JP2023170597A - Turbine stationary blade, blade segment and gas turbine - Google Patents

Abstract

To improve the cooling efficiency at a shroud of a turbine stationary blade.SOLUTION: A turbine stationary blade in at least one embodiment of the present disclosure comprises an airfoil part, and a shroud provided on at least one of one side or the other side of the airfoil part in a blade height direction. The shroud has a recess formed on a face opposite to the airfoil part across a gas path face. The recess includes a first region covered with an impingement plate, and a second region not covered with an impingement plate. The shroud includes: a first opening formed in the first region; a first side passage formed from a front edge side toward a rear edge side at one first side end in a circumferential direction and connected to the first opening at its one end; a second opening formed in the second region; and a second side passage formed from a front edge side toward a rear edge side of the other second side end in the circumferential direction, and connected to the second opening at its end.SELECTED DRAWING: Figure 3

Description

The present disclosure relates to turbine vanes, blade segments, and gas turbines.

For example, it is known that in a turbine stationary blade of a gas turbine, a shroud is provided on the outer and inner sides of the airfoil portion in the blade height direction. The surface of the shroud on which the airfoil is provided is a gas path surface and is exposed to hot combustion gases. Therefore, the shroud is designed to perform impingement cooling with cooling air on the opposite side of the airfoil across the gas path surface. By distributing cooling air after impingement cooling the shroud through a cooling passage provided in the shroud, the cooling air is efficiently used for cooling the shroud (see, for example, Patent Document 1).

Patent No. 6763636

From the perspective of improving gas turbine efficiency, it is desirable to use cooling air more effectively to cool the shroud more efficiently.

In view of the above circumstances, at least one embodiment of the present disclosure aims to improve cooling efficiency in a shroud of a turbine stator blade.

(1) A turbine stator blade according to at least one embodiment of the present disclosure includes:
an airfoil;
a shroud provided on at least one side or the other side of the airfoil in the blade height direction;
Equipped with
The shroud has a recess formed on a surface opposite to the airfoil portion across the gas path surface,
The recess includes a first region covered with an impingement plate and a second region not covered with the impingement plate,
The shroud is
a first opening formed in the first region;
a first side passage formed from a front edge side to a rear edge side at one first side end in the circumferential direction, and one end connected to the first opening;
a second opening formed in the second region;
a second side passage formed from the leading edge side to the trailing edge side at the other second side end in the circumferential direction, and having one end connected to the second opening;
has.

(2) A wing segment according to at least one embodiment of the present disclosure,
comprising a first segment including a turbine stationary blade having the configuration of (1) above, and a second segment;
The second side end of the first segment and the first side end of the second segment are bolted together.

(3) The gas turbine according to at least one embodiment of the present disclosure includes:
A compressor that compresses the air taken in,
a combustor that supplies fuel to and burns the compressed air compressed by the compressor;
a turbine that obtains rotational power from the combustion gas burned in the combustor;
has
The turbine includes turbine stationary blades having the configuration described in (1) above.

(4) The gas turbine according to at least one embodiment of the present disclosure includes:
A compressor that compresses the air taken in,
a combustor that supplies fuel to and burns the compressed air compressed by the compressor;
a turbine that obtains rotational power from the combustion gas burned in the combustor;
has
The turbine includes a blade segment having the configuration described in (2) above.

According to at least one embodiment of the present disclosure, cooling efficiency in a shroud of a turbine stator blade can be improved.

1 is a schematic diagram showing the overall configuration of a gas turbine as an example of a rotating machine. FIG. 3 is a cross-sectional view showing a gas flow path of the turbine. FIG. 2 is a radially inner view of a blade segment composed of two turbine stationary blades according to one embodiment. 4 is a sectional view taken along the line IV-IV in FIG. 3. FIG. FIG. 4 is a schematic cross-sectional view taken along arrow V in FIG. 3;

Hereinafter, some embodiments of the present disclosure will be described with reference to the accompanying drawings. However, the dimensions, materials, shapes, relative arrangements, etc. of the components described as embodiments or shown in the drawings are not intended to limit the scope of the present disclosure, and are merely illustrative examples. do not have.
For example, expressions expressing relative or absolute positioning such as "in a certain direction,""along a certain direction,""parallel,""orthogonal,""centered,""concentric," or "coaxial" are strictly In addition to representing such an arrangement, it also represents a state in which they are relatively displaced with a tolerance or an angle or distance that allows the same function to be obtained.
For example, expressions such as "same,""equal," and "homogeneous" that indicate that things are in an equal state do not only mean that things are exactly equal, but also have tolerances or differences in the degree to which the same function can be obtained. It also represents the existing state.
For example, expressions expressing shapes such as squares and cylinders do not only refer to shapes such as squares and cylinders in a strict geometric sense, but also include uneven parts and chamfers to the extent that the same effect can be obtained. Shapes including parts, etc. shall also be expressed.
On the other hand, the expressions "comprising,""comprising,""comprising,""containing," or "having" one component are not exclusive expressions that exclude the presence of other components.

FIG. 1 is a schematic diagram showing the overall configuration of a gas turbine as an example of a rotating machine, and FIG. 2 is a sectional view showing a gas flow path of the turbine.

In this embodiment, as shown in FIG. 1, the gas turbine 10 includes a compressor 11, a combustor 12, and a turbine 13 arranged coaxially with a rotor 14, and a generator 15 at one end of the rotor 14. connected. In the following description, the direction in which the axis of the rotor 14 extends is referred to as an axial direction Da, the circumferential direction centered on the axis of the rotor 14 is referred to as a circumferential direction Dc, and the direction perpendicular to the axis Ax of the rotor 14 is referred to as a radial direction. Dr. Note that the radial direction Dr is referred to as the blade height direction.

The compressor 11 generates high-temperature, high-pressure compressed air AC by compressing air AI taken in from an air intake port through a plurality of stationary blades and rotor blades. The combustor 12 supplies a predetermined fuel FL to the compressed air AC and burns it to generate high temperature and high pressure combustion gas FG. The turbine 13 drives and rotates a rotor 14 by passing high-temperature, high-pressure combustion gas FG generated in the combustor 12 through a plurality of stator blades and rotor blades, and drives a generator 15 connected to the rotor 14. do.

Further, as shown in FIG. 2, in the turbine 13, the turbine stationary blade (stator vane) 21 is configured such that the hub side of the airfoil portion 23 is fixed to the inner shroud 25, and the tip side is fixed to the outer shroud 27. There is. The turbine rotor blade (rotor blade) 41 is configured such that a base end portion of an airfoil portion 43 is fixed to a platform 45 . The outer shroud 27 and a split ring 51 disposed on the tip side of the rotor blade 41 are supported by the casing (turbine casing) 30 via a heat shield ring 53, and the inner shroud 25 is supported by the support ring 31. has been done. Therefore, the combustion gas flow path 32 through which the combustion gas FG passes is formed along the axial direction Da as a space surrounded by the inner shroud 25, the outer shroud 27, the platform 45, and the split ring 51.

FIG. 3 is a view of a blade segment 100 composed of two turbine stationary blades 21 according to an embodiment, viewed from inside the radial direction Dr.
FIG. 4 is a sectional view taken along the line IV-IV in FIG. 3.
FIG. 5 is a schematic cross-sectional view taken along the V arrow in FIG. 3, and shows one turbine stationary blade 21. As shown in FIG.
The turbine stationary blade 21 according to one embodiment constitutes a segment 101 in which one airfoil 23 is arranged for one outer shroud 27 and one inner shroud 25. In one embodiment, one wing segment 100 is formed by connecting two segments 101 with a bolted connection.

As shown in FIG. 3, the airfoil portion 23 is formed from a ventral wing surface 23c formed of a concave pressure surface and a dorsal wing surface 23d formed of a convex surface that is a suction surface. The ventral wing surface 23c and the dorsal wing surface 23d are connected by an upstream leading edge 23a and a downstream trailing edge 23b, forming an integrated airfoil portion 23.

Note that the inner shroud 25 and the outer shroud 27 function as gas path surface forming members. The gas path surface forming member has gas path surfaces 2a (25a, 27a) that partition the combustion gas flow path 32 and come into contact with the combustion gas FG. When there is no need to particularly distinguish between the inner shroud 25 and the outer shroud 27, the inner shroud 25 and the outer shroud 27 may be simply referred to as the shroud 2.

(About wing segment 100)
Wing segment 100 according to one embodiment includes two segments 101 that are bolted together, as described above. In the following description, for convenience of explanation, the segment 101 arranged such that the dorsal wing surface 23d faces the ventral wing surface 23c of the partner segment 101 will be referred to as the first segment 101A, and the ventral wing surface 23c will be referred to as the first segment 101A. The segment 101 disposed so as to face the dorsal wing surface 23d of the second segment 101 is referred to as a second segment 101B. In FIG. 3, the segment 101 on the left side of the drawing is the first segment 101A, and the segment on the right side of the drawing is the second segment 101B.
Therefore, FIGS. 4 and 5 each represent the first segment 101A.

The first segment 101A includes one first airfoil portion 23A and a first outer side provided at the tip side of the first airfoil portion 23A, that is, the outer end portion in the blade height direction (the outer end portion in the radial direction Dr). It includes a shroud 27A and a first inner shroud 25A provided on the hub side of the first airfoil portion 23A, that is, the inner end in the blade height direction (the inner end in the radial direction Dr).
Similarly, the second segment 101B includes one second airfoil 23B, a second outer shroud (not shown) provided on the distal end side of the second airfoil 23B, and a second outer shroud provided on the hub side of the second airfoil 23B. a second inner shroud 25B provided therein.
In addition, in the following description, there will be cases in which a configuration common to the first segment 101A and the second segment 101B will be explained, cases in which there is no need to make a particular distinction between the first segment 101A and the second segment 101B, and cases in which the first segment 101A and the second segment 101B will be explained. When referring to the second segment 101B generically, the letter B at the end of the code is omitted.

As shown in FIG. 3, the stator vane 21 according to some embodiments is arranged, for example, in the inner shroud 25 in the radial direction Dr inward on the opposite side of the airfoil portion 23 across the gas path surface 25a (see FIG. 4). A leading edge side retainer 61 and a trailing edge side retainer 63 are provided that extend toward the front edge. The leading edge retainer 61 is formed on the leading edge 23a side of the airfoil portion 23, and the trailing edge retainer 63 is formed on the trailing edge 23b side of the airfoil portion 23. The leading edge side retainer 61 and the trailing edge side retainer 63 are attached to the vehicle interior 30 via the support ring 31 (see FIG. 2).

Hereinafter, a cooling air passage in the inner shroud 25 of the stator vane 21 according to some embodiments will be described.
The inner shroud 25 of the stator vane 21 according to one embodiment has an inner region 255, which is a space in which cooling air CA supplied from the outside can be stored, inside the radial direction Dr, which is the surface opposite to the gas path surface 25a. Be prepared. The inner region 255 includes a peripheral edge of the inner shroud 25, that is, a first side end 251 on the ventral wing surface 23c side and a second side end on the dorsal wing surface 23d side, which form both ends of the inner shroud 25 in the circumferential direction Dc. 252, a front edge end portion 253 on the front edge 23a side in the axial direction Da, and a rear edge end portion 254 on the rear edge 23b side, and a space portion that is a concave portion recessed inward in the radial direction Dr. 257 and an impingement space 256 (described later). The inner region bottom surface 255a forming the bottom surface of the inner region 255 forms the inner surface in the radial direction Dr of the gas path surface 25a. That is, the space portion 257 and the impingement space 256 include an inner region bottom surface 255a, a first side end portion 251, which is an outer wall portion extending in the blade height direction (radial direction) from the inner region bottom surface 255a, and a second side end portion 251. This is a space formed from an end 252, a front edge 253, and a rear edge 254.
In the gas turbine 10 according to one embodiment, the cooling air CA is supplied to the space 257 from the outside in the radial direction of the stationary blade 21 through the through hole 24 that penetrates inside the airfoil portion 23 in the blade height direction. It is configured.

In the inner region 255, a region (first region R1) located on the ventral side of the airfoil portion 23 than the airfoil portion 23 when viewed from the blade height direction in the inner region bottom surface 255a is provided. An impingement plate 70 having a plurality of through holes 71 is arranged (see FIG. 5). In addition, in FIG. 3, the first region R1 is represented by hatching, and the illustration of the collision plate 70 is omitted.
A first region R1 of the inner region 255 forming the space 257 is divided by the collision plate 70 into a space 257 on the inside of the radial Dr and an impingement space 256 on the outside of the radial Dr. The space portion 257 and the impingement space 256 communicate with each other via the through hole 71 of the collision plate 70.
Of the inner region 255 forming the space 257, a region (second region R2) located on the dorsal side of the airfoil 23 than the airfoil 23 when viewed from the blade height direction is a region where the collision plate 70 Not covered with.

A part of the cooling air CA supplied to the space 257 is supplied to the impingement space 256 through the through hole 71, and performs impingement cooling (collision cooling) on the inner region bottom surface 255a of the first region R1. By impingement-cooling the inner region bottom surface 255a of the first region R1, overheating of the gas path surface 25a located on the ventral side of the airfoil portion 23 than the airfoil portion 23 due to the combustion gas is suppressed.
The cooling air CA after impingement-cooling the inner region bottom surface 255a of the first region R1 is supplied to the first side passage 131, which will be described later.

Further, a part of the cooling air CA supplied to the space 257 cools the inner region bottom surface 255a of the second region R2. By cooling the inner region bottom surface 255a of the second region R2, overheating of the gas path surface 25a located on the back side of the airfoil portion 23 rather than the airfoil portion 23 due to the combustion gas is suppressed.
The cooling air CA after cooling the inner region bottom surface 255a of the second region R2 is supplied to the second side passage 132, which will be described later.

A part of the cooling air CA supplied to the space 257 flows into the leading edge side flow passages 191 which are formed in plurality at intervals in the circumferential direction at the leading edge end part 253 and extend in the axial direction, and mainly flow into the leading edge side flow passages 191 that extend in the axial direction. The edge portion 253 is cooled. The cooling air CA flowing through the leading edge end 253 is discharged into the combustion gas passage 32 from an opening in the gas path surface 25a of the leading edge end 253.

The stationary blade 21 according to some embodiments includes an air passage 105 for causing the cooling air CA supplied to the space 257 to flow to the trailing edge end 254 of the inner shroud 25. The air passage 105 has a first side passage 131 , a second side passage 132 , a circumferential passage 135 , and a trailing edge passage 180 .
The first side passage 131 is an air passage formed from the leading edge 23a side to the trailing edge 23b side in the first side end 251 of the inner shroud 25. One end portion of the first side end portion 251 communicates with the impingement space 256 via the first opening portion 111 formed in the first region R1.
The other end of the first side passage 131 on the rear edge 23b side communicates with one end of a circumferential passage 135 in the circumferential direction Dc, which will be described later.

The second side passage 132 is an air passage formed from the leading edge 23a side to the trailing edge 23b side in the second side end 252 of the inner shroud 25. One end portion of the second side end portion 252 communicates with the space portion 257 via the second opening portion 112 formed in the second region R2.
The other end of the second side passage 132 on the rear edge 23b side communicates with the other end of the circumferential passage 135 in the circumferential direction Dc.

The circumferential passage 135 is a cooling passage extending in the circumferential direction Dc on the rear edge 23b side with respect to the space portion 257, and as described above, the circumferential passage 135 is a cooling passage that extends in the circumferential direction Dc on the rear edge 23b side of the first side passage 131 and the second end on the rear edge 23b side. The other end of the second side passage 132 on the rear edge 23b side is connected.

The trailing edge end passage 180 is a plurality of cooling passages arranged at intervals in the circumferential direction of the trailing edge end 254 , and has an upstream end 180 a connected to the circumferential passage 135 and a downstream end 180 b connected to the inner shroud 25 . It opens at the rear edge end surface 25c.

In the air passage 105 according to the embodiment configured in this way, a part of the cooling air CA in the space 257 is supplied from the first opening 111 to the first side passage 131 via the impingement space 256. A part of the cooling air CA in the space 257 is supplied to the second side passage 132 from the second opening 112.
The cooling air CA supplied to the first side passage 131 flows through the first side passage 131 from the leading edge 23 a side toward the trailing edge 23 b side, and mainly cools the first side end portion 251 .
The cooling air CA supplied to the second side passage 132 flows through the second side passage 132 from the leading edge 23 a side toward the trailing edge 23 b side, and mainly cools the second side end portion 252 .
The first opening 111 and the second opening 112 are arranged so that the cooling air CA flowing into the first side passage 131 and the second side passage 132 flows as long as possible along the axial direction Da. It is preferable to provide it near the front edge end 253.

The cooling air CA flowing through the first side passage 131 and the second side passage 132 toward the trailing edge 23b flows into the circumferential passage 135 from the first side passage 131 and the second side passage 132, It flows via circumferential passageway 135 into each of a plurality of trailing edge passageways 180 . The cooling air CA flowing into each of the plurality of trailing edge passages 180 flows from the upstream end 180a of the trailing edge passage 180 toward the trailing edge face 25c of the inner shroud 25, and mainly cools the trailing edge end 254. do. The cooling air CA is discharged into the combustion gas from the trailing edge end face 25c.

(Regarding the bolt connection between the first segment 101A and the second segment 101B)
In the first segment 101A of the wing segment 100 according to one embodiment, the second side end 252 on the dorsal wing surface 23d side of the first inner shroud 25A and the side end on the dorsal wing surface 23d side of the first outer shroud 27A A bolt hole 161 is formed to penetrate in the circumferential direction through the second side end portion 152, which is the second side end portion 152.
In the first segment 101A of the wing segment 100 according to one embodiment, one bolt hole 161 is formed in the second side end 252 of the first inner shroud 25A, as shown in FIG. A plurality of bolt holes 161 may be formed at intervals.
In the first segment 101A of the wing segment 100 according to one embodiment, a plurality of bolt holes 161 may be formed at intervals in the axial direction in the second side end 152 of the first outer shroud 27A. In the example shown in FIG. 4, there are three bolt holes 161, but there may be two or less, or four or more.

In the second segment 101B of the wing segment 100 according to one embodiment, the first side end portion 251 on the ventral wing surface 23c side of the second inner shroud 25B, and the first side end portion 251 on the ventral wing surface 23c side of the second outer shroud (not shown) A bolt hole (not shown) is formed to circumferentially penetrate a first side end (not shown) that is a side end.
In the second segment 101B of the wing segment 100 according to one embodiment, one bolt hole is preferably formed in the first side end 251 of the second inner shroud 25B. A plurality of bolt holes may be formed.
In the second segment 101B of the wing segment 100 according to one embodiment, a plurality of bolt holes may be formed in the first side end 151 of the second outer shroud 27B at intervals in the axial direction. As with the first outer shroud 27A, the number of bolt holes is preferably three, but may be two or less, or may be four or more.

The positions of the bolt hole 161 of the first segment 101A and the unillustrated bolt hole of the second segment 101B are set so that the bolt 171 can be inserted into the bolt hole 161 and the unillustrated bolt hole. There is.
In the wing segment 100 according to one embodiment, the first segment 101A and the second segment 101B are bolted together by inserting the bolt 171 into the bolt hole 161 and a bolt hole (not shown) and attaching the nut 172. .

In addition, in the turbine 13 of the gas turbine 10 according to one embodiment, a plurality of the blade segments 100 described above are arranged in the circumferential direction. Wing segments 100 adjacent in the circumferential direction are not bolted together. A seal plate (not shown) is disposed between the circumferentially adjacent blade segments 100 to prevent cooling air CA from leaking between the circumferentially adjacent blade segments 100.

In the turbine stationary blade 21 according to one embodiment, it is desirable that the cooling air CA that has cooled the first region R1 and the second region R2 is further used for cooling the inner shroud 25. In particular, since the cooling air CA after cooling the second region R2, which is a region where impingement cooling is not performed, has a relatively low temperature, it is desirable to effectively utilize it for cooling the inner shroud 25.

Therefore, the turbine stationary blade 21 according to one embodiment is configured as follows. That is, in the turbine stationary blade 21 according to one embodiment, the inner shroud 25 has the first opening 111 formed in the first region R1 and the leading edge 23a side at one first side end 251 in the circumferential direction Dc. A first side passage 131 is formed from the rear edge 23b to the rear edge 23b and has one end connected to the first opening 111. The inner shroud 25 is formed from the front edge 23a side to the rear edge 23b side at the second opening 112 formed in the second region R2 and the other second side end 252 in the circumferential direction Dc. 2. The second side passage 132 is connected to the second opening 112.

As a result, the cooling air CA after cooling the first region R1 where impingement cooling is performed and the cooling air CA after cooling the second region R2 where impingement cooling is not performed are further used to cool the inner shroud 25. Therefore, the cooling efficiency of the inner shroud 25 can be improved.

The gas turbine 10 according to one embodiment includes a compressor 11 that compresses taken-in air AI, a combustor 12 that supplies and combusts the compressed air AC compressed by the compressor 11, and a combustor 12 that combusts the compressed air AC. and a turbine 13 that obtains rotational power from the combustion gas FG. The turbine 13 includes the blade segment 100 (ie, the turbine stationary blade 21) according to the embodiment described above.

As a result, the cooling air CA after cooling the first region R1 where impingement cooling is performed and the cooling air CA after cooling the second region R2 where impingement cooling is not performed are further used to cool the inner shroud 25. Therefore, the cooling efficiency of the inner shroud 25 can be improved. Thereby, for example, excessive cooling and excessive use of cooling air CA can be suppressed, and the efficiency of the gas turbine 10 can be improved.

In the turbine stationary blade 21 according to one embodiment, the inner shroud 25 preferably has a circumferential passage 135 extending in the circumferential direction Dc closer to the trailing edge 23b than the first region R1. The other end of the first side passage 131 may be connected to one end of the circumferential passage 135, and the other end of the second side passage 132 may be connected to the other end of the circumferential passage 135.

Since cooling air CA from the second region R2 not covered by the collision plate 70 flows into the second side passage 132, cooling air CA from the first region R1 covered by the collision plate 70 flows into the second side passage 132. Compared to the first side passage 131, the pressure of the cooling air CA tends to be higher. Therefore, the flow rate of the cooling air CA flowing through the second side passage 132 tends to be large, and it is preferable to use the cooling air CA after flowing through the second side passage 132 to further cool the inner shroud 25. This is desirable from the point of view of efficient use.

According to the turbine stationary blade 21 according to one embodiment, the area around the circumferential passage 135 can be further cooled by the cooling air CA after flowing through the second side passage 132. Further, according to the turbine stationary blade 21 according to the embodiment, by further providing a cooling flow path (trailing edge end passage 180) on the downstream side of the circumferential passage 135, the first side passage 131 and the second side passage Since the cooling air CA that has flowed through the partial passage 132 can be made to flow through this cooling passage (trailing edge passage 180) to cool the inner shroud 25, the cooling air CA can be used efficiently.

In the turbine stationary blade 21 according to one embodiment, the inlet area S2 of the second opening 112 is preferably smaller than the inlet area S1 of the first opening 111.

As described above, since the cooling air CA from the second region R2 not covered by the collision plate 70 flows into the second side passage 132, the cooling air CA from the first region R1 covered by the collision plate 70 flows into the second side passage 132. The pressure of the cooling air CA tends to be higher than that in the first side passage 131 into which the air CA flows. Therefore, the amount of cooling air CA flowing into the circumferential passage 135 tends to be larger from the second side passage 132 than from the first side passage 131 .
According to the turbine stationary blade 21 according to the embodiment, since the flow rate of the cooling air CA flowing into the second side passage 132 can be suppressed by the second opening 112, the flow rate of the cooling air CA flowing into the second side passage 132 can be suppressed. can suppress the flow rate to an appropriate flow rate.

In the turbine stationary blade 21 according to one embodiment, a plurality of inner shrouds 25 are arranged in the circumferential direction Dc on the trailing edge 23b side, the upstream end 180a is connected to the circumferential passage 135, and the downstream end 180b is behind the inner shroud 25. It is preferable to have a trailing edge passage 180 that opens to the edge surface 25c.
Thereby, the cooling air CA after flowing through the first side passage 131 and the second side passage 132 is made to flow through the plurality of trailing edge end passages 180, and the area on the trailing edge 23b side of the inner shroud 25 (the trailing edge The end portion 254) can be cooled, and the cooling air CA can be used efficiently.

In the turbine stationary blade 21 according to one embodiment, the first region R1 may include a region located more ventral to the airfoil portion 23 than the airfoil portion 23 when viewed from the blade height direction. .

Since the combustion gas FG, which is the working fluid of the turbine 13, is a relatively high-temperature fluid, the combustion gas flow is higher in the region near the ventral side of the airfoil portion 23 in the combustion gas passage 32 than in the region near the back side. Since the flow velocity in the passage 32 tends to be higher, the temperature of the inner shroud 25 is also higher in the region located ventral to the airfoil 23 than in the airfoil 23. It tends to be higher than the ventrally located regions.
According to the turbine stationary blade 21 according to one embodiment, since the region where the temperature tends to become relatively high can be impingement-cooled, the temperature rise of the inner shroud 25 can be efficiently suppressed.

In the turbine stationary blade 21 according to one embodiment, the configuration regarding the air passage 105 described above is preferably provided in the inner shroud 25.
According to the turbine stationary blade 21 according to one embodiment, the first region R1 where impingement cooling is performed in the inner shroud 25 which is smaller in size than the outer shroud 27 and has a relatively small region in which the collision plate 70 is arranged. The cooling air CA after cooling the inner shroud 25 and the cooling air CA after cooling the second region R2 where impingement cooling is not performed can be used to further cool the inner shroud 25.

The blade segment 100 according to one embodiment includes a first segment 101A including a turbine stationary blade 21 according to one embodiment, and a second segment 101B. The second side end 252 of the first segment 101A and the first side end 251 of the second segment 101B are bolted together. Therefore, the second side end 252 of the first segment 101A and the first side end 251 of the second segment 101B are relatively difficult to cool because of the presence of bolts 171 for bolt connection.
According to the wing segment 100 according to one embodiment, the second side passage 132 is formed at the second side end 252 of the first segment 101A, and the first side passage 132 is formed at the first side end 251 of the second segment 101B. Since the passage 131 is formed, the second side end 252 of the first segment 101A and the first side end 251 of the second segment 101B, which are relatively difficult to cool, can be cooled.

In the wing segment 100 according to one embodiment, the first side passage 131 is formed by the bolts 171 and nuts 172 as fixing devices in the first side end 251 when viewed from the blade height direction of the airfoil 23. The second side passage 132 may overlap the connected portion in the axial direction Da, and the second side passage 132 is connected by a bolt 171 and a nut 172 as fixing devices in the second side end 252 when viewed from the blade height direction. It is preferable to overlap the portion shown in the axial direction Da.
Thereby, the second side end 252 of the first segment 101A and the first side end 251 of the second segment 101B, which are relatively difficult to cool, can be efficiently cooled.

The present disclosure is not limited to the embodiments described above, and also includes forms in which modifications are added to the embodiments described above, and forms in which these forms are appropriately combined.
For example, the configuration regarding the air passage 105 in the inner shroud 25 of the turbine stationary blade 21 of the embodiment described above may be provided in the outer shroud 27.

In the turbine stationary blade 21 of the embodiment described above, the cooling air CA that has entered the first side passage 131 and the second side passage 132 is flowed from the leading edge 23a side toward the trailing edge 23b side. The cooling air CA flowing into the first side passage 131 and the second side passage 132 may be made to flow from the trailing edge 23b side toward the leading edge 23a side.
In this case, the first opening 111 and the second opening 112 are arranged so that the cooling air CA flowing into the first side passage 131 and the second side passage 132 flows as long as possible along the axial direction Da. , is preferably provided near the trailing edge portion 254.
Further, in this case, it is preferable that the circumferential passage 135 is provided closer to the front edge 23a than the space 257, and the circumferential passage 135 is connected to the first side passage 131 and the second side passage 132.

In the turbine stationary blade 21 of the above-described embodiment, the cooling air CA flowing into the first side passage 131 can be controlled by appropriately setting the inlet area S1 of the first opening 111 and the inlet area S2 of the second opening 112. and the flow rate of the cooling air CA flowing into the second side passage 132 are adjusted as appropriate.
However, instead of setting the entrance area S1 of the first opening 111 and the entrance area S2 of the second opening 112 as appropriate, or By appropriately setting the area S2 and the passage cross-sectional area of the first side passage 131 and the second side passage 132, the flow rate of the cooling air CA flowing into the first side passage 131 and the second side passage The flow rate of the cooling air CA flowing into the side passage 132 may be adjusted as appropriate.

The contents described in each of the above embodiments can be understood as follows, for example.
(1) The turbine stationary blade 21 according to at least one embodiment of the present disclosure is provided on the airfoil portion 23 and at least one side or the other side of the airfoil portion 23 in the blade height direction (radial direction Dr). A shroud 2 is provided. The shroud 2 (inner shroud 25) has a concave portion (space portion 257) formed on a surface opposite to the airfoil portion 23 across the gas path surface 2a. The recessed portion (space portion 257) includes a first region R1 covered with an impingement plate (collision plate 70) and a second region R2 not covered with the impingement plate (collision plate 70). The shroud 2 (inner shroud 25) is formed from the front edge 23a side to the rear edge 23b side at the first opening 111 formed in the first region R1 and one first side end 251 in the circumferential direction Dc, The first side passage 131 has one end connected to the first opening 111. The shroud 2 (inner shroud 25) is formed from the front edge 23a side to the rear edge 23b side at the second opening 112 formed in the second region R2 and the other second side end 252 in the circumferential direction Dc, It has a second side passage 132 at one end connected to the second opening 112 .

According to the configuration (1) above, the cooling air CA after cooling the first region R1 where impingement cooling is performed and the cooling air CA after cooling the second region R2 where impingement cooling is not performed. can be further utilized for cooling the shroud 2 (inner shroud 25), so the cooling efficiency of the shroud 2 (inner shroud 25) can be improved.

(2) In some embodiments, in the configuration of (1) above, the shroud 2 (inner shroud 25) includes a circumferential passage 135 extending in the circumferential direction Dc on the rear edge 23b side of the first region R1; It is good to have. The other end of the first side passage 131 may be connected to one end of the circumferential passage 135, and the other end of the second side passage 132 may be connected to the other end of the circumferential passage 135.

According to the configuration (2) above, the area around the circumferential passage 135 can be further cooled by the cooling air CA after flowing through the second side passage 132. Further, according to the configuration (2) above, by further providing a cooling flow path (trailing edge end passage 180) on the downstream side of the circumferential passage 135, the first side passage 131 and the second side passage 132 The shroud 2 (inner shroud 25) can be cooled by allowing the cooling air CA that has flowed through the cooling flow path (trailing edge end passage 180) to cool the shroud 2 (inner shroud 25), so the cooling air CA can be used efficiently.

(3) In some embodiments, in the configuration of (2) above, the entrance area S2 of the second opening 112 may be smaller than the entrance area S1 of the first opening 111.

As described above, since the cooling air CA from the second region R2 that is not covered by the impingement plate (collision plate 70) flows into the second side passage 132, the cooling air CA flows into the second side passage 132. Compared to the first side passage 131 into which the cooling air CA from the covered first region R1 flows, the pressure of the cooling air CA tends to be higher. Therefore, the amount of cooling air CA flowing into the circumferential passage 135 tends to be larger from the second side passage 132 than from the first side passage 131 .
According to the configuration (3) above, since the flow rate of the cooling air CA flowing into the second side passage 132 can be suppressed by the second opening 112, the flow rate of the cooling air CA flowing into the second side passage 132 can be suppressed. The flow rate can be controlled to an appropriate level.

(4) In some embodiments, in the configuration of (2) or (3) above, a plurality of shrouds 2 (inner shrouds 25) are arranged in the circumferential direction Dc on the rear edge 23b side, and one end (upstream end 180a ) is connected to the circumferential passage 135, and the other end (downstream end 180b) preferably has a trailing edge end passage 180 that opens to the trailing edge end face 25c of the shroud 2 (inner shroud 25).

According to the configuration (4) above, the cooling air CA after flowing through the first side passage 131 and the second side passage 132 is made to flow through the plurality of trailing edge end passages 180 and the shroud 2 (inner shroud 25 ) can be cooled, and the cooling air CA can be used efficiently.

(5) In some embodiments, in any of the configurations (1) to (4) above, the first region R1 is located closer to the airfoil than the airfoil 23 when viewed from the blade height direction. It is preferable to include a region located on the ventral side of 23.

If the working fluid (combustion gas FG) of the turbine 13 is a high-temperature fluid, in the region near the ventral side of the airfoil portion 23 in the flow path (combustion gas flow path 32) of the working fluid (combustion gas FG), Since the flow velocity of the working fluid (combustion gas FG) tends to be higher than in the region near the dorsal side, the temperature of the shroud 2 (inner shroud 25) is also located closer to the ventral side of the airfoil section 23 than the airfoil section 23. The region located on the ventral side of the airfoil portion 23 tends to be higher than the region located on the ventral side of the airfoil portion 23.
According to the configuration (5) above, since it is possible to impingement-cool the region where the temperature tends to be relatively high, it is possible to efficiently suppress the temperature rise of the shroud 2 (inner shroud 25).

(6) In some embodiments, in any of the configurations (1) to (5) above, the shroud 2 (inner shroud 25) is provided inside the airfoil portion 23 in the blade height direction. Good too.

According to the configuration (6) above, in the inner shroud 25 which is smaller in size than the outer shroud 27 and has a relatively small area where the impingement plate (impingement plate 70) is arranged, the first The cooling air CA after cooling the region R1 and the cooling air CA after cooling the second region R2 where impingement cooling is not performed can be used to further cool the inner shroud 25.

(7) A blade segment according to at least one embodiment of the present disclosure includes a first segment 101A and a second segment 101B including the turbine stationary blade 21 having the configuration of any one of (1) to (6) above. The second side end 252 of the first segment 101A and the first side end 251 of the second segment 101B are bolted together.

The second side end 252 of the first segment 101A and the first side end 251 of the second segment 101B are relatively difficult to cool because of the presence of bolts 171 for bolt connection.
According to the configuration (7) above, the second side passage 132 is formed at the second side end 252 of the first segment 101A, and the first side passage 131 is formed at the first side end 251 of the second segment 101B. , the second side end 252 of the first segment 101A and the first side end 251 of the second segment 101B, which are relatively difficult to cool, can be cooled.

(8) In some embodiments, in the configuration of (7) above, the first side passage 131 is a fixture of the first side end 251 when viewed from the blade height direction of the airfoil 23. (Bolts 171 and nuts 172) may overlap in the axial direction Da, and the second side passage 132 is the fixing member (of the second side end 252) when viewed from the blade height direction. It is preferable that the portion overlap in the axial direction Da with the portion connected by the bolt 171 and nut 172).

According to the configuration (8) above, the second side end 252 of the first segment 101A and the first side end 251 of the second segment 101B, which are relatively difficult to cool, can be efficiently cooled.

(9) The gas turbine 10 according to at least one embodiment of the present disclosure includes a compressor 11 that compresses the taken-in air AI, and a combustor 12 that supplies fuel FL to the compressed air AC compressed by the compressor 11 and combusts it. and a turbine 13 that obtains rotational power from the combustion gas FG burned in the combustor 12. The turbine 13 includes a turbine stationary blade 21 having any of the configurations (1) to (6) above.

According to the configuration (9) above, the cooling air CA after cooling the first region R1 where impingement cooling is performed and the cooling air CA after cooling the second region R2 where impingement cooling is not performed. can be further utilized for cooling the shroud 2 (inner shroud 25), so the cooling efficiency of the shroud 2 (inner shroud 25) can be improved. Thereby, for example, excessive cooling and excessive use of cooling air CA can be suppressed, and the efficiency of the gas turbine 10 can be improved.

(10) The gas turbine 10 according to at least one embodiment of the present disclosure includes a compressor 11 that compresses taken-in air AI, and a combustor 12 that supplies fuel FL to the compressed air AC compressed by the compressor 11 and combusts it. and a turbine 13 that obtains rotational power from the combustion gas FG burned in the combustor 12. The turbine 13 includes the blade segments 100 having the configurations (7) to (8) above.

According to the configuration (10) above, the cooling air CA after cooling the first region R1 where impingement cooling is performed and the cooling air CA after cooling the second region R2 where impingement cooling is not performed. can be further utilized for cooling the shroud 2 (inner shroud 25), so the cooling efficiency of the shroud 2 (inner shroud 25) can be improved. Thereby, for example, excessive cooling and excessive use of cooling air CA can be suppressed, and the efficiency of the gas turbine 10 can be improved.

2 Shroud 2a Gas path surface 10 Gas turbine 11 Compressor 12 Combustor 13 Turbine 21 Turbine stator blade (Stator blade)
23 Airfoil section 23a Leading edge 23b Trailing edge 23c Ventral wing surface 23d Dorsal wing surface 25 Inner shroud 25a Gas path surface 27 Outer shroud 27a Gas path surface 32 Combustion gas flow path 70 Impingement plate
100 Wing segment 101 Segment 101A First segment 101B Second segment 105 Air passage 111 First opening 112 Second opening 131 First side passage 132 Second side passage 135 Circumferential passage 151 First side end 152 2nd side end 161 Bolt hole 171 Bolt 180 Rear edge end passage 251 1st side end 252 2nd side end 255 Inner region 255a Inner region bottom 257 Space (recess)

Claims (10)

an airfoil;
a shroud provided on at least one side or the other side of the airfoil in the blade height direction;
Equipped with
The shroud has a recess formed on a surface opposite to the airfoil portion across the gas path surface,
The recess includes a first region covered with an impingement plate having a plurality of through holes, and a second region not covered with the impingement plate,
The shroud is
a first opening formed in the first region;
a first side passage formed from a front edge side to a rear edge side at one first side end in the circumferential direction, and one end connected to the first opening;
a second opening formed in the second region;
a second side passage formed from the leading edge side to the trailing edge side at the other second side end in the circumferential direction, and having one end connected to the second opening;
has,
Turbine stationary blade. The shroud has a circumferential passage extending in the circumferential direction closer to the rear edge than the first region,
The other end of the first side passage is connected to one end of the circumferential passage,
The other end of the second side passage is connected to the other end of the circumferential passage.
The turbine stator blade according to claim 1. The entrance area of the second opening is smaller than the entrance area of the first opening.
The turbine stationary blade according to claim 2. The shroud has a plurality of trailing edge end passages arranged in the circumferential direction on the trailing edge side, one end connected to the circumferential passage, and the other end opening to the trailing edge end surface of the shroud.
The turbine stationary blade according to claim 2 or 3. The first region includes a region located more ventral to the airfoil than the airfoil when viewed from the blade height direction.
The turbine stationary blade according to any one of claims 1 to 3. The shroud is provided inside the airfoil in the blade height direction,
The turbine stationary blade according to any one of claims 1 to 3. A first segment including a turbine stationary blade according to any one of claims 1 to 3, and a second segment,
the second side end of the first segment and the first side end of the second segment are coupled with a fixture;
wing segment. The first side passage overlaps in the axial direction a portion of the first side end portion that is connected with the fixture when viewed from the blade height direction of the airfoil portion,
The second side passage overlaps in the axial direction a portion of the second side end portion that is connected with the fixture when viewed from the blade height direction.
A wing segment according to claim 7. A compressor that compresses the air taken in,
a combustor that supplies fuel to and burns the compressed air compressed by the compressor;
a turbine that obtains rotational power from the combustion gas burned in the combustor;
has
The turbine includes the turbine stationary blade according to any one of claims 1 to 3.
gas turbine. A compressor that compresses the air taken in,
a combustor that supplies fuel to and burns the compressed air compressed by the compressor;
a turbine that obtains rotational power from the combustion gas burned in the combustor;
has
The turbine comprises a blade segment according to claim 7.
gas turbine.

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