JPH10184310A - Gas turbine stationary blade - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a gas turbine stationary blades which prevent an end wall frame part of the stationary blade from melting when a gas turbine working gas temperature is set at high temperature aiming at a high efficiency and which can avoid crack due to oxidation or thermal stress and which has high reliability and durability. SOLUTION: A cooling passage in which a cooling medium flows from upstream to downstream parallel with a divided surface of a stationary blade segment is formed between a seal plate 52 and seal plate grooves 53a, 53b mounted in an end wall part 4 when respective stationary blade segments forming cyclic blade row are coupled in the end wall part 4 to each other by using the seal plate 52. A heating accelerator is disposed in the cooling passage to accelerate heat transfer.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

The present invention relates to a gas turbine stationary blade.

[0002]

2. Description of the Related Art A gas turbine generally generates a working fluid compressed to a high pressure by a compressor directly connected to the gas turbine, and adds fuel to the working fluid and burns the working fluid.
It is configured to drive a turbine by obtaining a high-temperature and high-pressure working fluid. The rotational energy of the turbine is converted to electrical energy by, for example, a generator coupled to the turbine.

At present, great expectations are placed on improving the performance and efficiency of a combined cycle combining a gas turbine and a steam turbine. In order to improve the performance and efficiency of the combined cycle, it is necessary to further increase the temperature and pressure of the working fluid of the gas turbine. Therefore, even if the temperature of the working fluid of the gas turbine is made higher than before, a high-performance and reliable cooling technology is required so that the members exposed to the high-temperature gas do not melt or crack due to oxidation or thermal stress. Becomes

When the working fluid of a gas turbine is heated to a high temperature, it is the first stage stationary blade that has the largest heat load and is difficult to cool, and not only the blade portion but also the end wall portion that constitutes the gas path flow path in the circumferential direction. Active cooling is required as well as the part. In a gas turbine, the vanes are circumferentially divided into a plurality of segments and arranged, and the vanes are connected to each other by using a seal plate. In most cases, the stationary blade endwall is impingement-cooled from the inner surface side, and the cooling air after the impingement cooling is blown out from the film hole to the mainstream gas side to film-cool the endwall surface. However, the portion of the frame portion of the stationary blade end wall where the seal plate is inserted cannot be subjected to impingement cooling, and is very difficult to perform cooling. In gas turbines where the working gas temperature is not very high, this part has not been actively cooled, but has been designed and operated with the aim of cooling by small leaks of cooling air from the seal plate. However, when the working gas temperature is increased in the future, such an uncertain cooling method cannot provide a highly reliable and durable gas turbine vane. In fact, some gas turbines having cracks in the vicinity of the seal plate insertion groove of the stator blade end wall have been observed in recent years in high-temperature gas turbines.

Japanese Patent Laid-Open No. 3-213602 discloses a cooling structure in the vicinity of a groove for inserting a stationary blade end wall seal plate.
FIG. 11 shows a cross-sectional shape of the seal plate insertion portion.
According to this, a small groove 38 is formed in the seal plate groove 30.
To form a flow in which the cooling air flows around the sealing member 40 from the outer surface 44 to the inner surface 42 to cool the inner wall 32. The cooling air that has cooled the inner wall 32 is blown into the mainstream gas.

[0006]

In the above cooling structure, the cooling air convectively cools the inner wall 32.
Considering the heat load of the first stage stationary blade of the high-temperature gas turbine, the speed of the cooling air flow required for cooling this portion is about several hundred meters per second. That is, a large amount of cooling air is discharged from this portion into the mainstream gas. In addition, the heat exchange distance of the cooling air is a very short distance about the depth of the seal plate groove, and the cooling air is released into the mainstream gas with almost no use of its own cooling potential. Become. That is, heat exchange by cooling is not performed so much that the gas is discharged into the mainstream gas at a low temperature.
If a large amount of low-temperature cooling air is released into the mainstream gas, it not only disturbs the flow of the mainstream gas but also lowers the temperature of the mainstream gas, which may lead to a reduction in the efficiency of the entire gas turbine.

Even if the operating gas temperature of the gas turbine is raised to increase the efficiency, if the flow rate of the cooling air is too large, it will only lead to a decrease in efficiency. The cooling efficiency will be increased and the flow rate of the cooling air will be reduced as much as possible. In addition, it is the key to high efficiency that the cooling air released into the mainstream gas uses up its cooling potential as much as possible, that is, heat exchange and release at a high temperature.

[0008]

That is, the present invention relates to a stationary cascade formed in an annular shape, wherein the cascade has end walls forming gas flow paths on an outer peripheral side and an inner peripheral side, respectively. The row is divided into a number of segments in the circumferential direction, and the vane segments adjacent to each other are connected to each other by inserting a seal plate into a seal plate groove provided on the end wall dividing surface. The intended purpose is achieved by forming a cooling flow channel for flowing a cooling medium in parallel with the end wall dividing surface between the seal plate groove and the seal plate with the blade.

[0009] Further, the stationary cascade is formed in an annular shape, and the cascade has end walls forming gas flow paths on an outer peripheral side and an inner peripheral side, respectively. The gas turbine vanes are divided into segments, and the adjacent vane segments are gas turbine vanes connected by inserting the seal plate into seal plate grooves provided on the end wall dividing surface, and the cross-sectional shape of the seal plate Is formed in a convex shape, so that a cooling channel for flowing a cooling medium in parallel with the end wall dividing surface is formed between the seal plate groove and the seal plate.

Further, a turbulence promoter is arranged in the cooling flow path.

[0011] Further, by forming a cross section of the seal plate in a convex shape, a cooling flow path for flowing a cooling medium in parallel with an end wall dividing surface is formed between the seal plate groove and the seal plate, A turbulence promoter is arranged on a valley surface of the seal plate that forms the cooling flow path.

[0012] Further, the stationary cascade is formed in an annular shape, and the cascade has end walls forming gas flow paths on an outer peripheral side and an inner peripheral side, respectively. The gas turbine vanes are divided into segments, and the adjacent vane segments are gas turbine vanes connected by inserting a seal plate into seal plate grooves provided on an end wall dividing surface. A film cooling hole is provided at an exposed portion of the gap existing between the wall dividing surfaces to the gas passage of the seal plate.

That is, with the gas turbine stationary blade formed as described above, it is possible to perform convection cooling from the inside of the end wall frame portion where the seal plate groove which is difficult to cool is formed, and film cooling from the outside. . Since the convection cooling flow path can be formed relatively long, the cooling air is exhausted into the mainstream gas after sufficient heat exchange, and cooling can be performed with a smaller amount of cooling air. Further, the film can be cooled by leaking a controlled amount of cooling air from the film hole. Therefore, even if the operating temperature of the gas turbine is set to a higher temperature, a highly reliable and durable gas turbine vane can be provided.

[0014]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS Hereinafter, the present invention will be described in detail based on an embodiment shown in the drawings. First, FIG. 1 shows a gas turbine stationary blade of the present invention. The stationary blade 1 is mainly composed of a wing portion 2, an outer peripheral end wall 3, and an inner peripheral end wall 4, and the mainstream gas forms a flow in the direction of arrow 6. In addition, the stationary blade 1
The seal plate grooves 53a and 53 are formed on the side surfaces of the outer peripheral end wall 3 and the inner peripheral end wall 4, respectively.
b, and are connected by inserting the outer peripheral side seal plate 52a and the inner peripheral side seal plate 52b into the seal plate grooves formed on the side surfaces of the end walls of the adjacent vanes, and are connected in the circumferential direction. An annular cascade is formed. The role of the seal plate is to prevent the cooling air from leaking into the mainstream gas from the gap between the connection of the stationary blades.

Here, cooling of the end wall portion will be described by taking the outer end wall 3 as an example. Generally, when cooling from the inside in order to keep the outer surface of the object to be cooled at a constant temperature, it is advantageous that the thickness of the object to be cooled is as thin as the strength allows. So for most endwalls,
Except for the edges that serve to form the seal plate grooves and maintain the strength, the thickness of the end wall is made as thin as the strength allows. As a result, a cavity for impingement cooling shown in 9 is formed, and the frame 13 surrounds the cavity. With the end wall configured in this way,
The cooling air 20 impingement-cools the end wall cooling surface 10 through the impingement holes 8 provided in the impingement plate 7. After the impingement cooling, the cooling air in the cavity 9 is removed from the end wall cooling surface 1.
The gas is discharged to the mainstream gas side through a film hole 12 penetrating the endwall mainstream gas side surface 11 to cool the endwall gas side surface 11. That is,
In order to cool the end wall, internal cooling is performed from the inside by impingement cooling, and film cooling is performed on the gas side surface to perform complex cooling. In a gas turbine in which the working gas is heated to a high temperature, cooling of either of them is indispensable.
However, the end wall frame 13 is a portion where it is difficult to perform internal cooling by impingement cooling. This is because, as described above, the frame portion 13 is configured to have a large thickness for maintaining strength and the seal plate 52b exists between the frame portion 13 and the gas side surface 11 to effectively perform impingement cooling. This is because an effective cavity cannot be formed.

Therefore, in the vane of the present invention, the seal plates 52a and 52b are formed with the seal plate grooves 53a and 53b.
The seal plates 52a and 5a are formed such that flow paths through which cooling air flows are formed between the seal plate 52a and the seal plate groove 53a and between the seal plate 52b and the seal plate groove 53b, respectively, in a state of being inserted into the seal plate 52a.
The sectional shape of 2b was changed from a conventional rectangular shape to a convex shape.
FIG. 2 shows the main part more clearly by taking the inner peripheral side end wall 4 as an example. Seal plate 52b
Is formed to have a convex cross-sectional shape, and has a cutout 60 at the end on the upstream side. On the other hand, a cooling air supply hole 61 is formed at the upstream end of the seal plate groove 53b, and a cooling air discharge hole 62 is formed at the downstream end. In this state, the seal plate 52b is inserted into the seal plate groove 53b in the direction of arrow 90. FIG. 3 shows a state where the seal plate 52b is inserted into the seal plate groove 53b. FIG. 4 is a sectional view thereof. In such a state, the cooling air 21 is supplied from the supply hole 61, passes through the seal plate cutting portion 60, and is guided to the cooling channel 54. In the cooling channel 54, the cooling air 21 flows from the upstream side to the downstream side to convectively cool the end wall frame 13 from the inside, and the cooled air is discharged from the discharge hole 62 into the mainstream gas. Here, in general, the stationary blade of the turbine serves to change the pressure of the mainstream gas into the flow velocity, so that the static pressure of the mainstream gas rapidly decreases from the upstream side to the downstream side of the stationary blade. On the other hand, the cooling air is supplied at a pressure higher than the static pressure of the mainstream gas on the upstream side of the stationary blade for the purpose of preventing backflow when a crack occurs and performing film cooling. Therefore, in this method, since the cooling air discharge hole 62 is provided on the downstream side where the static pressure is low, this pressure difference can be greatly utilized, and even in a narrow cooling passage having a large flow resistance such as 54, the cooling air can be cooled. A large flow velocity of air can be secured to promote heat transfer. In addition, unlike the cooling method shown in FIG.
Since the cooling air is discharged into the mainstream gas, the cooling capacity of the cooling air in the supply process can be sufficiently used, and the maximum cooling effect can be obtained with less cooling air consumption.

The stationary blade configured as described above solves the problem of cooling of the frame portion, which is a problem in end wall cooling, and has a highly reliable and durable gas that is free from the risk of melting, cracking due to oxidation or thermal stress. A turbine vane can be provided.

In the above description, the cooling channels are formed by making the cross-sectional shapes of the seal plates 51 and 52 convex, but as shown in FIG. Even if the cooling channel 55 is formed by processing the seal plate groove side, the obtained effect is the same and is not particularly limited.

Next, another embodiment of the present invention will be described. FIG.
The cross-sectional shape of the seal plate 52b is convex, and the turbulence promoting body 56
a is provided. FIG. 7 is a sectional view thereof. The cooling air supply, cooling and discharging mechanisms are the same as those shown in FIG. Generally, a turbulence promoter has a higher heat transfer promoting effect when it is provided on the cooling surface.However, even if it is provided on the opposite surface on the opposite side, the performance slightly decreases, but the effect of promoting heat transfer on the cooling surface is a model experiment. Has been confirmed. The model experiment has a rectangular channel with one side of 10 mm and the other side of 1.5 mm, and one of the two sides opposed to the long side of 10 mm is a non-heated surface, and the surface has a height of 0.3 mm and a width of 0 mm. .3mm turbulence enhancer is provided,
The experiment was performed using the other smooth surface as a heating surface and using air as a cooling medium. FIG. 8 is a characteristic diagram showing an experimental result. In FIG. 8, the horizontal axis represents the dimensionless Reynolds number Re representing the flow state of the cooling medium, and the vertical axis represents the dimensionless Nusselt number Nu representing the heat transfer characteristics.

[0020]

(Equation 1)

D: Equivalent diameter of cooling medium flow path v: Cooling medium flow rate α: Heat transfer coefficient ν: Kinematic viscosity coefficient of cooling medium λ: Thermal conductivity of cooling medium A turbulence promoter is provided on the surface facing the cooling surface. Even when provided, the heat transfer characteristic was approximately 1.8 times that of the smooth flow path. In the case where the turbulence promoting member is provided on the cooling surface, the heat transfer characteristic is about 4.4 times that of the smooth flow path. Therefore, in the example of the present invention, the turbulence promoting body 56a is provided on the seal plate 52b side where processing is easier. With such a structure, the heat transfer of the cooling air can be further promoted and the cooling effect can be increased as compared with the above-described structure. If the turbulence promoting body 56b is provided on the side of the seal plate groove 53b which is a cooling surface as shown in FIG. 9, a higher cooling effect can be expected.

The seal plate 52 shown in FIG.
The b is provided with a film cooling hole 57 for managing the cooling air 22 to leak and cool the film. As a result, the film can be cooled also in the vicinity of the end wall dividing surface, which has been difficult so far, and the cooling of the end wall frame is assisted. Further, it is possible to reduce film cooling unevenness on the entire end wall surface.

In all embodiments of the present invention, the cooling air discharge hole 62 may be anywhere on the end wall as long as the pressure difference for forming the cooling air flow is satisfied, and is not particularly limited. If the cooling air still has a cooling capacity at the stage when the cooling air reaches the downstream end of the seal plate groove, that is, if there is enough pressure and temperature, as shown in FIG. It is also possible to form the cooling holes 65 to guide the cooling air 21 and further discharge the cooling air 21 into the mainstream gas through the film cooling holes 58 communicating with the endwall mainstream gas side surface 11 from the cooling holes 65. In this manner, the downstream frame portion 14 of the end wall can be cooled, and the film air discharged from the film cooling hole 58 can be cooled immediately behind the stationary blade even if the film cooling of the stationary blade end wall is not effective. It can assist in film cooling of the located blade platform.

Although the embodiments of the present invention have been described above, the structure of the present invention can be applied to various cooling media such as steam and nitrogen, regardless of the cooling medium, and has an effect. However, by using the structure of the present invention, a highly reliable and durable stationary blade can be provided, and a highly reliable gas turbine can be provided.

[0025]

According to the present invention, a cooling flow path is formed between a seal plate of a stationary blade and a seal plate groove to enable effective cooling of a frame portion of an end wall, thereby increasing the operating temperature of a gas turbine. However, a highly reliable and durable stationary blade can be obtained.

[Brief description of the drawings]

FIG. 1 is a perspective view showing one embodiment of a gas turbine stationary blade according to the present invention.

FIG. 2 is a perspective view of a main part of FIG. 1;

FIG. 3 is a perspective view of a main part of FIG. 1;

FIG. 4 is a cross-sectional view of the cooling channel of FIG.

FIG. 5 is a sectional view of a cooling channel according to another embodiment of the present invention.

FIG. 6 is a perspective view of a main part showing another embodiment of the present invention.

FIG. 7 is a cross-sectional view of the cooling channel of FIG. 6;

FIG. 8 is an explanatory diagram of a result of a heat transfer model experiment.

FIG. 9 is a sectional view of a cooling channel according to another embodiment of the present invention.

FIG. 10 is a perspective view of a main part showing a modification of the present invention.

FIG. 11 is a cross-sectional view of an end wall frame portion cooling flow channel of a conventional gas turbine stationary blade.

[Explanation of symbols]

DESCRIPTION OF SYMBOLS 1 ... Gas turbine stationary blade, 2 ... Blade part, 3 ... Outer end wall, 4 ... Inner end wall, 6 ... Main gas flow direction, 7 ... Impingement plate, 8 ... Impinging hole
9: Cooling cavity, 10: End wall cooling surface, 1
DESCRIPTION OF SYMBOLS 1 ... Side wall of end wall main gas, 12 ... Film cooling hole, 13 ... End wall frame part, 20 ... Cooling air, 5
2a, 52b: seal plate, 53a, 53b: seal plate groove.

Claims (5)

[Claims] 1. A cascade formed in an annular shape, wherein the cascade has end walls forming gas flow paths on an outer peripheral side and an inner peripheral side, respectively, and the cascade includes a plurality of fins in a circumferential direction. The stator blade segments which are divided into segments, and which are adjacent to each other, are connected by inserting a seal plate into seal plate grooves provided on the end wall dividing surface, and between the seal plate groove and the seal plate, A gas turbine stationary blade characterized in that a cooling passage for flowing a cooling medium is formed in parallel with an end wall dividing surface. 2. A cascade formed in an annular shape, wherein the cascade has end walls forming gas flow paths on an outer peripheral side and an inner peripheral side, respectively. The stator vane segments which are divided into segments, and which are adjacent to each other, are connected by inserting the seal plate into the seal plate grooves provided on the end wall division surface, and the cross-sectional shape of the seal plate is made convex. A gas turbine vane, wherein a cooling passage for flowing a cooling medium is formed between the seal plate groove and the seal plate in parallel with the end wall dividing surface. 3. The gas turbine stationary blade according to claim 1, wherein a turbulence promoter is arranged in the cooling flow path. 4. A cooling channel for flowing a cooling medium in parallel with an end wall dividing surface is formed between the seal plate groove and the seal plate by making the cross-sectional shape of the seal plate convex. The gas turbine stationary blade according to claim 2, wherein a turbulence promoter is disposed on a valley surface of the seal plate that forms the cooling channel. 5. A cascade formed in an annular shape, wherein said cascade has end walls forming gas flow paths on an outer peripheral side and an inner peripheral side, respectively, and said cascade is formed in a circumferential direction by several The stator vane segments which are divided into segments and which are adjacent to each other are connected so as to insert the seal plate into the seal plate grooves provided on the end wall dividing surface and exist between the adjacent stator vane end wall dividing surfaces. A gas turbine vane, wherein a film cooling hole is provided in a portion of the gap exposed to the gas flow path of the seal plate.