previously assumed, and it was also found that the metal temperature of the inner shroud (122) of the third stage stator (33) was lower than previously assumed. Since a lower metal temperature of the turbine stator (100) than the previously assumed temperature causes a decrease in the efficiency of the gas turbine (1), it is desired that the metal temperature of the turbine stator (100) be close to the previously assumed temperature.

[0074] Accordingly, in some embodiments of the turbine stator (100), the number of outer passages (172) is greater than the number of inner passages (171).

[0075] Therefore, since the cooling capacity of the inner shroud (122) can be suppressed more than the cooling capacity of the outer shroud (121), excessive cooling of the inner shroud (122) can be suppressed, thereby allowing the cooling of the turbine stator (100) to be optimized. This allows for improved turbine efficiency.

[0076] More specifically, in some embodiments, the turbine stator (100) may be configured to satisfy any one of the following conditions (A), (B), or (C).

[0077] (A) The inner shroud (122) does not have an inner passage (171).

[0078] (B) The inner shroud (122) has one inner passage (171), and the cross-sectional area Si of the outlet end (171o) of the one inner passage (171) is smaller than the sum of the cross-sectional areas So of the outlet end (172o) of the at least one outer passage (172).

[0079] (C) The inner shroud (122) has two or more inner passages (171), and the value (Pi / Di) obtained by dividing the distance Pi between the inner passages (171) by the diameter Di of the inner passage (171) is greater than the value (Po / Do) obtained by dividing the distance Po between the outer passages (172) by the diameter Do of the outer passage (172).

[0080] The above condition (A) is explained.

[0081] The above condition (A) corresponds to a case where, for example as shown in FIGS. 3 and 4, an impingement plate (130) is not placed in the internal space (117), an inner passage (171) is not formed in the turbine stator (100), and the inner surface (122b) of the inner shroud (122) is not impingement cooled.

[0082] In the turbine stator (100) shown in FIGS. 3 and 4, the inner surface (122b) of the inner shroud (122) is not impingement cooled, so the cooling capacity of the inner shroud (122) can be suppressed compared to the cooling capacity of the outer shroud (121). By doing so, excessive cooling of the inner shroud (122) can be suppressed, thereby allowing the cooling of the turbine stator (100) to be optimized and the turbine efficiency to be improved.

[0083] The above condition (B) is explained.

[0084] The above condition (B) is a condition that applies when, for example as shown in FIG. 2, an impingement plate (130) is placed in the internal space (117), an inner passage (171) is formed in the turbine stator (100), and the inner surface (122b) of the inner shroud (122) is impingement coolable.

[0085] As in condition (B) above, if the inner shroud (122) has one inner passage (171) and the cross-sectional area Si of the outlet end (171o) of this one inner passage (171) is smaller than the sum of the cross-sectional areas So of the outlet end (172o) of the at least one outer passage (172), then the amount of cooling medium flowing in the inner passage (171) can be suppressed compared to the sum of the amounts of cooling medium flowing in all outer passages (172). By doing so, the cooling capacity of the inner shroud can be suppressed compared to the cooling capacity of the outer shroud, thereby suppressing excessive cooling of the inner shroud and allowing for proper cooling of the turbine stator blades. Accordingly, turbine efficiency can be improved.

[0086] The above condition (C) is explained.

[0087] The above condition (C) is a condition that applies when, for example as shown in FIG. 2, an impingement plate (130) is placed in the internal space (117), an inner passage (171) is formed in the turbine stator (100), and the inner surface (122b) of the inner shroud (122) is impingement coolable.

[0088] FIG. 5 is a view of the cross-section (125c) on the trailing edge side of the outer shroud (121) of the turbine stator (100) shown in FIG. 2 to 4, viewed from the downstream side.

[0089] FIG. 6 is a view of the cross-section (122c) on the trailing edge of the inner shroud (122) of the turbine stator (100) shown in FIG. 2, seen from the downstream side.

[0090] As in the above condition (C), if the inner shroud (122) has two or more inner passages (171), and the value (Pi / Di) obtained by dividing the distance Pi between the inner passages (171) by the diameter Di of the inner passage (171) is greater than the value (Po / Do) obtained by dividing the distance Po between the outer passages (172) by the diameter Do of the outer passage (172), then the total amount of cooling medium flow in all inner passages (171) can be suppressed compared to the total amount of cooling medium flow in all outer passages (172). By doing so, the cooling capacity of the inner shroud (122) can be suppressed compared to the cooling capacity of the outer shroud (121), thereby suppressing excessive cooling of the inner shroud (122) and allowing the cooling of the turbine stator (100) to be optimized. Thus, the turbine efficiency can be improved.

[0091] That is, in some embodiments, by configuring the turbine stator (100) to satisfy any one of the following conditions (A), (B), or (C), the cooling capacity of the inner shroud (122) can be suppressed compared to the cooling capacity of the outer shroud (121), thereby suppressing excessive cooling of the inner shroud (122) and allowing the cooling of the turbine stator (100) to be optimized. This allows for improved turbine efficiency.

[0092] In some embodiments, the value (ΣSi / (Sig+Sit)) obtained by dividing the sum of the cross-sectional area Si of the outlet end (171o) of the inner passage (171) by the sum of the area Sig of the gas pass surface (outer surface (122a)) of the inner shroud (122) and the area Sit of the cross-sectional area (122c) of the rear end of the inner shroud (122) is smaller than the value (ΣSo / (Sog+Sot)) obtained by dividing the sum of the cross-sectional area So of the outlet end (172o) of the outer passage (172) by the sum of the area Sog of the gas pass surface (outer surface (121a)) of the outer shroud (121) and the area Sot of the cross-sectional area (125c) of the rear end of the outer shroud (121).

[0093] In this case, when compared per unit area with respect to the total area (Sig, Sog) of the gas pass surface (outer surface (121a, 122a)) and the area (Sot, Sit) of the cross-section (125c, 122c) on the trailing side of the shroud (121, 122), the flow rate of the cooling medium flowing through the inner passage (171) is less than the flow rate of the cooling medium flowing through the outer passage (172). As a result, the cooling capacity of the inner shroud (122) can be suppressed compared to the cooling capacity of the outer shroud (121), thereby suppressing excessive cooling of the inner shroud (122) and allowing the cooling of the turbine stator (100) to be optimized.

[0094] In some embodiments, the sum of the cross-sectional area Si1 of the exit end (171o) of the inner passage (171) that is opened in the cross-section (122c) on the rear side of the inner shroud (122) (ΣSi1 / Sit) divided by the area Sit of the cross-section (122c) on the rear side of the inner shroud (122) is smaller than the sum of the cross-sectional area So1 of the exit end (172o) of the outer passage (172) that is opened in the cross-section (125c) on the rear side of the outer shroud (121) (ΣSo1 / Sot) divided by the area Sot of the cross-section (125c) on the rear side of the outer shroud (121).

[0095] In this case, when compared per unit area of ​​the cross-section (125c, 122c) on the trailing side of the shroud (121, 122), the flow rate of the cooling medium flowing through the inner passage (171) is less than the flow rate of the cooling medium flowing through the outer passage (172). As a result, the cooling capacity of the inner shroud (122) can be suppressed compared to the cooling capacity of the outer shroud (121), thereby suppressing excessive cooling of the inner shroud (122) and allowing the cooling of the turbine stator (100) to be optimized.

[0096] In some embodiments, the opening density of the inner impingement hole (114i) for impingement cooling the inner shroud (122) is smaller than the opening density of the outer impingement hole (114o) for impingement cooling the outer shroud (121).

[0097] Here, the opening density of the inner impingement holes (114i) is the sum of the opening areas of the inner impingement holes (114i) divided by the area of ​​the impingement plate (130) placed on the inner shroud (122).

[0098] Likewise, the opening density of the outer impingement holes (114o) is the sum of the opening areas of the outer impingement holes (114o) divided by the area of ​​the impingement plate (130) placed on the outer shroud (121).

[0099] Thus, the cooling capacity of the inner shroud (122) can be suppressed more than the cooling capacity of the outer shroud (121), thereby suppressing excessive cooling of the inner shroud (122) and allowing the cooling of the turbine stator (100) to be appropriate.

[0100] Additionally, the wording above, “the opening density of the inner impingement hole (114i) for impingement cooling the inner shroud (122) is smaller than the opening density of the outer impingement hole (114o) for impingement cooling the outer shroud (121),” includes cases where the opening density of the inner impingement hole (114i) for impingement cooling the inner shroud (122) is zero, that is, cases where the impingement plate (130) in which the inner impingement hole (114i) is formed is not mounted, as shown in FIG. 3 or FIG. 4.

[0101] In this case, by not performing impingement cooling on the inner shroud (122), excessive cooling of the inner shroud (122) is suppressed, and the cooling of the turbine stator (100) can be appropriately controlled.

[0102] (Regarding heat-insulating coatings)

[0103] In the turbine stator (100) shown in FIGS. 2 to 4, a heat-insulating coating (180) may be applied to the gas pass surface (outer surface (121a)) of the airfoil (110) or the outer shroud (121).

[0104] The application range of the heat-insulating coating (180) in the turbine stator (100) shown in FIGS. 2 to 4 is, for example, the surface (110s) of the airfoil (110), the outer surface (121a) of the outer shroud (121), the surface (175s) of the fillet portion (175) connecting the surface (110s) of the airfoil (110) and the outer surface (121a) of the outer shroud (121), and the surface (175s) of the fillet portion (175) connecting the surface (110s) of the airfoil (110) and the outer surface (122a) of the inner shroud (122) (see FIG. 2).

[0105] If there is sufficient cooling capacity of the inner shroud (122), it is thought that the manufacturing cost of the turbine stator (100) can be reduced by narrowing the application range of the heat-insulating coating (180), which has a relatively high construction cost.

[0106] That is, as described above, if the cooling of the inner shroud (122) by the cooling medium is suppressed, the temperature of the inner shroud (122) becomes a temperature relatively close to the combustion gas temperature of the hub side. In such a case, there is no need to apply a heat-insulating coating (180) to the outer surface (122a) of the inner shroud (122) to lower the temperature of the inner shroud (122) below the combustion gas temperature of the hub side.

[0107] Accordingly, in the turbine stator (100) shown in FIGS. 2 to 4, the heat-insulating coating (180) is not applied to the outer surface (122a) of the inner shroud (122). Additionally, the heat-insulating coating (180) may be applied to a part of the outer surface (122a) of the inner shroud (122), rather than the entire surface.

[0108] Here, the outer surface (122a) of the inner shroud (122) does not include the surface (175s) of the fillet portion (175) connecting the surface (110s) of the airfoil portion (110) and the outer surface (122a) of the inner shroud (122).

[0109] In this way, in some embodiments, the value (Sic / Sig) obtained by dividing the area of ​​the heat-insulating coating (180) applied to the gas pass surface (outer surface (122a)) of the inner shroud (122) by the area of ​​the gas pass surface (outer surface (122a)) of the inner shroud (122) may be smaller than the value (Soc / Sog) obtained by dividing the area of ​​the heat-insulating coating (180) applied to the gas pass surface (outer surface (121a)) of the outer shroud (121) by the area of ​​the gas pass surface (outer surface (121a)) of the outer shroud (121) by the area of ​​the gas pass surface (outer surface (121a)) of the outer shroud (121) by the area of ​​the outer shroud (121).

[0110] Additionally, the above-mentioned point where the value (Sic / Sig) is smaller than the value (Soc / Sog) includes the case where the construction area Sic of the heat-insulating coating (180) applied to the gas pass surface (outer surface (122a)) of the inner shroud (122) is zero.

[0111] That is, as described above, the inner shroud (122) does not need to have a heat-insulating coating (180) applied to the gas pass surface (outer surface (122a)) of the inner shroud (122).

[0112] In this way, by narrowing the application range of the heat-insulating coating on the gas pass surface (outer surface (122a)) of the inner shroud (122), the manufacturing cost of the turbine stator (100) can be reduced.

[0113] In the turbine stator (100) shown in FIGS. 2 to 4, the inner shroud (122) may have an outlet passage (118) that is in communication with the space (110i) inside the airfoil (110) and is not in communication with the first concave portion (inner space (117)).

[0114] Thus, when a cooling medium is supplied to the serpentine channel (115), which is the internal space (110i) of the airfoil (110), the cooling medium can be discharged to the outside of the turbine stator (100) through the outlet passage (118).

[0115] Additionally, the space (110i) inside the airfoil (110) described above may be a passage for a cooling medium formed inside the airfoil (110) in addition to the serpentine passage (115).

[0116] In some embodiments, the turbine stator (100) shown in FIGS. 2 to 4 may be a turbine three-stage stator (three-stage stator (33)) as described above.

[0117] Thus, as described above, when the combustion gas temperature on the hub side of the third stage stator (33) is lower than the prior assumption and the metal temperature of the inner shroud (122) of the third stage stator (33) is also lower than the prior assumption, the excessive cooling of the inner shroud (122) can be suppressed, thereby allowing the cooling of the turbine stator (100) (third stage stator (33)) to be appropriate.

[0118] The present disclosure is not limited to the embodiments described above, but also includes forms with modifications to the embodiments described above, or forms appropriately combined thereof.

[0119] For example, in some of the embodiments described above, the outlet end (171o) of the inner passage (171) is formed on the rear end (122c) of the inner shroud (122), but at least some of the outlet end (171o) may be formed on the outer surface (122a) of the inner shroud (122).

[0120] For example, in some of the embodiments described above, the exit end (172o) of the outer passage (172) is formed on the cross-section (125c) on the rear edge side of the outer shroud (121), but at least some of the exit end (172o) may be formed on the outer surface (121a) of the outer shroud (121).

[0121] For example, in some of the embodiments described above, the exit end (118o) of the exit passage (118) is formed on the rear end cross section (122c) of the inner shroud (122) or on the rear end cross section (161c) of the downstream rib (161b), but at least some of the exit end (118o) may be formed on the outer surface (122a) of the inner shroud (122).

[0122] The contents described in each of the above embodiments are understood, for example, as follows.

[0123] (1) A turbine stator (100) according to at least one embodiment of the present disclosure comprises an airfoil (110), an inner shroud (122) provided on the inner side of the airfoil (110), and an outer shroud (121) provided on the outer side of the airfoil (110). The inner shroud (122) has a first concave portion (internal space (117)) formed on the side opposite to the airfoil (110) with the gas pass surface (outer surface (122a)) of the inner shroud (122) in between. The outer shroud (121) has at least one outer passage (172) that communicates with the second recess (internal space (116)) formed on the side opposite to the airfoil (110) with the gas pass surface (outer surface (121a)) of the outer shroud (121) in between, and is also not in communication with the internal space (110i) (e.g., serpentine passage (115)) of the airfoil (110). The number of outer passages (172) is greater than the number of inner passages (171) that communicate with the first recess (internal space (117)) of the inner shroud (122) and are also not in communication with the internal space (110i) (e.g., serpentine passage (115)) of the airfoil (110).

[0124] According to the configuration of (1) above, the cooling capacity of the inner shroud (122) can be suppressed more than the cooling capacity of the outer shroud (121), thereby suppressing excessive cooling of the inner shroud (122) and allowing the cooling of the turbine stator (100) to be optimized. This allows for improved turbine efficiency.

[0125] (2) In some embodiments, the configuration of (1) above may satisfy any one of the following conditions (A), (B), or (C).

[0126] (A) The inner shroud (122) does not have an inner passage (171).

[0127] (B) The inner shroud (122) has one inner passage (171), and the cross-sectional area Si of the outlet end (171o) of the one inner passage (171) is smaller than the sum of the cross-sectional areas So of the outlet end (172o) of the at least one outer passage (172).

[0128] (C) The inner shroud (122) has two or more inner passages (171), and the value (Pi / Di) obtained by dividing the distance Pi between the inner passages (171) by the diameter Di of the inner passage (171) is greater than the value (Po / Do) obtained by dividing the distance Po between the outer passages (172) by the diameter Do of the outer passage (172).

[0129] According to the configuration of (2) above, if any one of the conditions of (A), (B), or (C) is satisfied, the cooling capacity of the inner shroud (122) can be suppressed compared to the cooling capacity of the outer shroud (121), thereby suppressing excessive cooling of the inner shroud (122) and allowing the cooling of the turbine stator (100) to be appropriate. This can improve turbine efficiency.

[0130] (3) In some embodiments, in the configuration of (2) above, the sum of the cross-sectional area Si of the outlet end (171o) of the inner passage (171) is divided by the sum of the area Sig of the gas pass surface (outer surface (122a)) of the inner shroud (122) and the area Sit of the cross-sectional area (122c) of the rear end of the inner shroud (122) (ΣSi / (Sig+Sit)), and the sum of the cross-sectional area So of the outlet end (172o) of the outer passage (172) is divided by the sum of the area Sog of the gas pass surface (outer surface (121a)) of the outer shroud (121) and the area Sot of the cross-sectional area (125c) of the rear end of the outer shroud (121) (ΣSo / (Sog+Sot)).

[0131] According to the configuration of (3) above, when compared per unit area with the total area of ​​the gas pass surface (outer surface (121a, 122a)) (Sog, Sig) and the cross-section (125c, 122c) on the trailing side of the shroud (121, 122), the flow rate of the cooling medium flowing through the inner passage (171) is less than the flow rate of the cooling medium flowing through the outer passage (172). As a result, the cooling capacity of the inner shroud (122) can be suppressed compared to the cooling capacity of the outer shroud (121), thereby suppressing excessive cooling of the inner shroud (122) and allowing the cooling of the turbine stator (100) to be appropriate.

[0132] (4) In some embodiments, in the configuration of (2) or (3) above, the sum of the cross-sectional area Si1 of the exit end (171o) of the inner passage (171) that is opened in the cross-sectional area (122c) of the inner shroud (122) is divided by the area Sit of the cross-sectional area (122c) of the inner shroud (122) that is opened in the exit end (171o) of the inner passage (171) is smaller than the sum of the cross-sectional area So1 of the exit end (172o) of the outer shroud (121) that is opened in the cross-sectional area (125c) of the outer shroud (121) is divided by the area Sot of the cross-sectional area (125c) of the outer shroud (121) that is opened in the exit end (172o) of the outer passage (172) is smaller than the sum of the cross-sectional area So1 of the exit end (172o) of the outer shroud (121) is divided by the area Sot of the cross-sectional area (125c) of the outer shroud (121) that is opened in the cross-sectional area (125c) that is opened in the exit end (172o) of the outer passage (172).

[0133] According to the configuration of (4) above, when compared per unit area of ​​the cross-section (125c, 122c) on the trailing side of the shroud (121, 122), the flow rate of the cooling medium flowing through the inner passage (171) is less than the flow rate of the cooling medium flowing through the outer passage (172). Thus, the cooling capacity of the inner shroud (122) can be suppressed compared to the cooling capacity of the outer shroud (121), thereby suppressing excessive cooling of the inner shroud (122) and allowing the cooling of the turbine stator (100) to be appropriate.

[0134] (5) In some embodiments, in any one of the configurations (1) to (4), the opening density of the inner impingement hole (114i) for impingement cooling the inner shroud (122) is smaller than the opening density of the outer impingement hole (114o) for impingement cooling the outer shroud (121).

[0135] According to the configuration of (5) above, the cooling capacity of the inner shroud (122) can be suppressed more than the cooling capacity of the outer shroud (121), so that the excessive cooling of the inner shroud (122) is suppressed, and thus the cooling of the turbine stator (100) can be appropriately controlled.

[0136] (6) In some embodiments, in the configuration of (5) above, the inner shroud does not have to be equipped with an impingement plate (130) having an inner impingement hole (114i) formed therein.

[0137] According to the configuration of (6) above, by not performing impingement cooling of the inner shroud (122), the excessive cooling of the inner shroud (122) is suppressed, and the cooling of the turbine stator (100) can be appropriately controlled.

[0138] (7) In some embodiments, in any one of the configurations (1) to (6), the value (Sic / Sig) obtained by dividing the area of ​​the heat-insulating coating (180) applied to the gas pass surface (outer surface (122a)) of the inner shroud (122) by the area of ​​the gas pass surface (outer surface (122a)) of the inner shroud (122) may be smaller than the value (Soc / Sog) obtained by dividing the area of ​​the heat-insulating coating (180) applied to the gas pass surface (outer surface (121a)) of the outer shroud (121) by the area of ​​the gas pass surface (outer surface (121a)) of the outer shroud (121) by the area of ​​the gas pass surface (outer surface (121a)) of the outer shroud (121) by the area of ​​the outer shroud (121).

[0139] If there is sufficient cooling capacity for the inner shroud (122), it is thought that the manufacturing cost of the turbine stator (100) can be reduced by narrowing the application range of the heat-insulating coating, which has a relatively high construction cost.

[0140] According to the configuration of (7) above, the manufacturing cost of the turbine stator (100) can be reduced by narrowing the application range of the heat-insulating coating (180) on the gas pass surface (outer surface (122a)) of the inner shroud (122).

[0141] (8) In some embodiments, in the configuration of (7) above, the inner shroud (122) does not have a heat-insulating coating (180) applied to the gas pass surface (outer surface (122a)) of the inner shroud (122).

[0142] If there is sufficient cooling capacity for the inner shroud, it is thought that the manufacturing cost of the turbine stator (100) can be reduced by narrowing the application range of the heat-insulating coating (180), which has a relatively high construction cost.

[0143] According to the configuration of (8) above, the manufacturing cost of the turbine stator (100) can be reduced by not applying a heat-insulating coating (180) to the gas pass surface (outer surface (122a)) of the inner shroud (122).

[0144] (9) In some embodiments, in any one of the configurations (1) to (8), the inner shroud (122) may have an exit passage (118) that is not in communication with the first recess (inner space (117)) and is in communication with the space (110i) (e.g., serpentine channel (115)) inside the airfoil (110).

[0145] According to the configuration of (9) above, when a cooling medium is supplied to the space (110i) (e.g., serpentine flow path (115)) inside the airfoil (110), the cooling medium can be discharged to the outside of the turbine stator (100) through the outlet passage (118).

[0146] (10) In some embodiments, in the configuration of (9) above, the space (110i) inside the airfoil (110) to which the exit passage (118) is connected may be a serpentine passage (115) for cooling the airfoil (110).

[0147] According to the configuration of (10) above, the cooling medium supplied to the serpentine channel (115) for cooling the airfoil (110) can be discharged to the outside of the turbine stator (100) through the outlet channel (118).

[0148] (11) A turbine stator (100) according to at least one embodiment of the present disclosure comprises an airfoil (110), an inner shroud (122) provided on the inner side of the airfoil (110), and an outer shroud (121) provided on the outer side of the airfoil (110). The opening density of the inner impingement hole (114i) for impingement cooling the inner shroud (122) is smaller than the opening density of the outer impingement hole (114o) for impingement cooling the outer shroud (121).

[0149] According to the configuration of (11) above, the cooling capacity of the inner shroud (122) can be suppressed more than the cooling capacity of the outer shroud (121), thereby suppressing excessive cooling of the inner shroud (122) and allowing the cooling of the turbine stator (100) to be appropriate. This can improve turbine efficiency.

[0150] (12) In some embodiments, in any one of the configurations (1) to (11), the turbine stator (100) may be a turbine 3rd stage stator (3rd stage stator (33)).

[0151] According to the configuration of (12) above, when the combustion gas temperature on the hub side of the turbine third stage stator (third stage stator (33)) is lower than the pre-determined value and the metal temperature of the inner shroud (122) of the turbine third stage stator (third stage stator (33)) is also lower than the pre-determined value, the excessive cooling of the inner shroud (122) can be suppressed, thereby allowing the cooling of the turbine stator (100) (third stage stator (33)) to be appropriate. Explanation of the symbols

[0152] 1 Gas turbine 33 Third stage stator (Turbine third stage stator) 100 turbine stator 110 Airfoil 110i space 114 through hole 114i inner impingement hole 114o outer impingement hole 116 Internal space (second depression) 117 Internal space (1st depression) 118 Exit Passage 121 Lateral Shroud 121a External surface 122 Medial Shroud 122a External surface 122c Section 125c cross-section 130 Impingement Plates 171 Medial passage 1710 Exit 172 Lateral passage 172o Exit 180 heat-insulating coating

Claims

Claim 1 The apparatus comprises an airfoil, an inner shroud provided on the inner circumference side of the airfoil, and an outer shroud provided on the outer circumference side of the airfoil; the inner shroud has a first concave portion formed on a surface opposite to the airfoil with the gas pass surface of the inner shroud in between; the outer shroud has a second concave portion formed on a surface opposite to the airfoil with the gas pass surface of the outer shroud in between, and at least one outer passage that communicates with the second concave portion and is not in communication with the internal space of the airfoil; the number of the outer passages is greater than the number of inner passages in the inner shroud that communicate with the first concave portion and are not in communication with the internal space of the airfoil; and the value (Sic / Sig) obtained by dividing the application area Sic of the heat-insulating coating applied to the gas pass surface of the inner shroud by the area Sig of the gas pass surface of the inner shroud is the gas pass surface of the outer shroud A turbine stator having a non-communication outlet passage in the first concave portion, wherein the application area Soc of the applied heat-insulating coating is smaller than the value (Soc / Sog) obtained by dividing the area Sog of the gas pass surface of the outer shroud, and the inner shroud communicates with the downstream end of the cooling path located at the downstream end of the space inside the airfoil. Claim 2 A turbine stator according to claim 1, satisfying any one of the following conditions (A), (B), or (C). (A) The inner shroud does not have the inner passage. (B) The inner shroud has one inner passage, and the cross-sectional area Si of the outlet end of the one inner passage is smaller than the sum of the cross-sectional areas So of the outlet ends of at least one outer passage, ΣSo. (C) The inner shroud has two or more inner passages, and the value obtained by dividing the distance Pi between the inner passages by the diameter Di of the inner passage (Pi / Di) is greater than the value obtained by dividing the distance Po between the outer passages by the diameter Do of the outer passage (Po / Do). Claim 3 A turbine stator according to claim 2, wherein the sum of the cross-sectional area Si of the outlet end of the inner passage, ΣSi, is divided by the sum of the area Sig of the gas pass surface of the inner shroud and the cross-sectional area Sit of the trailing edge of the inner shroud (ΣSi / (Sig+Sit)), which is smaller than the sum of the cross-sectional area So of the outlet end of the outer passage, ΣSo, is divided by the sum of the area Sog of the gas pass surface of the outer shroud and the cross-sectional area Sot of the trailing edge of the outer shroud (ΣSo / (Sog+Sot)). Claim 4 A turbine stator according to claim 2 or claim 3, wherein the sum of the cross-sectional areas Si1 of the outlet ends of the inner passage that are open at the cross-section of the trailing edge of the inner shroud (ΣSi1 / Sit) is divided by the cross-sectional area Sit of the cross-section of the trailing edge of the inner shroud, and the sum of the cross-sectional areas So1 of the outlet ends of the outer passage that are open at the cross-sectional area of ​​the trailing edge of the outer shroud (ΣSo1 / Sot) is smaller than the sum of the cross-sectional areas So1 of the outlet ends of the outer passage that are open at the cross-sectional area Sot of the cross-sectional area of ​​the trailing edge of the outer shroud. Claim 5 A turbine stator according to any one of claims 1 to 3, wherein the opening density of the inner impingement hole for impingement cooling the inner shroud is smaller than the opening density of the outer impingement hole for impingement cooling the outer shroud. Claim 6 In claim 5, the inner shroud is not equipped with an impingement plate having the inner impingement hole formed therein. Claim 7 In any one of claims 1 to 3, the inner shroud is a turbine stator in which a heat-insulating coating is not applied to the gas pass surface of the inner shroud. Claim 8 delete Claim 9 In any one of claims 1 to 3, the space inside the airfoil portion to which the outlet passage communicates is a turbine stator that is a serpentine flow path for cooling the airfoil portion. Claim 10 The airfoil portion, an inner shroud provided on the inner circumference side of the airfoil portion, and an outer shroud provided on the outer circumference side of the airfoil portion are provided; the opening density of the inner impingement holes for impingement cooling the inner shroud is smaller than the opening density of the outer impingement holes for impingement cooling the outer shroud; the value obtained by dividing the application area Sic of the heat-insulating coating applied to the gas pass surface of the inner shroud by the area Sig of the gas pass surface of the inner shroud (Sic / Sig) is smaller than the value obtained by dividing the application area Soc of the heat-insulating coating applied to the gas pass surface of the outer shroud by the area Sog of the gas pass surface of the outer shroud (Soc / Sog); the inner shroud communicates with the downstream end of a cooling channel located on the downstream side of the internal space of the airfoil portion, and a first concave portion formed on the surface opposite to the airfoil portion with the gas pass surface of the inner shroud in between Turbine stator having a non-combustible exit passage. Claim 11 In claim 10, the inner shroud is a turbine stator in which a heat-insulating coating is not applied to the gas pass surface of the inner shroud. Claim 12 In any one of claims 1 to 3, the turbine stator is a turbine stator that is a turbine three-stage stator.

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