JP7232035B2 - Gas turbine stator blades and gas turbines - Google Patents

Description

The present disclosure relates to gas turbine vanes and gas turbines.

In a gas turbine having a plurality of combustors, acoustic propagation between the combustors may cause combustion oscillations in the vicinity of the exits of the combustors. Such combustion oscillation can be a factor that hinders stable operation of the gas turbine. Therefore, measures have been taken to reduce the combustion oscillation that occurs near the exit of the combustor.

For example, US Pat. A gas turbine is disclosed in which an annular gas flow path is divided circumferentially.

Japanese Patent No. 5726267

As described above, in the gas turbine described in Patent Document 1, the annular gas flow path formed between the combustor outlet and the stationary blade is formed with a gap that at least partially connects in the circumferential direction. . Therefore, the formation of gaps between adjacent combustors may limit the effect of suppressing acoustic propagation.

In view of the above circumstances, it is an object of at least one embodiment of the present invention to provide a gas turbine stator vane and a gas turbine capable of reducing combustion oscillations caused by acoustic propagation between outlets of multiple combustors. and

(1) A gas turbine stator vane according to at least one embodiment of the present invention,
A stator vane for a gas turbine,
Equipped with a hollow wing body,
the upstream end of the blade body in the axial direction of the gas turbine includes a convex portion projecting upstream in the axial direction or a concave portion recessed downstream in the axial direction;
The concave portion or the convex portion includes at least one purge hole communicating between the outer space and the inner space of the blade body.

According to the above configuration (1), the concave portion or the convex portion positioned upstream of the stator vane is replaced by a mating member on the combustor side (for example, a combustor or a seal member provided between the combustor and the stator vane). ), the pair of side wall surfaces of the concave or convex portion of the stationary blade and the mating member can overlap in the axial direction. As a result, the gap between the combustor and the stator vane is formed as a small concave-convex gap, and combustion oscillation caused by acoustic propagation between the exit portions of the plurality of combustors can be reduced.

Further, according to the above configuration (1), the cooling medium is supplied from the internal space of the blade through the purge hole to the gap between the recess or protrusion of the stationary blade and the mating member on the combustor side. be able to. In this way, the upstream end of the stationary blade is cooled by the flow of the cooling medium supplied to the gap between the concave or convex portion of the stationary blade and the mating member, and thermal damage to the upstream end is prevented. is prevented, and stable operation of the gas turbine becomes possible.

(2) In some embodiments, in the configuration of (1) above,
The convex portion or the concave portion extends from the tip end side to the base end side of the wing body along the wing height direction,
The purge holes include a plurality of purge holes arranged along the blade height direction from the tip end side to the base end side.

According to the above configuration (2), the cooling medium is discharged to the outer space of the blade through a plurality of purge holes arranged from the tip end to the base end of the blade, thereby allowing combustion to occur in the inner space of the blade. It is possible to suppress backflow of gas, reduce the temperature of the combustion gas outside the purge hole, and suppress thermal damage to the upstream end of the blade body.

(3) In some embodiments, in the configuration of (1) or (2) above,
The blade body includes at least one side purge hole adjacent to the purge hole in the circumferential direction of the gas turbine.

According to the above configuration (3), the cooling medium is discharged from the side purge hole to the outer space of the blade body, thereby suppressing the backflow of the combustion gas to the inner space of the blade body, thereby suppressing the combustion outside the side purge hole. By lowering the temperature of the gas, damage to the axially upstream end of the blade body can be suppressed.

(4) In some embodiments, in the configuration of any one of (1) to (3) above,
The blade body connects the protrusion or the recess and the trailing edge, and connects the suction surface wall portion on the suction surface side where the blade surface forms a convex surface and the protrusion or the recess and the trailing edge, a pressure surface wall portion on the pressure surface side in which the blade surface forms a concave surface;
The outer wall surface of the upstream end portion includes a pair of flat plate portions extending to the suction surface wall portion and the pressure surface wall portion with the convex portion or the concave portion sandwiched in the circumferential direction.

According to the configuration (4) above, it becomes easy to process the upstream end of the blade body.

(5) In some embodiments, in the configurations of (1) to (4) above,
further comprising a cylindrical first insert extending along the blade height direction in the internal space of the blade body;
the first insert includes an opening forming portion having an opening extending along the wing height direction;
The opening is positioned opposite the at least one purge hole.

According to the above configuration (5), the high-pressure cooling medium that has flowed out of the first insert from the internal space of the first insert through the opening of the opening forming portion is smoothly removed from the recesses or protrusions of the stationary blade through the purge holes. It can be fed directly into the gap between the part and the mating member on the combustor side. As a result, it is possible to suppress the backflow of the combustion gas through the aforementioned gap while cooling the stator blades.

(6) In some embodiments, in the configuration of (5) above,
The first insert includes a cylindrical body and an insert collar provided at one end of the cylindrical body in the shape of a flange,
The insert collar has a notch portion arranged corresponding to the position of the opening forming portion.

According to the configuration (6) above, since the insert collar has the cutout portion in the portion corresponding to the position of the opening of the opening forming portion, a part of the cooling medium supplied from the outside is transferred to the inner wall surface of the blade body and the first cutout portion. Since the cooling medium can be directly supplied to the gap between the outer peripheral surface of the 1 insert, the high-pressure cooling medium can be supplied to the purge hole while suppressing the pressure loss of the cooling medium as much as possible.

(7) In some embodiments, in the configuration of (5) or (6) above,
A first impingement space is formed between the inner wall surface of the wing body and the outer peripheral surface of the first insert,
A portion of the first insert excluding the opening forming portion includes an impingement wall portion formed with a plurality of impingement holes communicating the internal space of the first insert and the first impingement space.

According to the above configuration (7), the cooling medium is discharged from the internal space of the first insert into the first impingement space through the impingement hole formed in the impingement wall, thereby Impingement cooling can be provided.

(8) In some embodiments, in the configuration of (7) above,
A pair of first seal dams extending in the blade height direction are formed on the inner wall surface of the upstream end of the blade body so as to separate the first impingement space from the first insert. ,
The pair of first seal dams are arranged to sandwich the purge hole in the circumferential direction of the gas turbine.

According to the above configuration (8), the communication space, which is included in the impingement space and connects the opening of the opening forming portion and the purge hole so as to communicate with each other, is set with respect to the first impingement space by the pair of first seal dams. can be partitioned. Therefore, the pressure of the cooling medium supplied from the internal space of the blade body to the first insert through the purge hole is applied to the gap between the concave or convex portion of the stationary blade and the mating member on the combustor side. High pressure cooling medium can be effectively supplied without lowering. As a result, the first impingement space is edge-separated from the communicating space, so that effective impingement cooling can be achieved according to the state of the blade surface.

(9) In some embodiments, in the configuration of (8) above,
the upstream end includes a flat plate portion extending in a circumferential direction of the gas turbine;
The convex portion or the concave portion is formed in the flat plate portion,
The pair of first shield dams are formed on the inner wall surface of the flat plate portion.

According to the above configuration (9), by forming the pair of first seal dams on the inner wall surface of the flat plate portion, the leakage flow between the first seal dams and the first insert is effectively suppressed, and effective impingement is achieved. cooling can be achieved.

(10) In some embodiments, in the configuration of (8) or (9) above,
The first insert includes a pair of ribs arranged on both sides of the opening in the circumferential direction on the outer peripheral surface of the first insert,
The pair of ribs extend along the blade height direction and are fitted to the pair of first seal dams, respectively.

According to the configuration (10) above, effective impingement cooling of the blade body can be achieved in the first impingement space.

(11) In some embodiments, in the configuration of (10) above,
The opening forming portion of the first insert and the impingement wall portion are provided adjacent to each other with the rib interposed therebetween.

According to the above configuration (11), the cooling medium discharged to the external space of the blade through the purge hole and the cooling medium injected into the first impingement space through the impingement hole are used to effectively clean the blade. can be cooled effectively.

(12) In some embodiments, in the configuration of any one of (8) to (11) above,
further comprising a cylindrical second insert provided adjacent to the downstream side of the first insert in the axial direction of the gas turbine and extending along the blade height direction;
The blade body includes a suction surface wall portion on the suction surface side connecting the projection or the recess and the trailing edge and connecting the projection or the recess and the trailing edge to the blade. n (n is an integer equal to or greater than 1) extending so as to connect the suction surface wall portion and the pressure surface wall portion inside the blade body on the pressure surface side, the surface forming a concave surface; a partition wall of
The first insert and the second insert include a front partition wall, which is the partition wall farthest from the trailing edge among the n partition walls, and the rear edge of the outer wall of the wing with the front partition wall interposed therebetween. provided in the front space defined by the front outer wall portion, which is the portion opposite to the
A second impingement space is formed between the inner wall surface of the wing body and the outer peripheral surface of the second insert,
The stator blade includes a partition that partitions the first impingement space and the second impingement space,
The number of seal dams partitioning the first impingement space with the first insert is one pair, and the number of seal dams partitioning the second impingement space with the second insert is one pair or less.

According to the above configuration (12), the number of seal dams partitioning the first impingement space with the first insert is one pair, and the number of seal dams partitioning the second impingement space with the second insert is , less than or equal to one pair. Therefore, even when the pressure in the front space is divided into three or more pressures, the first insert and the second insert can be easily inserted into the front space, and an increase in the amount of leakage flow at the partition is suppressed. can do.

(13) In some embodiments, in the configuration of (12) above,
The partition section includes a partition wall that partitions the first insert and the second insert from the front outer wall to the front partition wall.

According to the configuration (13) above, it is possible to completely prevent leakage flow between the first impingement space and the second impingement space located on both sides of the partition wall.

(14) In some embodiments, in the configuration of (12) above,
The partition includes a pair of second seal dams extending in the blade height direction so as to partition the second impingement space from the second insert.

According to the configuration (14) above, it is possible to suppress leakage flow between the first impingement space and the second impingement space located on both sides of the second seal dam with a simple configuration.

(15) In some embodiments, in the configuration of (14) above,
The pair of second seal dams are formed on the inner wall surface of the front partition wall and the inner wall surface of the front outer wall.

According to the above configuration (15), it is possible to partition between the inner wall surface of the front partition wall and the second insert, and between the front outer wall portion and the second insert, with a simple configuration.

(16) In some embodiments, in the configuration of (14) or (15) above,
The second insert has a pair of ribs on the outer peripheral surface of the second insert,
The pair of ribs of the second insert extend along the blade height direction and are fitted to the pair of second seal dams, respectively.

According to the configuration (16) above, a leak flow between the first impingement space and the second impingement space located on both sides of the pair of second seal dams can be suppressed with a simple configuration.

(17) A gas turbine according to at least one embodiment of the present invention,
a plurality of circumferentially arranged combustors having outlets including radial walls along the radial direction;
a stator vane according to any one of (1) to (16) above, which is positioned downstream of the pair of radial wall portions facing each other among the outlet portions of the combustors adjacent in the circumferential direction;
Prepare.

According to the configuration (17) above, since the stator blade according to any one of the above (1) to (16) is provided, combustion vibration caused by acoustic propagation between the exit portions of the plurality of combustors can be effectively suppressed. can be reduced. As a result, the frequency of combustion oscillations requiring countermeasures is reduced, and stable operation of the gas turbine becomes possible.

SUMMARY OF THE INVENTION At least one embodiment of the present invention provides a gas turbine vane and gas turbine capable of reducing combustion oscillations at the exit of a plurality of combustors.

1 is a schematic configuration diagram of a gas turbine according to one embodiment; FIG. 1 is a diagram showing a schematic configuration of an outlet 52 of a combustor 4 and an inlet of a turbine 6 of a gas turbine 1 according to one embodiment; FIG. 1 is a diagram showing a schematic configuration of an outlet 52 of a combustor 4 and an inlet of a turbine 6 of a gas turbine 1 according to one embodiment; FIG. FIG. 4 is a schematic cross-sectional view (cross-sectional view taken along line VI-VI shown in FIG. 3) around a first-stage stator vane 23 according to one embodiment; FIG. 3 is a schematic cross-sectional view of another aspect around the first stage stator vane 23 according to the embodiment (cross-sectional view taken along line VI-VI shown in FIG. 3); FIG. 5 is an enlarged cross-sectional view showing an example of a specific configuration of the first-stage stator vane 23A shown in FIG. 4; FIG. 7 is a diagram showing a cross section taken along line AA in FIG. 6; 6 is an enlarged cross-sectional view showing another example of a specific configuration of the first-stage stator vane 23A shown in FIG. 5; FIG. FIG. 9 is a view showing a CC cross section in FIG. 8; It is a figure showing an example of a schematic structure of the 1st insert 114 concerning one embodiment. It is a figure showing an example of a schematic structure of the 1st insert 114 concerning other embodiments. FIG. 16 is a cross-sectional view showing a range in which an insert collar 124 is provided in the first insert 114 shown in FIG. 15; FIG. 3 is a schematic cross-sectional view showing a configuration example of a first shield dam 142; 5 is an enlarged cross-sectional view showing another example of a specific configuration of the first-stage stator vane 23A shown in FIG. 4; FIG. FIG. 11 is a view showing a BB cross section in FIG. 10; 6 is an enlarged cross-sectional view showing another example of a specific configuration of the first-stage stator vane 23A shown in FIG. 5; FIG. FIG. 13 is a diagram showing a DD cross section in FIG. 12; FIG. 17 is a perspective view of an insert obtained by combining the first insert and the second insert shown in FIG. 14 or FIG. 16; FIG. 10 is an installation view of a common shared insert collar for inserts that combine the first insert and the second insert; FIG. 4 is a schematic cross-sectional view showing a configuration example of a second shield dam 166; FIG. 11 is a schematic cross-sectional view showing another configuration example of the second shield dam 166; 5 is an enlarged cross-sectional view showing another example of a specific configuration of the first-stage stator vane 23A shown in FIG. 4; FIG. FIG. 22 is a view showing the EE cross section in FIG. 21; 4 is a schematic perspective view showing the configuration of an outer shroud 62 according to one embodiment; FIG. FIG. 11 is a schematic cross-sectional view around a first-stage stator vane 23 according to another embodiment; FIG. 11 is a schematic cross-sectional view around a first-stage stator vane 23 according to another embodiment;

Several embodiments of the present invention will now 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 invention, and are merely illustrative examples. do not have.

FIG. 1 is a schematic configuration diagram of a gas turbine according to one embodiment.
As shown in FIG. 1, a gas turbine 1 includes a compressor 2 for generating compressed air, a combustor 4 for generating combustion gas using the compressed air and fuel, and a combustion gas that is rotationally driven. a turbine 6 configured to: In the case of the gas turbine 1 for power generation, the turbine 6 is connected with a generator (not shown).

The compressor 2 includes a plurality of stator blades 16 fixed to the compressor casing 10 side, and a plurality of rotor blades 18 implanted in the rotor 8 so as to be alternately arranged with respect to the stator blades 16. .
Air taken in from an air intake port 12 is sent to the compressor 2, and this air passes through a plurality of stationary blades 16 and a plurality of moving blades 18 and is compressed to produce a high temperature and high pressure. of compressed air.

Fuel and compressed air generated by the compressor 2 are supplied to the combustor 4 , and the fuel is combusted in the combustor 4 to generate combustion gas, which is a working fluid for the turbine 6 . be done. As shown in FIG. 1 , the gas turbine 1 has a plurality of combustors 4 arranged in a casing 20 along the circumferential direction around the rotor.

The turbine 6 has a combustion gas flow path 28 defined by the turbine casing 22 and includes a plurality of stator vanes 24 and rotor blades 26 provided in the combustion gas flow path 28 . The stationary blades 24 are supported from the turbine casing 22, and a plurality of stationary blades 24 arranged along the circumferential direction of the rotor 8 form a row of stationary blades. Further, the rotor blades 26 are implanted in the rotor 8, and a plurality of rotor blades 26 arranged along the circumferential direction of the rotor 8 form a rotor blade cascade. The row of stationary blades and row of moving blades are alternately arranged in the axial direction of the rotor 8 . Among the plurality of stator vanes 24, the stator vane 24 provided on the most upstream side, that is, the stator vane 24 provided at a position closest to the combustor 4 is a first stage stator vane 23 (hereinafter simply referred to as "stationary vane 23"). ).

In the turbine 6, the combustion gas from the combustor 4 that has flowed into the combustion gas flow path 28 passes through the plurality of stationary blades 24 and the plurality of rotor blades 26, thereby driving the rotor 8 to rotate, and generating power coupled to the rotor 8. The machine is driven to generate electricity. Combustion gas after driving the turbine 6 is discharged to the outside through an exhaust chamber 30 .

The combustor 4 is arranged in an annular shape around the rotor 8 , and the combustion gas generated by combustion of the fuel passes through the outlet 52 of the combustor 4 located at the downstream end of the transition piece 50 to the turbine 6 . flow into the stationary blade 23 of the

Hereinafter, the axial direction of the gas turbine 1 (the axial direction of the rotor 8) is simply referred to as the “axial direction”, the radial direction of the gas turbine 1 (the radial direction of the rotor 8) is simply referred to as the “radial direction”, and the gas The circumferential direction of the turbine 1 (the circumferential direction of the rotor 8) is simply referred to as "circumferential direction". Further, regarding the flow direction of the combustion gas in the combustion gas passage 28, the upstream side in the axial direction is simply referred to as the "upstream side", and the downstream side in the axial direction is simply referred to as the "downstream side".

2 and 3 are diagrams showing schematic configurations of an outlet portion 52 of the combustor 4 and an inlet portion of the turbine 6, respectively, of the gas turbine 1 according to one embodiment. 2 is a cross-sectional view along the circumferential direction, and FIG. 3 is a cross-sectional view along the radial direction. FIG. 4 is a diagram showing a schematic configuration of a combustor 4 (combustor assembly 100) in which the outlet portion 52 and the stator blades 23 are combined according to one embodiment.

As shown in FIGS. 2 and 3 , the gas turbine 1 includes a plurality of combustors 4 arranged in the circumferential direction, and stator vanes 23 located downstream of the exit portion 52 of the combustor 4 .

A plurality of circumferentially arranged combustors 4 constitute a combustor assembly 100 according to some embodiments. As shown in FIGS. 2 and 3 , each of the plurality of combustors 4 has an outlet section 52 located at the downstream end of the combustor 4 , the outlet section 52 of each combustor 4 extending radially along the radial wall portions 54a, 54b extending along the radial direction and circumferential wall portions 56a, 56b extending along the circumferential direction. Here, among the outlet portions 52 of the combustors 4 adjacent in the circumferential direction, the radial wall portion 54 of one combustor 4 and the radial wall portions 54a and 54b of the other combustor 4 face each other. A pair of radial walls 54a and 54b (see FIG. 2).

As shown in FIG. 2, the plurality of stator vanes 23 arranged along the circumferential direction includes a stator vane 23A provided downstream of the above-described pair of radial wall portions 54a and 54b.
In some embodiments, as shown in FIG. 2, the plurality of stator vanes 23 further includes another stator vane 23B provided at a circumferential position between a pair of circumferentially adjacent stator vanes 23A, 23A. . As shown in FIG. 2, the stationary blade 23A extends upstream from the leading edges of the other stationary blades 23B. In FIG. 2, the position of the leading edge of another stationary blade 23B in the axial direction is indicated by a dashed line L1. In the exemplary embodiment shown in FIG. 2, the plurality of circumferentially arranged stator vanes 23 includes stator vanes 23A and other stator vanes 23B that are alternately arranged in the circumferential direction.

In some embodiments, as shown in FIG. 3, the stator vane 23A includes a hollow airfoil body 70, an inner shroud 60 provided at a tip portion 107 radially inward of the airfoil body 70, and the stator vane 23A. and an outer shroud 62 provided at the radially outer proximal end 105 . The stationary blade 23 is supported by the turbine casing 22 via the outer shroud 62 .

In some embodiments, for example, as shown in FIG. 4, the axially upstream end 82 of the wing-body 70 includes a protrusion 86 that projects axially upstream. The blade body 70 connects the convex portion 86 and the trailing edge 76, connects the suction surface wall portion 75 on the suction surface side where the blade surface forms a convex surface, connects the convex portion 86 and the trailing edge 76, and connects the convex portion 86 and the trailing edge 76. and a pressure surface wall portion 73 on the pressure surface side forming a concave surface. The convex portion 86 includes at least one purge hole 87 communicating between the outer space 94 and the inner space 109 of the wing body 70 .

A pair of flat plate portions 148 perpendicular to the axial direction and extending in the circumferential direction are formed on the upstream end face of the upstream end portion 82 in the axial direction. Both circumferential ends of the portion 148 are coupled to the blade surface of the suction surface wall portion 75 and the blade surface of the pressure surface wall portion 73 . The convex portion 86 extends axially upstream from between the pair of flat plate portions 148 and has a groove shape formed in the mating member 55 arranged axially upstream so as to face the upstream end portion 82 . is formed so as to be fitted into the mating member concave portion 55a. Circumferentially facing surfaces of the projection 86 form a pair of side wall surfaces 88 . Between the outer surface of the convex portion 86 and the inner surface of the mating member concave portion 55a, a gap is formed to the extent that the cooling medium discharged from the purge hole 87 flows. The axial length from the flat plate portion 148 to the top surface 86a, which is the axially upstream end surface of the projection 86, is greater than the width of the projection 86 in the circumferential direction defined by the distance between the pair of side wall surfaces 88, It is larger than the axial gap 95 formed between the flat plate portion 148 of the upstream end portion 82 and the axial downstream end face 55 b of the mating member 55 .

According to the stationary blade 23A, when the convex portion 86 positioned at the upstream end portion 82 of the stationary blade 23A is fitted with the mating member 55 on the combustor 4 side, the convex portion 86 of the stationary blade 23A is The pair of side wall surfaces 88 and the mating member 55 can be overlapped axially and circumferentially.

As a result, the axial gap 95a between the axial downstream end face 55b of the mating member 55 and the flat plate portion 148 of the stationary blade 24 on the pressure surface wall portion 73 side and the axial downstream end face 55b of the mating member 55 are formed. and the flat plate portion 148 on the suction surface wall portion 75 side of the stationary blade 24 are not connected in a straight line in the circumferential direction, but the pair of side wall surfaces 88 of the convex portion 86 By combining the side member 55 with the counterpart member concave portion 55a, a structure is formed in which the convex portion 86 and the counterpart member 55 overlap in the axial direction and the circumferential direction. and the gap 95b are connected. Therefore, the gap 95 that connects the outlets 52 of the plurality of adjacent combustors 4 is formed as a small groove-shaped gap that bends in an uneven shape, and the combustion vibration caused by acoustic propagation in the outlets 52 of the plurality of combustors 4 is suppressed. can be reduced.

Further, according to the above configuration, the cooling medium is supplied from the internal space 109 of the blade body 70 to the gap 95 between the convex portion 86 of the stationary blade 23A and the mating member 55 on the combustor 4 side through the purge hole 87. can supply. In this way, the upstream end 82 of the stationary blade 23A is cooled by the flow of the cooling medium supplied to the gap 95 between the protrusion 86 of the stationary blade 23A and the mating member. heat damage is prevented.

This enables stable operation of the gas turbine 1 . For example, compressed air generated by the compressor 2 may be supplied to the internal space 109 of the wing body 70 as a cooling medium.

FIG. 5 is a diagram showing a schematic configuration in which the outlet portion 52 and the stator blades 23 of the combustor 4 (combustor assembly 100) are combined according to one embodiment.
In some embodiments, for example, as shown in FIG. 5 , the axially upstream end 82 of the wing-body 70 includes an axially downstream recess 84 . The blade body 70 connects the recess 84 and the trailing edge 76, connects the recess 84 and the trailing edge 76, and connects the recess 84 and the trailing edge 76 to a suction surface wall portion 75 on the suction surface side in which the blade surface forms a convex surface, and the blade surface forms a concave surface. and a pressure surface wall portion 73 on the pressure surface side to be formed. Recess 84 includes at least one purge hole 87 communicating between outer space 94 and inner space 109 of wing body 70 .

A pair of flat plate portions 148 perpendicular to the axial direction and extending in the circumferential direction are formed on the upstream end face in the axial direction of the upstream end portion 82 . 148 are coupled to the suction surface wall portion 75 and the pressure surface wall portion 73 at both circumferential ends thereof. The recessed portion 84 is recessed axially downstream from between the pair of flat plate portions 148 and is formed in the mating member 55 arranged axially upstream so as to face the upstream end portion 82 . It is formed so that the mating member convex portion 55c is fitted. Circumferentially facing surfaces of the recess 84 form a pair of side wall surfaces 88 . Between the inner surface of the concave portion 84 and the outer surface of the mating member convex portion 55c, a gap is formed to the extent that the cooling medium discharged from the purge hole 87 flows. The axial depth of the recess 84 is greater than the circumferential width defined by the pair of side wall surfaces 88 of the recess 84 , and the flat plate portion 148 of the upstream end 82 and the axial downstream end surface 55 b of the mating member 55 is greater than the axial gap 95 formed between
According to the stator vane 23A, when the recess 84 positioned at the upstream end 82 of the stator vane 23A is fitted with the mating member 55 on the combustor 4 side, the pair of recesses 84 of the first stage stator vane 23A The side wall surface 88 and the mating member 55 can overlap axially and circumferentially.

As a result, the axial gap 95a between the axial downstream end face 55b of the mating member 55 and the flat plate portion 148 of the stationary blade 24 on the pressure surface wall portion 73 side and the axial downstream end face 55b of the mating member 55 are formed. and the flat plate portion 148 on the suction surface wall portion 75 side of the stationary blade 24 are not connected linearly in the circumferential direction, but are formed by the pair of side wall surfaces 88 of the concave portion 84 and the opposite side wall surface 88. The combination of the member 55 and the mating member convex portion 55c forms a structure that overlaps in the axial direction and the circumferential direction, and the gap 95a and the gap 95b in the axial direction are connected via the overlapping structure. Therefore, the gap 95 that connects the outlets 52 of the plurality of adjacent combustors 4 is formed as a groove-shaped small gap that bends in an uneven shape. Vibration can be reduced.

Further, according to the above configuration, the cooling medium is supplied from the internal space 109 of the blade body 70 to the gap 95 between the recessed portion 84 of the stationary blade 23A and the mating member 55 on the combustor 4 side through the purge hole 87. can do. Thus, the stationary blade 23A can be cooled by the flow of the cooling medium supplied to the gap 95 between the concave portion 84 of the stationary blade 23A and the mating member.

This enables stable operation of the gas turbine 1 . For example, compressed air generated by the compressor 2 may be supplied to the internal space 109 of the wing body 70 as a cooling medium.

FIG. 6 is an enlarged cross-sectional view showing an example of a specific configuration of the stationary blade 23A shown in FIG. 4. As shown in FIG. FIG. 7 is a diagram showing the AA section in FIG. FIG. 8 is an enlarged cross-sectional view showing another example of the specific configuration of the stationary blade 23A shown in FIG. FIG. 9 is a diagram showing a CC cross section in FIG.

In some embodiments, the wing body 70 includes at least one side purge hole 89 circumferentially adjacent to the purge hole 87, for example, as shown in FIGS. In the illustrated form, the blade body 70 includes a pair of side purge holes 89 provided on both sides of the purge hole 87 in the circumferential direction.

According to this configuration, the cooling medium is discharged from the side purge holes 89 to the external space 94 of the blade body 70 , thereby suppressing the backflow of the combustion gas to the internal space 109 of the blade body 70 . By lowering the temperature of the combustion gas, thermal damage to the upstream end portion 82 of the blade body 70 can be suppressed.

In some embodiments, for example, as shown in FIGS. 6 and 8 , the outer wall surface 112 of the upstream end 82 is formed into a suction surface wall portion 75 and a pressure surface wall portion 75 with the protrusion 86 or the recess 84 circumferentially sandwiched therebetween. It includes a pair of flat plate portions 148 formed of flat surfaces extending to 73 . Such a configuration facilitates machining of the upstream end portion 82 of the blade body 70 .

In some embodiments, as shown, for example, in FIGS. 7 and 9, the protrusions 86 or recesses 84 extend along the blade height direction from the tip side of the wing body 70 (the inner shroud 60 side in FIG. 3) toward the base. Extending toward the end side (outer shroud 62 side in FIG. 3), the purge holes 87 include a plurality of purge holes 87 arranged from the tip side to the base end side of the blade body 70 along the blade height direction. The convex portion 86 or the concave portion 84 may extend over half or more of the length between the base end portion 105 and the tip portion 107 of the stationary blade 23A in the blade height direction. In addition, as shown in FIGS. 7 and 9, the cooling medium supplied to the blade body 70 from the outside has a one-sided supply structure in which the cooling medium is supplied from one direction on the base end portion 105 side.

According to such a configuration, the cooling medium is discharged to the outer space 94 of the blade body 70 through the plurality of purge holes 87 arranged from the tip end side to the base end side of the blade body 70 , thereby allowing the cooling medium to flow into the inner space 109 of the blade body 70 . , the temperature of the combustion gas outside the purge hole 87 can be lowered, and thermal damage to the upstream end portion 82 of the blade body 70 can be suppressed.

The basic concept of the wing structure common to several embodiments will be described below.
As shown in FIGS. 4 and 5, the wing structure common to some embodiments is such that the internal space of the wing body 70 from the upstream end 82 of the wing body 70 to the trailing edge 76 is separated by a partition wall 150. The cooling medium that is partitioned into a plurality of spaces and supplied to the spaces cools the blade body 70 in the process of being discharged into the combustion gas through a plurality of cooling holes 136 formed in the blade surfaces 67 and 69. there is In particular, the front outer wall portion 158 forming the front space 160 closest to the upstream end portion 82 is arranged on the most upstream side of the blade surface in the flow direction of the combustion gas. Appropriate cooling structure is required.

Moreover, the combustion gas flowing on the suction surface wall portion 75 side of the blade body has a faster flow velocity than the combustion gas flowing on the pressure surface wall portion 73 side, and the combustion gas pressure decreases significantly from the pressure surface wall portion 73 side. Therefore, the discharge amount of the cooling medium changes depending on the pressure difference between the front space 160 adjacent to the suction surface wall portion 75 side of the blade body 70 and the combustion gas flow. Since the consumption flow rate of the cooling medium affects the thermal efficiency of the gas turbine 1, appropriate management of the consumption amount of the cooling medium is required. In other words, the blade structure is such that the pressure difference between the front side space 160 of the blade body 70 and the combustion gas is kept as uniform as possible in the direction along the blade surface where the combustion gas flows, and the consumption of the cooling medium is appropriate. requested. A blade structure that changes the pressure in the front space 160 according to the pressure change of the combustion gas flowing along the blade surface 69 on the suction surface wall portion 75 side will be described below.

In some embodiments, for example, as shown in FIGS. 6 and 8, the airfoil 70 of the vane 23A has a suction surface wall portion 75 inside the airfoil 70 to partition the interior space of the airfoil into a plurality of compartments. and the pressure surface wall portion 73, and includes n (n is an integer equal to or greater than 1) partition walls 150 extending linearly, and the internal space of the upstream end portion 82 includes a first Insert 114 and second insert 152 are positioned. The first insert 114 and the second insert 152 are composed of a front partition wall 154, which is the partition wall 150 farthest from the trailing edge 76 (see FIGS. 4 and 5) of the n partition walls 150, and an outer wall 156 of the wing body 70. It is provided in a front space 160 defined by a front outer wall portion 158 which is a portion on the opposite side of the rear edge 76 across the front partition wall 154 . A low-pressure impingement space 176 (second impingement space) is formed between the inner wall surface 130 of the wing body 70 and the outer peripheral surface 132 of the second insert 152 . The second insert 152 includes an impingement wall 198 having a plurality of impingement holes 194 communicating between an interior space 196 of the second insert 152 and the low pressure impingement space 176 . In the illustrated form, the impingement wall portion 198 is provided opposite the front outer wall portion 158 . Wing body 70 includes upstream end 82, partition 150, front partition 154, front outer wall 158, and front space partition 170 (described below).

That is, as shown in FIGS. 6 and 8 , the stationary blade 23A includes a front space partition wall 170 (partition) that partitions the high pressure impingement space 134 and the low pressure impingement space 176 . In the blade structure shown in FIGS. 6 and 8, the pressure of the combustion gas flowing along the blade surface is almost the same on the pressure surface wall portion 73 side and the suction surface wall portion 75 side, without a large difference. However, the pressure of the combustion gas flow on the side of the suction surface wall portion 75 sharply decreases toward the trailing edge 76 with the position of the front space partition wall 170 as a boundary. That is, the position on the blade surface 69 side where the axially upstream partition wall of the front space partition wall 170 connects to the blade body 70 is a branching point where the pressure of the combustion gas flow on the suction surface wall portion 75 side decreases.

The blade structure shown in FIG. 6 or 8 is a structure in which the front space 160 is divided into two spaces by the front space partition wall 170 according to the pressure change of the combustion gas on the suction surface wall portion 75 side, as described above. . The front space 160 is divided into a space close to the convex portion 86 or the concave portion 84 where the combustion gas pressure on the suction surface wall portion 75 side including a portion of the pressure surface wall portion 73 side is high, and a space near the concave portion 84 on the suction surface wall portion 75 side where the combustion gas pressure is high. It is a structure in which the first insert 114 and the second insert 152 are arranged in the respective spaces, which are divided into a low convex portion 86 and a space far from the concave portion 84 . Although details will be described later, the impingement walls 140 and 198 forming the first insert 114 and the second insert 152 form a high pressure impingement space 134 between the inner wall surface 130 of the wing body 70 and the first insert 114. A low pressure impingement space 176 is formed between the inner wall surface 130 of the wing body 70 and the second insert 152 . A plurality of impingement holes 138, 194 are formed in the impingement wall portions 140, 198, and the cooling medium supplied from the outside to the internal spaces 109, 196 is ejected from the impingement holes 138, 194 to enter the high-pressure impingement space. 134 and the inner wall surface 130 facing the low pressure impingement space 176 is impingement cooled (impingement cooled). The cooling medium after impingement cooling film-cools the blade surfaces 67 and 69 of the blade body 70 in the process of being discharged into the combustion gas from the cooling holes 136 formed in the blade surfaces.

The high pressure impingement space 134, in which the first insert 114 is located, faces both the pressure side wall 73 and the suction side wall 75, but as described above, the pressure side wall 73 and the suction side wall 75 The pressure of the combustion gas is approximately the same pressure. Therefore, regarding the specifications (hole diameter, opening area, arrangement, etc.) of the plurality of impingement holes 138 formed in the impingement wall portion 140 constituting the first insert 114, each of the pressure surface wall portion 73 and the suction surface wall portion 75 has It is desirable that the impingement holes 138 of the opposing impingement wall portions 140 have the same specifications. As a result, the pressure difference between the high-pressure impingement space and the combustion gas becomes uniform on both the pressure surface wall portion 73 side and the suction surface wall portion 75 side, and the flow rate of the cooling medium discharged into the combustion gas from the cooling holes 136 becomes uniform. become.

On the other hand, in the case of the low-pressure impingement space 176 in which the second insert 152 is arranged, the pressure of the combustion gas on the side of the suction surface wall portion 75 facing the low-pressure impingement space 176 rapidly decreases along the blade surface 69. . Therefore, maintaining the pressure in the low pressure impingement space 176 at the same pressure as the pressure in the high pressure impingement space 134 increases the pressure difference between the low pressure impingement space 176 and the combustion gas. In other words, the flow rate of the coolant discharged into the combustion gas from the low-pressure impingement space 176 through the cooling holes 136 exceeds an appropriate value, causing an excessive amount of coolant to flow out, resulting in a decrease in the thermal efficiency of the gas turbine. In order to suppress excessive outflow of the cooling medium from the low-pressure impingement space 176 into the combustion gas, a specification different from the specification of the impingement holes 138 formed in the impingement wall portion 140 of the first insert 114 should be selected. is desirable. For example, the flow path of the cooling medium to be discharged can be adjusted by reducing the diameter or opening area of the impingement holes 194 formed in the impingement wall portion 198 of the second insert 152 or by increasing the arrangement pitch of the impingement holes 194. is desirable.

In this way, the specifications such as the appropriate hole diameter, opening area, arrangement, etc. of the impingement holes 138 and 194 formed in the impingement wall portions 140 and 198 of the first insert 114 and the second insert 152 are appropriately selected so that high pressure can be achieved. It is desirable to select impingement hole specifications such that the pressure differential between impingement space 134 and low pressure impingement space 176 and the combustion gases is approximately the same. As a result, the impingement cooling of the inner wall surface 130 of the blade body 70 by the impingement holes 138 and 194 and the film cooling effect of the cooling holes 136 properly cool the blade body 70 . Furthermore, by properly managing the flow rate of the coolant discharged into the combustion gases from the first insert 114 and the second insert 152, coolant consumption is reduced and the thermal efficiency of the turbine is improved.

On the other hand, as described above, the communication space 172 has substantially the same pressure as the internal space 109 and is the space with the highest pressure in the front space 160 . Therefore, in order to seal between the communication space 172 and the high-pressure impingement space 134, a pair of seal dams 142 whose structural details will be described later are arranged.

Next, the structures of the first insert and the second insert will be described.
FIG. 10 is a diagram showing an example of a schematic configuration of the first insert 114 according to one embodiment. FIG. 11 is a diagram showing an example of a schematic configuration of a first insert 114 according to another embodiment. FIG. 12 is a cross-sectional view showing an example of a range in which the insert collar 124 is provided in the first insert 114 shown in FIG. 11. As shown in FIG. In FIG. 12, only the insert collar 124 is shown in solid lines, and the wing body 70 and the first insert 114 and the second insert 152 are shown in phantom lines.

As shown in FIGS. 6-9, the first insert 114 includes an opening formation 118 having an opening 116 extending along the blade height direction, the opening 116 axially facing the purge hole 87 . are arranged as follows.

In some embodiments, for example, as shown in FIGS. 7, 9, 10 and 11, the first insert 114 includes a tubular body 120 and a proximal end 105 side of the tubular body 120 . an insert collar 124 provided in the shape of a brim at the end of the tube 120; In the first insert 114 shown in FIG. 10, the insert collar 124 is provided over the entire circumference of the end portion of the tubular body 120 on the base end portion 105 side. A lid portion 121 (see, for example, FIGS. 7 and 9) having an annular guide 121a formed on the inner surface of the cylindrical body 120 is disposed on the distal end portion 107 side of the cylindrical body 120 . The lid portion 121 is fixed by welding or the like to the end face of the wing body 70 on the tip portion 107 side. An annular guide 121 a formed on the lid portion 121 is fitted into the inner peripheral surface 120 a of the tubular body 120 of the first insert 114 . The outer peripheral surface of the guide 121a and the inner peripheral surface 120a of the cylinder 120 are fitted so as to be slidable in the blade height direction, and absorb the difference in thermal expansion of the cylinder 120 in the blade height direction. It is possible. The reinforcing member 123 is located on the tip portion 107 side of the inner peripheral surface 120a of the cylindrical body 120 of the first insert 114 and is closest to the tip portion 107 side of the purge holes 87 formed in the blade body 70 in the blade height direction. It is desirable that the purge hole 87 be located closer to the tip 107 side than the purge hole 87 located at the bottom, and is fixed to the inner surface of the cylindrical body 120 by welding or the like. By providing the reinforcing member 123, the rigidity of the tip portion 107 side of the tubular body 120 is increased, and the occurrence of vibration of the tubular body 120 is suppressed.

In another embodiment of the first insert 114 shown in FIG. 11, the insert collar 124 has cutouts 126 positioned corresponding to the location of the aperture formations 118 in the circumferential direction. For example, as shown in FIG. 12, the outer peripheral edge 124a of the insert collar 124 extends from the barrel 120 to the wing body 70, and on the proximal end 105 side of the barrel 120, the inner wall surface 130 of the wing body 70 and the front partition. The wing body 70 and the front partition wall 154 and the front space partition wall are arranged so as to seal the gap (high pressure impingement space 134) between the wall 154 and the wall surface 190 of the front space partition wall 170 and the outer peripheral surface 132 of the first insert 114. It is fixed to the end face of 170 by welding, brazing, or the like. 7 and 9, the cooling medium supplied to the first insert 114 is supplied from one direction (single-side supply structure) on the base end portion 105 side. The first insert 114 is fixed via an insert collar 124 to the end surface of the base end portion 105 side of the wing body 70 shown in FIG. 10 or 11 .

According to the above configuration, since the insert collar 124 has the cutout portion 126 in the portion corresponding to the position of the opening 116 of the opening forming portion 118 in the circumferential direction, part of the cooling medium supplied from the outside is cut into the blade body 70. Since the cooling medium can be directly supplied to the gap between the inner wall surface 130 and the outer peripheral surface 132 of the first insert 114 , the high-pressure cooling medium can be supplied to the purge hole 87 while suppressing the pressure loss of the cooling medium as much as possible.

In some embodiments, for example, as shown in FIGS. 6 and 8, a cylindrical second insert 114 is provided axially adjacent and downstream of the first insert 114 and extends along the blade height direction. An insert 152 is provided. Similarly to the first insert 114, the second insert 152 also has a tubular body 120 formed in a tubular shape, an insert collar 124 provided at the end of the tubular body 120 on the proximal end 105 side, and a distal end. It includes a lid portion 121 provided at the end on the 107 side and a reinforcing member 123 fixed to the inner surface of the cylinder 120 . However, since the second insert 152 is arranged closer to the trailing edge 76 than the first insert 114 , the inner space 196 of the second insert 152 does not need to be directly connected to the communicating space 172 . Therefore, the notch 126 is not formed in the insert collar 124 of the second insert 152 . Other structures are the same as those of the first insert 114 shown in FIG. The attachment of the second insert 152 to the wing body 70, the front partition wall 154, and the front space partition wall 170 is fixed by welding or brazing through the insert collar 124 in the same manner as the first insert 114. be. As a result, the gaps between the inner wall surface 130 and the front partition wall 154 of the wing body 70 and the wall surface 190 of the front space partition wall 170 and the outer peripheral surface 132 of the second insert 152 are sealed.

According to the configuration of the first insert described above, the high-pressure cooling medium flows from the internal space 109 of the first insert 114 into the high-pressure impingement space 134 through the impingement holes 138 formed in the impingement wall portion 140 of the first insert 114 . is discharged. Also, the cooling medium is discharged from the internal space 196 of the second insert 152 to the low-pressure impingement space 176 through the impingement holes 194 formed in the impingement wall portion 198 of the second insert 152 . As a result, in the front space 160, the impingement holes 138, 194 of the first insert 114 and the second insert 152 are properly adjusted in accordance with changes in the pressure of the combustion gas flowing along the blade surface 69 on the suction surface wall portion 75 side. By selecting the specifications, the impingement cooling of the inner wall surface 130 of the wing body 70 can be effectively performed.

According to this configuration, the high-pressure cooling medium that has flowed out of the first insert 114 from the internal space 109 of the first insert 114 through the opening 116 of the opening forming portion 118 is smoothly discharged from the purge hole 87 to the stationary blade 23A. It can be directly supplied to the gap 95 between the convex portion 86 or the concave portion 84 and the mating member 55 (see FIGS. 4 and 5) on the combustor 4 side. As a result, the convex portion 86 or the concave portion 84 of the stationary blade 23A can be cooled, and thermal damage to the convex portion 86 or the concave portion 84 can be prevented.

Also, on the inner wall surface 130 of the upstream end 82 of the blade body 70 , a pair of first seal dams 142 each extending in the blade height direction so as to separate a high-pressure impingement space 134 from the first insert 114 . is formed. The pair of first seal dams 142 are arranged to sandwich the purge hole 87 in the circumferential direction. That is, one of the pair of first seal dams 142 is formed on the inner wall surface 130 of the suction surface wall portion 75 of the upstream end portion 82 , and the other of the pair of first seal dams 142 is formed on the upstream end portion 82 . It is formed on the inner wall surface 130 of the pressure surface wall portion 73 .

FIG. 13 is a cross-sectional view showing an example of the configuration of the first shield dam 142. As shown in FIG.
As shown in FIG. 13, the first insert 114 has a pair of ribs 178 on the outer peripheral surface 132 of the first insert 114, which are arranged to sandwich the opening 116 in the circumferential direction. and are fitted to the pair of first shield dams 142 respectively. In the illustrated exemplary embodiment, one of the pair of first seal dams 142 is a projection 200 formed on the inner wall surface 130 of the upstream end 82 of the blade body 70 and extending in the blade height direction. A groove portion 202 extending in the blade height direction is formed on the top surface of the portion 200 , and the tip of the rib 178 engages with the groove portion 202 . The other of the pair of first seal dams 142 is a projection 200 formed on the inner wall surface 130 of the upstream end 82 of the blade body 70 and extending in the blade height direction. A groove portion 202 extending in the blade height direction is formed, and the tip of the rib 178 engages with the groove portion 202 . As shown in FIG. 13, the pair of seal dams 142 are arranged symmetrically on both sides in the circumferential direction with the purge hole 87 as the center.

6 and 8, the upstream end 82 includes a pair of flat plate portions 148 extending in the circumferential direction and forming a flat surface, and the protrusion 86 or the recess 84 is located between the flat plate portions 148. The pair of first seal dams 142 are formed on the inner wall surface 130 of the flat plate portion 148 so as to protrude axially upstream or be recessed axially downstream. According to such a configuration, the provision of the flat plate portion 148 facilitates processing of the upstream end portion of the blade body.

As described above, the combination of the division of the front space 160 by the front space partition wall 170 and the arrangement of the pair of first shield dams 142 allows the front space 160 to be divided into the internal spaces 109 and 196, the communication space 172, and the high pressure impingement space 134. The low pressure impingement space 176 is divided into five spaces. However, the internal spaces 109 and 196 and the communication space 172 are spaces having a pressure close to the supply pressure of the cooling medium having almost the same pressure. Therefore, the space between the inserts (the first insert 114 and the second insert 152) and the wing body 70 consists of a communication space 172 with the highest pressure and an intermediate pressure lower than the communication space 172 and higher than the low pressure impingement space 176. and a low-pressure impingement space 176 with the lowest pressure.

As described above, by selecting the appropriate impingement holes 139 and 194 of the first insert 114 and the second insert 152 and appropriately arranging the spaces with different pressures, the high pressure impingement space 134 and the low pressure impingement space The pressure difference between 176 and the combustion gas flow on the negative pressure surface wall portion 75 side becomes substantially the same, and the pressure difference is leveled. Therefore, it is possible to suppress the phenomenon that the cooling medium flows excessively from the cooling holes of the blade surface 69 or the metal temperature of the blade surface rises due to insufficient flow rate of the cooling medium. As a result, thermal damage to the blade body is prevented, and the consumption of the cooling medium is appropriately managed, thereby improving the thermal efficiency of the gas turbine.

Moreover, the combination of the protrusions 200 of the pair of first seal dams 142 and the ribs 178 inserted into the grooves 202 formed in the protrusions 200 creates a communication space 172 that communicates the opening 116 and the purge hole 87 with a high-pressure impingement. A space 134 is sealed. Therefore, leakage of the cooling medium from the communication space 172 to the high-pressure impingement space 134 via the pair of first seal dams 142 is suppressed. Therefore, the purge hole 87 from the internal space 109 of the blade body 70 passes through the gap 95 (see FIGS. 4 and 5) between the convex portion 86 or concave portion 84 of the stationary blade 23A and the mating member on the combustor 4 side. A high-pressure cooling medium can be supplied to the purge hole 87 without lowering the pressure of the cooling medium supplied to the first insert 114 through the . As a result, the stationary blade 23A can be cooled, and thermal damage to the convex portion 86 or the concave portion 84 can be prevented.

Next, another aspect of the structure for partitioning the space of the wing body provided with the first insert 114 and the second insert 152 shown in FIGS. 14 and 16 will be described below.
FIG. 14 is an enlarged cross-sectional view showing an example of another aspect of the specific configuration of the stationary blade 23A shown in FIG. FIG. 15 is a diagram showing a BB section in FIG. FIG. 16 is an enlarged cross-sectional view showing another example of the specific configuration of the stationary blade 23A shown in FIG. 17 is a diagram showing a DD cross section in FIG. 16. FIG. FIG. 18 is a perspective view showing one embodiment of an insert that combines first insert 114 and second insert 152 and shared insert collar 125 . FIG. 19 is an installation diagram of the shared insert collar showing the range of installation of the shared insert collar 125. Only the shared insert collar 125 is shown in solid lines, and other configurations such as the wing bodies and inserts are shown in phantom lines. there is

In some embodiments, for example, as shown in FIGS. 14 and 16, a partition structure for the space of the front space 160 in which the first insert 114 and the second insert 152 are located, compared to the structure of FIGS. is a structure in which a partition wall such as the front space partition wall 170 is not provided, and the impingement space is partitioned by a combination of a plurality of pairs of seal dams. That is, a pair of seal dams 142 are arranged between the outer peripheral surface 132 of the first insert 114 near the upstream end 82 and the inner wall surface 130 of the wing body 70 . A pair of second seal dams 166 (partitions) are arranged between the outer peripheral surface 132 of the second insert 152 and the inner wall surface 130 of the wing body 70 . This structure is different from the structures shown in FIGS.

In some embodiments, the wing body 70 includes at least one side purge hole 89 circumferentially adjacent to the purge hole 87, for example as shown in FIGS. In the illustrated form, the blade body 70 includes a pair of side purge holes 89 provided on both sides of the purge hole 87 in the circumferential direction.

According to this configuration, the cooling medium is discharged from the side purge holes 89 to the external space 94 of the blade body 70 , thereby suppressing the backflow of the combustion gas to the internal space 109 of the blade body 70 . By lowering the temperature of the combustion gas, damage to the upstream end portion 82 of the blade body 70 can be suppressed.

In some embodiments, for example, as shown in FIGS. 14 to 17, the stationary blade 23A has a tubular first insert 114 extending in the blade height direction in the internal space 109 of the blade body 70. Prepare. The first insert 114 includes an opening forming portion 118 having an opening 116 extending along the blade height direction, the opening 116 being arranged opposite the purge hole 87 .

As shown in FIGS. 15 and 17, the convex portion 86 or the concave portion 84 extends along the blade height direction from the tip side (the inner shroud 60 side in FIG. 3) of the blade body 70 to the proximal side (the outer shroud side in FIG. 3). 62 side), and the purge holes 87 include a plurality of purge holes 87 arranged from the tip side to the base end side of the blade body 70 along the blade height direction. The convex portion 86 or the concave portion 84 may extend over half or more of the length between the base end portion 105 and the tip portion 107 of the stationary blade 23A in the blade height direction.

According to this configuration, the high-pressure cooling medium that has flowed out of the first insert 114 from the internal space 109 of the first insert 114 through the opening 116 of the opening forming portion 118 flows from the tip side of the blade body 70 to the base end side. By discharging the cooling medium to the outer space 94 of the blade body 70 from the plurality of purge holes 87 arranged along the By reducing the temperature of the gas, damage to the upstream end 82 of the blade body 70 can be suppressed. In addition, the protrusions 86 or the recesses 84 of the stationary blade 23A can be cooled to prevent thermal damage to the protrusions 86 or the recesses 84 .

Due to the configuration of the multiple pairs of seal dams described above, the front space 160 is divided into five spaces: the internal spaces 109 and 196 , the communication space 172 , the high pressure impingement space 134 and the low pressure impingement space 176 . However, the internal spaces 109 and 196 and the communication space 172 are spaces having a pressure close to the supply pressure of the cooling medium having almost the same pressure. Therefore, when the space between the insert (first insert 114, second insert 152) and the blade body 70 is divided by the difference in pressure, the communication space 172 and the internal spaces 109 and 196 with the highest pressure, the communication space 172 and the internal space A high pressure impingement space 134 having an intermediate pressure lower than the pressures 109 and 196 and higher than the low pressure impingement space 176, and a low pressure impingement space 176 having the lowest pressure. be.

Also in this embodiment, impingement holes 138 and 194 are formed in the impingement wall portions 140 and 198 of the first insert 114 and the second insert 152, as shown in FIGS. A high-pressure impingement space 134 formed between the outer peripheral surface 132 of the first insert 114 and the inner wall surface 130 of the wing body 70 is formed through an impingement hole 138 formed in the impingement wall portion 140 of the first insert 114. , and communicates with the internal space 109 of the first insert 114 . Also, the low-pressure impingement space 176 formed between the outer peripheral surface 132 of the second insert 152 and the inner wall surface 130 of the wing body 70 is the impingement hole 194 formed in the impingement wall portion 198 of the second insert 152 . is communicated with the internal space 196 of the second insert 152 via the . In addition, a plurality of cooling holes 136 are formed on the pressure surface wall portion 73 side and the suction surface wall portion 75 side of the blade body 70, and a high pressure impingement space 134, a low pressure impingement space 176, and combustion on the blade surfaces 67, 69 side. It communicates the gas space.

Also in this embodiment, the pressures of the combustion gas flow on the pressure surface wall portion 73 side and the suction surface wall portion 75 side of the first insert 114 are substantially the same. The pressure of the combustion gas on the suction surface wall portion 75 side is higher than the position of the blade surface of the suction surface wall portion 75 corresponding to the position of the space gap 160a where the first insert 114 and the second insert 152 face each other, on the trailing edge 76 side. drop sharply. Therefore, as in the embodiment shown in FIGS. 6 and 8, the specifications of the impingement holes 138, 194 of the first insert 114 and the second insert 152 are appropriately selected to provide the high pressure impingement space 134 and the low pressure impingement space. It is desirable that the pressure difference between 176 and the combustion gas flow on the side of the suction surface wall portion 75 be approximately the same. For example, the impingement of the impingement wall portion 140 of the first insert 114 is such that the flow rate of the cooling medium discharged from the high pressure impingement space 134 side is approximately the same as the flow rate of the coolant discharged from the low pressure impingement space 176 side. Selecting the proper specification of the impingement holes 194 in the impingement wall 198 of the second insert 152 as compared to the holes 138 to select the proper amount of coolant discharged from the second insert 152 into the combustion gases. It is desirable to

By arranging the spaces having different pressures, the pressure difference between the high-pressure impingement space 134 and the low-pressure impingement space 176 and the combustion gas flow flowing along the blade surface 69 on the suction surface wall portion 75 side becomes substantially the same. Differences are smoothed out. Therefore, it is possible to suppress the phenomenon that the cooling medium flows out excessively from the cooling holes of the blade surface 69 or the metal temperature of the blade surface rises due to the shortage of the discharge flow rate of the cooling medium. As a result, thermal damage to the blade body is prevented, and the consumption of the cooling medium is appropriately managed, thereby improving the thermal efficiency of the gas turbine.

Also, on the inner wall surface 130 of the upstream end 82 of the blade body 70 , a pair of first seal dams 142 each extending in the blade height direction so as to separate a high-pressure impingement space 134 from the first insert 114 . is formed. The pair of first seal dams 142 are arranged symmetrically with the purge hole 87 interposed therebetween in the circumferential direction. That is, one of the pair of first seal dams 142 is formed on the inner wall surface 130 of the suction surface wall portion 75 of the upstream end portion 82 , and the other of the pair of first seal dams 142 is formed on the upstream end portion 82 . It is formed on the inner wall surface 130 of the pressure surface wall portion 73 . The detailed structure of the first shield dam 142 is the same as the structure shown in FIG.

A gap 95 (see FIGS. 4 and 5) is formed between the convex portion 86 or concave portion 84 of the stationary blade 23A and the mating member on the combustor 4 side. Cooling medium supplied to the first insert from the interior space 109 of the wing body 70 through the purge holes 87 can be supplied to the gap 95 without lowering the pressure of the cooling medium. As a result, the stationary blade 23A can be cooled, and thermal damage to the convex portion 86 or the concave portion 84 can be prevented.

Also, a pair of first seal dams 142 are formed on the inner wall surface 130 of the flat plate portion 148 and combined with ribs 178 fixed to the outer peripheral surface 132 of the first insert 114, so that the first seal dams 142 and the first insert 114 are bonded together. It can effectively suppress the leakage flow between and achieve effective impingement cooling.

In some embodiments, the insert shown, for example, in Figures 18 and 19 includes a first insert 114 and a second insert 152 and a shared insert collar 125 shared by both inserts. The first insert 114 and the second insert 152 each include a tubular body 120 formed in a tubular shape. In addition, the shared insert collar 125 shared by the first insert 114 and the second insert 152 is fixed to the end of the cylindrical body 120 on the base end 105 side by welding or the like, and the cylindrical body 120 and the shared insert collar 125 are integrated. It is The outer peripheral end 125a of the shared insert collar 125 seals the gap (high pressure impingement space 134) between the inner wall surface 130 of the wing body 70 and the front partition wall 154 and the outer peripheral surface 132 of the first insert 114 and the second insert 152. , it is fixed by welding or the like to the end face of the wing body 70 including the front partition wall 154 on the base end portion 105 side. 15 and 17, the cooling medium supplied to the first insert 114 and the second insert 152 is supplied from one direction (single-side supply structure) on the base end portion 105 side, and the common insert The collar 125 is arranged on the end face in the blade height direction on the base end portion 105 side shown in FIGS. Other internal structures of the first insert 114 and the second insert 152 (cover portion 121, guide 121a, reinforcing member 123, etc.) are the same as those shown in FIG.

In addition, as shown in FIG. 19, the shared insert collar 125 indicated by the solid line has a base end portion that matches the planar cross-sectional shape of the internal space 109 of the first insert 114 and the internal space 196 of the second insert 152 . An opening 122 is formed on the 105 side. By attaching the first insert 114 and the second insert 152 with the shared insert collar 125 and lid portion 121 to the blade body 70, both ends of the high pressure impingement space 134 and the low pressure impingement space 176 in the blade height direction are closed. be.

In the first insert 114 shown in FIGS. 18 and 19, the shared insert collar 125 has cutouts 126 positioned corresponding to the location of the opening formations 118 in the circumferential direction.

According to the above configuration, since the shared insert collar 125 has the notch portion 126 at a position corresponding to the position of the opening 116 of the opening forming portion 118 in the circumferential direction, part of the cooling medium supplied from the outside is Since the cooling medium can be directly supplied to the gap between the inner wall surface 130 of the first insert 114 and the outer peripheral surface 132 of the first insert 114 , the high-pressure cooling medium can be supplied to the purge hole 87 while minimizing the pressure loss of the cooling medium. Note that the notch 126 is not required for the shared insert collar 125 near the second insert 152 that does not face the opening 116 of the opening forming portion 118 .

The above-described shared insert collar 125 has been described in a mode of applying the integrally formed shared insert collar 125 common to the first insert 114 and the second insert 152, but from the viewpoint of ease of insert installation, the first insert A structure in which insert collars 124 are individually attached to 114 and first insert 152 may also be used. In that case, the first insert 114 and the second insert 152 to which the insert collar 124 is attached are individually inserted into the front space 160, and the outer peripheral end 124a of each insert collar 124 is positioned at the blade height of the wing body 70 and the front partition wall 154. Fixed to the vertical end face. After that, the end portion of the insert collar 124 of the first insert 114 adjacent to the second insert 152 and the end portion of the insert collar 124 of the second insert 152 adjacent to the first insert 114 are welded or the like. They may be joined together to form a shared insert collar 125 . According to such a structure, the gap between the wall surface 190 of the wing body 70 and the front partition wall 154 and the outer peripheral surface 132 of the first insert 114 and the second insert 152, and the adjacent second The gap between the opposing outer peripheral surface 132 on the insert side and the opposing outer peripheral surface 132 on the adjacent first insert side of the second insert 152 is sealed, and the insert can be easily attached to the front space 160. .

In some embodiments, for example, as shown in FIGS. 14 and 16, the forward space 160 in which the second insert 152 is located includes a wing to partition the low pressure impingement space 176 therebetween. A pair of second shield dams 166 are provided that extend in the height direction. In the illustrated form, a pair of second seal dams 166 are formed on the wall surface 190 of the front partition wall 154 and the inner wall surface 130 of the front outer wall portion 158, respectively. 14 and 16, the number of seal dams separating the high pressure impingement space 134 from the first insert 114 is one pair and the low pressure impingement space 176 from the second insert 152. The number of shield dams partitioning is one pair.

In some embodiments, the second insert 152 includes a pair of ribs 182 on the outer peripheral surface 132 of the second insert 152, such as shown in FIG. It extends along the direction and is fitted to a pair of second shield dams 166 respectively. In the illustrated exemplary form, one of the pair of second seal dams 166 is a protrusion 204 extending in the wing height direction formed on the wall surface 190 of the front partition wall 154, and the top surface of the protrusion 204 is formed with a groove 206 extending in the blade height direction, and the tip of one of the pair of ribs 182 engages with the groove 206 . As shown in FIG. 20B , the other of the pair of second seal dams 166 is a protrusion 204 formed on the inner wall surface 130 of the front outer wall 158 and extending in the blade height direction. A groove portion 206 extending in the blade height direction is formed in the groove portion 206 , and the tip of the other rib 182 of the pair of ribs 182 engages with the groove portion 206 .

According to this configuration, the combination of the protrusions 204 of the pair of second seal dams 166 and the ribs 182 inserted into the grooves 206 formed in the protrusions 204 allows the pair of second seal dams 166 to be positioned on both sides of the pair of second seal dams 166 . A leak flow between the high-pressure impingement space 134 and the low-pressure impingement space 176 can be suppressed with a simple configuration.

Such a configuration enables effective impingement cooling of the blade body 70 in the high-pressure impingement space 134 .

In some embodiments, for example, as shown in FIG. 14, the opening forming portion 118 of the first insert 114 and the impingement wall portion 140 are adjacent to each other with the rib 178 therebetween.

According to this configuration, the cooling medium discharged to the external space 94 of the blade body 70 through the purge hole 87 and the cooling medium jetted into the high-pressure impingement space 134 through the impingement holes 138 cause the stationary blade 23A to be cooled. The wing body 70 can be effectively cooled.

Next, the structure of another embodiment which combines one insert shown in FIG. 21 and a plurality of pairs of seal dams to partition the airfoil space will be described below. FIG. 21 is an enlarged cross-sectional view showing an example of another embodiment of the specific configuration of the stationary blade 23A shown in FIG. FIG. 22 is a diagram showing the EE cross section in FIG.

In some embodiments, as shown in FIGS. 21 and 22, the front space 160 is provided with one third insert 164 and multiple pairs of third seal dams 168a, 168b, and the space of the front space 160 is divided into five spaces. It is a structure that divides into That is, between the outer peripheral surface 132 of the third insert 164 near the upstream end 82 and the inner wall surface 130 of the wing body 70, a pair of third seal dams 168a are arranged so that the outer peripheral surface 132 of the third insert 164 A pair of third seal dams 168b are arranged between the inner wall surface 130 of the wing body 70 or the wall surface 190 of the front partition wall 154 and closer to the trailing edge 76 side than the pair of third seal dams 168a. .

In some embodiments, the wing-body 70 includes at least one side purge hole 89 circumferentially adjacent to the purge hole 87, for example, as shown in FIGS. In the illustrated form, the blade body 70 includes a pair of side purge holes 89 provided on both sides of the purge hole 87 in the circumferential direction.

According to such a configuration, the cooling medium is discharged from the side purge holes 89 to the external space 94 of the blade body 70 , thereby suppressing the backflow of the combustion gas to the internal space 197 of the blade body 70 , thereby reducing By lowering the temperature of the combustion gas, damage to the upstream end portion 82 of the blade body 70 can be suppressed.

In some embodiments, for example, as shown in FIGS. 21 and 22, the stationary blade 23A has a cylindrical third insert 164 extending in the blade height direction in the internal space 197 of the blade body 70. Prepare. The third insert 164 includes an opening forming portion 118 having an opening 116 extending along the blade height direction, the opening 116 being arranged opposite the purge hole 87 .

According to this configuration, the high-pressure cooling medium that has flowed out of the third insert 164 from the internal space 197 of the third insert 164 through the opening 116 of the opening forming portion 118 is smoothly discharged from the purge hole 87 to the counterpart member 55 . can be directly supplied to the gap 95 between the As a result, the protrusions 86 of the stationary blade 23A can be cooled and thermal damage to the protrusions 86 can be prevented.

In some embodiments, for example, as shown in FIGS. 21 and 22, the convex portion 86 extends along the blade height direction from the distal end portion 107 side of the blade body 70 to the proximal end portion 105 side to provide a purge gas. The holes 87 include a plurality of purge holes 87 arranged from the tip portion 107 to the base end portion 105 of the blade body 70 along the blade height direction. The convex portion 86 may extend over half or more of the length between the base end portion 105 and the tip portion 107 of the stationary blade 23A in the blade height direction.

According to this configuration, the cooling medium is discharged from the plurality of purge holes 87 arranged from the distal end portion 107 to the proximal end portion 105 of the blade body 70 into the external space 94 of the blade body 70 , thereby restoring the internal space of the blade body 70 . By suppressing the reverse flow of the combustion gas to 197, the temperature of the combustion gas outside the purge hole 87 can be lowered, and damage to the upstream end portion 82 of the blade body 70 can be suppressed.

Due to the configuration of the plurality of pairs of third seal dams 168a, 168b described above, the front space 160 in which the third insert 164 is arranged consists of the internal space 197, the communication space 172, the high pressure impingement spaces 134a, 134b, and the low pressure impingement space 176. Divided into 5 spaces. However, the internal space 197 and the communication space 172 are spaces having a pressure close to the supply pressure of the cooling medium having almost the same pressure.

Also in this embodiment, impingement holes 195 are formed in the impingement wall portion 199 of the third insert 164, as shown in FIG. The high-pressure impingement spaces 134a and 134b formed between the outer peripheral surface 132 of the third insert 164 and the inner wall surface 130 of the wing body 70 are the impingement holes 195a formed in the impingement wall portion 199 of the third insert 164. , communicates with the internal space 197 of the third insert 164 via the . The low-pressure impingement space 176 formed between the outer peripheral surface 132 of the third insert 164 and the inner wall surface 130 of the wing body 70 is the impingement hole 195b formed in the impingement wall portion 199 of the third insert 164. , communicates with the internal space 197 of the third insert 164 via the . In addition, a plurality of cooling holes 136 are formed on the pressure surface wall portion 73 side and the suction surface wall portion 75 side of the blade body 70, and a high pressure impingement space 134, a low pressure impingement space 176, and combustion on the blade surfaces 67, 69 side. The gas space communicates through cooling holes 136 .

In the case of this embodiment, the cooling medium supplied to the high pressure impingement spaces 134a and 134b and the low pressure impingement space 176 is supplied from the internal space 197 which is one common space. On the other hand, the high pressure impingement space 134a is arranged on the suction surface wall portion 75 side close to the upstream end portion 82, and the high pressure impingement space 134b is arranged on the pressure surface wall portion 73 side close to the upstream end portion 82. . The low-pressure impingement space 176 is located on the negative pressure surface wall portion 75 side and closer to the trailing edge 76 side than the high-pressure impingement space 134a.

The position of the third seal dam 168b that separates the high pressure impingement space 134a on the side of the suction surface wall 75 from the low pressure impingement space 176 corresponds to the position where the pressure of the combustion gas flow on the side of the suction surface wall 75 drops sharply. there is That is, the pressure of the combustion gas flow on the side of the suction surface wall portion 75 near the upstream end portion 82 on the upstream side in the axial direction from the position of the third seal dam 168b is It is roughly the same as the pressure of the stream. Further, the pressure of the combustion gas on the suction surface wall portion 75 side decreases from the position of the third seal dam 168b on the suction surface wall portion 75 side toward the trailing edge 76 .

Therefore, in the front space 160 , the internal space 197 and the communication space 172 have the highest pressure, and the high pressure impingement spaces 134 a and 134 b have lower pressure than the internal space 197 . In addition, the specifications of the impingement holes 195 of the impingement wall portion 199 are selected so that the pressure of the low-pressure impingement space 176 is lower than that of the high-pressure impingement spaces 134a and 134b and the pressure of the front-side space 160 is the lowest. There is a need to. That is, the front space 160 is divided into three pressure regions. Therefore, in order to equalize the flow rate of the coolant discharged into the combustion gas from the high-pressure impingement spaces 134a, 134b and the low-pressure impingement space 176 through the plurality of cooling holes 136 formed in the blade body 70, The specifications of the impingement holes 195 formed in the impingement wall portion 199 facing each impingement space are changed according to the pressure change in each impingement space, and each impingement space and the suction surface wall portion 75 or the pressure surface wall portion are changed. It is desirable to equalize the pressure difference with 73 . That is, since the high-pressure impingement spaces 134a and 134b have substantially the same pressure, the impingement holes 195a having the same specifications can be selected. However, on the blade surface 69 on the suction surface wall portion 75 side, the pressure drop of the combustion gas flow is large. It is desirable to select a specification different from that of the mounting hole 195 to restrict the flow rate of the cooling medium. The above configuration reduces consumption of excessive cooling medium.

The third insert 164 applied to this embodiment has substantially the same structure as the first insert 114 shown in FIGS. 10 and 11, for example. As described above, the opening forming portion 118 in which the opening 116 is formed is arranged on the outer peripheral surface 132 of the third insert 164 facing the position of the purge hole 87 formed in the convex portion 86 of the upstream end portion 82 . ing. An insert collar 167 corresponding to the insert collar 124 of the first insert 114 is formed at the end of the third insert 164 on the base end 105 side. The insert collar 167 is fixed and integrated with the outer peripheral end of the base end portion 105 of the cylindrical body 120 of the third insert 164 by welding or the like. The outer peripheral end 167a of the insert collar 167 is fixed by welding or the like to the end faces of the blade body 70 and the front partition wall 154 on the base end portion 105 side in the blade height direction, as shown by the dashed line in FIG. An opening 122 is formed in the central portion of the end portion of the insert collar 167 on the base end portion 105 side. An insert collar 167 is attached to the wing body 70, the front partition wall 154, and the end face of the base end 105 side of the third insert 164 in the blade height direction. , both radial ends of the high pressure impingement spaces 134a, 134b and the low pressure impingement space 176 are closed.

A notch portion 126 is formed at an end portion of the insert collar 167 corresponding to the position of the opening forming portion 118 in the wing height direction. 21 and 22, the cooling medium supplied to the third insert 164 is supplied from one direction (single-side supply structure) on the side of the base end 105, and the insert collar 167 is arranged on the side of the base end 105. It is a structured structure.

According to the above configuration, since the insert collar 167 has the notch portion 126 in the portion corresponding to the position of the opening 116 of the opening forming portion 118 in the circumferential direction, part of the cooling medium supplied from the outside is cut into the blade body 70. Since the coolant can be directly supplied to the gap between the inner wall surface 130 and the outer peripheral surface 132 of the third insert 164, the high-pressure coolant can be supplied to the purge hole 87 while minimizing the pressure loss of the coolant.

As described above, in the front space 160, the space between the inner wall surface 130 of the wing body 70 and the wall surface 190 of the front partition wall 154 and the outer peripheral surface 132 of the third insert 164 includes the plurality of pairs of third seal dams 168a, 168b. , and is divided into three spaces with different pressures, the communication space 172, the high pressure impingement spaces 134a and 134b, and the low pressure impingement space 176, as described above. The configuration of the multiple pairs of third shield dams 168a, 168b is shown in FIG. The plurality of pairs of third seal dams 168a and 168b have the same structure as the pair of first seal dams 142 shown in FIG. 202. A pair of ribs 178 having the same structure as shown in FIG. 13 are fixed to the outer peripheral surface 132 of the third insert 164 . The communication space 172, the high-pressure impingement spaces 134a and 134b, and the low-pressure impingement space 176 are sealed by a combination of the grooves 202 formed in the protrusion 200 and the ribs 178, forming a pair of third seal dams 168a, A leakage flow of the cooling medium generated between the spaces via 168b is suppressed.

FIG. 23 is a schematic perspective view showing the configuration of the outer shroud 62 according to one embodiment.
As shown in FIG. 23 , the outer shroud 62 includes a plate-like portion 208 that constitutes a flow passage wall, and a front edge side of the plate-like portion 208 that protrudes radially outward from the plate-like portion 208 , and that the gas turbine 1 is stationary. and a leading edge outer hook portion 210 that is secured to a wall (not shown). The front edge outer hook portion 210 extends in the circumferential direction. The leading edge outer hook portion 210 is formed with a lightening portion 212 that is concave toward the upstream side, and the bottom surface 214 of the lightening portion 212 is formed on the inner wall surface 130 of the upstream end portion 82 of the wing body 70 . It is connected and formed on the same plane as the inner wall surface 130 .

According to such a configuration, it is possible to smoothly insert the first insert 114 into the wing body 70 while suppressing interference by the leading edge outer hook portion 210 .

Although the embodiments of the present invention have been described above, the present invention is not limited to the above-described embodiments, and includes modifications of the above-described embodiments and modes in which these modes are combined as appropriate.

For example, in some embodiments shown in FIGS. 4 and 5, the upstream end 82 of the wing body 70 has a pair of flat plate portions 148. However, as shown in FIG. The end portion 82 may not have a pair of flat plate portions. In the form shown in FIG. 24 , side wall surfaces 88 on both sides in the circumferential direction of the convex portion 86 of the upstream end portion 82 are connected to smoothly curved blade surfaces 67 and 69 , respectively. Combustion oscillation can be similarly suppressed in such a configuration. In the embodiment shown in FIG. 24, the downstream end surface 55b of the mating member 55 is formed into a curved surface shape in accordance with the shape of the blade surface.

The embodiment shown in FIG. 24 has a different shape of the upstream end 82 compared to the embodiment shown in FIGS. It has the same structure as , and the action and effect due to that structure are also the same.

In another embodiment, for example, as shown in FIG. 25, the pair of flat plate portions 148 may be formed along oblique directions intersecting the axial direction and the circumferential direction. In the form shown in FIG. 25, each of the pair of flat plate portions 148 is inclined away from the purge hole 87 toward the downstream side in the axial direction. Combustion oscillation can be similarly suppressed in such a configuration. Further, machining of the upstream end portion 82 of the wing body 70 and machining of the downstream end face 55b of the mating member 55 are easy.

The embodiment shown in FIG. 25 differs from the embodiment shown in FIGS. 4 to 22 in the shape of the flat plate portion 148 of the upstream end 82, but the other structures are the same as in the embodiment shown in FIGS. It has the same structure as the form, and the action and effect due to that structure are also the same.

As used herein, expressions such as "in a certain direction", "along a certain direction", "parallel", "perpendicular", "center", "concentric" or "coaxial", etc. express relative or absolute arrangements. represents not only such arrangement strictly, but also the state of being relatively displaced with a tolerance or an angle or distance to the extent that the same function can be obtained.
For example, expressions such as "identical", "equal", and "homogeneous", which express that things are in the same state, not only express the state of being strictly equal, but also have tolerances or differences to the extent that the same function can be obtained. It shall also represent the existing state.
Further, in this specification, expressions representing shapes such as a quadrilateral shape and a cylindrical shape not only represent shapes such as a quadrilateral shape and a cylindrical shape in a geometrically strict sense, but also within the range in which the same effect can be obtained. , a shape including an uneven portion, a chamfered portion, and the like.
Moreover, in this specification, the expressions “comprising”, “including”, or “having” one component are not exclusive expressions excluding the presence of other components.

1 gas turbine 2 compressor 4 combustor 4a downstream end 6 turbine 8 rotor 10 compressor casing 12 inlet 16 stator vanes 16, 24 stator vanes 18, 26 rotor blade 20 casing 22 turbine casing 23 (23A, 23B) stator vane (1st stage stationary blade)
28 combustion gas flow path 30 exhaust chamber 50 transition piece 52 outlet portion 54 radial wall portion 54a downstream end 55 mating member 56 circumferential wall portion 56a downstream end 60 inner shroud 62 outer shroud 67, 69 blade surface 70 blade body 73 pressure surface wall portion 75 suction surface wall portion 76 trailing edge 82 upstream end portion 84 recessed portion 86 convex portion 87 purge hole 88 side wall surface 89 side purge hole 94 outer space 95 gaps 200, 204 protruding portion 100 combustor assembly 105 base end portion 107 tip portions 109, 196, 197 inner space 112 outer wall surface 114 first inserts 116, 122 opening 118 opening forming portion 120 cylindrical body 121 lid portion 121a guide 123 reinforcing members 124, 167 insert collars 124a, 167a outer peripheral edge 125 shared insert collar 125a outer peripheral end 126 notch 130 inner wall surface 132 outer peripheral surface 133 inner peripheral surface 134 high pressure impingement space 136 cooling holes 138, 194, 195a, 195b impingement holes 140, 198, 199 impingement wall portion 142 first seal dam 148 flat plate Part 150 Partition wall 152 Second insert 154 Front partition wall 156 Outer wall 158 Front outer wall part 160 Front space 160a Space gap 162 Partition part 164 Third insert 166 Second seal dam 168 Third seal dam 170 Front space partition walls 178, 182 Rib 172 Communication Space 176 Low-pressure impingement space 190 Wall surface 202, 206 Groove 208 Plate-like portion 210 Front edge outer hook portion 212 Lightening portion 214 Bottom surface

Claims (17)

A stator vane for a gas turbine,
a hollow wing body ;
an inner shroud provided at the radially inner end of the wing body;
an outer shroud provided at the radially outer end of the blade;
with
an upstream end of the blade body in the axial direction of the gas turbine includes a convex portion projecting upstream from the inner shroud and the outer shroud in the axial direction;
A stationary blade of a gas turbine, wherein the convex portion includes at least one purge hole that communicates an external space and an internal space of the blade body. The convex portion extends from the tip end side to the base end side of the wing body along the wing height direction,
2. The gas turbine stator vane according to claim 1, wherein said purge holes include a plurality of purge holes arranged along said blade height direction from said tip side to said base end side. 3. The gas turbine stator vane according to claim 1, wherein said blade body includes at least one side purge hole adjacent to said purge hole in the circumferential direction of said gas turbine. A stator vane for a gas turbine,
Equipped with a hollow wing body,
the upstream end of the blade body in the axial direction of the gas turbine includes a convex portion projecting upstream in the axial direction or a concave portion recessed downstream in the axial direction;
the concave portion or the convex portion includes at least one purge hole communicating between the outer space and the inner space of the blade;
The blade body includes a suction surface wall portion on the suction surface side connecting the projection or the recess and the trailing edge and connecting the projection or the recess and the trailing edge to the blade. a pressure surface wall portion on the pressure surface side, the surface forming a concave surface;
the outer wall surface of the upstream end portion includes a pair of flat plate portions extending to the suction surface wall portion and the pressure surface wall portion with the convex portion or the concave portion sandwiched in the circumferential direction;
The convex portion or the concave portion is
a first circumferential end continuous with one of the flat plate portions belonging to the negative pressure surface wall;
a second circumferential end portion connected to the other flat plate portion belonging to the pressure surface wall portion;
including
Gas turbine vanes. further comprising a cylindrical first insert extending along the blade height direction in the internal space of the blade body;
the first insert includes an opening forming portion having an opening extending along the wing height direction;
5. The gas turbine stator vane according to claim 4 , wherein said opening is arranged opposite said at least one purge hole. The first insert includes a cylindrical body and an insert collar provided at one end of the cylindrical body in the shape of a flange,
6. The gas turbine stator vane according to claim 5, wherein said insert collar has a notch portion arranged corresponding to the position of said opening forming portion. A first impingement space is formed between the inner wall surface of the wing body and the outer peripheral surface of the first insert,
A portion of the first insert excluding the opening forming portion includes an impingement wall portion formed with a plurality of impingement holes communicating between the internal space of the first insert and the first impingement space, A stator vane for a gas turbine according to claim 5 or 6. A pair of first seal dams extending in the blade height direction are formed on the inner wall surface of the upstream end of the blade body so as to separate the first impingement space from the first insert. ,
8. The gas turbine stator vane according to claim 7, wherein said pair of first seal dams are arranged to sandwich said purge hole in the circumferential direction of said gas turbine. the upstream end includes a flat plate portion extending in a circumferential direction of the gas turbine;
The convex portion or the concave portion is formed in the flat plate portion,
9. The gas turbine stator vane according to claim 8, wherein said pair of first seal dams are formed on inner wall surfaces of said flat plate portion. The first insert includes a pair of ribs arranged on both sides of the opening in the circumferential direction on the outer peripheral surface of the first insert,
10. The gas turbine stator vane according to claim 8, wherein said pair of ribs extend along said blade height direction and are fitted into said pair of first seal dams respectively. 11. The gas turbine stator vane according to claim 10, wherein said opening forming portion of said first insert and said impingement wall portion are provided adjacent to each other with said rib interposed therebetween. further comprising a cylindrical second insert provided adjacent to the downstream side of the first insert in the axial direction of the gas turbine and extending along the blade height direction;
The blade body includes a suction surface wall portion on the suction surface side connecting the projection or the recess and the trailing edge and connecting the projection or the recess and the trailing edge to the blade. n (n is 1 or more) extending linearly so as to connect the pressure surface wall portion on the pressure surface side, the surface forming a concave surface, and the suction surface wall portion and the pressure surface wall portion inside the blade body; an integer of ) partition walls and
The first insert and the second insert include a front partition wall, which is the partition wall farthest from the trailing edge among the n partition walls, and the rear edge of the outer wall of the wing with the front partition wall interposed therebetween. provided in the front space defined by the front outer wall portion, which is the portion opposite to the
A second impingement space is formed between the inner wall surface of the wing body and the outer peripheral surface of the second insert,
The stator blade includes a partition that partitions the first impingement space and the second impingement space,
The number of seal dams partitioning the first impingement space with the first insert is one pair, and the number of seal dams partitioning the second impingement space with the second insert is one pair or less. Item 12. The gas turbine stator vane according to any one of Items 8 to 11. 13 . The gas turbine stator vane according to claim 12 , wherein the partition portion includes a partition wall that partitions the first insert and the second insert from the front side wall portion to the front side partition wall. 13. The gas turbine static generator according to claim 12, wherein said partition includes a pair of second seal dams extending in said blade height direction so as to partition said second impingement space from said second insert. wings. 15. The gas turbine stator vane according to claim 14, wherein the pair of second seal dams are formed on the inner wall surface of the front partition wall and the inner wall surface of the front outer wall portion. The second insert has a pair of ribs on the outer peripheral surface of the second insert,
16. The gas turbine stator vane according to claim 14, wherein said pair of ribs of said second insert extend along said blade height direction and are fitted into said pair of second seal dams respectively. a plurality of circumferentially arranged combustors having outlets including radial walls along the radial direction;
17. The stator vane according to any one of claims 1 to 16, positioned downstream of a pair of said radial wall portions facing each other among said outlet portions of said combustors adjacent in said circumferential direction;
A gas turbine with a

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