JP2010077868A - Rim seal structure of gas turbine - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a rim seal structure of a gas turbine capable of securing a sufficient seal function and reducing a purge air volume required for preventing the intrusion of mainstream high-temperature gas in a disk cavity. <P>SOLUTION: The rim seal structure of the gas turbine reduces the intrusion of the mainstream high-temperature gas supplied from a combustion section into a turbine body through the disk cavity 11 formed between a rotor blade 20 and a stator blade 30. The structure includes a first shock absorbing cavity 22 formed between the stator blade 30 and a thin part formed at the lower surface of the platform 21 of the rotor blade 20, and a second shock absorbing cavity 14 formed inward of the first shock absorbing cavity 22 in a rotor radial direction. The first shock absorbing cavity 22 is positioned in an intrusion path upstream of a portion where a mainstream high-temperature gas flow reaches the second shock absorbing cavity 14, and has a shape directing the velocity component of the mainstream high-temperature gas flow intruding into the first shock absorbing cavity 22 to the peripheral direction. <P>COPYRIGHT: (C)2010,JPO&INPIT

Description

  The present invention relates to a rim seal structure for a gas turbine.

Conventionally, a mainstream high-temperature gas (also referred to as “combustion gas” or “hot gas”) that is supplied from a combustion section (combustor) and drives the turbine is upstream of the stationary blade of the gas turbine. A sealing mechanism called a rim seal is installed for the purpose of preventing the turbine from entering the turbine body. This rim seal prevents the mainstream hot gas from entering by ejecting a purge fluid purge air from a disk cavity formed between the moving blade and the stationary blade.
The purge air used here is a part of compressed air introduced from an outside blade cooling air supply source such as a compressor, and the inside blade air flow path in the stationary blade connected to the outside blade cooling air supply source. , And is injected into the disk cavity from a seal air supply hole provided through an appropriate wall surface of the stationary blade component on the stationary blade tip (axial center) side.
The main reason for the necessity of the rim seal structure is to prevent the blade root of the rotor blade and the rotor disk that receive a large load during operation from becoming hot.

As the rim seal upstream of the stationary blade described above, for example, a single overlap seal structure (see, for example, Patent Document 1) having a configuration in which a stationary blade rim is disposed on a moving blade seal fin, or a seal fin is used. There is a double overlap seal structure (see, for example, Patent Documents 2, 3, and 4) that are arranged twice.
Japanese Patent Laid-Open No. 10-259703 (see FIG. 2) Japanese Patent Laid-Open No. 8-311983 (see FIG. 5) US Pat. No. 6,506,016 (see FIG. 1) US Pat. No. 6,884,028 (see FIG. 1) JP 2001-115801 A JP 2007-85340 A

In the turbine section of a gas turbine, the main cause of hot gas entrainment in which mainstream high-temperature gas enters the turbine body from the gap between the rotor blades and the stationary blades is largely due to the circumferential static pressure distribution of the gas path. Such a circumferential static pressure distribution is caused by the wake of the upstream trailing edge (rear edge) and the potential of the downstream blade, which is an unavoidable problem in current gas turbines.
In the rim seal structure that prevents the entrainment of hot gas, it is desired to reduce the amount of purge air that affects the gas turbine performance. Also, since the purge air uses a part of the compressed air supplied from the compressor, the reduction of the purge air amount means that the proportion of the compressed air amount that can be used for combustion in the combustor increases. Contributes to performance improvement.

From such a background, in the rim seal structure of a gas turbine, it is desired to secure a sufficient sealing function and reduce the amount of purge air that is ejected into the disk cavity to prevent the mainstream high-temperature gas from entering.
The present invention has been made in view of the above circumstances, and its object is to ensure a sufficient sealing function and reduce the amount of purge air necessary for preventing mainstream hot gas from entering the disk cavity. An object of the present invention is to provide a rim seal structure for a gas turbine.

  The gas turbine rim seal structure according to the present invention is a gas turbine rim seal structure that reduces the mainstream high-temperature gas supplied from the combustion section from entering a disk cavity formed between the rotor blades and the stationary blades. A first buffer cavity is formed between a thin wall provided on the lower surface of the platform and the stationary blade side end surface, and a second buffer cavity is formed on the inner side in the rotor radial direction from the first buffer cavity. The mainstream hot gas flow is positioned in the intrusion path before reaching the second buffer cavity, and the mainstream hot gas flow entering the first buffer cavity has a shape in which a velocity component is directed in the circumferential direction. have.

  According to such a rim seal structure of a gas turbine, a first buffer cavity formed between a thin fillet provided on the platform lower surface of the rotor blade and a stationary blade side end surface, and a rotor radially inner side from the first buffer cavity. The first buffering cavity is located in an intrusion path before the flow of the mainstream hot gas reaches the second buffering cavity, and the mainstream hot gas entering the first buffering cavity is included in the first buffering cavity. Since the flow component has a shape that directs the velocity component in the circumferential direction, when a part of the mainstream hot gas flowing through the gas path flows into the first buffer cavity of the disk cavity due to the influence of the circumferential static pressure distribution, The static pressure in the circumferential direction approaches uniformly in the first buffer cavity at the upstream position before reaching the two buffer cavities. That is, the mainstream high-temperature gas that has flowed into the intrusion path of the disk cavity from the inlet opening formed between the rotor blade and the stationary blade is sufficiently mixed in the circumferential direction by the first buffer cavity formed by the meat fillet and heads toward the rim seal portion. The pressure distribution in the circumferential direction of the flow can be relaxed.

In the above invention, since the fluid supply hole for allowing the seal fluid to flow out from the inside of the stationary blade toward the rim seal portion is provided, the sealing performance of the rim seal portion can be improved.
In this case, the fluid supply hole is preferably provided in a plug that can be attached to and detached from the stationary blade, whereby the flow rate of the sealing fluid that flows out to the rim seal portion can be easily changed.

  A gas turbine according to the present invention includes a compression unit that compresses combustion air, a combustion unit that injects fuel into high-pressure air sent from the compression unit to generate a mainstream high-temperature gas, and a downstream of the combustion unit. And a turbine section that is driven by the mainstream hot gas, and the turbine section includes the rim seal structure according to any one of claims 1 to 3.

  According to such a gas turbine, since the turbine portion includes the rim seal structure according to any one of claims 1 to 3, the mainstream high temperature that has flowed into the intrusion path of the gap formed between the moving blades and the stationary blades Since the gas is sufficiently mixed in the circumferential direction by the cavity space formed by the fillet, the pressure distribution in the radial direction of the flow toward the rim seal portion can be relaxed.

According to the rim seal structure of the gas turbine of the present invention described above, when a part of the mainstream hot gas flowing through the gas path flows into the first buffer cavity from the inlet opening between the moving blade / stator blade, the second buffer cavity is reached. Since the static pressure in the circumferential direction approaches uniformly in the first buffer cavity in the previous upstream position, the flow velocity and pressure of the flow from the first buffer cavity to the second buffer cavity are reduced in the disk cavity. As a result, in the rim seal portion of the gas turbine provided with the first buffer cavity and the second buffer cavity, the flow rate of the seal fluid required for sealing the mainstream high temperature gas can be reduced.
Therefore, in the rim seal portion of the gas turbine, it is possible to secure a sufficient sealing function and reduce the flow rate of the sealing fluid that is ejected into the disk cavity to prevent the mainstream high-temperature gas from entering. It has a remarkable effect on efficiency improvement. In other words, the gas turbine employing the above-described rim seal structure of the present invention improves operational efficiency by ensuring both a sufficient sealing function and a reduction in the flow rate of the sealing fluid necessary to prevent the inflow of mainstream high-temperature gas. Can be made.

Hereinafter, an embodiment of a rim seal structure for a gas turbine according to the present invention will be described with reference to the drawings.
As shown in FIG. 4, the gas turbine 1 includes a compression unit (compressor) 2 that compresses combustion air, and injects and burns fuel into the high-pressure air sent from the compression unit 2 to perform high-temperature combustion. A combustion section (combustor) 3 that generates gas, and a turbine section (turbine) 4 that is located on the downstream side of the combustion section 3 and is driven by high-temperature combustion gas (mainstream high-temperature gas) that exits the combustion section 3 The main element.

FIG. 1 is a cross-sectional view of an essential part showing a first embodiment of a rim seal structure provided in a turbine part 4 of a gas turbine 1.
The illustrated rim seal 10 reduces the mainstream high-temperature gas supplied from the combustion unit 3 from entering the gap (hereinafter referred to as “disk cavity”) 11 formed between the moving blade 20 and the stationary blade 30. For this purpose, the seal structure is installed on the upstream side of the stationary blade 30. The upstream side in this case is based on the flow direction of the mainstream hot gas indicated by the arrow G in the drawing. That is, in the rim seal 10, the mainstream hot gas flows in the axial direction (see arrow G shown in the drawing) in the gas path 12 in which the moving blade 20 and the stationary blade 30 are disposed, and thus a part of the rim seal 10 is separated from the mainstream hot gas. (Refer to the arrow g shown in the figure) is caught in the disk cavity 11 and is prevented or suppressed from entering the blade root of the moving blade 20 and the vicinity of the rotor.
In the following explanation regarding the mainstream hot gas, the entire mainstream hot gas flowing through the gas path 12 is referred to as “mainstream hot gas G”, and a part of the hot mainstream gas that has been divided is referred to as “hot branch gas g”. To do.

  As shown in FIGS. 1 to 3, the platform 21 of the moving blade 20 includes a first buffer cavity 22 provided on the lower surface. The first buffer cavity 22 is upstream of a stationary blade rim 31 provided on a member surface of the platform 21 on the downstream side of the moving blade 20 in the flow direction of the mainstream hot gas G, that is, on the stationary blade 30 side located on the downstream side. It is a concave portion of the meat thinning formed on the surface facing the side end surface 31a.

  The first buffer cavity 22 is formed by forming a hollow space in the region closer to the upstream side than the downstream end surface 21a of the platform 21 facing the gas path 12. On the lower side, rotor blade outer peripheral seal fins 23 are formed. In other words, the recessed space of the first buffer cavity 22 is formed on the upstream side of the second buffer cavity 14 described later in the rim seal formed between the rotor blade outer peripheral seal fin 23 and the stationary blade rim 31. . That is, the first buffer cavity 22 is divided between the flow direction of the high-temperature diverted gas g that diverts from the gas path 12 toward the center of the rotation axis and enters the disk cavity 11, and the flow direction of the mainstream high-temperature gas G that is substantially parallel to the rotation axis. In both directions, it is located upstream of the second buffer cavity 14 constituting the rim seal.

  Accordingly, the concave space of the first buffer cavity 22 is located upstream of the second buffer cavity 14 in the intrusion path where the hot diverted gas g branched from the flow of the mainstream hot gas G reaches the disk cavity 11, and The flow of the high-temperature diverted gas g that has entered the buffer cavity 22 has a shape that directs the velocity component in the circumferential direction. The moving blade 20 having the first buffer cavity 22 is usually a cast part integrally formed with the platform 21. Therefore, the first buffer cavity 22 is also formed by providing a thin portion at the time of casting. It can be easily manufactured by machining.

  In the intrusion path of the high-temperature diverted gas g that diverts from the gas path 12 to the disk cavity 11 side, the first buffer cavity 22 described above is a recessed space formed at the entrance of the intrusion path immediately after the diversion. In the disk cavity 11 on the downstream side of the first buffer cavity 22 in the intrusion path, the rotor blade outer peripheral seal fins 23 and the rotor blade inner peripheral seal fins 25 are arranged in duplicate to form a rim seal portion.

A second buffer cavity 14 is formed further on the inner peripheral side than the rotor blade outer peripheral seal fin 23, that is, on the rotation shaft center side (lower side in the drawing). The second buffer cavity 14 faces the stationary blade inner peripheral fin 32 and is surrounded by the inner peripheral surface of the stationary blade rim 31, the inner peripheral surface of the moving blade outer peripheral seal fin 23, and the outer peripheral surface of the moving blade inner peripheral seal fin 25. Space. The second buffer cavity 14 is located downstream of the first buffer cavity 22. In other words, the space portion of the second buffer cavity 14 is formed on the downstream side of the rim seal portion formed between the moving blade inner peripheral seal fin 25 and the stationary blade rim 31.
Therefore, the space portion of the second buffer cavity 14 is located downstream of the first buffer cavity 22 in the intrusion path where the high-temperature diverted gas g branched from the flow of the mainstream high-temperature gas G reaches the disk cavity 11, and the second buffer cavity 14. The flow of the hot diverted gas g that has entered the buffer cavity 14 is sufficiently mixed.

  The first-stage overlap seal portion (hereinafter referred to as “first-stage seal portion”) 13 reduces the cross-sectional area of the flow path and increases the flow path resistance toward the center of the rotation axis. The rotor blade outer peripheral seal fin 23 and the stationary blade rim 31 are disposed at a position downstream of the first buffer cavity 22 in the flow direction. The illustrated outer peripheral seal fin 23 of the moving blade extends the lower surface portion in which the first buffer cavity 22 is formed in the platform 21 to the stationary blade 30 side, and protrudes outward in the circumferential direction (upward on the paper surface) of the turbine unit 4 from the tip portion. A fin is formed. The rotor blade outer peripheral seal fins 23 are arranged so as to overlap with a stator blade rim 31 protruding from the stator blade 30 toward the rotor blade 20 in a state having a predetermined gap in the circumferential direction.

  The disk cavity 11 further downstream from the first stage seal portion 13 is provided with a second buffer cavity 14 for entraining the flow of the hot diverted gas g that has passed through the first stage seal portion 13. A second-stage overlap seal portion (hereinafter referred to as “second-stage seal portion”) 15 is provided at a position passing through the second buffer cavity 14. The second-stage seal portion 15 is located at a position downstream of the second buffer cavity 14 in the flow direction of the high-temperature diverted gas g so as to increase the flow path resistance toward the rotation axis center side by narrowing the flow path cross-sectional area. The moving blade inner peripheral seal fin 25 and the stationary blade inner peripheral fin 32 are provided. The illustrated moving blade inner peripheral seal fin 25 protrudes from the wake seal plate 24 on the moving blade 20 side attached to the lower side of the platform 21 to the stationary blade 30 side, and forms a fin protruding outward in the circumferential direction from the tip. It is a thing. The rotor blade inner peripheral seal fin 25 is disposed so as to overlap with the stator blade inner peripheral fin 32 protruding from the stator blade 30 toward the rotor blade 20 in a state having a predetermined gap in the circumferential direction.

  Reference numeral 35 in the figure denotes a buffer plate, which receives the bias of the coil spring 36 to bring the seal portion 37 into close contact with a stationary blade constituent member such as the seal ring holding ring 33 and the like. Is forming. Since the blade inner passage 38 communicates with the discharge side of the compression section 2 via an inner blade passage (not shown) formed inside the stationary blade 30, a part of the compressed air can be introduced. It can be done.

In the rim seal portion of the double overlap seal structure configured as described above, a part of the compressed air is introduced as a seal fluid through the blade flow path formed in the stationary blade 30 and the like. In the following description, a part of the compressed air introduced from the compression unit 2 as a sealing fluid will be referred to as “cooling air” for use as the purge air for the disk cavity 11.
The cooling air introduced into the blade flow path 38 in the stationary blade 30 penetrates through the wall surface of a member constituting the stationary blade 30 such as a seal ring holding ring 33 used for attaching a labyrinth seal (not shown). Jetted as purge air from the provided seal air supply hole 34. This cooling air is diverted after flowing out from the seal air supply hole 34, and flows in the seal gap 11a as indicated by a broken line arrow Ca in the drawing. The cooling air is diverted, and one of the diverted cooling air is sucked and flows out to the inner side of the rotor, and the other cooling air flows into the disk cavity 11 through the seal gap 11a and becomes purge air. Therefore, the cooling air injected from the stationary blade 30 into the disk cavity 11 is purged, so that the disk cavity 11 is maintained at a pressure higher than the circumferential static pressure distribution when there is no purge air. In addition, cooling with cooling air having a low temperature is also performed.

  On the other hand, the mainstream hot gas G flowing in the gas path 12 may be partially higher than the pressure of the purge air in the radial direction. 20 enters the disk cavity 11 through an inlet opening 11b formed between the blade 20 and the stationary blade 30. Such intrusion of the high temperature diverted gas g can be prevented or suppressed by setting the purge air pressure in the disk cavity 11 high. However, in order to set the purge air pressure in the disk cavity 11 high, it is necessary to increase the amount of cooling air introduced from the compression unit 2, and as a result, the temperature of the mainstream hot gas G decreases and the combustion air is supplied to the combustion unit 3. It becomes a factor which reduces the operating efficiency of the gas turbine 1 by the reduction | decrease in the amount of compressed air for combustion etc. which is possible. Therefore, regarding the setting of the purge pressure, it is preferable to prioritize the operation efficiency of the gas turbine 1 in consideration of various conditions.

  The hot diverted gas g that has entered the disk cavity 11 from the inlet opening 11 b formed between the rotor blade 20 and the stationary blade 30 first flows into the first buffer cavity 22, and is more fully contained in the first buffer cavity 22. Stir and mix to relieve pressure distribution in the circumferential direction. Such mixing is facilitated by the fact that the first buffer cavity 22 is located on the upstream side of the first-stage seal portion 13 having a high flow resistance, and a relatively large volume space that easily forms a flow along the wall surface. It is because it has.

  For this reason, the high-temperature diverted gas g directed to the first stage seal portion 13 has a reduced radial pressure distribution compared to the case where the first buffer cavity 22 is not provided, and therefore a disk required to obtain the same sealing performance. The purge air pressure in the cavity 11 can be reduced. Therefore, it is possible to maintain the sealing performance of the rim seal 10 and reduce the amount of cooling air required for the purge air.

  Therefore, in the above-described double overlap seal structure, the main flow shunt gas g reaching the inlet opening 11b of the disk cavity 11 formed between the rotor blade 20 and the stationary blade 30 is first due to the presence of the rotor blade outer peripheral fins 23. The velocity of the high-temperature diverted gas g diffused in the buffer cavity 22 (see FIG. 3) and passing through the first-stage seal portion 13 is reduced. Further, the high-temperature diverted gas g is diffused d2 (see FIG. 3) into the second buffer cavity 14 due to the presence of the rotor blade inner peripheral fins 25, and the speed of the high-temperature diverted gas g passing through the second stage seal portion 15 is Even smaller. That is, the main flow shunt gas g reaching the disk cavity 11 formed between the moving blade 20 and the stationary blade 30 enters the first buffer cavity 22 provided on the outer side in the circumferential direction of the rotor shaft, The flow velocity component is diffused in the circumferential direction by being guided by the wall surface extending in the direction, and the downward speed of passing through the rotor blade outer peripheral fins 23 and the rotor blade inner peripheral fins 25 in the radial direction of the high-temperature diverted gas g is reduced. Can do.

Next, a second embodiment of the rim seal structure provided in the turbine portion 4 of the gas turbine 1 will be described with reference to FIG. FIG. 5 is a cross-sectional view of the main part showing the rim seal structure. The same reference numerals are given to the same parts as those in the above-described embodiment, and the detailed description thereof will be omitted.
The rim seal 10 </ b> A described in this embodiment is provided with a fluid supply hole 40 through which the cooling air Ca of the sealing fluid flows out from the inside of the stationary blade 30 toward the rim seal portion. In the illustrated configuration example, a plurality of fluid supply ports 40 communicating with the in-blade channel 38 of the stationary blade 30 are formed in the wall surface connecting the stationary blade rim 31 of the stationary blade 30 and the stationary blade inner peripheral fin 32. Has been. The fluid supply hole 40 is provided so that a part of the cooling air Ca that flows toward the seal air supply hole 34 through the blade inner flow path 38 flows toward the first stage seal portion 13.

  For this reason, since the cooling air Ca flows out in the first stage seal portion 13 in a direction that prevents the flow of the high-temperature diverted gas g entering the second buffer cavity 14, the sealing performance of the rim seal portion is improved. be able to. In particular, it is extremely difficult for the hot diverted gas g whose flow velocity and pressure have decreased in the first buffer cavity 22 to pass through the first-stage seal portion 13 against the flow of cooling air flowing out from the fluid supply hole 40. Therefore, it is effective for improving the sealing performance.

Moreover, FIG. 6 is sectional drawing of the principal part which shows the modification of the fluid supply hole 40 about 2nd Embodiment mentioned above. In addition, the same code | symbol is attached | subjected to the part similar to embodiment mentioned above, and the detailed description is abbreviate | omitted.
In this modification, a plug 50 is provided in the fluid supply hole 40. The plug 50 is detachably attached to a wall surface connecting the stationary blade rim 31 of the stationary blade 30 and the stationary blade inner peripheral fin 32 by, for example, screwing or the like. Accordingly, the flow rate of the cooling air Ca flowing out to the rim seal portion can be easily changed and made variable by preparing a plurality of plugs 50 having different diameters of the fluid supply holes 40. That is, the amount of cooling air that flows out to the rim seal portion can be appropriately adjusted according to various conditions such as installation environment and use conditions, or as adjustment during manufacturing.

FIG. 7 is a diagram for explaining the operation of the fluid supply hole 40 described above, and shows the relationship between the cavity temperature and the total seal air amount (cooling air amount).
In FIG. 7, a solid line A in the figure indicates the cavity temperature of the disk cavity 11 (point A in FIGS. 5 and 6) on the downstream side (circumferential inner side) from the second stage seal portion 15. A solid line B indicates the cavity temperature of the second buffer cavity 14 (point B in FIGS. 5 and 6) on the downstream side (circumferential inner side) from the first stage seal portion 13.
The cavity temperature t1 is the temperature set at the cavity point A in the disc cab 11 that is on the inner side in the circumferential direction, and the cavity temperature t2 is the temperature that is set at the cavity point B in the second buffering cavity 14 on the outer side in the circumferential direction. In the configuration in which the supply hole 40 is not provided, the amount of sealing air is set to Q2 in order to set the solid lines A and B to stable cavity temperatures below the set temperatures t1 and t2.
The rotor blade outer peripheral seal fin 23 can withstand a higher temperature because it is made of a high temperature member constituting the rotor blade. Therefore, in order to maintain the cavity temperature t2 in the outer circumferential direction, What is necessary is to supply at least the sealing air amount Q3 corresponding to the point P1.

On the other hand, when the fluid supply hole 40 is provided, the cooling air supplied from the fluid supply hole 40 is directly supplied with the cooling air in the vicinity at the point B, and therefore the temperature at the cavity point B of the second buffer cavity 14. Can be made lower (see broken line B ').
At this time, the amount of cooling air supplied from the fluid supply hole 40 is Q _B , the amount of cooling air supplied from the rotor shaft side is Q _A, and the fluid supply hole for the total seal air amount Q (Q = Q _A + Q _B ) In the same manner as in the case of not providing 40, the cooling air amount Q _A decreases compared to the case where the fluid supply hole 40 is not provided, and the temperature at the cavity point A on the inner side in the circumferential direction slightly increases (dashed line A). However, the total sealing air amount Q can be reduced by the direct cooling effect at the point B.
Therefore, in order to maintain the cavity temperature t2 or less at the cavity point B and maintain the cavity temperature t1 or less at the cavity point A, the total seal air amount Q1 corresponding to the point P2 in the figure may be supplied. As a result, the total seal air amount Q can be reduced by the difference between Q3 and Q1.

According to the above-described present invention, when a part of the mainstream hot gas G flowing through the gas path 12 becomes the high-temperature diverted gas g and flows into the first buffer cavity 22, the first buffer cavity 22 in the upstream position of the rim seal portion. Since the static pressure in the circumferential direction approaches uniformly, the flow velocity and pressure of the flow toward the rim seal portion are reduced, and the flow rate of the cooling air Ca required for sealing the high temperature diverted gas g can be reduced in the rim seal portion.
Therefore, it is possible to secure a sufficient sealing function and reduce the flow rate of the cooling air Ca that is ejected into the disk cavity 11 to prevent the high temperature diverted gas g from entering, so that the efficiency of the gas turbine 1 is improved. be able to. That is, by adopting the above-described rim seal structure of the present invention, it is possible to achieve both a sufficient sealing function and a reduction in the flow rate of cooling air necessary for preventing the inflow of the mainstream high-temperature gas, and a gas turbine 1 with high operating efficiency. Can be provided.
In addition, this invention is not limited to embodiment mentioned above, In the range which does not deviate from the summary of this invention, it can change suitably.

It is principal part sectional drawing which shows 1st Embodiment about the rim seal structure of the gas turbine which concerns on this invention. It is a perspective view of the principal part which shows the buffer cavity provided in the platform lower surface of a moving blade. It is a longitudinal cross-sectional view of the gas turbine which shows the whole positional relationship of a 1st buffer cavity and a 2nd buffer cavity. It is a section perspective view showing the outline of a gas turbine. It is principal part sectional drawing which shows 2nd Embodiment about the rim seal structure of the gas turbine which concerns on this invention. It is principal part sectional drawing which shows the modification of FIG. 5 about the rim seal structure of the gas turbine which concerns on this invention. It is a figure explaining the effect | action of a fluid supply hole, and the relationship between cavity temperature and the amount of seal air is shown.

Explanation of symbols

DESCRIPTION OF SYMBOLS 1 Gas turbine 2 Compression part 3 Combustion part 4 Turbine part 10,10A Rim seal 11 Disc cavity 12 Gas path 13 First stage overlap seal part (first stage seal part)
14 Second buffer cavity 15 Second stage overlap seal part (second stage seal part)
20 Rotor 21 Platform 22 First buffer cavity 23 Rotor outer peripheral seal fin 25 Rotor inner peripheral seal fin 30 Stator vane 31 Stator vane rim 40 Fluid supply hole 50 Plug

Claims (4)

In a rim seal structure of a gas turbine for reducing mainstream hot gas supplied from a combustion section from entering a disk cavity formed between a moving blade and a stationary blade,
A first buffering cavity formed between a fillet provided on the lower surface of the platform of the moving blade and the stationary blade side end surface, and a second buffering cavity formed on the inner side in the rotor radial direction from the first buffering cavity, The first buffer cavity is located in an intrusion path before the flow of the mainstream hot gas reaches the second buffer cavity, and the flow of the mainstream hot gas that has entered the first buffer cavity in the circumferential direction. A gas turbine rim seal structure having a shape for directing a velocity component.   The rim seal structure for a gas turbine according to claim 1, wherein a fluid supply hole for allowing a seal fluid to flow from the inside of the stationary blade toward the rim seal portion is provided.   The rim seal structure for a gas turbine according to claim 2, wherein the fluid supply hole is provided in a plug detachable from the stationary blade. A compression section for compressing combustion air; a combustion section for injecting fuel into high-pressure air sent from the compression section to generate mainstream hot gas; and the mainstream hot gas located downstream of the combustion section A gas turbine comprising a rim seal structure according to any one of claims 1 to 3.

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