JP2003129803A - Gas turbine - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a gas turbine restraining burnout of an end portion of a segment. SOLUTION: A turbine split ring, a turbine stator blade shroud, and other turbine passage walls are formed of a plurality of segments 200 annularly joined, and constitute a passage wall for combustion gas. A seal plate 400 is inserted into seal grooves 211, 211 formed on joint surfaces 210, 210 facing between the segments 200, 200, thereby crossing between the joint surfaces 210, 210 and sealing off a gap 300. Furthermore, air grooves 1, which permit cooling air to flow from the side of a back surface 240 of the turbine passage wall to the side of a front surface 250 of the seal plate 400, are provided on inner surfaces of the seal grooves 211.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine, and more particularly, to a gas turbine provided with a separable turbine flow path wall such as a turbine rotor blade split ring, a turbine rotor blade platform, and a turbine vane shroud. Regarding

[0002]

2. Description of the Related Art A split annular wall is generally used for a split ring of a turbine rotor blade that serves as a flow path wall for combustion gas, a platform of the turbine rotor blade, and a shroud of a turbine stator blade. FIG. 20 is an overall configuration diagram showing a turbine vane of a conventional gas turbine. The turbine vane includes blade 100 and
Segments 200, 2 supporting the wing 100 at both ends
00 is included. The segment 200 is a cast plate-shaped member, and is joined in a plurality of annular shapes to form an inner shroud and an outer shroud of the turbine vane. Figure 21
FIG. 21 is an assembled perspective view showing a joint portion of the segment 200 described in FIG. 20. In consideration of thermal expansion during operation of the gas turbine, the segment 200 has opposing joint surfaces 210, 2
It is installed with a predetermined gap 300 between 10 pieces. However, this gap 300 is not completely closed when the turbine is in operation, leaving a width of about 1 to 2 mm. Therefore, in the conventional gas turbine, the seal groove 211 extending from the front edge 220 to the rear edge 221 is formed on the joint surface 210, and the seal plate 400 made of a Hastelloy member is inserted into the seal groove 211 to form the gap 300. It was sealed. The seal plate 400 is inserted into the seal groove 211 from the front edge 220 side after the segment 200 is assembled and bolted (see FIG. 21).

In this prior art, high temperature combustion gas 5
00 flows along the shroud surface to form the joint surface 200,
It is wound in the gap 300 between the 200 (see FIG. 22). Then, the edge portion 230 of the segment 200 is burnt out 510, and the edge portion 230 is thinned to cause various serious problems such as dropping of the seal plate 400. Therefore, research and development have been actively conducted in the related art for solving the above problems.
As such a technique, a cavity (not shown) for storing cooling air is formed on the rear surface 240 side of the segment 200,
A technique is known in which this cooling air is discharged to the front surface 250 side to cool the joint surface 210 edge portion 230. Hereinafter, these problems will be described in detail while sequentially showing these conventional techniques and solving problems.

[0004]

As a first conventional technique, a technique described in Japanese Utility Model Laid-Open No. 6-60702 is known. FIG. 23 is a front sectional view showing a joint portion of a turbine vane according to Conventional Technique 1, and FIG. 24 is a perspective view showing a seal plate shown in FIG. The gas turbine of the related art 1 is characterized in that a seal plate 410 having a step portion 411 is used in the gap 300 between the joint surfaces 210, 210. The step portion 411 is formed by cutting out the conventional seal plate 400 at regular intervals in the thickness direction and the width direction (see FIG. 24). As a result, the step portion 411 of the seal plate 410 receives the cooling air 600 in the segment 200 while being fitted in the seal groove 211.
A passage is formed from the rear surface 240 side to the front surface 250 side (see FIG. 23). In the conventional technique 1, the cooling air 6
00 is the step portion 4 from the rear surface 240 side of the segment 200.
It is discharged to the front surface 240 side of the seal plate 410 through the passage created by 11, and fills the gap 300 between the joint surfaces 210, 210 to cool the edge portion 230 thereof. As a result, the joint surface 2
Burnout 510 of the tenth edge portion 230 is suppressed.

Here, the seal plate 400 has a thin structure that is elastically deformable. However, it is not easy to cut out the thin seal plate 400 and form the step portion 411. In addition, if the thin seal plate 400 is cut out, the strength thereof is significantly reduced. On the other hand, if the thickness of the seal plate 400 is increased, the thickness of the segment 200 is also increased accordingly, the seal plate 400 cannot be deformed due to thermal deformation of the segment 200 during turbine operation, and the segment 200 is damaged or the seal plate 400 comes off. There is also a problem that occurs. In addition, the segment 20
If 0 is made thick, the heat capacity also increases, and there is a problem that excessive cooling is required. Therefore, the present invention has been made in view of the above, and burns 5 that occurs in the edge portion 230 of the joint surface 210 while using the existing seal plate 400.
A first object of the present invention is to provide a gas turbine capable of suppressing 10

As a second conventional technique, a structure in which a hole is formed in the seal plate 400 is known. 25 is a front sectional view showing a joint portion of a turbine vane according to Conventional Technique 2, FIG. 26 is a perspective view showing a seal plate shown in FIG. 25, and FIG. 27 is shown in FIG. It is a side sectional view of the joined portion. In the gas turbine of the related art 2, a plurality of through holes 42 are formed in the center of the conventional seal plate 400.
A seal plate 420 in which 1 is opened vertically is used (see FIG. 26). The seal plate 420 is the segment 200.
Are inserted into the seal grooves 211, 211 and fitted. In the prior art 2, the cooling air 600 is supplied to the seal plate 4
It passes through the through holes 421 of 20 and jets from the shroud back surface 240 side to the front surface 250 side. As a result, the cooling air 60
0 cools the edges 230 of the segment 200 to prevent its burnout 510.

However, in the gas turbine of the prior art 2 described above, since the through hole 421 of the seal plate 420 is opened perpendicularly to the plate surface, the cooling air 600 is jetted vertically from the through hole 421 and the segment 200. Edge 230
And immediately mix with the flow of combustion gas 500 (see FIG. 27). Therefore, there is a problem that the edge portion 230 cannot be cooled effectively. There is also a method of increasing the flow rate of the cooling air 600 by increasing the diameter of the through hole 421 to enhance the cooling effect. However, in such a method, since the amount of the cooling air 600 mixed with the combustion gas 500 also increases, there is a problem that the pressure loss of the combustion gas 500 is significant and the gas turbine efficiency is reduced. Therefore, a second object of the present invention is to provide a gas turbine capable of suppressing the burnout 510 of the edge portion 230 of the segment 200 and reducing the mixing loss of the cooling air 600 and the combustion gas 500.

Further, in the gas turbine of the above-mentioned Prior Art 2, since the through hole 421 penetrates the seal plate 420 straightly, the resistance that the cooling air 600 receives when passing through the seal plate 420 is small. Therefore, the sealing plate 420 of the related art 2 has a low function of sealing the joint surfaces 210, 210, and causes the cooling air 600 to pass more than necessary, which causes a problem that the pressure loss of the combustion gas 500 is large. It was
Therefore, a third object of the present invention is to provide a gas turbine capable of effectively suppressing burnout 510 occurring at the edge portion 230 of the segment 200 while ensuring the sealing function of the seal plate 420.

A third conventional technique is Japanese Patent Laid-Open No.
The technique described in Japanese Patent Laid-Open No. 00-25744 is known. FIG. 29 is a perspective view showing a main portion of a turbine vane according to Conventional Technique 3, and FIG. 30 is a front sectional view showing a joint portion of the segment shown in FIG. The gas turbine of the related art 3 includes a front edge 220 to a rear edge 2 along the front surface 250 side end portion 251 of one segment 200.
21 has a plurality of blowing holes 430 for the cooling air 600 (see FIG. 29). The blowing hole 430 is formed so as to be inclined toward the edge portion 230 of the segment 200 and penetrate the segment 200 from the rear surface 240 side to the front surface 250 side (see FIG. 30). In the conventional technique 3, the cooling air 6
00 blows out from the rear surface 240 side of the segment 200 through the blowing hole 430 to the front surface 250 side. Thereby, the cooling air 600 flows into the gap 300 between the joint surfaces 210, 210 and cools the edge portion 230 of the segment 200.

However, in the gas turbine of the prior art 3 described above, no study has been made on the effective and efficient arrangement of the blowout holes 430. Therefore, the present invention has been made in view of the above, and a fourth object thereof is to provide a gas turbine capable of cooling the edge portion 230 of the segment 200 more effectively.

[0011]

In order to achieve the above first object, a gas turbine according to the present invention is formed by joining a plurality of segments in an annular shape and constitutes a passage wall for combustion gas. The turbine split ring, the turbine rotor blade platform, the turbine vane shroud, and other turbine channel walls are fitted into the seal grooves formed on the facing joint surfaces between the segments, and the joint is provided between the joint surfaces. With a seal plate that seals the gap between the surfaces,
An air groove is provided on the inner surface of the seal groove for passing cooling air from the back side of the turbine passage wall to the front side of the seal plate.

In the present invention, the cooling air passes through the air groove formed on the inner surface of the seal groove and is jetted from the rear surface side of the turbine passage wall to the front surface side of the seal plate to cool the joint surface of the segment. As a result, burning of the edge of the segment is suppressed.

Further, in order to achieve the first object,
The gas turbine according to the present invention (claim 2) is defined by claim 1.
Further, in the gas turbine according to the aspect 1, the air grooves are provided at different positions in plan view on the front surface side and the rear surface side with respect to the turbine flow path wall.

In the present invention, the cooling air enters the seal groove from the air groove on the back side, moves in the seal groove in the longitudinal direction, and is ejected from the outlet of the air groove on the front side. Here, since the cooling air is decelerated by the crank-shaped flow path formed by the air groove and the seal groove, the jet flow rate thereof is suppressed. As a result, the segment edge can be cooled while maintaining the sealing function of the seal plate. It should be noted that the “planarly different positions” include, for example, the case where the air grooves are alternately formed in a zigzag pattern on the front inner surface and the back inner surface of the seal groove.

In order to achieve the second object,
A gas turbine according to the present invention (claim 3) is formed by joining a plurality of segments in an annular shape and forms a combustion gas passage wall, a turbine split ring, a turbine moving blade platform, and a turbine stationary blade shroud. Other turbine channel walls, and having a seal plate that is fitted into seal grooves formed on the facing joint surfaces between the segments and that extends between the joint surfaces and seals the gap between the joint surfaces, In addition, the sealing plate has a through hole that is inclined toward the trailing edge side of the turbine flow path wall.

In the present invention, since the through hole of the seal plate is inclined toward the trailing edge side, the cooling air is ejected while inclining toward the trailing edge side on the front side of the sealing plate. As a result, the cooling air is ejected more smoothly than in the case where the through holes are formed perpendicularly to the seal plate, and a cooling air layer is formed on the front surface of the seal plate.

In order to achieve the second object,
The gas turbine according to the present invention (claim 4) is defined by claim 3
In the gas turbine according to claim 1, the inclination angle of the through hole is such that the cooling air jetted out is larger than a lower limit value that reaches the edge portion of the turbine passage wall, and the edge portion of the turbine passage wall. It is characterized in that the angle is set so as to be smaller than the upper limit value that exceeds the range and mixes into the combustion gas.

In the present invention, the cooling air cools the edges of the segments to prevent their burnout. In addition, the lower limit value that reaches the edge is that the jetted cooling air does not reach the edge,
It refers to the critical value of the angle that flows downstream without cooling the edges. The upper limit value mixed in the combustion gas means a critical value at which the jetted cooling air immediately mixes with the combustion gas beyond the edge and does not form a cooling air layer near the edge. Specifically, these critical values are determined by three-dimensional fluid analysis according to the specifications, structure, etc. of each gas turbine.

Further, in order to achieve the third object,
A gas turbine according to the present invention (claim 5) is formed by joining a plurality of segments in an annular shape and forms a combustion gas passage wall, a turbine split ring, a turbine moving blade platform, and a turbine stationary blade shroud. Other turbine flow path walls, having a seal plate that is fitted into seal grooves formed on the facing joint surfaces between the segments and that extends between the joint surfaces to seal a gap between the joint surfaces, Further, the seal plate is characterized in that cutout portions are provided at side portions of upper and lower surfaces thereof, which are two-dimensionally different on the upper surface side and the lower surface side with respect to the turbine passage wall.

In the present invention, the cooling air enters the seal groove from the cutout portion on the lower surface side of the seal plate, passes the side of the seal plate, and is blown out from the cutout portion on the upper surface side. Here, the cooling air is decelerated by the crank-shaped flow path formed by the cutout portion and the seal groove, and the ejection flow rate thereof is suppressed. As a result, the segment edge can be cooled while maintaining the sealing function of the seal plate.

By the way, according to the research conducted by the inventors regarding the above-mentioned Prior Art 3, the flow of the combustion gas 500 is not uniform on the surface of the shroud, and the streamline direction changes depending on the location. FIG. 31 is a plan sectional view showing a shroud of the turbine first stage stationary blade shown in FIG. 29. Segment 200
The joining surface 210 of is inclined toward the blade 100b side from the leading edge 220 toward the trailing edge 221 and is formed obliquely. In the prior art 3, the combustion gas 500 flows substantially vertically from the front edge 220 side and flows on the shroud surface, and the gap 300
The front edge 220 side of the wing 100a
It corresponds to the belly 110. Then, the combustion gas 500 is transferred to the blade 1
00a to change the streamline direction, and to pass through the trailing edge 221 side of the clearance 300 from the direction of the blade 100a and pass through the back 120 of the blade 100b to the trailing edge 221. That is, the combustion gas 500 is emitted from the shroud 200 on the front edge 220 side and the rear edge 22 side.
The first side straddles the gap 300 between the joint surfaces 210 from different directions.

Here, the blow-out hole 430 is formed in the joint surface 21.
The cooling air 600 is not supplied to the gap 300 unless it is on the upstream side of the gap 300 of 0. Further, since the gas turbine according to the related art 3 has the blowing hole 430 at the end portion 251 of the segment 200 on the blade 100a side, the combustion gas 500
Spans the front edge 220 side of the gap 300 from the blade 100b side. Therefore, in the gas turbine of the related art 3 described above,
The blow-out hole 430 on the front edge 220 side is located downstream of the gap 300 and is not cooled. Further, such a problem exists not only in the turbine first-stage vane but also in the turbine rotor blade and the turbine vane in the later stage. Therefore, the present invention has been made in view of the above-mentioned new problems by the inventors' research, and the combustion gas 5
A fourth object is to provide a gas turbine that can more effectively cool the edge portion 230 of the segment 200 with respect to the change in the direction in which 00 crosses the gap 300.

In order to achieve the fourth object, the gas turbine of the present invention (claim 6) is a turbine rotor blade,
Turbine stator blades and other blades, and a plurality of segments supporting the blades are annularly joined to each other to form a combustion gas passage wall, a turbine split ring, a turbine rotor blade platform, and a turbine stator blade shroud. Other turbine channel walls and the gap between the facing joint surfaces of the segments as a reference, with the vicinity where the flow direction of the combustion gas across the gap changes as a boundary, from the end of one segment to the end of the other segment And a plurality of cooling air blowout holes formed by switching to the section.

In the present invention, since the gas turbine is formed by switching the formation positions of the blowing holes according to the streamline direction of the combustion gas, a larger amount of cooling air is blown as compared with the gas turbine of the prior art 3. The hole can be formed upstream of the gap between the joint surfaces. As a result, the number of blow-out holes that do not have a cooling action is reduced, and the cooling effect of the edge portion of the segment can be improved. It should be noted that the boundary near the point where the flow direction of the combustion gas changes is different for each turbine flow passage wall and also at the start of operation of the turbine and after a certain period of time elapses. By the method,
It is preferable to appropriately change and determine each turbine. Further, in this vicinity, the streamline direction of the combustion gas straddling the gap may be parallel to the gap, so the “boundary” may have a certain range and does not necessarily mean one point.

In order to achieve the above-mentioned fourth object,
A gas turbine according to the present invention (claim 7) is formed by annularly joining a turbine moving blade, a turbine stationary blade and other blades and a plurality of segments supporting the blade, and a combustion gas passage wall. A turbine split ring, a turbine rotor blade platform, a turbine vane shroud and other turbine flow path walls, and a plurality of cooling air blowing holes formed at the vent side end of the blade of the segment. Having, and the gap between the facing joint surfaces of the segment, with the vicinity of the vicinity of the flow direction of the combustion gas straddling the gap changing, the leading edge side is refracted to the wing side with its back to the gap. The blow-out hole is formed so as to be located on the upstream side of the gap.

In the present invention, the gas turbine is formed by refracting the leading edge side of the gap so that the blowing hole is located upstream of the gap between the joint surfaces. More cooling air blowing holes can be formed on the upstream side with respect to the gap between the joint surfaces. Thereby, the burnout of the segment edge can be suppressed more effectively. In addition, since the boundary in the vicinity of the combustion gas that changes the flow direction across the gap is different for each turbine flow path wall, specifically, by a three-dimensional flow path analysis and other methods known to those skilled in the art,
It is preferably determined for each individual turbine channel wall.
Further, in this vicinity, the streamline direction of the combustion gas straddling the gap may be parallel to the gap, so that the “boundary” may have a certain range and does not necessarily mean one point.

[0027]

BEST MODE FOR CARRYING OUT THE INVENTION The present invention will be described below in detail with reference to the drawings. The present invention is not limited to this embodiment. In addition, constituent elements of the embodiments described below include those that can be usually modified by those skilled in the art.

(First Embodiment) FIG. 1 is a perspective view showing a main part of a gas turbine according to a first embodiment of the present invention. 2 is a cross-sectional view taken along the line AA showing the joint portion of the segment shown in FIG. 1, and FIG. 3 is a cross-sectional view seen from the line B showing the shroud joint surface shown in FIG. Also, FIG.
FIG. 4 is a perspective view showing a joint surface of the segment shown in FIGS. 1 to 3. In these figures, the same components as those of the conventional gas turbine described above are designated by the same reference numerals, and the description thereof will be omitted. This gas turbine is not processed into the seal plate 400 as compared with the gas turbine of the above-mentioned Prior Art 1,
The difference is that the air groove 1 is formed on the joint surface 210 side of the segment 200. That is, the seal plate 400 of this gas turbine has a simple plate shape, and does not have the step portion 411 (see FIG. 24) unlike the seal plate 410 of the related art 1. The seal plate 400 is inserted and fitted into the seal groove 211 of the joint surface 210, and the joint surfaces 210, 210
The gap 300 between them is sealed.

Here, a constant space 211 is secured on the side of the seal plate 400 in order to absorb thermal expansion of the segment 200 during operation of the turbine. In addition, the air groove 1 having a rectangular cross section is formed on the inner surface of the seal groove 211 by cutting out to the depth of the seal groove 211 in the direction perpendicular to the joint surface 210 (see FIGS. 1 to 4). In addition, the air grooves 1 are formed in a zigzag pattern from the front edge 220 to the rear edge 221 of the segment 200 at the inner surface of the seal groove 211 alternately vertically and at equal intervals (see FIGS. 3 and 4). As a result, the air groove 1 forms a passage for the cooling air 600 that is bent in a crank shape when the seal plate 400 is installed (see FIGS. 2 and 3).

In the first embodiment, the cooling air 60
0 enters the space 212 on the side of the seal plate 400 from the air groove 1a on the back surface 240 side, flows in the space 212 along the seal plate 400 toward the trailing edge 230, and enters the air groove 1b on the front surface 250 side. Then, the cooling air 600 is
The air is ejected from the outlet 3 of the air groove 1b into the gap 300 between the joint surfaces 210, 210 (see FIGS. 2 and 3). Thereby, the cooling air 600 cools the edge portion 230 of the segment 200 and suppresses the burnout 510 thereof.

According to the first embodiment, the cooling air 60
Since the passage of 0 (the air groove 1) is formed on the segment 200 side, the seal plate 400 having higher strength can be adopted as compared with the seal plate 410 of the prior art 1. Further, the sealing plate 410 of the above-mentioned conventional technique 1 for processing the thin sealing plate 400 by processing the segment 200 side.
The passage of the cooling air 600 can be formed more easily as compared with. Further, since the air grooves 1 are formed vertically and alternately at equal intervals, the edge portion 230 of the segment 200 can be cooled uniformly and evenly. Further, since the air groove 1 is formed at a position different from the shroud in plan view, the above-mentioned conventional technique 1
As described above, the flow velocity of the cooling air 600 can be further reduced as compared with the case where the inlet and the outlet of the cooling air 600 are formed at the same position in the plane by the step portion 411. Thereby, the sealing property between the joint surfaces 210, 210 is improved, and the pressure loss of the combustion gas due to the excessive ejection of the cooling air 600 can be suppressed.

In the first embodiment, the air groove 1 has a rectangular cross-sectional shape, but it is not limited to this and may have a semi-circular shape, a triangular shape, or another cross-sectional shape (see FIG. 5).
Further, the width and depth of the air groove 1 are not particularly limited, but in consideration of the original role of the seal plate 400, it is preferable that the size is set to the extent that burnout 510 can be suppressed sufficiently and sufficiently. If the cross-sectional area of the air groove 1 is excessively increased, the cooling air 600
This is because the pressure loss of the combustion gas 500 occurs due to an increase in the flow rate of

Further, in the first embodiment, the air groove 1 is cut to the inner side of the seal groove 211, but not limited to this, it may have a length communicating with the space 212 on the side of the seal plate 400. Good. Further, the number of the air grooves 1 and the interval 4 are not particularly limited. However, from the viewpoint of ensuring the sealing property between the segments 200, 200, it is preferable to set the number or interval 4 at which burnout 510 can be suppressed sufficiently and sufficiently.

Further, in the first embodiment, the air grooves 1 are formed in a zigzag pattern alternately in the vertical direction, but the present invention is not limited to this.
For example, it may be concentratedly formed on a significant portion of the burnout 510 (see FIG. 6). According to this configuration, the cooling air 600
Edge 200 of the segment 200 while suppressing the flow rate of
The burnout 510 of 30 can be effectively suppressed.

In the first embodiment, the air groove 1b may be formed by inclining the outlet 3 toward the downstream side of the combustion gas 500 or the trailing edge 221 side of the shroud. FIG. 7 is a plan view showing a joint portion of the segment 200 having the inclined air groove 1b. In this configuration, the cooling air 600 is supplied to the outlet 3 of the inclined air groove 1b.
From the front edge 220 by ejecting from the gap 300 between the joint surfaces 210.
Flows smoothly from the side toward the trailing edge 221. As a result, as compared with the case where the air groove 1 is provided perpendicularly to the joint surface 210, the cooling air 600 can flow more smoothly, so that the burnout 510 of the edge portion 230 of the segment 200 can be suppressed more effectively.

(Second Embodiment) FIG. 8 is a perspective view showing a main part of a gas turbine according to a second embodiment of the present invention, and FIG. 9 shows a joint portion of the segment described in FIG. FIG. In these figures, the same components as those of the conventional gas turbine described above are designated by the same reference numerals, and the description thereof will be omitted. This gas turbine is different from the gas turbine of the conventional technique 2 in that a seal plate 10 having a through hole 11 inclined toward the trailing edge 221 is employed. The through-hole 11 is formed by inclining the jet port of the cooling air 600 toward the trailing edge 221 side by 60 degrees, and is provided in plural at the center of the seal plate 10 from the leading edge 220 side to the trailing edge 221 side (see FIG. 9). The seal plate 10 is formed by forming a through hole 11 in the conventional seal plate 400.

In the second embodiment, the cooling air 60
0 spouts from the back surface 240 side of the shroud through the through hole 11 of the seal plate 10 and into the gap 300 on the front surface 250 side.
At this time, the cooling air 600 is jetted to the trailing edge 221 side with an inclination of about 60 degrees to form a layer in the gap 300 between the joint surfaces 210, 210, and cools the edge portion 230 of the segment 200 to prevent its burnout 510. To do.

According to the second embodiment, since the through hole 11 is inclined toward the trailing edge 221 side, the cooling air 600 is less likely to be mixed with the combustion gas 500, as compared with the case of the above-mentioned prior art 2. The gap 300 between the joint surfaces 210, 210 is filled, and the edge portion 230 thereof is effectively cooled. Also, the edge 23
Since the cooling of 0 can be effectively performed, the flow rate of the cooling air 600 can be suppressed and the pressure loss due to the mixing of the cooling air 600 and the combustion gas 500 can be reduced.

In the second embodiment, the inclination angle of the through hole 11 is about 60 degrees (preferably 55 to 65 degrees).
However, the present invention is not limited to this and may be inclined toward the trailing edge 230 side at an angle of 40 degrees or more and 75 degrees or less with respect to the plane of the shroud. Here, the lower limit of 40 degrees or more is that if the inclination angle is less than this, the cooling air 600 ejected from the through holes 11 does not reach the edge portion 230 of the segment 200 and flows along the gap 300 toward the trailing edge 221 side. Based on the inventors' knowledge that the edge portion 230 cannot be cooled effectively (see FIG. 28). In addition, the upper limit of 75 degrees or more is that the cooling air 600 passes through the edge portion 230 and immediately mixes with the combustion gas 500 at an inclination angle larger than this, so that the edge portion 230 cannot be effectively cooled. Based on findings (see FIG. 27). It should be noted that the existence of such a range of the tilt angle has not hitherto been known, and will be newly disclosed based on the research of the inventors. Further, the range of this inclination angle is the flow velocity of the combustion gas, the diameter of the through hole 11, the cooling air 600.
Since it changes depending on the flow rate and other factors, it is preferable to appropriately change the value within the range obvious to those skilled in the art, using the above numerical value as a guide. It should be noted that the angle of 60 degrees is a numerical value that is suitable for the existing gas turbine.

(Third Embodiment) FIG. 10 is a perspective view showing a main part of a gas turbine according to a third embodiment of the present invention, and FIG. 11 is a connecting portion of the segments shown in FIG. FIG. 12 is shown in FIG.
FIG. 13 is a cross-sectional view taken along the line EE showing the joint portion of the segment shown in FIG. 1, and FIG. 13 is a perspective view showing a main portion of the sealing plate shown in FIGS. 10 to 12. In these figures, the same components as those of the conventional gas turbine described above are designated by the same reference numerals, and the description thereof will be omitted. This gas turbine is characterized in that a seal plate 20 having a notch portion 21 is adopted, as compared with the gas turbine of the above-mentioned Prior Art 3. A plurality of notches 21 are formed alternately on the upper and lower surfaces and both side edges in the longitudinal direction of the seal plate 20 (see FIG. 13).
Therefore, the cutout portions 20 are formed at different positions in plan view on the upper and lower surfaces, and are formed at positions on both side edges that do not face each other (see FIGS. 11 to 13). The seal plate 20 is fitted by inserting both edges into the seal grooves 211, 211 of the joint surface 210 and passed between the joint surfaces 210, 210 to seal the gap 300.

In the third embodiment, the cooling air 60
0 passes through the notch 21a on the lower surface of the seal plate 20 from the back surface 240 side of the shroud, and the space 2 on the side of the seal plate 20.
Enter 12 (see FIG. 11). Next, the cooling air 60
0 flows into this space 212 toward the trailing edge 221 due to the pressure difference and enters the notch 21b on the upper surface of the seal plate 20 (see FIG.
2), and jets from the outlet 22 into the gap 300 between the joint surfaces 210, 210 to cool the edge portion 230 of the segment 200 (see FIG. 11).

According to the third embodiment, the seal plate 20
Is a notch 21a on the upper surface side and a notch 2 on the lower surface side.
Since it has a position different from that of 1b in a plane, a crank-shaped passage of the cooling air 600 is formed. As a result, as compared with the seal plate 410 (see FIG. 24) of the related art 1, the flow rate of the cooling air 600 is reduced, the flow rate is suppressed, and the sealing function of the seal plate 20 is ensured. Further, since the cutout portions 21 are formed at positions not facing each other on both side edges of the seal plate 20, the strength of the seal plate 20 is improved as compared with the seal plate 410 of Prior Art 1 (see FIG. 24). Further, since the cutout portion 21 is formed in the seal plate 20, when processing the joint surface 210 of the segment 200 (see FIGS.
Compared with (see FIG. 4), there is also an advantage that posterior design changes such as increase or decrease in the number of notches 21 can be adjusted on the side of the seal plate, which has a low cost.

In the third embodiment, the shape, the cross-sectional shape, the cross-sectional area, the interval, etc. of the cutout portion 21 are not particularly limited as long as the burnout 510 of the edge portion 230 of the segment 200 can be suppressed. Good. Therefore, it is desirable to appropriately change the design of the shape of the cutout portion 21 within the scope obvious to those skilled in the art. Also, the notch 21
Is not limited to staggered formation, and for example, burnout 5
It may be configured to be intensively installed in 10 remarkable areas (not shown).

(Fourth Embodiment) FIG. 14 is a turbine 1 of a gas turbine according to a fourth embodiment of the present invention.
It is a perspective view showing a step vane. 15 is a plan sectional view showing the shroud of the turbine first stage stationary blade shown in FIG. 14, and FIG. 16 is a sectional view taken along line F-F showing the joint portion of the segment shown in FIG. In these figures, the same components as those of the conventional gas turbine described above are designated by the same reference numerals, and the description thereof will be omitted. In this gas turbine, as compared with the gas turbine of the related art 3, before and after the boundary in which the streamline direction of the combustion gas 500 straddling the blowout hole 30 of the cooling air 600 across the gap 300 between the joint surfaces 210, 210 is reversed. And the end 251 of the segment 200 on the different side
It is characterized by being switched and formed. As a result, in this gas turbine, as compared with the gas turbine according to the related art 3 described above, more blowing holes 30 are arranged on the upstream side of the gap 300. Specifically, the blowing hole 30 is
Front edge portion 1 of wing 100a that directs belly 110 toward gap 300
30 and a trailing edge portion 140 of the wing 100b facing away from each other are connected by a straight line, and an intersection 31 (hereinafter, referred to as a "switching point") of the straight line and the gap 300 is used as a reference, and the opposite segment 200 side end portion 251. Formed by switching to (Fig. 1
5). Then, the blowing hole 30 has a switching point 31.
On the leading edge 220 side, the segment 20 on the blade 100b side
Is formed on the 0 edge portion 230, and the trailing edge 22 from the switching point 31
On the first side, the edge portion 230 of the segment 200 on the blade 100a side
Is formed. In addition, these blowing holes 30 penetrate the segment 200 from the back surface 240 to the front surface 250, and are formed to be inclined toward the edge portion 230 (see FIG. 16).

In the fourth embodiment, the combustion gas 50
0 flows almost vertically from the leading edge 220 side and flows on the shroud surface, and the leading edge 220 side of the clearance 300 is placed on the blade 100b.
It straddles from the direction and hits the belly 110 of the wing 100a. Then, the combustion gas 500 is guided by the blade 100a to change the streamline direction, straddles the trailing edge 221 side of the gap 300 from the blade 100a direction, and extends along the spine 120 of the blade 100b at the trailing edge 221.
Exit to. That is, the combustion gas 500 is the shroud 2
00 The front edge 220 side and the rear edge 221 side straddle the gap 300 between the joint surfaces 210 from different directions. Here, the switching point 31 is located near the boundary where the combustion gas 500 changes the direction in which it crosses the gap 300. As a result, the blow-out hole 30 allows the cooling air 600 to flow from the upstream side of the gap 300 to cool the edge portion 230 of the segment 200 (see FIG. 17).

According to the fourth embodiment, the blowing hole 3
Since 0 is formed on the upstream side of the gap 300 by switching the formation position, the cooling air 600 can be more effectively flowed into the gap 300 as compared with the gas turbine of the conventional technique 4. Thereby, the loss of the cooling air 600 is reduced and the burnout 5 of the edge portion 230 of the segment 200 is efficiently performed.
10 can be suppressed. Further, since the number of the blowing holes 30 is almost the same as that of the conventional technique 4, the cooling air 600 blown out is
Becomes almost the same, and the pressure loss due to mixing with the combustion gas 500 does not increase.

In the fourth embodiment, the switching point 31 of the blowing hole 30 is the leading edge portion 1 of the blade 100a.
Although it is uniquely defined with reference to 30 and the trailing edge 140 of the blade 100b, it should be appropriately changed to a suitable position by a design change obvious to those skilled in the art. In addition, the fourth embodiment
The position of the switching point 31 disclosed in 1) is based on the inventor's knowledge that it can be generally applied to an existing gas turbine if it is set near such a position.

Further, in the fourth embodiment, the switching points 31 may be provided at a plurality of positions, and an appropriate number is selected by analyzing the flow path of the combustion gas in the actual machine by fluid analysis or other methods known to those skilled in the art. It may be provided at any position. The switching point 31 is located at the downstream side of the gap 300 with respect to the nonuniform combustion gas 500 flow.
This is because it should be determined for the purpose of reducing the number of 0s.

Further, in the fourth embodiment, the blow-out hole 30 penetrates the segment 200 while being inclined toward the edge portion 230. However, the present invention is not limited to this, and if the ejection port is at the end portion 251 of the segment 200, it is inclined. You don't have to. As a result, the cooling air 600 is supplied to the gap 30 between the joint surfaces 210.
Because it can be poured into 0. However, such an inclined blowing hole 30 is preferable in that the cooling air 600 is smoothly blown to the front surface 250 of the segment 200.

(Fifth Embodiment) FIG. 18 shows a turbine 1 of a gas turbine according to a fifth embodiment of the present invention.
It is a plane sectional view showing a shroud of a step vane. In the figure, the same components as those of the conventional gas turbine are designated by the same reference numerals, and the description thereof will be omitted. In this gas turbine, as compared with the gas turbine of the conventional technique 3, the front edge 220 side of the gap 40 of the joint surface 210 is refracted to the blade 100b side so that the blowout hole 42 is located upstream of the gap 40. It has a feature in that it is formed in. At the refraction point 41 of the gap 40, the combustion gas 500 is generated in the gap 30 of the prior art 3.
It is determined near the boundary where the streamline direction that crosses 0 is changed. Specifically, the gap 300 of the conventional technique 3 and the belly 1 in the gap 300
10 and the leading edge 41a of the wing 100a and the gap 3
It is an intersection with a straight line connecting the trailing edge portion 130 of the wing 100b with the spine 120 directed to 00. The gap 40 has this inflection point 41.
Is used as a reference, the front edge 220 side is bent toward the blade 100b side with respect to the inflow direction of the combustion gas 500. Also, wings 10
The leading edge 220 is attached to the end portion 251 of the segment 200 on the 0a side.
From this to the trailing edge 221, the blowing hole 42 of the cooling air 600 is uniformly formed. A pair of seal plates 400 are inserted between the joint surfaces 210 from both sides of the front edge 220 side and the rear edge 221 side to seal the gap 40.

In the fifth embodiment, the combustion gas 50
A part of 0 flows in almost vertically from the leading edge 220 of the shroud, and enters the gap 40 on the leading edge 220 side from the inflection point 41 on the leading edge 1 side.
Straddle from the 00a side. At this time, the blowing hole 42 is
Since it is at the end portion 251 of the segment 200 on the 00a side, the cooling air 600 is blown from the upstream side to the gap 40.
Accordingly, the cooling air 600 flows into the gap 40 while forming a film cooling layer at the end portion 251 of the segment 200, and cools the edge portion 230 of the segment 200. The other part of the combustion gas 500 is guided by the antinode 100 of the blade 100a and straddles the gap 40 from the blade 100a side on the trailing edge 221 side of the inflection point 41. At this time, the blowing hole 42
Exists in the end portion 251 of the segment 200 on the blade 100a side, the cooling air 600 is blown from the upstream side to the gap 40, and the segment 200 is formed in the same manner as the front edge 220 side.
Cool the rim 230.

According to the fifth embodiment, the joint surface 210
In the gap 40, the front edge 220 side is refracted from the refraction point 41, and the blowing hole 42 is positioned upstream of the gap 40. Therefore, the blown cooling air 600 is poured into the gap 40 without waste and the edge portion 230 of the segment 200 is discharged. Cooling. Thereby, the cooling efficiency can be improved. In addition, the blowing hole 42
Since the number of the combustion gas is approximately the same as that of the gas turbine of the related art 3, the combustion gas 500 due to the increase in the flow rate of the cooling air 600
The pressure loss of can be suppressed.

In the fifth embodiment, the position of the refraction point 41 is determined by the leading edge portion 130 of the blade 100a and the blade 1.
Although it is uniquely defined based on the trailing edge portion 140 of 00b, it should be appropriately changed to a suitable position by a design change obvious to those skilled in the art. The position of the refraction point 41 disclosed in the fifth embodiment is based on the knowledge of the inventor that it can be generally applied to the existing gas turbine if it is set near such a position.

Further, in the fifth embodiment, the refraction points 41 may be provided at a plurality of points, and a suitable number is selected by analyzing the flow path of the combustion gas by an actual machine by fluid analysis or other methods known to those skilled in the art. It may be provided at any position. Gap 40
Is refracted by the blowing holes 42 located downstream of the gap 40 with respect to the non-uniform combustion gas 500 flow.
Because it should be decided for the purpose of reducing the number of.

In the fifth embodiment, the gap 4
0 is bent with the leading edge 220 side inclined with respect to the inflow direction of the combustion gas 500, but the leading edge side 40a is formed in the same direction as the inflow direction of the combustion gas 500 (see FIG. 19), and the leading edge 220 of the segment 200 is formed. The cooling structure according to the first to third embodiments may be adopted on the side. In this variation,
The edge portion 230 on the front edge 220 side of the segment 200 is cooled by the cooling structure according to the first to third embodiments. Further, the trailing edge 221 side is cooled by the cooling air 600 blown out from the blowing hole 42. According to this modified example, by using the cooling structure according to the first to third embodiments in combination with the fifth embodiment, the combustion gas 500 enters the gap 40 of the joint surface 210 in parallel from the front edge 220 side. Also in this case, the burnout 510 of the edge portion 230 of the segment 200 can be effectively suppressed.

As described above, according to the gas turbine of the present invention (Claim 1), the air groove as the passage for the cooling air is provided on the segment side. Since it is not necessary to provide a stepped portion on the seal plate, a thinner seal plate can be adopted.

Further, according to the gas turbine of the present invention (Claim 2), the cooling air passage is formed in a crank shape by providing the air grooves at different positions in plan view with respect to the turbine passage wall. Therefore, it is possible to suppress the flow rate of the cooling air to be ejected and maintain the sealing function of the seal plate while securing the cooling effect of the edge portion of the segment.

Further, according to the gas turbine of the present invention (claim 3), since the through hole of the seal plate is formed so as to be inclined toward the trailing edge side, the cooling air makes the through hole perpendicular to the seal plate. Compared with the case of forming the cooling air layer, the cooling air layer is formed on the front surface of the seal plate by jetting smoothly.

Further, according to the gas turbine of the present invention (claim 4), since the through hole has an inclination angle at which the cooling air can appropriately cool the edge portion of the segment, burnout of the edge portion of the segment is effective. Can be suppressed to.

Further, according to the gas turbine of the present invention (claim 5), since the cooling air passages are formed in a crank shape by providing the notches at different positions in plan view, the segment edge portion While securing the cooling effect, it is possible to suppress the flow rate of the jetting cooling air and maintain the sealing function of the seal plate.

Further, according to the gas turbine of the present invention (claim 6), as compared with the gas turbine of the prior art 3, more cooling air blowing holes are provided on the upstream side with respect to the gap between the joint surfaces. Therefore, the burnout of the edge of the segment can be suppressed more effectively.

Further, according to the gas turbine of the present invention (claim 7), as compared with the gas turbine of the prior art 3, more cooling air blowing holes are provided on the upstream side with respect to the gap between the joint surfaces. Therefore, the burnout of the edge of the segment can be suppressed more effectively.

[Brief description of drawings]

FIG. 1 is a perspective view showing a main part of a gas turbine according to a first embodiment of the present invention.

2 is a view showing a joint portion of the segment shown in FIG.
FIG.

FIG. 3 is a sectional view of the shroud joint surface shown in FIG. 1 as viewed from B.

FIG. 4 is a perspective view showing a joint surface of the segment shown in FIGS.

5 is a side sectional view showing a modification of the gas turbine shown in FIG.

FIG. 6 is a side sectional view showing a modified example of the gas turbine shown in FIG. 1.

FIG. 7 is a plan view showing a modified example of the gas turbine shown in FIG.

FIG. 8 is a perspective view showing a main part of a gas turbine according to a second embodiment of the present invention.

9 is a cross-sectional view taken along the line C of FIG. 8 showing the joint portion of the segment shown in FIG.

FIG. 10 is a perspective view showing a main part of a gas turbine according to a third embodiment of the present invention.

FIG. 11 is a cross-sectional view taken along line DD of the segment shown in FIG.

FIG. 12 E showing the joint of the segment shown in FIG.
It is a -E sectional view.

13 is a perspective view showing a main part of the seal plate shown in FIGS. 10 to 12. FIG.

FIG. 14 is a perspective view showing a turbine first stage stationary blade of a gas turbine according to a fourth embodiment of the present invention.

FIG. 15 is a plan sectional view showing a shroud of the turbine first stage stationary blade shown in FIG. 14.

16 is a cross-sectional view taken along line F-F of the joint portion of the segment shown in FIG.

FIG. 17 is a plan sectional view showing a shroud of the turbine first stage vane shown in FIG. 14.

FIG. 18 is a plan sectional view showing a shroud of a turbine first stage stationary blade of a gas turbine according to a fifth embodiment of the present invention.

19 is a plan sectional view showing a modified example of the gas turbine shown in FIG.

FIG. 20 is an overall configuration diagram showing a turbine vane of a conventional gas turbine.

21 is an assembled perspective view showing a joint portion of the segment shown in FIG. 20. FIG.

FIG. 22 is a front sectional view showing a main part of a conventional gas turbine.

FIG. 23 is a front sectional view showing a turbine vane of a gas turbine of Prior Art 1.

FIG. 24 is a perspective view showing the seal plate shown in FIG. 23.

FIG. 25 is a front cross-sectional view showing a turbine stationary blade of a gas turbine of Prior Art 2.

FIG. 26 is a perspective view showing the seal plate shown in FIG. 25.

FIG. 27 is a side sectional view of the joint section shown in FIG. 25.

FIG. 28 is a side sectional view showing a case where a through hole of a seal plate included in the gas turbine shown in FIG. 9 has an inclination angle of 30 degrees.

FIG. 29 is a perspective view showing a main part of a turbine vane according to Conventional Technique 3.

FIG. 30 is a front sectional view showing a joint portion of the segment shown in FIG. 29.

FIG. 31 is a plan sectional view showing a shroud of the turbine first stage stationary blade shown in FIG. 29.

[Explanation of symbols]

1 air groove 10 Seal plate 11 through holes 20 seal plate 21 Notch 30 blowout holes 31 Switching point 40 gap 41 refraction point

Claims (7)

[Claims] 1. A turbine split ring, a turbine rotor blade platform, a turbine vane shroud, and other turbine channel walls that are formed by joining a plurality of segments in an annular shape and that form a channel wall for combustion gas. A seal plate that is fitted into seal grooves formed on the facing joint surfaces between the segments and that seals a gap between the joint surfaces across the joint surfaces; and an inner surface of the seal groove The gas turbine is characterized in that an air groove for passing cooling air is provided from the back side of the turbine flow path wall to the front side of the seal plate. 2. The gas turbine according to claim 1, wherein the air groove is provided at a position that is different in plan view between the front surface side and the rear surface side with respect to the turbine flow path wall. 3. A turbine dividing wall, which is formed by joining a plurality of segments in an annular shape and constitutes a passage wall for combustion gas, a platform of turbine blades, a turbine vane shroud, and other turbine passage walls, While being fitted in a seal groove formed in the facing joint surface between the segments, a seal plate that spans between the joint surfaces to seal a gap between the joint surfaces, and the seal plate, A gas turbine having a through hole inclined toward the trailing edge of the turbine flow path wall. 4. The inclination angle of the through hole is set so that the jetting cooling air is larger than a lower limit value reaching the edge of the turbine passage wall, and exceeds the edge of the turbine passage wall. The gas turbine according to claim 3, wherein the angle is set so as to be smaller than an upper limit value mixed into the combustion gas. 5. A turbine dividing wall, which is formed by joining a plurality of segments in an annular shape, and constitutes a passage wall for combustion gas, a turbine blade platform, a turbine vane shroud, and other turbine passage walls, While being fitted in a seal groove formed in the joint surface between the segments facing each other, a seal plate that extends between the joint surfaces to seal a gap between the joint surfaces, and the seal plate, A gas turbine characterized in that cutouts are provided at side portions of the upper and lower surfaces thereof, which are two-dimensionally different on the upper surface side and the lower surface side with respect to the turbine flow path wall. 6. A turbine split ring, a turbine, which is formed by joining a turbine moving blade, a turbine vane, and other blades, and a plurality of segments supporting the blade in an annular shape, and which forms a flow path wall for combustion gas. A blade platform, a turbine vane shroud or other turbine channel wall, and a segment between one of the segments that crosses the flow direction of the combustion gas across the gap based on the gap between the facing joint surfaces of the segment. A plurality of cooling air blowing holes formed by switching from one end to the other segment end. 7. A turbine split ring, a turbine, which is formed by joining a turbine moving blade, a turbine vane, and other blades, and a plurality of segments supporting the blade in an annular shape, and which constitutes a flow path wall for combustion gas. A turbine blade platform, a turbine vane shroud and other turbine flow path walls, and a plurality of cooling air blowing holes formed at the end of the segment located on the ventral side of the blade, and The gap between the joint surfaces facing each other, with the vicinity of the vicinity where the flow direction of the combustion gas across the gap changes, as a boundary, the front edge side is refracted to the wing side with its back to the gap,
A gas turbine characterized in that the blow-out hole is formed upstream of the gap.