Turbine vane and mating structure, gas turbine, and method for adjusting the amount of cooling air for the mating structure

Abstract

Provided is a turbine static blade composed of a blade body and shrouds respectively formed on both ends in the blade height direction of the blade body. The shroud includes: a main member having a heat-resistance coat film section which extends from the axial upstream end to the axial downstream end of a gas passing surface; and a protruding section which protrudes from a sided end section on the front edge side of the main member toward the axial upstream side. The protruding section comprises a plurality of cooling paths which are disposed in an annular shape at predetermined intervals in the circumferential direction, and connect an inlet opening which is formed in an inner circumferential surface of the protruding section and an inner surface of the main member, and an outlet opening which is formed in an outer circumferential surface of the protruding section.

Description

[Technical field]

[0001] The present disclosure relates to a cooling aisle A turbine vane having Fitting Structure and gas turbine using the same and a method for adjusting the amount of cooling air in a fitting structure Regarding. This application claims priority to Japanese Patent Application No. 2022-000793, filed on January 6, 2022, the contents of which are incorporated herein by reference. [Background technology]

[0002] In a gas turbine, air compressed by a compressor is mixed with fuel in a combustor and combusted to generate high-temperature combustion gas. The combustion gas rotates a rotor in which turbine vanes and turbine moving blades are arranged alternately, and the turbine moving blades are fixed, generating electric power. A seal member is provided between the combustor and the turbine vane to seal the gap between the combustor and the turbine vane. Furthermore, at a connection between the seal member provided between the combustor and the turbine vane and the turbine vane, a fixed seal is provided on the seal member to prevent cooling air in the casing from leaking into the combustion gas flow path through the gap between the seal member and the turbine vane, and the fixed seal comes into contact with the turbine vane to seal the gap between the seal member and the turbine vane. The turbine vane, combustor, and seal member are exposed to high-temperature combustion gas and are therefore susceptible to thermal damage. For this reason, cooling air passages are provided in these components to supply cooling air to cool the components and prevent thermal damage. An example of such a cooling structure is disclosed in Patent Document 1. [Prior art documents] [Patent documents]

[0003] [Patent Document 1] Patent Publication No. 2021-32082 Summary of the Invention [Problem to be solved by the invention]

[0004] However, in the case of gas turbines that have been in operation for a long time, the contact surface between the stationary seal and the turbine vane wears away due to vibrations, etc., resulting in a breakdown of the turbine vane and the seal member. Fitting If the gap between the turbine vane and the seal member at the portion where the turbine vane is inserted widens, the combustion gas may be caught in the gap, and the high-temperature combustion gas may cause thermal damage to the components of the turbine vane.

[0005] The object of the present disclosure is to provide a method for manufacturing a turbine stator blade and a seal member. Fitting When combustion gas is caught in the gap in the Fitting Turbine stator vane with cooling structure capable of suppressing thermal damage to components in the turbine vane and Fitting The present invention aims to provide a structure and a gas turbine equipped with the same. [Means for solving the problem]

[0006] One aspect of the present disclosure is a turbine vane comprising a vane body and a shroud formed on both ends of the vane body in a vane height direction, the shroud comprising a main member having a heat-resistant coating portion extending from an axial upstream end to a downstream end of a gas path surface, and a leading edge side of the main member. Side and a protrusion protruding axially upstream from the end portion, the protrusion being arranged in a ring shape at a predetermined interval in the circumferential direction and comprising a plurality of cooling passages connecting an inlet opening formed on an inner surface of the protrusion or an inner surface of the main member and an outlet opening formed on an outer circumferential surface of the protrusion. Effect of the Invention

[0007] According to the turbine vane of the present disclosure, even if combustion gas is caught in the gap at the fitting portion between the turbine vane and the seal member, cooling of the protruding portion of the turbine vane is enhanced and thermal damage to the protruding portion is suppressed, thereby improving the reliability of the gas turbine. [Brief description of the drawings]

[0008] [Figure 1] FIG. 1 is a schematic configuration diagram of a gas turbine according to a first embodiment of the present disclosure. [Diagram 2]FIG. 2 is a diagram showing a configuration around a combustor of the first embodiment according to the present disclosure. As shown in FIG. [Diagram 3] FIG. 3 is a diagram showing a configuration around a turbine vane and a seal member according to the first embodiment of the present disclosure. [Figure 4] FIG. 4 is a structural diagram of a combination of a turbine vane and a seal member according to the first embodiment of the present disclosure. [Diagram 5] FIG. 5 is a plan layout diagram of a turbine vane provided with a cooling air passage according to the first embodiment of the present disclosure. [Figure 6] FIG. 6 shows a modification of the cooling air passage of the first embodiment according to the present disclosure. [Figure 7] FIG. 7 is a diagram showing the relationship between the gap amount and the amount of cooling air. [Figure 8] FIG. 8 is a diagram showing the relationship between the amount of cooling air and the combustion gas temperature at the fitting portion. [Figure 9] FIG. 9 is a structural diagram of a combination of a turbine vane and a seal member according to the second embodiment of the present disclosure. [Figure 10] FIG. 10 is a structural diagram of a combination of a turbine vane and a seal member according to the third embodiment of the present disclosure. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0009] Hereinafter, an embodiment of the present disclosure will be described with reference to the drawings.

[0010] <<Gas Turbine Configuration>> One embodiment of a gas turbine according to the present disclosure will be described with reference to Fig. 1. Fig. 1 is a schematic configuration diagram showing a gas turbine 1 according to this embodiment.

[0011] 1, a gas turbine 1 according to this embodiment includes a compressor 2 for generating compressed air A, a combustor 4 for generating combustion gas G by using the compressed air A and fuel FL, and a turbine 6 that is rotationally driven by the combustion gas G. In the case of the gas turbine 1 for power generation, a generator (not shown) is connected to the turbine 6, and electricity is generated by the rotational energy of the turbine 6.

[0012] The compressor 2 includes a compressor casing 10, an intake chamber 12 provided on the inlet side of the compressor casing 10 for taking in air, a rotor 8 provided so as to penetrate both the compressor casing 10 and a turbine casing 22 described later, and various blades arranged in the compressor casing 10. The various blades include an inlet guide vane 14 provided on the intake chamber 12 side, a plurality of compressor stator vanes 16 fixed to the compressor casing 10, and a plurality of compressor rotor blades 18 planted on the rotor 8 so as to be arranged alternately in the axial direction with respect to the compressor stator vanes 16. In such a compressor 2, the air taken in from the intake chamber 12 is compressed in the process of passing through the plurality of compressor stator vanes 16 and the plurality of compressor rotor blades 18, and compressed air A is generated. The compressed air A is discharged from the compressor 2 into a space 21 in a casing 20 on the axial downstream side, and is supplied to a combustor 4 arranged in the casing 20.

[0013] As shown in Fig. 1, a plurality of combustors 4 are arranged in an annular shape around a rotor 8 inside a casing 20. Fuel FL and compressed air A generated by the compressor 2 are supplied to the combustors 4, and the fuel FL is combusted to generate high-temperature combustion gas G, which is a working fluid for a turbine 6. The generated combustion gas G is sent from the combustors 4 to the turbine 6 located downstream in the axial direction.

[0014] The turbine 6 includes a turbine casing 22 and various turbine blades arranged inside the turbine casing 22. The turbine blades include a plurality of turbine stator vanes 40 fixed to the turbine casing 22 side, and a plurality of turbine rotor blades 26 implanted in the rotor 8 so as to be arranged alternately with respect to the turbine stator vanes 40 in the axial direction. In the turbine 6, the rotor 8 extends in the axial direction, and the combustion gas G discharged from the turbine casing 22 is discharged to an exhaust casing 28 downstream in the axial direction. In FIG. 1, the left side of the figure is the axial upstream side, and the right side of the figure is the axial downstream side. In the following description, the term "radial direction" simply refers to a direction perpendicular to the rotor 8. In addition, the term "circumferential direction" refers to the rotation direction of the rotor 8.

[0015] The turbine rotor blades 26 are configured to generate a rotational driving force from the high-temperature, high-pressure combustion gas G flowing together with the turbine stator vanes 40 inside the turbine casing 22. This rotational driving force is transmitted to the rotor 8, and drives a generator (not shown) connected to the rotor 8.

[0016] An exhaust chamber 29 is connected to the axially downstream side of the turbine casing 22 via an exhaust casing 28. The combustion gas G after driving the turbine 6 is discharged to the outside through the exhaust casing 28 and the exhaust chamber 29.

[0017] Combustor surrounding structure FIG. 2 shows a schematic structure around the combustor 4 of the gas turbine 1 in one embodiment. FIG. 3 shows a structure around the turbine vane 40 and the seal member 60. As shown in FIG. 2, a plurality of combustors 4 are arranged in an annular shape in the casing 20 with the rotor 8 at the center. The combustor 4 has a plurality of fuel nozzles 30 that supply fuel FL to the combustor 4, and a combustion tube 32 that mixes and burns the fuel FL and compressed air A. The combustion tube 32 has an inner tube 33 that burns the fuel FL and the compressed air A to generate combustion gas G, and a transition piece 34 that supplies the combustion gas G to the turbine 6. A flange 35 that connects to the turbine vane 40 via the seal member 60 is arranged at the axial downstream end of the transition piece 34. The flange 35 is formed on the entire outer periphery of the transition piece 34 that forms the combustion gas flow path 37.

[0018] 3, a predetermined gap is provided between the flange 35 of the transition piece 34 and the turbine stator vane 40 connected on the axial downstream side to absorb the difference in thermal expansion in the axial direction, and a removable seal member 60 is inserted therein. A plurality of seal members 60 are arranged in an annular shape around the rotor 8. The seal member 60 is a combination of an inner seal member 61 arranged radially inside with respect to the combustion gas flow path 37 and an outer seal member 62 arranged radially outside, constituting a set of seal members 60. Although a set of seal members 60 is arranged on the axial downstream side corresponding to a set of combustors 4, a set of seal members 60 may be arranged corresponding to a plurality of combustors 4. The seal member 60 is connected, at its axial upstream side, to a transition piece 34 that forms a combustion gas flow path 37 via a flange 35, and is detachably connected, at its axial downstream side, to a shroud 44 of a turbine stator blade 40. Fitting A radial outer surface 61 a of the inner seal member 61 and a radial inner surface 62 a of the outer seal member 62 are in contact with the combustion gas flow path 37 .

[0019] First Embodiment Fig. 4 shows an example of a combined structure of the turbine stator vane 40 and the seal member 60 according to the first embodiment, in which the turbine stator vane 40 is combined with an inner seal member 61. The combined structure of the turbine stator vane 40 and the outer seal member 62 has a similar structure. Fig. 5 shows an example of a plan cross-sectional view of the turbine stator vane 40 as viewed from the blade height direction. 3 to 5 is formed of a blade body 42 extending in the blade height direction, and a shroud 44 connected to the blade body 42 at both ends in the blade height direction of the blade body 42. The shroud 44 is composed of an inner shroud 45 formed at the inner end in the blade height direction with respect to the blade body 42, and an outer shroud 46 formed at the outer end in the blade height direction. An outer surface 45a of the inner shroud 45 and an outer surface 46a of the outer shroud 46 form a gas path surface 44a and are in contact with the combustion gas flow path 37.

[0020] As shown in Fig. 4 and Fig. 5, the shroud 44 (inner shroud 45) of the turbine stator blade 40 is composed of a main member 47 constituting the main body of the shroud 44, a plate-shaped protruding portion 48 that protrudes axially upstream from an axially upstream side end portion 47a of the main member 47 and extends annularly in the circumferential direction, and an impingement plate 50 having a plurality of through holes 50a, the details of which will be described later. 45The main member 47 is composed of a bottom portion 47c forming the bottom surface of the main member 47, and an outer wall portion 47d formed on an outer edge 47b that is the outer periphery of the bottom portion 47c and protruding in the opposite direction to the combustion gas flow path 37 in the blade height direction with respect to the gas path surface 44a. The main member 47 and the protruding portion 48 are formed integrally. Inside the main member 47, a cavity 51 is formed that is surrounded by the outer wall portion 47d and recessed from the outer wall end portion 47e in the direction approaching the gas path surface 44a in the blade height direction. An impact plate 50 having a plurality of through holes 50a is arranged in the cavity 51, and the cavity 51 is divided by the impact plate 50 into an inner cavity 51a formed on the side approaching the gas path surface 44a in the blade height direction and an outer cavity 51b formed adjacent to the gas path surface 44a in the direction away from the gas path surface 44a in the blade height direction. The inner cavity 51 a and the outer cavity 51 b communicate with each other via a through hole 50 a of the collision plate 50 .

[0021] A cooling passage 49, the details of which will be described later, is formed in the protruding portion 48. The protruding portion 48 protrudes toward the leading edge 43a of the shroud 44 in the axial direction, and extends in the circumferential direction to the positions of the flanges 35 of the transition piece 34 formed at both ends in the circumferential direction. The outer surface of the bottom portion 47c in contact with the combustion gas G forms a gas path surface 44a.

[0022] A heat-resistant coating 52 is formed on the gas path surface 44a of the shroud 44 to suppress thermal damage from the high-temperature combustion gas G flowing through the combustion gas flow path 37. The heat-resistant coating 52 is formed so as to cover almost the entire surface of the gas path surface 44a of the shroud 44, and is formed on the entire surface ranging from the position of the coating portion upstream end 52a on the axial upstream side of the main member 47 and the protruding portion 48 in contact with the combustion gas G to the downstream end 47f of the main member 47 on the axial downstream side, and from the positive pressure surface end 47g to the negative pressure surface end 47h in the circumferential direction.

[0023] <Sealing material> The seal member 60 (inner seal member 61) shown in FIG. 4 is disposed between the transition piece 34 of the combustor 4 and the turbine stator blade 40 in the axial direction. The seal member 60 is composed of a plate-shaped first body portion 60a extending from the axial upstream side to the axial downstream side and formed in a ring shape in the circumferential direction, a plate-shaped second body portion 60b connected to the axial upstream end of the first body portion 60a, extending inward in the blade height direction, and formed in a fan shape in the circumferential direction, a plate-shaped third body portion 60c protruding in the axial downstream direction from a middle position in the blade height direction of the second body portion 60b and formed in a ring shape in the circumferential direction, and a connection portion 60d connected at an inner end portion in the inner direction in the blade height direction of the second body portion 60b, extending in the axial upstream direction, and extending outward in the blade height direction from the axial upstream end. The first body portion 60a, the second body portion 60b, the third body portion 60c, and the connection portion 60d are integrally formed. A gap formed in the blade height direction between the second body portion 60b and the connecting portion 60d has a structure in which the flange 35 of the transition piece 34 can be inserted from the outer side in the blade height direction. In addition, a recess 63 recessed in the axial direction from the axial downstream end of the first body portion 60a and the third body portion 60c to the axial upstream side is formed between the first body portion 60a and the third body portion 60c, and the protruding portion 48 of the turbine stator blade 40 is inserted into the recess 63. Fitting The sealing member 60 has a structure that can be used for cooling the combustion gas flow passage 37. The sealing member 60 includes a cooling passage 60e arranged in the first body portion 60a and the second body portion 60b to avoid thermal damage due to heat input from the combustion gas flow passage 37. The cooling passages 60e are arranged in a plurality of rows at predetermined intervals in the circumferential direction of the first body portion 60a and the second body portion 60b. The cooling passage 60e extends in the second body portion 60b in the outward direction in the blade height direction, extends in the axial direction in the first body portion 60a from the position where the first body portion 60a and the second body portion 60b are connected, and is formed so as to open to the combustion gas flow passage 37. The compressed air A in the space 21 surrounded by the casing 20 is used as the cooling air supplied to the cooling passage 60e from the inner end of the second body portion 60b in the blade height direction.

[0024] In the above description, the inner seal member 61 has been taken as an example, but the outer seal member 62 has the same structure as the inner seal member 61. However, the inner seal member 61 is a seal member 60 disposed between the transition piece 34 of the combustor 4 and the inner shroud 45 of the turbine stator vane 40. On the other hand, the outer seal member 62 differs from the inner seal member 61 in that it is disposed between the transition piece 34 of the combustor 4 and the outer shroud 46 of the turbine stator vane 40. Therefore, in the case of the outer seal member 62, in the above description of the inner seal member 61, the outward direction in the blade height direction should be read as the inward direction in the blade height direction, and the inward direction in the blade height direction should be read as the outward direction in the blade height direction.

[0025] <Interlocking structure> 3 and 4, the protruding portion 48 of the turbine stator vane 40 is inserted into the recessed portion 63 of the seal member 60 from the axial downstream side, and a fitting structure 70 consisting of the seal member 60 and the protruding portion 48 of the turbine stator vane 40 is formed. The fitting structure 70 is composed of the protruding portion 48 of the turbine stator vane 40 and the first main body portion 60a and third main body portion 60c of the seal member 60 in combination. Fitting The structure 70 is fitted between the protrusion 48 and the seal member 60 while maintaining a slight gap in the blade height direction, and is a structure that is detachable in the axial direction.

[0026] As shown in FIG. 4, the position of the coating upstream end 52a on the axial upstream side of the heat-resistant coating 52 formed on the gas path surface 44a of the shroud 44 is axially different from the position of the axial downstream end 60aa of the first main body portion 60a of the seal member 60. matchIt is preferable that the heat-resistant coating 52 is formed on the entire outer circumferential surface 48a of the protruding portion 48, or slightly downstream in the axial direction from the position of the axial downstream end 60aa. If the heat-resistant coating 52 is formed on the entire outer circumferential surface 48a of the protruding portion 48, the outer circumferential surface 48a of the protruding portion 48 may come into contact with the inner circumferential surface 60ba of the first main body portion 60a of the seal member 60 due to mechanical vibrations generated during the operation of the gas turbine 1, and the heat-resistant coating 52 may be damaged by the contact. Therefore, in order to avoid damage to the heat-resistant coating 52, it is preferable that the position of the coating upstream end 52a of the heat-resistant coating 52 formed on the gas path surface 44a of the shroud 44 is at least the same position as the position of the axial downstream end 60aa of the first main body portion 60a of the seal member 60, or is axially downstream from the position of the axial downstream end 60aa of the first main body portion 60a. That is, when viewed from the blade height direction, it is desirable not to form the heat-resistant coating 52 in an area upstream of the upstream end 52a of the heat-resistant coating 52, which is an area of ​​the outer circumferential surface 48a of the protruding portion 48 that overlaps with the seal member 60. Note that the position of the axial downstream end 60aa of the first main body portion 60a of the seal member 60 is the axially most downstream position of the seal member 60, and therefore, unless otherwise specified, the position of the axial downstream end 60aa of the first main body portion 60a of the seal member 60 may be read as the position of the axial downstream end of the seal member 60.

[0027] Fixed seal As shown in FIG. 4, a plate-shaped stationary seal 64 extending in the circumferential direction along the outer peripheral surface 60ca is disposed on the outer peripheral surface 60ca of the third main body portion 60c of the seal member 60 (inner seal member 61). The stationary seal 64 is joined to the outer peripheral surface 60ca of the third main body portion 60c by welding or the like. The outer peripheral surface 64a of the stationary seal 64 joined to the third main body portion 60c in the blade height direction faces the recessed portion 63 and contacts the inner inner peripheral surface 48c of the protruding portion 48 of the turbine stator blade 40 in the blade height direction. During operation of the gas turbine 1, the outer peripheral surface 64a of the stationary seal 64 and the inner inner peripheral surface 48c of the protruding portion 48 are constantly in contact with each other, and a seal surface is formed between the outer peripheral surface 64a of the stationary seal 64 and the inner inner peripheral surface 48c of the protruding portion 48. By forming a sealing surface, high-pressure compressed air A accumulated in the space 21 inside the casing 20 is prevented from leaking into the combustion gas flow path 37 through the gap between the protrusion 48 and the sealing member 60.

[0028] As described above, the protrusion 48 of the turbine stator blade 40 and the first body portion 60a and the third body portion 60c of the seal member 60 are FittingIn the structure 70, the gap 71 formed between the outer peripheral surface 48a of the protruding portion 48 and the inner peripheral surface 60ba of the first main body portion 60a in the blade height direction is desirably made as small as possible in order to suppress the entrainment of the combustion gas G from the combustion gas flow path 37. On the other hand, between the protruding portion 48 and the third main body portion 60c of the turbine stator vane 40, the inner peripheral surface 48c of the protruding portion 48 and the outer peripheral surface 60ca of the third main body portion 60c are in contact with each other via the fixed seal 64, so no gap is generated in the blade height direction. The range in which the gap 71 is formed is the range in which the first main body portion 60a of the seal member 60 and the protruding portion 48 overlap in the axial direction when viewed from the blade height direction, and is the range from the position of the upstream end 48b of the protruding portion 48 to the gap inlet 71a of the gap 71. The position of the gap inlet 71a of the gap 71 corresponds to the position of the downstream end 60aa of the first main body portion 60a in the axial direction. The overlapping range of the gaps 71 in the circumferential direction may be the range from both circumferential ends of the turbine stator vane 40 to both circumferential ends of the protruding portion 48. Also, it may be the range up to the flanges 35 at both circumferential ends of the axial downstream end of the transition piece 34. In the following description, several embodiments of the cooling passage 49 disposed in the protruding portion 48 will be described.

[0029] First Embodiment The cooling passage 49 according to the first embodiment will be described with reference to Figs. 4 and 5. The protruding portion 48 of the turbine stator blade 40 includes a cooling passage 49 that connects an outlet opening 49a formed on an outer peripheral surface 48a of the protruding portion 48 on the outer side in the blade height direction to an inlet opening 49b formed on an inner peripheral surface 48c of the protruding portion 48 on the inner side in the blade height direction, and penetrates the protruding portion 48 in the blade height direction. A plurality of cooling passages 49 are arranged at predetermined intervals in the circumferential direction of the protruding portion 48. It is preferable that the outlet opening 49a of the cooling passage 49 formed on the outer peripheral surface 48a of the protruding portion 48 is axially upstream of the position of the coating portion upstream end 52a, which is the axially upstream end of the heat-resistant coating portion 52 formed on the gas path surface 44a of the shroud 44, and is formed in a gap portion 71 in a range from the position of the upstream end 48b of the outer peripheral surface 48a of the protruding portion 48 to the coating portion upstream end 52a. Further, it is more preferable that the outlet opening 49a be disposed at a position closer to the coating portion upstream end 52a than to the axially upstream end of the outer circumferential surface 48a of the protrusion 48.

[0030] 《Entrance Opening》 The inlet opening 49b of the cooling passage 49 is formed on the inner peripheral surface 48c of the protruding portion 48, and is desirably formed at a position close to the side end portion 47a on the leading edge 43a side of the shroud 44. Between the turbine stator vane 40 and the seal member 60, a predetermined gap is formed in the axial direction to absorb the thermal expansion difference between them in the axial direction. In order to prevent the compressed air A in the space 21 of the casing 20 from flowing out to the combustion gas flow path 37 due to the thermal expansion difference between them, a seal surface is formed between the outer peripheral surface 64a of the stationary seal 64 fixed to the third main body portion 60c and the inner peripheral surface 48c of the protruding portion 48. The relative positions of the protruding portion 48 of the turbine stator vane 40 and the seal member 60 in the axial direction change while maintaining the seal surface. Therefore, even if the third main body portion 60c of the seal member 60 approaches the side end 47a of the shroud 44 of the turbine vane 40 due to the thermal expansion difference between the protruding portion 48 of the turbine vane 40 and the seal member 60, it is preferable to arrange the position of the inlet opening 49b at a position close to the side end 47a so that the inlet opening 49b of the cooling passage 49 is not blocked by the fixed seal 64. In order to provide the inlet opening 49b of the cooling passage 49 at a position close to the side end 47a on the leading edge 43a side of the turbine vane 40 on the axial downstream side, it is preferable that the cooling passage 49 is an inclined passage that faces the downstream side in the flow direction of the cooling air and inclines toward the upstream side in the axial direction. When the protruding portion 48 is viewed from the blade height direction, it is preferable that the position of the inlet opening 49b of the cooling passage 49 is arranged at a position closer to the side end 47a on the leading edge 43a side of the turbine vane 40 than the position of the outlet opening 49a, and arranged at a position closer to the side end 47a on the leading edge 43a side than the axial downstream end of the fixed seal 64.

[0031] Circumferential arrangement of cooling passages 5, the cooling passages 49 arranged in the protruding portion 48 are arranged at predetermined intervals in the circumferential direction over the entire circumferential width L1 range between the pressure surface side end 47g and the suction surface side end 47h of the protruding portion 48. When the cooling passages 49 are viewed from the outside in the blade height direction, an outlet opening 49a formed in an outer peripheral surface 48a of the protruding portion 48 is disposed axially upstream of an inlet opening 49b formed in an inner peripheral surface 48c of the protruding portion 48. Moreover, it is preferable that the direction in which the cooling passages 49 extend in the leading edge-trailing edge direction be an inclined passage having a predetermined circumferential inclination with respect to the axial direction, and it is even more preferable that the cooling passages 49 be inclined toward the suction surface side end 47h from the direction in which the pressure surface side end 47g extends. By making the cooling passages 49 into inclined passages, the cooling area is increased, the cooling of the protrusion 48 is strengthened, and the amount of cooling air discharged from the cooling passages 49 can be reduced.

[0032] <Modification of Cooling Passage> 6 shows a modified example of the cooling passages 49 arranged in the protruding portion 48. The cooling passages 49 shown as the modified example are arranged in a circumferential direction not in the range of the entire circumferential width L1 (FIG. 5) from the pressure surface side end 47g to the suction surface side end 47h of the protruding portion 48, but in a range of a partial circumferential width L2, and the position of the midpoint Q of the arranged width L2 is arranged in a position circumferentially closer to the pressure surface side end 47g side by a length L3 from the position of the leading edge 43a of the blade surface 43. The gas path surface 44a of the shroud 44 of the turbine stator blade 40 is more likely to be overheated on the gas path surface 44a on the pressure surface side end 47g side than on the gas path surface 44a on the suction surface side end 47h side by the combustion gas G flowing down along the blade surface 43 from the leading edge 43a side to the trailing edge 43b side. This tendency is the same for the protrusion 48, and it is effective to arrange the cooling passages 49 in the protrusion 48 in a range that strengthens cooling on the pressure surface end 47g side rather than the suction surface end 47h side, based on the position of the leading edge 43a. By arranging the cooling passages 49 in a circumferential range in a partial range closer to the pressure surface end 47g side than the leading edge 43a, the protrusion 48 is effectively cooled. Also, compared to the arrangement of the cooling passages 49 shown in FIG. 5, the amount of cooling air is further reduced.

[0033] 《Effect of cooling passage》 Here, the operation and effect of providing the cooling passage 49 in the protruding portion 48 will be described. As described above, in FIG. 4, the compressed air A accumulated in the space 21 of the casing 20 is supplied to the cooling passage 49 as cooling air. The cooling air is supplied to the cooling passage 49 from the inlet opening 49b of the protruding portion 48, and convection-cools the protruding portion 48. After convection-cooling the protruding portion 48, the cooling air flows out of the outlet opening 49a of the cooling passage 49 into the cooling passage 49 formed between the protruding portion 48 and the seal member 60. Fitting The cooling air is discharged into a gap 71 of the structure 70. The cooling air discharged into the gap 71 collides with an inner circumferential surface 60ba of the first main body portion 60a of the seal member 60 arranged opposite the outer circumferential surface 48a of the protrusion 48 in the blade height direction, and impinges and cools the inner circumferential surface 60ba of the first main body portion 60a. The cooling air after impingement-cooling the inner circumferential surface 60ba of the first main body portion 60a flows axially downstream inside the gap 71 and is discharged into the combustion gas flow path 37 from a gap inlet 71a of the gap 71 formed between the downstream end 60aa of the first main body portion 60a, which is the axial downstream end of the seal member 60, and the outer circumferential surface 48a of the protrusion 48. However, the action and effect of the cooling air differs depending on whether the position of the outlet opening 49a through which the cooling air is discharged from the cooling passage 49 is into the gap 71 upstream in the axial direction from the gap inlet 71a, or into the combustion gas flow path 37 downstream in the axial direction from the gap inlet 71a and downstream from the position of the axially downstream end 60aa of the seal member 60. The difference will be described below.

[0034] 7 shows the relationship between the gap amount CB of the gap 71 and the amount of cooling air AV discharged from the cooling passage 49 when the cooling air is discharged from the cooling passage 49 to the gap 71, with the horizontal axis showing the gap amount CB and the vertical axis showing the amount of cooling air AV. The gap amount CB means the distance in the blade height direction between the outer peripheral surface 48a of the protruding portion 48 in the gap 71 and the inner peripheral surface 60ba of the first main body portion 60a of the seal member 60 arranged opposite to each other in the blade height direction (radial direction). In FIG. 7, in the first region R1 where the gap amount CB is smaller than the point P1, the gap amount CB is small, so that the pressure loss generated while the cooling air in the space 21 of the casing 20 is discharged to the combustion gas flow path 37 is mostly the discharge side pressure loss of the cooling air in the section from the outlet opening 49a of the cooling passage 49 to the gap inlet 71a, and is approximately equivalent to the pressure difference between the combustion gas flow path 37 and the space 21 of the casing 20. Here, the discharge side pressure loss of the cooling air refers to the pressure loss of the cooling air in the gap 71 that occurs when the cooling air discharged from the outlet opening 49a of the cooling passage 49 collides with the inner circumferential surface 60ba of the first main body portion 60a, changes its flow direction to the axial downstream side, flows down the gap 71 to the axial downstream side, and is discharged from the gap inlet 71a to the combustion gas flow path 37. loss Therefore, the pressure loss of the cooling air in the cooling passage 49 between the space 21 of the casing 20 and the outlet opening 49a of the cooling passage 49 is small, and the amount of cooling air AV flowing through the cooling passage 49 is limited to an extremely small amount. That is, the amount of cooling air AV in the first region R1 is limited by the gap amount CB of the gap portion 71 at the P1 point or smaller than the P1 point, and the amount of cooling air AV discharged from the cooling passage 49 cannot be increased unless the gap amount CB is increased to a value exceeding the P1 point. Conversely, as long as at least the gap amount CB of the P1 point or a gap amount CB smaller than the P1 point is maintained, the cooling air in the space 21 of the casing 20 hardly flows out to the combustion gas flow path 37 side. In other words, the first region R1 is a gap rate-determining region in which the amount of cooling air AV is limited by the gap amount CB of the gap portion 71.

[0035] On the other hand, in the second region R2 where the gap amount CB increases beyond point P1, the cooling air amount AV increases as the gap amount CB increases. In the second region R2, the discharge pressure loss of the gap 71 relatively decreases as the gap amount CB increases, so that the pressure difference between the inlet opening 49b and the outlet opening 49a of the cooling passage 49 increases in response to the decrease in the discharge pressure loss, and the cooling air amount AV flowing through the cooling passage 49 increases linearly. However, in the second region R2, the cooling air amount AV increases as the gap amount CB increases only up to the gap amount CB at point P2.

[0036] Next, in the third region R3 where the gap amount CB further increases from point P2, the gap amount is larger than in other regions, the discharge side pressure loss decreases, and the pipe pressure loss in the cooling passage 49 increases. Therefore, in the third region R3, even if the gap amount CB increases, the cooling air amount AV does not increase, and the cooling air discharged from the cooling passage 49 maintains approximately the same cooling air amount AV. The pressure loss of the cooling air in the third region R3 is mostly the pipe pressure loss between the inlet opening 49b and the outlet opening 49a of the cooling passage 49, and exceeds the discharge side pressure loss of the gap 71. In other words, the third region R3 is a cooling passage rate-determining region where the cooling air amount AV is limited by the opening area of ​​the cooling passage 49. In the second region R2, as the gap amount CB of the gap 71 increases, the discharge side pressure loss of the gap 71 decreases, discharge The pressure difference between the inlet opening 49b and the outlet opening 49a of the cooling air passage increases by an amount commensurate with the reduction in the exhaust side pressure loss, and the amount of cooling air AV flowing through the cooling passage 49 can be increased. However, the increase in the amount of cooling air AV associated with the reduction in the exhaust side pressure loss of the gap 71 only occurs up to the gap amount CB at the position of point P2. In the third region R3, unlike the second region R2, the reduction in the exhaust side pressure loss of the gap 71 associated with the enlargement of the gap amount CB is small. In this region, an increase in the amount of cooling air AV associated with an increase in the gap amount CB cannot be expected unless the opening area of ​​the cooling passage 49 is increased by increasing the hole diameter or number of holes in the cooling passage 49, and the pressure loss inside the cooling passage 49 is reduced.

[0037] In the first region R1 below the gap amount CB at point P1, the amount of cooling air AV discharged from the cooling passage 49 is limited only by the predetermined gap amount CB, regardless of the opening area of ​​the cooling passage 49, and only a small amount of cooling air is discharged to the combustion gas flow path 37. On the other hand, in the third region R3, the amount of cooling air AV discharged from the cooling passage 49 is limited by the specifications of the cooling passage 49 (hole diameter, number of holes, opening area, etc.). In the third region R3, regardless of the size of the gap amount CB of the gap portion 71, only a certain amount of cooling air AV is discharged unless the opening area of ​​the cooling passage 49 is changed. From another perspective, in the third region R3, by selecting a cooling passage 49 with appropriate specifications, the discharge of excessive cooling air is suppressed and the amount of cooling air is reduced. The second region R2 between the first region R1 and the third region R3 is an intermediate region transitioning from the first region R1 to the third region R3, and is a region in which the amount of cooling air AV discharged increases as the gap amount CB increases.

[0038] Ambient gas temperature 8 is a diagram showing the relationship between the amount of cooling air in the gap 71 and the ambient gas temperature in relation to the position of the outlet opening 49a of the cooling air discharged from the cooling passage 49. Case 1 is a case where the outlet opening 49a of the cooling passage 49 is located on the outer circumferential surface 48a of the protruding portion 48 and is disposed at a position close to the gap inlet 71a in the gap 71. Case 2 is a case where the outlet opening 49a of the cooling passage 49 is disposed at a position facing the combustion gas flow passage 37 axially downstream of the position of the gap inlet 71a, outside the outlet opening 60ea of ​​the cooling passage 60e of the seal member 60 in the blade height direction, and close to the combustion gas flow passage 37.

[0039] As shown in Fig. 4, the combustion gas G entrained in the gap 71 from the gap inlet 71a is a part of the combustion gas G flowing through the combustion gas flow passage 37, and is the combustion gas G that remains in a region between a downstream end 60aa of the first main body portion 60a of the seal member 60 (inner seal member 61) and the gas path surface 44a on the leading edge 43a side of the turbine stator blade 40. This region is a region on the axial downstream side of the axial downstream end 60aa of the first main body portion 60a where the outlet opening 60ea of ​​the cooling flow passage 60e arranged in the seal member 60 is arranged. Therefore, the ambient gas temperature in this region is diluted by the cooling air discharged from the outlet opening 60ea of ​​the seal member 60, and is lower than the temperature of the combustion gas G flowing through the combustion gas flow passage 37.

[0040] In case 1, the ambient gas temperature in the gap 71 decreases as the amount of cooling air AV discharged from the cooling passage 49 increases. As described above, the combustion gas G entrained in the gap 71 from the gap inlet 71a is diluted by the cooling air discharged from the cooling passage 60e, and its temperature is lower than that of the combustion gas G flowing through the combustion gas passage 37. Furthermore, after the temperature is lowered, the combustion gas G entrained in the narrow space of the gap 71 is further diluted by the cooling air from the cooling passage 49 discharged into the gap 71, and the ambient gas temperature in the gap 71 from the axial upstream end 48b of the protrusion 48 to the gap inlet 71a is further lowered.

[0041] As a result, thermal damage to the outer peripheral surface 48a of the protrusion 48 and the inner peripheral surface (inner peripheral surface 60ba of the first main body portion 60a) of the seal member 60 facing the outer peripheral surface 48a in the blade height direction is suppressed. In other words, the protrusion 48 is cooled by the convection cooling effect of the cooling air flowing through the cooling passage 49 and the dilution effect of the ambient gas in the gap 71 by the cooling air discharged from the cooling passage 49. eveningThe suppression effect of heat input from the ambient gas to the protruding portion 48 suppresses thermal damage to the protruding portion 48. Similarly, with regard to an inner circumferential surface 60ba of the first main body portion 60a of the seal member 60 facing the outer circumferential surface 48a of the protruding portion 48 in the blade height direction, thermal damage to the seal member 60 is avoided due to the effect of impingement cooling of the cooling air discharged from the outlet opening 49a and the suppression effect of heat input from the ambient gas to the inner circumferential surface 60ba of the first main body portion 60a due to the dilution effect of the ambient gas in the gap portion 71 by the cooling air.

[0042] In particular, when the outlet opening 49a of the cooling passage 49 is disposed at a position close to the gap inlet 71a, which is the axial downstream end of the seal member 60 in the gap 71, rather than at a position close to the upstream end 48b on the axial upstream side of the protrusion 48, the drop in the ambient gas temperature relative to the cooling air volume AV is remarkable. That is, when the outlet opening 49a of the cooling passage 49 is disposed in the gap 71 close to the gap inlet 71a, the combustion gas G can be diluted early before the high-temperature combustion gas G entrained in the gap 71 from the combustion gas flow path 37 diffuses into the gap 71. Therefore, it is possible to reduce the ambient gas temperature in the gap 71 in a short time with a small amount of cooling air, and the dilution effect by the cooling air is large, and excessive consumption of the cooling air volume is suppressed, thereby reducing the cooling air volume. In addition, since the ambient gas temperature in the gap 71 is reduced early, thermal damage to the protrusion 48 and the seal member 60 can be avoided.

[0043] On the other hand, in the case of Case 2, an example of which is shown in Patent Document 1, the position of the outlet opening 49a of the cooling passage 49 is provided on the outer side in the blade height direction than the position of the outlet opening 60ea of ​​the cooling flow passage 60e of the seal member 60, on the side approaching the combustion gas flow passage 37. In the case of Case 2, the combustion gas G that remains in the region sandwiched between the downstream end 60aa of the first main body portion 60a of the seal member 60 near the gap inlet 71a and the gas path surface 44a on the leading edge 43a side of the turbine stator blade 40 is diluted by the cooling air discharged from the outlet opening 49a of the cooling passage 49 before reaching this region, and is further diluted by the cooling air discharged from the outlet opening 60ea of ​​the cooling flow passage 60e of the seal member 60, so that the combustion gas G is diluted by the cooling air in two stages. However, in the case of Case 2, the area downstream of the gap inlet 71a where the combustion gas G remains is a large space compared to the space in the gap 71, and even if the temperature is once lowered by dilution with the cooling air, the effect of dilution with the cooling air is limited due to the combustion gas G surrounding this area in the circumferential direction. The combustion gas G in this state is entrained into the gap 71 from the gap inlet 71a. Therefore, while the amount of cooling air AV discharged from the cooling passage 49 is small, the ambient gas temperature in the gap 71 hardly drops, and the dilution effect by the cooling air is not observed. As shown in FIG. 8, a drop in the ambient gas temperature in the gap 71 is observed when a predetermined amount or more of cooling air is supplied.

[0044] However, the dilution effect of the cooling air is smaller than that of case 1, but when the position of the outlet opening 49a of the cooling passage 49 is disposed on the outer peripheral surface 48a of the protruding portion 48 on the axial downstream side near the gap inlet 71a with respect to the structure of case 1, the dilution effect of the cooling air is higher than that of case 2. That is, in this case, the outlet opening 49a of the cooling passage 49 is disposed inside the position of the outlet opening 60ea of ​​the cooling flow passage 60e of the seal member 60 in the blade height direction. In this case, the combustion gas G diluted by the cooling air discharged from the cooling passage 49 is the combustion gas G diluted by the cooling air discharged from the cooling flow passage 60e of the seal member 60, and is diluted at a location close to the gap inlet 71a. In this region, there is less wraparound of the surrounding combustion gas G, and the temperature of the combustion gas G drawn in from the gap inlet 71a is kept lower than that of case 2. Therefore, the cooling air dilution effect is greater than in case 2, and the ambient gas temperature in the gap 71 is reduced with a smaller amount of cooling air than in case 2, suppressing thermal damage to the protrusion 48 and the seal member 60.

[0045] As described above, in the case of the gas turbine 1 after a long period of operation, the seal member 60 and the turbine vane 40 Fitting In the structure 70, the fixed seal 64 that seals the gap between the space 21 of the casing 20 and the combustion gas flow path 37 may wear out on the outer circumferential surface 64a due to mechanical vibrations caused by long-term operation, and the gap amount CB of the gap 71 may increase, which may increase the amount of combustion gas G that is entrained in the gap 71. Fitting According to the structure 70, as shown in FIG. 7, the gap amount CB of the gap portion 71 becomes smaller than P due to wear of the stationary seal 64. 1When the gap 71 is expanded beyond the gap amount CB at the point, the amount of cooling air AV discharged from the cooling passage 49 increases rapidly in a linear manner along the line shown in the second region R2 as the gap amount CB increases. However, even if the gap amount CB of the gap 71 is expanded and the combustion gas G flowing into the gap 71 from the gap inlet 71a increases, the amount of cooling air AV increases as the gap amount CB increases, so that the entrained combustion gas G is immediately diluted and the ambient gas temperature decreases. As a result, thermal damage to the seal member 60 and the turbine vane 40 forming the gap 71 is suppressed. In particular, if the position of the outlet opening 49a of the cooling passage 49 formed in the gap 71 is close to the gap inlet 71a, the dilution effect on the high-temperature combustion gas entrained in the gap 71 is achieved with a smaller amount of cooling air. However, as described above, the amount of cooling air AV supplied from the cooling passage 49 increases with an increase in the gap amount CB only up to the position of point P2, and the amount of cooling air AV in the third region R3 thereafter is restricted by the cooling passage 49 and is limited to a constant amount (cooling passage rate-limited). Therefore, even if the combustion gas G is entrained in the gap 71 due to wear of the fixed seal 64, the amount of cooling air required for dilution is immediately supplied, and the discharge of an excessive amount of cooling air exceeding the required amount is suppressed.

[0046] In this embodiment Fitting According to the structure 70, the specifications of the cooling passage 49 (hole diameter, number of holes, opening area) and By appropriately selecting the gap amount CB of the gap portion 71, Fitting Even if the combustion gas G is caught in the gap 71 of the structure 70, the protrusion 48 of the turbine stator vane 40 and the seal member 60 are properly cooled, and thermal damage is avoided. In addition, the gap amount CB of the gap 71 is kept small, and an appropriate combination of the gap amount CB and the cooling passage 49 is selected, so that excessive loss of cooling air is suppressed, the amount of cooling air is reduced, the efficiency of the gas turbine is improved, and the reliability of the gas turbine 1 is increased. improvement do.

[0047] Second Embodiment 9 shows a structure of a second embodiment in which a turbine stator vane 40 is combined with a seal member 60. The combined structure of the turbine stator vane 40 and the seal member 60 shown in this embodiment is the same as that of the first embodiment, except that the position of the inlet opening 49b of the cooling passage 49 is different. 9, the present embodiment is the same as the first embodiment in that an outlet opening 49a of a cooling passage 49 formed in a protruding portion 48 is formed on an outer circumferential surface 48a of the protruding portion 48. However, the present embodiment differs from the first embodiment in that a position of an inlet opening 49b is formed not on an inner circumferential surface 48c of the protruding portion 48 but on an inner surface 47i of an outer wall portion 47d, which is the inner surface of a main member 47 of the shroud 44.

[0048] 9, the cooling passage 49 in this embodiment penetrates the protruding portion 48 of the turbine stator blade 40 and the outer wall portion 47d of the shroud 44 in the axial direction, and the outlet opening 49a of the cooling passage 49 is connected to the gap portion 71. The inlet opening 49b of the cooling passage 49 opens into an outer cavity 51b arranged on the inside in the blade height direction away from the gas path surface 44a with respect to the impingement plate 50 arranged in the shroud 44. The circumferential arrangement of the cooling passages 49 is similar to that of the first embodiment, and may be arranged within the range of the entire circumferential width L1 shown in FIG. 5, or may be arranged within the range of a partial circumferential width L2 as shown in the modified example of FIG. 6.

[0049] According to the cooling passage 49 of the present embodiment, by arranging the cooling passage 49 within the range of the entire width L1 in the circumferential direction, the protruding portion 48 is properly cooled even if the combustion gas G is caught in the gap portion 71. Moreover, similarly to the modified example of the cooling passage 49 of the first embodiment shown in FIG. 6, even if the cooling passage 49 is arranged within a partial range of width L2 in the circumferential direction, the protruding portion 48 is properly cooled and the amount of cooling air is reduced. In the case of the shroud 44 of the present embodiment, the cooling air is supplied directly from the space 21 in the casing 20 to the outer cavity 51b of the shroud 44 and then supplied to the cooling passage 49.

[0050] The relationship between the gap amount CB of the gap portion 71 and the cooling air amount AV in the first embodiment shown in FIG. 7 can also be applied to this embodiment. That is, in FIG. 7, in the first region R1 where the gap amount CB is smaller than the point P1, the discharged cooling air amount AV is limited only by the gap amount CB regardless of the size of the hole diameter of the cooling passage 49, and only a small amount of cooling air is discharged, which is a gap rate-limited region. On the other hand, in the third region R3 where the gap amount CB is larger than the point P2, the discharged cooling air amount AV is limited by the specifications of the opening area of ​​the cooling passage 49, etc., and is a cooling passage rate-limited region where a constant amount of cooling air is discharged regardless of the size of the gap amount CB. The second region R2 between the first region R1 and the third region R3 is an intermediate region that transitions from the first region R1 to the third region R3, and is a region where the cooling air amount increases as the gap amount CB increases. In this embodiment, too, the combination structure of the turbine stator vane 40 and the seal member 60 is used. Fitting By forming an outlet opening 49a of the cooling passage 49 in the gap portion 71 of the structure 70, the protrusion 48 of the turbine vane 40 and the seal member 60 are properly cooled, excessive loss of cooling air is suppressed, and the amount of cooling air is reduced.

[0051] In addition, the amount of cooling air and the gap for different positions of the outlet opening 49a of the cooling passage 49 shown in FIG. Department The relationship of the ambient gas temperature 71 can also be applied to this embodiment.

[0052] Third Embodiment 10 shows a structure of a third embodiment in which a turbine stator vane 40 is combined with a seal member 60. The combined structure of the turbine stator vane 40 and the seal member 60 shown in this embodiment is the same as the first and second embodiments, except that the position of the inlet opening 49b of the cooling passage 49 is different. The inlet opening 49b of the cooling passage 49 in this embodiment has the same structure as that in the second embodiment in that it is formed in the outer wall portion 47d of the shroud 44. However, the position of the inlet opening 49b of the cooling passage 49 in this embodiment is different from that in the second embodiment in that it communicates with the inner cavity 51a arranged on the side closer to the gas path surface 44a with respect to the collision plate 50 arranged in the shroud 44, in that it communicates with the outer cavity 51b.

[0053] In the case of this embodiment, the position of the inlet opening 49b of the cooling passage 49 is different from that of the second embodiment, and as a result, the present embodiment has different actions and effects from those of the second embodiment. That is, in the case of the shroud 44 of this embodiment, the cooling air is supplied directly from the space 21 in the casing 20 to the outer cavity 51b of the shroud 44. The cooling air supplied to the outer cavity 51b is discharged into the inner cavity 51a through the multiple through holes 50a of the collision plate 50. The cooling air supplied to the inner cavity 51a becomes a jet and collides with the inner circumferential surface of the bottom portion 47c forming the gas path surface 44a, thereby impinging and cooling the inner circumferential surface. The cooling air after impinging and cooling the inner circumferential surface of the bottom portion 47c is supplied to the inlet opening 49b formed in the outer wall portion 47d on the leading edge 43a side.

[0054] Therefore, in the case of the cooling passage 49 in this embodiment, the cooling air supplied from outside the shroud 44 impinges and cools the bottom 47c of the shroud 44, and then convectively cools the protruding portion 48 while flowing through the cooling passage 49. This allows the cooling air to be reused, and the cooling efficiency of the gas turbine 1 is increased, compared to the second embodiment. On the other hand, the cooling air in the outer cavity 51b is reduced in pressure due to pressure loss while flowing into the inner cavity 51a through the through hole 50a of the collision plate 50 and impinging on the bottom 47c. Therefore, the pressure in the inner cavity 51a is lower than the pressure in the outer cavity 51b. That is, the cooling air supplied to the cooling passage 49 is supplied at a higher pressure in the structure of the second embodiment than in the structure of this embodiment, and the allowable width of pressure loss is increased, so that the cooling effect on the protruding portion 48 is greater in the second embodiment.

[0055] In this embodiment, the same arrangement as that of the circumferential cooling passages 49 in the first embodiment shown in Fig. 5 is applied. By arranging the cooling passages 49 within the range of the entire circumferential width L1, the protruding portion 48 is properly cooled even if the combustion gas G is caught in the gap portion 71. Moreover, as in the modified example of the cooling passages 49 in the first embodiment shown in Fig. 6, even if the cooling passages 49 are arranged within a partial range of width L2 in the circumferential direction, the protruding portion 48 is properly cooled and the amount of cooling air is reduced.

[0056] The relationship between the gap amount CB of the gap portion 71 and the cooling air amount AV in the first embodiment shown in FIG. 7 can also be applied to this embodiment. In FIG. 7, the first region R1 where the gap amount CB is smaller than the point P1 is a gap rate-limited region in which the amount of cooling air discharged AV is limited only by the gap amount CB regardless of the size of the hole diameter of the cooling passage 49, and only a small amount of cooling air is discharged. The third region R3 where the gap amount CB is larger than the point P2 is a cooling passage rate-limited region in which the amount of cooling air discharged AV is limited by the specifications of the opening area of ​​the cooling passage 49, and a constant amount of cooling air is discharged regardless of the size of the gap amount CB. The second region R2 between the first region R1 and the third region R3 is an intermediate region that transitions from the first region R1 to the third region R3, and is a region in which the amount of cooling air discharged AV increases as the gap amount CB increases.

[0057] The dimensions, materials, shapes, and relative arrangements of the components described in the above embodiments or shown in the drawings are merely illustrative examples and are not intended to limit the scope of the invention. For example, expressions expressing relative or absolute arrangements, such as "in a certain direction," "along a certain direction," "parallel," "orthogonal," "center," "concentric," or "coaxial," not only strictly express such arrangements, but also express a state in which there is a relative displacement with a tolerance or an angle or distance to the extent that the same function is obtained. For example, expressions expressing that things are in an equal state, such as "same," "equal," and "homogeneous," not only strictly express a state in which there is an equality, but also express a state in which there is a tolerance or a difference to the extent that the same function is obtained. For example, an expression expressing a shape such as a square or a cylinder not only expresses a square or a cylinder in the strict geometric sense, but also expresses a shape including uneven portions, chamfered portions, etc., to the extent that the same effect is obtained. On the other hand, an expression such as "comprises", "includes", "has", "includes", or "has" a component is not an exclusive expression that excludes the presence of other components.

[0058] The contents described in each of the above embodiments can be understood as follows. (1) A turbine vane according to a first aspect is a turbine vane comprising a blade body and a shroud formed on both ends of the blade body in the blade height direction, the shroud including a main member having a heat-resistant coating extending from an axially upstream end to a downstream end of a gas path surface, and a protrusion protruding axially upstream from a side end portion on a leading edge side of the main member, the protrusion being arranged in a ring shape at predetermined intervals in the circumferential direction and including a plurality of cooling passages connecting an inlet opening formed on an inner surface of the protrusion or an inner surface of the main member to an outlet opening formed on an outer peripheral surface of the protrusion.

[0059] According to the turbine vane described in (1) above, when combustion gas is caught in the gap at the fitting portion between the turbine vane and the seal member, the cooling passage arranged in the protrusion enhances cooling of the protrusion, suppresses thermal damage to the protrusion, and improves the reliability of the gas turbine.

[0060] (2) A turbine vane according to a second aspect is the turbine vane according to (1), wherein the outlet opening of the cooling passage is formed axially upstream of a position of an upstream end of the heat-resistant coating portion formed on the main member in the axial direction. .

[0061] According to the turbine vane described in (2) above, cooling of the protruding portion, which is exposed to high-temperature combustion gas and on which it is difficult to form a heat-resistant coating, is strengthened.

[0062] (3) A turbine vane according to a third aspect is a turbine vane according to (2), wherein the outlet opening of the cooling passage is positioned axially downstream of the upstream end of the protrusion in the axial direction and close to the upstream end of the coating portion.

[0063] According to the turbine vane described in (3) above, the outlet opening of the cooling passage is positioned close to the upstream end of the coating portion, so that it is possible to dilute the combustion gas entrained in the gap with a smaller amount of cooling air, thereby reducing the amount of cooling air.

[0064] (4) A turbine vane according to a fourth aspect comprises any one of (1) to (3). Either one The inlet opening formed on the inner circumferential surface of the protruding portion of the cooling passage is formed close to the side end portion on the leading edge side of the main member.

[0065] According to the turbine vane described in (4) above, the inlet opening of the cooling passage can be located close to the side end portion on the leading edge side of the turbine vane, so that the inlet opening can be prevented from being blocked.

[0066] (5) A turbine vane according to a fifth aspect comprises any one of (1) to (3). Either one A turbine vane, wherein the shroud includes a cavity formed by a bottom portion forming a bottom surface of the main member and an outer wall portion formed at an outer edge of the shroud and protruding in a blade height direction, and the inlet opening formed in the inner surface of the main member communicates with the cavity.

[0067] According to the turbine vane described in (5) above, by providing the inlet opening of the cooling passage on the outer wall portion of the shroud, clogging of the inlet opening can be avoided.

[0068] (6) A turbine vane according to a sixth aspect is the turbine vane according to (5), wherein the shroud has a collision wall that divides the cavity into an inner cavity formed on the gas path surface side in the blade height direction and an outer cavity disposed adjacent to the inner cavity in the blade height direction in the opposite direction to the gas path surface. board, the inlet opening of the cooling passage is formed in an inner surface of the outer wall portion on a leading edge side of the inner cavity.

[0069] According to the turbine vane described in (6) above, the inlet opening of the cooling passage is formed on the inner surface of the inner cavity of the shroud, and the cooling air supplied to the inner cavity impinges and cools the bottom of the shroud, and is then supplied to the cooling passage and convection-cools the protruding portion. This allows the cooling air to be reused, improving cooling efficiency.

[0070] (7) A turbine vane according to a seventh aspect comprises any one of (1) to (6). Either one A turbine stator vane, wherein the multiple cooling passages are arranged over a range of a partial width rather than the entire circumferential width from the suction surface end to the pressure surface end of the shroud, and the circumferential center position of the range in which the multiple cooling passages are arranged is closer to the pressure surface end than the leading edge of the blade body.

[0071] According to the turbine vane described in (7) above, the cooling passage formed in the protrusion is arranged over a partial range rather than the entire width of the protrusion, and cooling of the pressure surface end side is enhanced, so that the protrusion is properly cooled and the amount of cooling air is reduced.

[0072] (8) A turbine vane according to an eighth aspect comprises any one of (1) to (7). Either one A turbine vane, wherein the cooling passage approaches the exit opening from the inlet opening and is inclined axially upstream.

[0073] According to the turbine vane described in (8) above, the cooling passage is a passage that approaches the outlet opening and is inclined axially upstream, so that clogging of the inlet opening is avoided.

[0074] (9) A turbine vane according to a ninth aspect comprises any one of (1) to (8). Either one In a turbine vane, the cooling passage forms an inclined flow path approaching the outlet opening from the inlet opening and approaching a suction surface side end of the shroud in the axial direction.

[0075] According to the turbine vane described in (9) above, by forming the cooling passage as an inclined passage, the cooling area is enlarged and the amount of cooling air is reduced.

[0076] (10) A turbine vane according to a tenth aspect comprises any one of (1) to (9). Either one A turbine stator vane, the shroud comprising an outer shroud formed at an outer end of the vane body in the vane height direction, and an inner shroud formed at an inner end of the vane body in the vane height direction.

[0077] (11) The eleventh aspect Fitting The structure includes a combustor, a turbine vane including a vane body and a shroud, and a seal member disposed between the turbine vane and the combustor and facing a combustion gas flow path. Fitting a plurality of cooling passages each of which is arranged at a predetermined interval in the circumferential direction and connects an inlet opening formed on an inner surface of the protrusion or an inner surface of the main member to an outlet opening formed on an outer peripheral surface of the protrusion, the outlet opening being upstream of a position of the axial downstream end of the sealing member and opening into a gap portion extending in the circumferential direction, the gap portion being formed between the inner surface of the sealing member and the outer peripheral surface of the protrusion.

[0078] The above (11) Fitting According to the structure, the turbine is composed of a combustor, a turbine vane, and a seal member. Fitting A cooling passage is formed in the protruding portion of the structure, and an outlet opening of the cooling passage is Fitting Since the cooling passage is open to the gap of the structure, the gap size is increased, and even if the combustion gas is caught in the gap, it is immediately diluted by the small amount of cooling air discharged from the cooling passage, and the ambient gas temperature in the gap is lowered. Fitting Thermal damage to the structure is suppressed.

[0079] (12) The twelfth aspect Fitting The structure is (11) Fitting The structure includes a first body portion in contact with the combustion gas flow path and extending in the axial direction, a second body portion connected to the first body portion at an axially upstream end and extending in the blade height direction, and a third body portion extending in the axially downstream direction from a middle position in the blade height direction of the second body portion, and a plate-shaped fixed seal is arranged on an outer peripheral surface of the third body portion and extends in the circumferential direction along the outer peripheral surface, and a seal surface is formed between the outer peripheral surface of the fixed seal and the inner peripheral surface of the protrusion.

[0080] The above (12) FittingAccording to this structure, a plate-shaped fixed seal is arranged on the outer peripheral surface of the third main body portion, and a sealing surface is formed between the outer peripheral surface of the fixed seal and the inner peripheral surface of the protrusion, thereby preventing cooling air from leaking into the combustion gas flow path through the gap between the seal member and the protrusion.

[0081] (13) The fitting structure according to the thirteenth aspect is (11) or (12). Fitting In the structure, the shroud comprises an outer shroud formed at an outer end of the blade body in the blade height direction, and an inner shroud formed at an inner end of the blade body in the blade height direction.

[0082] (14) A gas turbine according to a fourteenth aspect comprises any one of (1) to (10). Either one Equipped with turbine vanes.

[0083] (15) A gas turbine according to a fifteenth aspect comprises any one of (11) to (13). Either one of them fits It has a structure.

[0084] (16) The sixteenth aspect Fitting The method for adjusting the amount of cooling air in a structure includes a turbine vane including a vane body and a shroud, a combustor, and a seal member facing a combustion gas flow path and disposed between the turbine vane and the combustor in the axial direction. Fitting a first main body portion extending in the axial direction and having an outer circumferential surface in contact with a combustion gas flow path; a plurality of the protruding portions are arranged at predetermined intervals in the circumferential direction, and the protruding portions include a plurality of cooling passages connecting an inlet opening formed on an inner circumferential surface of the protruding portion and an outlet opening formed on the outer circumferential surface of the protruding portion; and a gap portion formed in the blade height direction between the outer circumferential surface of the protruding portion and the inner circumferential surface of the sealing member. Fitting A method for adjusting the amount of cooling air in a structure, comprising: Gap amountIn a first region where the gap is small, the amount of cooling air discharged is limited by the gap amount of the gap and is kept small. In a second region where the gap amount is larger than that of the first region, the amount of cooling air discharged increases as the gap amount increases. In a third region where the gap amount is larger than that of the second region, the amount of cooling air discharged increases. aisle The opening area is limited to reduce the amount of cooling air discharged.

[0085] The above (16) Fitting According to a method for adjusting the amount of cooling air in a structure, a turbine stator blade, a combustor, and a seal member disposed between the turbine stator blade and the combustor are provided. Fitting A method for adjusting cooling air in a structure, in which while the gap amount of the gap is small, the amount of cooling air is limited to the gap amount of the gap, and the amount of cooling air discharged into the gap is kept small. Fitting The structure is properly cooled and thermal damage is suppressed, improving the reliability of the gas turbine. In addition, the amount of cooling air required is reduced, improving the cooling efficiency of the gas turbine. [Explanation of symbols]

[0086] 1. Gas turbine 2. Compressor 4. Combustor 6. Turbine 8 Rotor 10 Compressor compartment 12 Intake chamber 14 Inlet guide vane 16 Compressor vane 18 Compressor blade 20 Casing 21 Space 22 Turbine casing 26 Turbine blade 28 Exhaust compartment 29 Exhaust chamber 30 fuel nozzle 32 Combustion tube 33 Inner cylinder 34 Tailpiece 35 Flange 37 Combustion gas flow path 40 Turbine vane 42 Wing body 43 Wing surface 43a leading edge 43b Trailing edge 44 Shroud (inner shroud 45, outer shroud 46) 44a Gas Pass Surface 47 Main member 47a Side end 47b Outer edge 47c bottom 47d Exterior wall 47e Exterior wall end 47f downstream end 47g Pressure side end 47h Negative pressure side end 47i Inner surface 48 Protrusion 48a Outer surface 48b Upstream end 48c Inner surface 49 Cooling passage 49a Exit opening 49b Entrance opening 50 Collision plate 50a through hole 51 Cavity (inner cavity 51a, outer cavity 51b) 52 Heat-resistant coating part 52a Upstream end of coating part 60 sealing member (inner sealing member 61, outer sealing member 62) 60a First body part 60aa downstream end 60ba inner surface 60b 2nd body part 60c 3rd body part 60ca outer surface 60d Connection 60e Cooling Channel 60ea exit opening 61a Radial outer surface 62a Radial inner surface 63 Recess 64 Fixed seal 64a Outer surface 70 Fitting structure 71 Gap 71a Gap entrance R1 1st area R2 2nd area R3 3rd area A. Compressed air CB Gap AV Cooling Air Volume

Claims

1. The wing body and A shroud is formed at each end of the blade body in the blade height direction; A turbine vane for a gas turbine comprising: The shroud includes: a main member having a heat resistant coating portion extending from an upstream end to a downstream end of a gas path surface in an axial direction; a protruding portion protruding axially upstream from a side end portion on a leading edge side of the main member; Including, a seal member is disposed between a combustor of the gas turbine and the turbine vane, The protrusion is The seal member is configured to be fitted into a recess formed in the seal member and recessed toward the upstream side in the axial direction. A plurality of the bearings are arranged in an annular shape at predetermined intervals in the circumferential direction, a plurality of cooling passages connecting an inlet opening formed on an inner circumferential surface of the protrusion or an inner surface of the main member and an outlet opening formed on an outer circumferential surface of the protrusion, The outlet opening is located upstream of a downstream end of the seal member in the axial direction, and opens into a gap portion that is formed between an inner peripheral surface of the seal member and an outer peripheral surface of the protruding portion and extends in a circumferential direction. Turbine vanes.

2. The outlet opening of the cooling passage is The heat-resistant coating is formed on the main member on the axially upstream side of the upstream end of the heat-resistant coating. The turbine vane of claim 1 .

3. The outlet opening of the cooling passage is disposed axially downstream of the upstream end of the protrusion and close to the upstream end of the coating portion. The turbine vane of claim 2 .

4. The inlet opening formed on the inner circumferential surface of the protruding portion of the cooling passage is formed close to the side end portion on the leading edge side of the main member. A turbine vane according to any one of claims 1 to 3.

5. The shroud includes: A cavity is formed by a bottom portion forming a bottom surface of the main member and an outer wall portion formed on an outer edge of the shroud and protruding in a blade height direction, The inlet opening formed in the inner surface of the main member communicates with the cavity. A turbine vane according to any one of claims 1 to 3.

6. The shroud includes: The cavity is divided into an inner cavity formed on the gas path surface side in the blade height direction and an outer cavity arranged adjacent to the inner cavity in the opposite direction to the gas path surface in the blade height direction, The inlet opening of the cooling passage is formed in an inner surface of the outer wall portion on a leading edge side of the inner cavity. The turbine vane of claim 5 .

7. the plurality of cooling passages are arranged in a range of a partial width, not an entire width, in a circumferential direction from a suction surface side end to a pressure surface side end of the shroud, The circumferential center position of the range in which the plurality of cooling passages are arranged is closer to the pressure surface end portion side than the leading edge position of the blade body. The turbine vane of claim 1 .

8. The cooling passage approaches the outlet opening from the inlet opening and is inclined axially upstream. The turbine vane of claim 1 .

9. The cooling passage forms an inclined flow passage approaching the outlet opening from the inlet opening and approaching the suction surface side end of the shroud. The turbine vane of claim 1 .

10. The shroud includes: an outer shroud formed at an outer end of the blade body in the blade height direction; an inner shroud formed at an inner end of the blade body in the blade height direction; 2. The turbine vane of claim 1.

11. A combustor; A turbine vane including a vane body and a shroud; a seal member that faces a combustion gas flow path and is disposed between the turbine vane and the combustor; A fitting structure comprising: The shroud includes: a main member extending axially from an upstream end to a downstream end; a protruding portion protruding axially upstream from a side end portion on a leading edge side of the main member, The protruding portions are arranged at predetermined intervals in the circumferential direction, an inlet opening formed on an inner circumferential surface of the protrusion or an inner surface of the main member; and an outlet opening formed on an outer circumferential surface of the protrusion; a plurality of cooling passages connecting the the seal member has a recess formed therein that is recessed toward the upstream side in the axial direction, and the protruding portion of the turbine stator blade is configured to be able to fit into the recess, The outlet opening is located upstream of the downstream end of the seal member in the axial direction, and opens into a gap portion that is formed between an inner peripheral surface of the seal member and an outer peripheral surface of the protruding portion and extends in a circumferential direction. Interlocking structure.

12. A combustor; A turbine vane including a vane body and a shroud; a seal member that faces a combustion gas flow path and is disposed between the turbine vane and the combustor; A fitting structure comprising: The shroud includes: a main member extending axially from an upstream end to a downstream end; a protruding portion protruding axially upstream from a side end portion on a leading edge side of the main member, The protruding portions are arranged at predetermined intervals in the circumferential direction, an inlet opening formed on an inner circumferential surface of the protrusion or an inner surface of the main member; and an outlet opening formed on an outer circumferential surface of the protrusion; a plurality of cooling passages connecting the the outlet opening is located upstream of a downstream end of the seal member in the axial direction, and opens into a gap portion that is formed between an inner peripheral surface of the seal member and an outer peripheral surface of the protruding portion and extends in a circumferential direction; The sealing member is a first body portion in contact with the combustion gas flow path and extending in an axial direction; A second body portion connected to the first body portion at an upstream end in the axial direction and extending in a blade height direction; a third body portion extending in an axial downstream direction from a middle position in a blade height direction of the second body portion, A plate-shaped stationary seal is disposed on an outer circumferential surface of the third main body portion and extends circumferentially along the outer circumferential surface, A fitting structure in which a seal surface is formed between an outer peripheral surface of the stationary seal and an inner peripheral surface of the protrusion.

13. The shroud includes: An outer shroud formed at an outer end of the blade body in the blade height direction; an inner shroud formed at an inner end of the blade body in the blade height direction; 13. A fitting structure according to claim 11 or 12.

14. A gas turbine comprising the turbine vane according to claim 1.

15. A gas turbine comprising the fitting structure according to claim 11.

16. A fitting structure including a turbine vane including a vane body and a shroud, a combustor, and a seal member facing a combustion gas flow path and disposed between the turbine vane and the combustor in an axial direction, The shroud includes a protruding portion protruding axially upstream from a side end portion on a leading edge side, The seal member includes a first body portion extending in an axial direction and having an outer circumferential surface in contact with the combustion gas flow path, The protrusion is A plurality of the rollers are arranged at predetermined intervals in the circumferential direction, a plurality of cooling passages connecting an inlet opening formed on an inner circumferential surface of the protrusion and an outlet opening formed on an outer circumferential surface of the protrusion; A method for adjusting the amount of cooling air in the fitting structure having a gap formed in a blade height direction between an outer circumferential surface of the protrusion and an inner circumferential surface of the seal member, comprising: In the first region where the gap amount of the gap is small, the gap amount of the gap is limited, and the amount of cooling air discharged is suppressed to a small amount, In the second region where the gap amount is larger than that in the first region, the amount of cooling air discharged increases as the gap amount increases, In the third region where the gap is larger than that in the second region, the amount of cooling air discharged is restricted by the opening area of ​​the cooling passage, thereby suppressing the amount of cooling air discharged. A method for adjusting the amount of cooling air in a mating structure.

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