CN117090642A - Turbine stator blade, blade segment body and gas turbine - Google Patents

Abstract

A turbine vane, a blade segment body, and a gas turbine improve cooling efficiency in a shroud of the turbine vane. The turbine vane according to at least one embodiment of the present disclosure includes: wing-shaped parts; and a shroud provided on at least one of one side and the other side of the wing in the blade height direction. The shroud has a concave portion formed on a surface of the shroud opposite to the wing portion with the gas passage surface interposed therebetween. The recess includes: a first area covered by an impingement plate; and a second area not covered by the impingement plate. The shield has: a first opening formed in the first region; a first side passage formed from the leading edge side to the trailing edge side at a first side end in the circumferential direction of the shroud, the first side passage having one end connected to the first opening; a second opening formed in the second region; and 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 of the shroud, and having one end connected to the second opening.

Description

Turbine stator blade, blade segment body and gas turbine

Technical Field

The present disclosure relates to turbine vanes, blade segment bodies (segments), and gas turbines.

Background

For example, in a turbine vane of a gas turbine, a technique is known in which shrouds are provided on the outer and inner sides of an airfoil in the blade height direction. The surface of the shroud on which the wings are provided is a gas passage surface and is exposed to high temperature combustion gases. Therefore, the shroud is impingement cooled by the cooling air on the surface opposite to the airfoil across the gas passage surface. Then, cooling air after impingement cooling of the shroud is circulated through a cooling passage provided in the shroud, whereby the cooling air is efficiently used for cooling the shroud (for example, refer to patent document 1).

Prior art literature

Patent literature

Patent document 1: japanese patent No. 6763636

From the standpoint of improving the efficiency of the gas turbine, it is desirable to more effectively utilize the cooling air in order to more efficiently cool the shroud.

Disclosure of Invention

In view of the above, it is an object of at least one embodiment of the present disclosure to improve cooling efficiency in a shroud of a turbine vane.

(1) The turbine vane according to at least one embodiment of the present disclosure includes:

wing-shaped parts; and

a shield arranged on at least one of one side and the other side of the vane height direction of the wing part,

the shroud has a concave portion formed on a surface of the shroud opposite to the wing portion across a gas passage surface,

the recess includes: a first area covered by an impingement plate; and a second area not covered by the impingement plate,

the shield has:

a first opening formed in the first region;

a first side passage formed from a leading edge side to a trailing edge side at a first side end in a circumferential direction of the shroud, the first side passage having one end connected to the first opening;

a second opening formed in the second region; and

and a second side passage formed from the leading edge side to the trailing edge side at the other second side end portion in the circumferential direction of the shroud, and having one end portion connected to the second opening portion.

(2) The blade segment according to at least one embodiment of the present disclosure includes the first segment and the second segment of the turbine vane having the structure of (1) above,

the second side end of the first section body and the first side end of the second section body are bolted.

(3) The gas turbine according to at least one embodiment of the present disclosure includes:

a compressor for compressing the taken-in air;

a combustor that supplies fuel to the compressed air compressed by the compressor and combusts the fuel; and

and (2) a turbine that obtains rotational power from the combustion gas obtained by combustion in the combustor, the turbine including the turbine vane having the structure of (1) above.

(4) The gas turbine according to at least one embodiment of the present disclosure includes:

a compressor for compressing the taken-in air;

a combustor that supplies fuel to the compressed air compressed by the compressor and combusts the fuel; and

a turbine that obtains rotational power by combustion gas obtained by combustion in the combustor,

the turbine includes the blade segment body having the structure of (2) above.

Effects of the invention

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

Drawings

Fig. 1 is a schematic view showing an overall configuration of a gas turbine as an example of a rotary machine.

Fig. 2 is a cross-sectional view showing a gas flow path of the turbine.

Fig. 3 is a view of a blade segment body composed of two turbine vanes of an embodiment, viewed from the radially inner side.

Fig. 4 is an IV-IV view in cross-section of fig. 3.

Fig. 5 is a schematic V-view cross-section of fig. 3.

Reference numerals illustrate:

a shield;

2a. a gas passage surface;

gas turbine;

a compressor;

a burner;

a turbine;

turbine vanes (vanes);

wing-shaped part;

front edge;

trailing edge;

ventral blade surface;

back side blade face;

inner shield;

gas passage surface;

outside shield;

gas passage surface;

combustion gas flow path;

crash panel (impingement panel);

blade segment body;

segment body;

first segment;

second segment;

air passage;

first opening;

a second opening;

first side passage;

second side passage;

circumferential passages;

151. a first side end;

second side end;

161. bolt holes;

bolts;

trailing edge end passage;

251. the first side end;

second side end;

medial region;

inner zone floor;

257. space (recess).

Detailed Description

Several embodiments of the present invention will be described below with reference to the accompanying drawings. The dimensions, materials, shapes, relative arrangements, and the like of the constituent members described as the embodiments or shown in the drawings are not intended to limit the scope of the present invention thereto, but are merely illustrative examples.

For example, expressions such as "in a certain direction", "along a certain direction", "parallel", "orthogonal", "center", "concentric" or "coaxial" indicate relative or absolute arrangement, and indicate a state in which the relative or absolute arrangement is relatively displaced by an angle or distance having a tolerance or a degree that the same function can be obtained, as well as such an arrangement in a strict sense.

For example, the expressions "identical", "equal", and "homogeneous" and the like indicate states in which things are equal, and indicate not only exactly equal states but also states in which there are tolerances or differences in the degree to which the same function can be obtained.

For example, the expression of the shape such as a quadrangular shape and a cylindrical shape means not only the shape such as a quadrangular shape and a cylindrical shape in a geometrically strict sense, but also the shape including a concave-convex portion, a chamfer portion, and the like within a range where the same effect can be obtained.

On the other hand, the expression "comprising," "including," or "having" one component is not an exclusive expression excluding the presence of other components.

Fig. 1 is a schematic view showing the overall structure of a gas turbine as an example of a rotary machine, and fig. 2 is a cross-sectional view showing a gas flow path of the turbine.

In the present embodiment, as shown in fig. 1, the gas turbine 10 is configured by coaxially disposing a compressor 11, a combustor 12, and a turbine 13 with a rotor 14, and a generator 15 is connected to one end of the rotor 14. In the following description, the direction in which the axis of the rotor 14 extends is referred to as the axial direction Da, the circumferential direction with respect to the axial center of the rotor 14 is referred to as the circumferential direction Dc, and the direction perpendicular to the axis Ax of the rotor 14 is referred to as the radial direction Dr. The radial direction Dr is referred to as the blade height direction.

In the compressor 11, the air AI taken in from the air intake port is compressed by a plurality of vanes and blades, and high-temperature and high-pressure compressed air AC is generated. The combustor 12 supplies a predetermined fuel FL to the compressed air AC and burns the fuel, thereby generating a high-temperature and high-pressure combustion gas FG. The turbine 13 drives the rotor 14 to rotate by passing the high-temperature and high-pressure combustion gas FG generated by the combustor 12 through a plurality of vanes and blades, and drives the generator 15 connected to the rotor 14.

As shown in fig. 2, in the turbine 13, the turbine vane (vane) 21 is configured such that the hub side of the airfoil 23 is fixed to the inner shroud 25 and the tip side is fixed to the outer shroud 27. The turbine bucket (bucket) 41 is configured such that a base end portion of the airfoil 43 is fixed to the platform 45. The outer shroud 27 and the split ring 51 disposed on the tip end side of the bucket 41 are supported by the casing (turbine casing) 30 via the heat insulating ring 53, and the inner shroud 25 is supported by the support ring 31. Accordingly, 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 body 100 composed of two turbine vanes 21 according to an embodiment, as viewed from the inside in the radial direction Dr.

Fig. 4 is an IV-IV view in cross-section of fig. 3.

Fig. 5 is a schematic V-view cross-section of fig. 3, showing one turbine vane 21.

The turbine vane 21 of one embodiment constitutes a segment 101 in which one airfoil 23 is disposed with respect to one outer shroud 27 and one inner shroud 25. In one embodiment, one blade segment 100 is formed by joining two segments 101 by bolting.

As shown in fig. 3, the airfoil 23 is formed with an airfoil 23 formed by connecting and integrating an upstream-side front edge 23a and a downstream-side rear edge 23b in the axial direction Da, the airfoil 23 being formed by a concave surface as a pressure surface and a rear-side blade surface 23d formed by a convex surface as a negative pressure surface.

The inner shroud 25 and the outer shroud 27 function as gas passage surface forming members. The gas passage surface forming member is a member that defines the combustion gas flow path 32 and has a gas passage surface 2a (25 a, 27 a) that contacts the combustion gas FG. When the inner shroud 25 and the outer shroud 27 do not need to be particularly distinguished, the inner shroud 25 and the outer shroud 27 may be simply referred to as the shroud 2.

(with respect to blade segment body 100)

As described above, the blade segment body 100 of an embodiment includes two segment bodies 101 that are bolted together. In the following description, for convenience of explanation, the segment 101 in which the back-side blade surface 23d is arranged to face the back-side blade surface 23c of the opposite-side segment 101 is referred to as a first segment 101A, and the segment 101 in which the back-side blade surface 23c is arranged to face the back-side blade surface 23d of the opposite-side segment 101 is referred to as a second segment 101B. In fig. 3, the left-hand segment 101 is shown as a first segment 101A and the right-hand segment is shown as a second segment 101B.

Accordingly, fig. 4 and 5 are diagrams showing the first segment 101A, respectively.

The first segment 101A includes: a first wing 23A; a first outer shroud 27A provided at the tip end side of the first airfoil portion 23A, that is, at the outer end portion in the blade height direction (outer end portion in the radial direction Dr); and a first inner shroud 25A provided on the hub side of the first airfoil 23A, that is, on the inner end portion in the blade height direction (inner end portion in the radial direction Dr).

Similarly, the second segment 101B includes: a second wing 23B; a second outer shroud, shown in the drawing, provided on the front end side of the second wing 23B; and a second inner shroud 25B provided on the hub side of the second wing 23B.

In the following description, when the structure in which the first segment 101A and the second segment 101B are shared is described, it is not necessary to particularly distinguish between the first segment 101A and the second segment 101B, and when the first segment 101A and the second segment 101B are collectively referred to, the last letter of the reference numeral is omitted.

As shown in fig. 3, the vane 21 according to several embodiments includes, for example, a leading edge side retainer 61 and a trailing edge side retainer 63 in the inner shroud 25, and the leading edge side retainer 61 and the trailing edge side retainer 63 extend in the radial direction Dr on the opposite side of the airfoil 23 across the gas passage surface 25a (see fig. 4). The leading edge side holder 61 is formed on the leading edge 23a side of the airfoil 23, and the trailing edge side holder 63 is formed on the trailing edge 23b side of the airfoil 23. The leading edge side holder 61 and the trailing edge side holder 63 are attached to the casing 30 via the support ring 31 (see fig. 2).

The passages of cooling air in the inner shroud 25 of the vane 21 according to several embodiments will be described below.

The inner shroud 25 of the vane 21 according to the embodiment includes an inner region 255 that is a space capable of storing cooling air CA supplied from the outside, on the inner side in the radial direction Dr, which is a surface opposite to the gas passage surface 25a. The inner region 255 is a region surrounded by a peripheral edge portion of the inner shroud 25, that is, a first side end portion 251 on the side of the web-side blade surface 23c, a second side end portion 252 on the side of the back-side blade surface 23d, a leading edge end portion 253 on the leading edge 23a side in the axial direction Da, and a trailing edge end portion 254 on the side of the trailing edge 23b, which form concave portions 257 and impact spaces 256 (described later) recessed inward in the radial direction Dr. The inner region bottom surface 255a forming the bottom surface of the inner region 255 forms a surface inside the gas passage surface 25a in the radial direction Dr. That is, the space portion 257 and the impact space 256 are spaces formed by the inner region bottom surface 255a, and the first side end portion 251, the second side end portion 252, the leading edge end portion 253, and the trailing edge end portion 254, which are outer wall portions extending from the inner region bottom surface 255a in the blade height direction (radial direction Dr).

In the gas turbine 10 according to the embodiment, the cooling air CA is supplied from the outside in the radial direction Dr of the vane 21 to the space portion 257 via the through hole 24 penetrating the inside of the airfoil portion 23 in the blade height direction.

In the inner region 255, a collision plate (impact plate) 70 (see fig. 5) having a plurality of through holes 71 is disposed so as to cover a region (first region R1) located on the ventral side of the airfoil 23 from the airfoil 23 when viewed from the blade height direction within the inner region bottom surface 255a. In fig. 3, the first region R1 is hatched, and the description of the collision plate 70 is omitted.

The first region R1 in the inner region 255 forming the space portion 257 is divided by the collision plate 70 into a space portion 257 inside in the radial direction Dr and an impact space 256 outside in the radial direction Dr. The space portion 257 communicates with the impact space 256 via the through hole 71 of the collision plate 70.

When viewed from the blade height direction within the inner region 255 forming the space portion 257, a region (second region R2) located further to the back side of the airfoil portion 23 than the airfoil portion 23 is not covered with the impingement plate 70.

A part of the cooling air CA supplied to the space portion 257 is supplied to the impingement space 256 via the through hole 71, and impingement-cools (collision-cools) 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 by the combustion gas FG of the gas passage surface 25a located on the ventral side of the airfoil 23 with respect to the airfoil 23 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 described below.

In addition, a part of the cooling air CA supplied to the space portion 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 by the combustion gas FG of the gas passage surface 25a located further to the back side of the airfoil 23 than the airfoil 23 is suppressed.

The cooling air CA cooled on the inner region bottom surface 255a of the second region R2 is supplied to the second side passage 132 described later.

Part of the cooling air CA supplied to the space portion 257 flows into the leading edge side flow path 191 which is formed in the leading edge end portion 253 at a plurality of intervals in the circumferential direction Dc and extends in the axial direction Da, and mainly cools the leading edge end portion 253. The cooling air CA flowing through the leading edge portion 253 is discharged to the combustion gas flow path 32 from an opening formed in the gas flow path surface 25a of the leading edge portion 253.

The stator vanes 21 of several embodiments include an air passage 105, and the air passage 105 is configured to circulate the cooling air CA supplied to the space 257 through 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 end passage 180.

The first side passage 131 is an air passage formed from the front edge 23a side to the rear edge 23b side at the first side end 251 of the inner shroud 25, and one end of the first side passage 131 on the front edge 23a side communicates with the impact space 256 at the first side end 251 via the first opening 111 formed in the first region R1.

The other end portion of the first side passage 131 on the trailing edge 23b side communicates with one end portion in the circumferential direction Dc of a circumferential passage 135 described later.

The second side passage 132 is an air passage formed from the front edge 23a side to the rear edge 23b side at the second side end 252 of the inner shroud 25, and one end of the second side passage 132 on the front edge 23a side communicates with the space 257 at the second side end 252 via the second opening 112 formed in the second region R2.

The other end portion on the trailing edge 23b side of the second side passage 132 communicates with the other end portion of the circumferential direction Dc of the circumferential passage 135.

The circumferential passage 135 is a cooling passage extending in the circumferential direction Dc at a position closer to the trailing edge 23b than the space portion 257, and connects the other end portion of the first side passage 131 on the trailing edge 23b side with the other end portion of the second side passage 132 on the trailing edge 23b side as described above.

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

In the air passage 105 according to the embodiment configured as described above, a part of the cooling air CA in the space portion 257 is supplied from the first opening 111 to the first side passage 131 via the impingement space 256, and a part of the cooling air CA in the space portion 257 is supplied from the second opening 112 to the second side passage 132.

The cooling air CA supplied to the first side passage 131 flows from the front edge 23a side toward the rear edge 23b side in the first side passage 131, and mainly cools the first side end 251.

The cooling air CA supplied to the second side passage 132 flows from the front edge 23a side toward the rear edge 23b side in the second side passage 132, and mainly cools the second side end 252.

The first opening 111 and the second opening 112 are preferably provided near the leading edge 253 so that the cooling air CA flowing into the first side passage 131 and the second side passage 132 flows in a longer section along the axial direction Da as much as possible.

The cooling air CA flowing toward the trailing edge 23b side in the first side passage 131 and the second side passage 132 flows from the first side passage 131 and the second side passage 132 into the circumferential passage 135, and flows into the plurality of trailing edge end passages 180 through the circumferential passage 135, respectively. The cooling air CA flowing into the plurality of trailing edge end passages 180 flows from the upstream end 180a of the trailing edge end passage 180 toward the trailing edge end surface 25c of the inner shroud 25, primarily cooling the trailing edge end 254. The cooling air CA is discharged from the trailing edge face 25c into the combustion gas.

(bolt-coupling of the first segment 101A and the second segment 101B)

In the first segment body 101A of the blade segment body 100 according to the embodiment, the bolt holes 161 penetrating in the circumferential direction Dc are formed in the second side end 252 on the back side blade surface 23d side in the first inner shroud 25A and the second side end 152 on the back side blade surface 23d side in the first outer shroud 27A.

In the first segment 101A of the blade segment 100 according to the embodiment, as shown in fig. 4, one bolt hole 161 is formed in the second side end 252 of the first inner shroud 25A, but a plurality of bolt holes 161 may be formed at intervals in the axial direction Da.

In the first segment 101A of the blade segment 100 according to the embodiment, a plurality of bolt holes 161 are preferably formed in the second side end 152 of the first outer shroud 27A at intervals in the axial direction Da. In the example shown in fig. 4, the number of bolt holes 161 is three, but may be two or less, or four or more.

In the second segment body 101B of the blade segment body 100 according to the embodiment, unillustrated bolt holes penetrating in the circumferential direction Dc are formed in a first side end 251 on the ventral blade surface 23c side in the second inner shroud 25B and in an unillustrated first side end on the ventral blade surface 23c side in the second outer shroud.

In the second segment 101B of the blade segment 100 according to the embodiment, one bolt hole is preferably formed in the first side end 251 of the second inner shroud 25B, but a plurality of bolt holes may be formed at intervals in the axial direction Da.

In the second segment 101B of the blade segment 100 according to the embodiment, a plurality of bolt holes are preferably formed in the first side end 151 of the second outer shroud 27B at intervals in the axial direction Da, and for example, three bolt holes are preferably formed in the same manner as in the first outer shroud 27A, but two or less, or four or more may be formed.

The bolt hole 161 of the first block 101A and the bolt hole, not shown, of the second block 101B are set in respective positions so that the bolt 171 can be inserted into the bolt hole 161 and the bolt hole, not shown.

In the blade segment body 100 according to the embodiment, the first segment body 101A and the second segment body 101B are bolt-coupled by inserting bolts 171 through the bolt holes 161 and bolt holes not shown and assembling nuts 172.

In the turbine 13 of the gas turbine 10 according to the embodiment, a plurality of the blade segment bodies 100 are arranged in the circumferential direction Dc. Adjacent blade segment bodies 100 in the circumferential direction Dc are not bolted to each other. A seal plate, not shown, for preventing leakage of cooling air CA from between the blade segment bodies 100 adjacent to each other in the circumferential direction Dc is disposed between the blade segment bodies 100 adjacent to each other in the circumferential direction Dc.

In the turbine vane 21 according to the embodiment, the cooling air CA cooled in the first region R1 and the second region R2 is preferably further used for cooling the inner shroud 25. In particular, the temperature of the cooling air CA after cooling the second region R2, which is a region where impingement cooling is not performed, is relatively low, and therefore, it is preferable to effectively use the cooling of the inner shroud 25.

Therefore, the turbine vane 21 of the embodiment is configured as follows. That is, in the turbine vane 21 of one embodiment, the inner shroud 25 includes: a first opening 111 formed in the first region R1; and a first side passage 131 having one first side end 251 in the circumferential direction Dc of the inner shroud 25 formed from the front edge 23a side to the rear edge 23b side and one end connected to the first opening 111. The inner shroud 25 has: a second opening 112 formed in the second region R2; and a second side passage 132 having a second side end 252 formed from the front edge 23a side to the rear edge 23b side in the circumferential direction Dc of the inner shroud 25, and having one end connected to the second opening 112.

As a result, the cooling air CA after cooling the first region R1 that is subjected to the impingement cooling and the cooling air CA after cooling the second region R2 that is not subjected to the impingement cooling can be further used for cooling the inner shroud 25, and 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 the intake air AI; a combustor 12 that supplies and combusts a fuel FL to the compressed air AC compressed by the compressor 11; and a turbine 13 that obtains rotational power by the combustion gas FG obtained by combustion in the combustor 12. The turbine 13 includes the blade segment body 100 (i.e., the turbine vane 21) according to the above-described embodiment.

As a result, the cooling air CA after cooling the first region R1 that is subjected to the impingement cooling and the cooling air CA after cooling the second region R2 that is not subjected to the impingement cooling can be further used for cooling the inner shroud 25, and therefore the cooling efficiency of the inner shroud 25 can be improved. This can suppress, for example, excessive cooling and excessive use of the cooling air CA, thereby improving the efficiency of the gas turbine 10.

In the turbine vane 21 of an embodiment, the inner shroud 25 preferably has a circumferential passage 135, and the circumferential passage 135 extends in the circumferential direction Dc at a position closer to the trailing edge 23b than the first region R1. Preferably, the other end of the first side passage 131 is connected to one end of the circumferential passage 135, and the other end of the second side passage 132 is connected to the other end of the circumferential passage 135.

Since the cooling air CA flows into the second side passage 132 from the second region R2 not covered by the collision plate 70, the pressure of the cooling air CA tends to be higher than that of the first side passage 131 into which the cooling air CA flows from the first region R1 covered by the collision plate 70. Therefore, the flow rate of the cooling air CA flowing through the second side passage 132 is easily increased, and from the viewpoint of efficiently using the cooling air CA, it is preferable to further use the cooling air CA flowing through the second side passage 132 for cooling the inner shroud 25.

According to the turbine vane 21 of an embodiment, the periphery of the circumferential passage 135 can be further cooled by the cooling air CA flowing in the second side passage 132. In addition, according to the turbine vane 21 of the embodiment, since the cooling flow path (the trailing edge end passage 180) is further provided on the downstream side of the circumferential passage 135, the cooling air CA flowing through the first side passage 131 and the second side passage 132 can circulate through the cooling flow path (the trailing edge end passage 180) to cool the inner shroud 25, and therefore the cooling air CA can be efficiently used.

In the turbine vane 21 of the 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 flows into the second side passage 132 from the second region R2 not covered by the collision plate 70, the pressure of the cooling air CA tends to be higher than that of the first side passage 131 into which the cooling air CA flows from the first region R1 covered by the collision plate 70. Therefore, the cooling air CA flowing into the circumferential passage 135 is more likely to come from the second side passage 132 than the cooling air CA coming from the first side passage 131.

According to the turbine vane 21 of the embodiment, the flow rate of the cooling air CA flowing into the second side passage 132 can be suppressed at the second opening 112, and therefore the flow rate of the cooling air CA flowing into the second side passage 132 can be suppressed to an appropriate flow rate.

In the turbine vane 21 of one embodiment, it is preferable that the inner shroud 25 has a plurality of trailing edge end passages 180 arranged in the circumferential direction Dc on the trailing edge 23b side, and that the upstream end 180a is connected to the circumferential passage 135 and the downstream end 180b is opened at the trailing edge end surface 25c of the inner shroud 25.

In this way, the cooling air CA flowing through the first side passage 131 and the second side passage 132 can flow through the plurality of trailing edge end passages 180 to cool the region (trailing edge end 254) on the trailing edge 23b side of the inner shroud 25, and the cooling air CA can be efficiently used.

In the turbine vane 21 of the embodiment, the first region R1 preferably includes a region located on the ventral side of the airfoil 23 from the airfoil 23 when viewed in the blade height direction.

Since the combustion gas FG, which is the working fluid of the turbine 13, is a relatively high-temperature fluid, in the region near the ventral side of the airfoil 23 in the combustion gas flow path 32, the flow velocity of the combustion gas flow path 32 tends to be higher than that in the region near the back side, and therefore the temperature of the inner shroud 25 tends to be higher in the region located closer to the ventral side of the airfoil 23 than in the airfoil 23.

According to the turbine vane 21 of the embodiment, since the region in which the temperature tends to be relatively high can be subjected to the impingement cooling, the temperature rise of the inner shroud 25 can be effectively suppressed.

In the turbine vane 21 according to one embodiment, the structure related to the air passage 105 is preferably provided in the inner shroud 25.

According to the turbine vane 21 of the embodiment, in the inner shroud 25 having a smaller size than the outer shroud 27 and a smaller area where the impingement plate 70 is disposed, the cooling air CA after cooling the first region R1 that is subjected to impingement cooling and the cooling air CA after cooling the second region R2 that is not subjected to impingement cooling can be further used for cooling the inner shroud 25.

The blade segment 100 according to one embodiment includes a first segment 101A and a second segment 101B including the turbine vane 21 according to one embodiment. The second side end 252 of the first section 101A is bolted to the first side end 251 of the second section 101B. Therefore, the second side end 252 of the first segment 101A and the first side end 251 of the second segment 101B have bolts 171 for bolt-coupling, and thus are relatively difficult to cool.

According to the blade segment 100 of the embodiment, 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, so that 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 blade segment body 100 according to the embodiment, it is preferable that the first side passage 131 overlaps with a portion of the first side end portion 251 that is coupled by the bolt 171 and the nut 172 as the fixing members when viewed from the blade height direction, and the second side passage 132 overlaps with a portion of the second side end portion 252 that is coupled by the bolt 171 and the nut 172 as the fixing members when viewed from the blade height direction.

This makes it possible to efficiently cool the second side end 252 of the first block 101A and the first side end 251 of the second block 101B, which are difficult to cool.

The present disclosure is not limited to the above-described embodiments, and includes modifications to the above-described embodiments and combinations of these modes as appropriate.

For example, the structure related to the air passage 105 in the inner shroud 25 of the turbine vane 21 according to the above embodiment may be provided in the outer shroud 27.

In the turbine vane 21 of the above-described embodiment, the cooling air CA flowing into the first and second side passages 131 and 132 is caused to flow from the front edge 23a side toward the rear edge 23b side, but the cooling air CA flowing into the first and second side passages 131 and 132 may also be caused to flow from the rear edge 23b side toward the front edge 23a side.

In this case, the first opening 111 and the second opening 112 may be provided near the trailing edge 254 so that the cooling air CA flowing into the first side passage 131 and the second side passage 132 flows in a longer section along the axial direction Da as much as possible.

In this case, it is preferable that the circumferential passage 135 is provided on the front edge 23a side of the space portion 257, and the circumferential passage 135 is connected to the first side passage 131 and the second side passage 132.

In the turbine vane 21 according to the embodiment described above, the flow rate of the cooling air CA flowing into the first side passage 131 and the flow rate of the cooling air CA flowing into the second side passage 132 are appropriately adjusted by appropriately setting the inlet area S1 of the first opening 111 and the inlet area S2 of the second opening 112.

However, instead of appropriately setting the inlet area S1 of the first opening 111 and the inlet area S2 of the second opening 112, or appropriately setting the inlet area S1 of the first opening 111 and the inlet area S2 of the second opening 112, the flow rate of the cooling air CA flowing into the first side passage 131 and the flow rate of the cooling air CA flowing into the second side passage 132 may be appropriately adjusted by appropriately setting the passage cross-sectional areas of the first side passage 131 and the second side passage 132.

The contents described in the above embodiments are grasped as follows, for example.

(1) The turbine vane 21 according to at least one embodiment of the present disclosure includes: wing 23; and a shroud 2 provided on at least one of one side and the other side in the blade height direction (radial direction Dr) of the wing portion 23. The shroud 2 (inner shroud 25) has a concave portion (space portion 257) formed on a surface opposite to the wing portion 23 across the gas passage surface 2a. The concave portion (space portion 257) includes: a first region R1 covered by an impact plate (collision plate 70); and a second region R2 that is not covered by the impact plate (collision plate 70). The shroud 2 (inner shroud 25) includes: a first opening 111 formed in the first region R1; and a first side passage 131 formed from the front edge 23a side to the rear edge 23b side at one first side end 251 in the circumferential direction Dc of the shroud 2, and having one end connected to the first opening 111. The shroud 2 (inner shroud 25) includes: a second opening 112 formed in the second region R2; and a second side passage 132 formed from the front edge 23a side to the rear edge 23b side at the other second side end 252 in the circumferential direction Dc of the shroud 2, and having one end connected to the second opening 112.

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

(2) In several embodiments, in addition to the structure of (1) above, it is preferable that the shroud 2 (inner shroud 25) has a circumferential passage 135 extending in the circumferential direction Dc at a position closer to the trailing edge 23b than the first region R1. The other end of the first side passage 131 is connected to one end of the circumferential passage 135, and the other end of the second side passage 132 is connected to the other end of the circumferential passage 135.

According to the configuration of (2) above, the periphery of the circumferential passage 135 can be further cooled by the cooling air CA flowing through the second side passage 132. In addition, according to the configuration of (2) above, by further providing the cooling flow path (the trailing edge end path 180) on the downstream side of the circumferential path 135, the cooling air CA flowing through the first side path 131 and the second side path 132 can circulate through the cooling flow path (the trailing edge end path 180) to cool the shroud 2 (the inner shroud 25), and therefore the cooling air CA can be efficiently used.

(3) In several embodiments, in addition to the structure of (2) above, it is preferable that the inlet area S2 of the second opening 112 is smaller than the inlet area S1 of the first opening 111.

As described above, since the cooling air CA flows into the second side passage 132 from the second region R2 not covered by the impingement plate (impingement plate 70), the pressure of the cooling air CA tends to be higher than that of the first side passage 131 from the first region R1 covered by the impingement plate (impingement plate 70). Therefore, the cooling air CA flowing into the circumferential passage 135 is more likely to come from the second side passage 132 than the cooling air CA coming from the first side passage 131.

According to the configuration of (3) above, the flow rate of the cooling air CA flowing into the second side passage 132 can be suppressed at the second opening 112, and therefore the flow rate of the cooling air CA flowing into the second side passage 132 can be suppressed to an appropriate flow rate.

(4) In several embodiments, in addition to the structure of (2) or (3) above, it is preferable that the shroud 2 (inner shroud 25) has a plurality of trailing edge end passages 180 arranged in the circumferential direction Dc on the trailing edge 23b side, one end (upstream end 180 a) is connected to the circumferential passage 135, and the other end (downstream end 180 b) is opened at the trailing edge end face 25c of the shroud 2 (inner shroud 25).

According to the configuration of (4) above, the cooling air CA flowing through the first side passage 131 and the second side passage 132 can be circulated through the plurality of trailing edge end passages 180 to cool the region on the trailing edge 23b side of the shroud 2 (inner shroud 25), and the cooling air CA can be efficiently used.

(5) In several embodiments, in addition to any one of the structures (1) to (4) above, it is preferable that the first region R1 includes a region located on the ventral side of the wing 23 from the wing 23 when viewed in the blade height direction.

If the working fluid (combustion gas FG) of the turbine 13 is a high-temperature fluid, the flow velocity of the working fluid (combustion gas FG) tends to be higher in the region near the ventral side of the airfoil 23 than in the region near the back side in the flow path (combustion gas flow path 32) of the working fluid (combustion gas FG), and therefore the temperature of the shroud 2 (inner shroud 25) tends to be higher in the region located closer to the ventral side of the airfoil 23 than in the airfoil 23.

According to the configuration of (5), since the region in which the temperature tends to be relatively high can be subjected to the impingement cooling, the temperature rise of the shroud 2 (inner shroud 25) can be effectively suppressed.

(6) In several embodiments, in addition to any one of the structures (1) to (5) above, it is preferable that the shroud 2 (inner shroud 25) is provided on the inner side of the wing 23 in the blade height direction.

According to the configuration of (6) above, in the inner shroud 25 having a smaller size than the outer shroud 27 and a smaller area where the impingement plate (impingement plate 70) is disposed, the cooling air CA after cooling the first region R1 that is impingement-cooled and the cooling air CA after cooling the second region R2 that is not impingement-cooled can be further used for cooling the inner shroud 25.

(7) The blade segment of at least one embodiment of the present disclosure includes the first segment 101A and the second segment 101B including the turbine vane 21 having any one of the structures (1) to (6) described above. The second side end 252 of the first section 101A is bolted to the first side end 251 of the second section 101B.

The second side end 252 of the first segment 101A and the first side end 251 of the second segment 101B have bolts 171 for bolt coupling, and therefore are relatively difficult to cool.

According to the configuration of (7) above, the second side passage 132 is formed in the second side end 252 of the first segment 101A, and the first side passage 131 is formed in the first side end 251 of the second segment 101B, so that the second side end 252 of the first segment 101A and the first side end 251 of the second segment 101B, which are difficult to cool, can be cooled.

(8) In several embodiments, in addition to the structure of (7) above, it is preferable that the first side passage 131 overlaps with a portion of the first side end portion 251 that is coupled by the fastener (bolt 171 and nut 172) when viewed from the blade height direction of the airfoil portion 23 in the axial direction Da, and it is preferable that the second side passage 132 overlaps with a portion of the second side end portion 252 that is coupled by the fastener (bolt 171 and nut 172) when viewed from the blade height direction in the axial direction Da.

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

(9) The gas turbine 10 of at least one embodiment of the present disclosure has: a compressor 11 that compresses the intake air AI; a combustor 12 that supplies fuel FL to compressed air AC compressed by the compressor 11 and burns the fuel FL; and a turbine 13 that obtains rotational power by the combustion gas FG obtained by combustion in the combustor 12. The turbine 13 includes the turbine vane 21 having any one of the configurations (1) to (6) above.

According to the configuration of (9) above, the cooling air CA after cooling the first region R1 that is subjected to the impingement cooling and the cooling air CA after cooling the second region R2 that is not subjected to the impingement cooling can be further used for cooling the shroud 2 (inner shroud 25), and therefore the cooling efficiency of the shroud 2 (inner shroud 25) can be improved. This can suppress, for example, excessive cooling and excessive use of the cooling air CA, thereby improving the efficiency of the gas turbine 10.

(10) The gas turbine 10 of at least one embodiment of the present disclosure has: a compressor 11 that compresses the intake air AI; a combustor 12 that supplies fuel FL to compressed air AC compressed by the compressor 11 and burns the fuel FL; and a turbine 13 that obtains rotational power by the combustion gas FG obtained by combustion in the combustor 12. The turbine 13 includes the blade segment body 100 having the structure of (7) or (8) described above.

According to the configuration of (10) above, the cooling air CA after cooling the first region R1 that is subjected to the impingement cooling and the cooling air CA after cooling the second region R2 that is not subjected to the impingement cooling can be further used for cooling the shroud 2 (inner shroud 25), and therefore the cooling efficiency of the shroud 2 (inner shroud 25) can be improved. This can suppress, for example, excessive cooling and excessive use of the cooling air CA, thereby improving the efficiency of the gas turbine 10.

Claims (10)

1. A turbine vane, wherein,

the turbine vane includes:

wing-shaped parts; and

a shield arranged on at least one of one side and the other side of the vane height direction of the wing part,

the shroud has a concave portion formed on a surface of the shroud opposite to the wing portion across a gas passage surface,

the recess includes: a first region covered by an impingement plate having a plurality of through holes; and a second area not covered by the impingement plate,

the shield has:

a first opening formed in the first region;

a first side passage formed from a leading edge side to a trailing edge side at a first side end in a circumferential direction of the shroud, the first side passage having one end connected to the first opening;

a second opening formed in the second region; and

and a second side passage formed from the leading edge side to the trailing edge side at the other second side end portion in the circumferential direction of the shroud, and having one end portion connected to the second opening portion.

2. The turbine vane of claim 1, wherein,

the shroud has a circumferential passage extending in the circumferential direction at a position closer to the trailing 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.

3. The turbine vane of claim 2, wherein,

the inlet area of the second opening is smaller than the inlet area of the first opening.

4. The turbine vane of claim 2 or 3, wherein,

the shroud has a trailing edge end passage, a plurality of which are arranged in the circumferential direction on the trailing edge side, one end of which is connected to the circumferential passage, and the other end of which is open to the trailing edge end surface of the shroud.

5. The turbine vane of any one of claims 1 to 3, wherein,

the first region includes a region located on a ventral side of the airfoil when viewed from the blade height direction.

6. The turbine vane of any one of claims 1 to 3, wherein,

the shroud is provided on the inner side of the wing in the blade height direction.

7. A blade segment body, wherein,

the blade segment is provided with a first segment and a second segment comprising the turbine vane of any one of claims 1 to 6,

the second side end of the first section body is coupled to the first side end of the second section body by a fastener.

8. The blade segment body of claim 7, wherein,

the first side passage overlaps in the axial direction with a portion of the first side end portion joined by the fixing piece as viewed from the blade height direction of the airfoil portion,

the second side passage overlaps in the axial direction with a portion of the second side end portion joined by the fixing piece when viewed from the blade height direction.

9. A gas turbine, wherein,

the gas turbine has:

a compressor for compressing the taken-in air;

a combustor that supplies fuel to the compressed air compressed by the compressor and combusts the fuel; and

a turbine that obtains rotational power by combustion gas obtained by combustion in the combustor,

the turbine is provided with the turbine vane of any one of claims 1 to 6.

10. A gas turbine, wherein,

the gas turbine has:

a compressor for compressing the taken-in air;

a combustor that supplies fuel to the compressed air compressed by the compressor and combusts the fuel; and

a turbine that obtains rotational power by combustion gas obtained by combustion in the combustor,

the turbine is provided with a blade segment body according to claim 7 or 8.