JPS6242121B2 - - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION The present invention relates to a stationary blade of a gas turbine mainly used in a high-temperature gas turbine or the like.

In recent years, gas turbines have become increasingly hotter in order to improve their performance and increase their output.As a result, gas turbine blades are exposed to high temperatures, and currently there are materials that have strength at such high temperatures. Since there is no airfoil, a method of cooling the wings is used.

Stator blades (hereinafter referred to as blades for convenience in this explanation) used in conventional gas turbines are shown in Fig. 1-A,
As shown in the examples of Figure 1-B, Figure 1-C, and Figure 1-D, the blade 1 is formed hollow, cooling air is guided, and the inside is cooled by convection as shown in Figure 1-A. A core 4 is provided inside the hollow wing 1, and the core 4
Many pores 5 at the tip of the core 4 lead cooling air into the core.
The air is blown out toward the inner surface of the blade, locally increasing heat transfer, and forcedly cooled as shown in Figure 1-B, and the cooling air is further guided into the hollow blade 1.
Air is blown out of the blade from a large number of pores 6 at the leading edge of the blade, and the blade 1
There are some types of gas turbines, such as the one shown in Figure 1-C, which is covered with a cooling air layer to isolate heat from the high-temperature combustion gas and cooled with a film. Wing 1 of figure D
It has reached this point.

In addition, as shown in the conventional example shown in FIG.
There is also a wing 1 that is film cooled.

In addition, in the said FIG. 1-A to FIG. 1-E, the same parts are shown by the same part number.

Here, the leading edge of the gas turbine blade 1 that is exposed to the combustion gas and reaches the highest temperature is the leading edge of the blade 1 where the mainstream gas is dammed up, so cooling this leading edge is the most important. As the temperature of the turbine increases, film cooling is also used, and the amount of cooling air required to cool this part is also increasing.

However, if the blade 1 is film-cooled and the amount of cooling air required for this increases, the amount of cooling air mixed with the mainstream gas will increase accordingly, and the average mainstream gas temperature will decrease, thereby reducing the cycle efficiency of the gas turbine. will decrease.

Cooling air for cooling the blades 1 is usually obtained by extracting air compressed by a compressor 8 driven by a turbine section 10 of a gas turbine before a combustor 9, as shown in the system diagram in FIG. , casing or
It is supplied into the blade 1 through piping etc. connected to this.

Therefore, if the amount of cooling air increases, the power required for compression by the compressor 8 will increase, and the efficiency and output of the gas turbine 10 will decrease by this amount.

In addition, in order to completely cool the film,
The pressure difference between the cooling air and the pressure of the mainstream gas needs to be appropriate. If this pressure difference is small, not only will blowing not occur locally, but the mainstream gas may also flow back into the blade, resulting in poor cooling performance. On the other hand, if the pressure difference is too large, the cooling air will blow out forcefully, making it difficult to form a cooling air layer along the wing surface, and even aerodynamic performance will be impaired.

Generally, the pressure of the mainstream gas is only slightly lower than the pressure of the cooling air, so in order to ensure complete blowing, the main stream gas is routed from the outlet of the compressor 8 in the mainstream gas system to the blade row of the turbine section 10 of the gas turbine. In some cases, a throttle resistor or the like is provided between these steps to lower the pressure of the mainstream gas.

In this way, lowering the pressure of the mainstream gas is
Since this amount is not involved in work, it becomes a loss and the output decreases.

In addition, when cooling air is blown out from each part of the blade 1 to perform film cooling, there is a pressure distribution in the mainstream gas along the blade surface, and the blade structure is required to blow out a predetermined amount of cooling air to each location. It's getting complicated.

Further, providing cooling air blow-off holes requires a lot of processing time, increases costs, reduces strength, and shortens blade life.

As mentioned above, in the structure of conventional cooled blades,
As the temperature of gas turbines increases, film cooling is performed from the leading edge of the blade, and the amount of cooling air required for this is also increasing, resulting in a decrease in gas turbine thermal efficiency due to mainstream gas cooling and a loss in the required power of the compressor. In addition, there were problems such as a reduction in output due to lowering the mainstream gas pressure, and countermeasures against this problem were strongly desired.

Therefore, the present invention solves the above-mentioned conventional problems,
This was done with the aim of making it possible to improve the efficiency of gas turbines.

That is, the present invention separates the head of a stator blade of a gas turbine from other stator blade structures and forms it from ceramic, and also provides holes or grooves in the platform and shroud of the stator blade, and The upper and lower ends of the head are mounted in the holes or grooves of the platform, and the section from the front end of the mainstream gas passage surface of the platform to the rear end of the head is made of ceramic.

Embodiments of the present invention will be described below with reference to the drawings. FIG. 3 is a sectional view of the stator blade of a gas turbine in one embodiment of the present invention, and FIG.
A cross-sectional view along the camber line of the stationary blade shown in the figure.
FIG. 5 is a sectional view of the wing head in FIG.
The same parts as in the conventional example shown in FIGS. A to 1-D are indicated by the same part numbers.

First, in the blade 1 of the present invention applied to the turbine section 10 of a gas turbine similar to that explained in the conventional example of FIG.
is a hollow tip, 15 is a partition, 16 is a cooling air passage, 17 is a cooling air outlet at the tip, 18 is a platform, 19 is a shroud, and 20 is a cap.

Next, in this blade 1, the head 12 is made of ceramic, the main body 13, the platform 18, and the shroud 19 are made of metal, that is, a heat-resistant alloy, and the mainstream gas passage surface of the platform 18 is made of metal. The area from the front end of the head 12 to the rear end of the head 12 is covered with a ceramic member 23 as shown in the side view of FIG.

Here, the range of the head 12 is defined as the range where the mainstream gas is dammed up or the range where the heat transfer coefficient is high.

Further, the head 12 is separated from the main body 13 by a curve such that the main body 13 side is convex or by a broken line so that the angle at which the dividing line contacts the outer surface of the wing is large.

In addition, the main body 13 and the platform 18
and Shroud 19 are integrated.

Further, the platform 18 is provided with a hole 21 that is slightly larger than the head 12, and the shroud 19 is also provided with a groove 22 that is slightly larger than the head 12.
The head 12 is inserted into the groove 22 through the hole 21, the cap 20 is placed in the hole 21, and the upper end of the cap 20 is welded all around.

The main body part 13 has a distal end part 14 by a partition 15.
A hollow section is provided, which is divided into a rear edge section 2 and a rear edge section 2. A large number of small cooling air blowing holes 17 are provided at the tip of the hollow section, and cooling air passages 16 are provided at the outer surface of the hollow section, that is, the surface where it meets the head section 12. The hollow part of the trailing edge part 2 has an internal convection cooling structure.

Note that the cooling air blowing holes 24 on the passage surface of the platform 18 are provided after the rear end of the head 12.

The present invention is constructed as described above, and the cooling air is guided to both the platform 18 and the shroud 19, and is blown into the passage from the cooling air blowing hole 24.
The cooling air guided to the platform 18 side is guided to the hollow part of the tip part 14 and the rear edge part 2 of the main body part 13, and
The cooling air guided into the hollow part of 4 is blown out from the cooling air blowing hole 17 at the tip of the main body part 13 to the cooling air passage 16 between the head part 12 and the main body part 13, and then passed through the cooling air passage 16. The air is then blown out of the blade, and the main body 13 is covered with a layer of cooling air to cool the film.

Further, the cooling air guided into the hollow portion of the trailing edge portion 2 convects the inside of the main body portion 13 and is blown out of the blade from the cooling air blowing hole 3 at the trailing edge.

Note that here, the shroud 19 may be provided with the hole 21 and the platform 18 may be provided with the groove 22, or both may be provided with holes.

As described above, in the present invention, the head 12 of the wing 1 is
It is separated from other wing structural parts, namely, the main body part 13, the platform 18, the shroud 19, etc., and the wing 1
The structural strength of the head 12 is maintained by the latter, and the aerodynamic force applied to the head 12 is also supported by the main body, so the head 12 does not require structural strength.

In addition, although the blade 1 is affected by the thermal expansion of the turbine casing or deforms due to its own thermal expansion, there are holes 21 in the platform 18 for attaching the head 12 to these, and holes 21 in the shroud. Since the groove 22 of No. 19 is larger than the head 12 and there is a gap between the groove 22 and the head 12, this force will not be applied to the head 12 even if the wing 1 is deformed.

That is, if the wing 1 is not deformed, the rear surface of the head 12 is in contact with the tip of the main body 13 and the rear surfaces of the holes 21 and grooves 22 due to aerodynamic force, and no force is applied from the wing 1. , if wing 1 is deformed, hole 2
1 and the groove 22 may be misaligned, the main body portion 13 may be bent, or the main body portion 13 may protrude, and force acts on the head portion 12.

Here, if there is no gap between the hole 21 and the groove 22, when the blade 1 deforms, all the force will also act on the head 12, but since there is a gap between the hole 21 and the groove 22, the blade 1 will deform. However, the head 12 moves within the hole 21 and groove 22 and no large force is applied.

Therefore, the gap between the hole 21 and the groove 22 needs to be larger than the amount of deformation of the blade 1, specifically 0.1 to 0.15
mm is enough.

Note that when the entire blade 1 expands due to thermal expansion, the center lines of the holes 21 and grooves 22 do not shift, and the main body 13 does not protrude. need not be considered.

For this reason, the head 12 can be made of ceramic, which could not be used to construct the blade 1 conventionally due to insufficient reliability in terms of structural strength.

The reason why the cap 20 is welded to the platform 18 all the way around is to prevent the mainstream gas from leaking out of the mainstream gas passage through the gap between the holes 21.

For this reason, in the present invention, the mainstream gas is dammed up, and the leading edge, or head, which is the highest temperature of the blade, is made of ceramic, which has higher heat resistance than metal, so it can be used at a higher temperature than metal. There is no need to blow out cooling air from the leading edge to cool the film.

In addition, although the metal main body requires cooling, it is not a part that gets particularly hot as a wing, but the tip is a head made of ceramic that blocks the heat of the combustion gas, and the head and main body are connected. In between, cooling air flows through the cooling air passage and prevents heat transfer from the head, so it does not reach high temperatures like the leading edge of a conventional wing, and cools the tip of the main body. The amount of cooling air needed for this purpose is small.

On the other hand, since the side surface of the main body is exposed to combustion gas, cooling air is blown out from the cooling air passage between the head and the main body to cool the film.

Here, the cooling air blown out from the tip of the main body has two functions: it prevents heat transfer from the head to the main body, and it also cools the side surface of the main body as a film.

In this way, if the head does not need to be cooled and the amount of cooling air used to cool the main body is reduced, the amount of cooling air required for the entire wing will be reduced.

Cooling air is blown into the mainstream gas after cooling the blades or for film cooling, but it mixes with the mainstream gas and cools it, so the smaller the amount of cooling air, the lower the average mainstream gas temperature. The thermal efficiency of the gas turbine is improved.

Furthermore, if the amount of cooling air is reduced, the amount of air compressed by the compressor that is extracted as cooling air is reduced, and this amount can be extracted as output, thereby improving gas turbine efficiency and output.

In addition, in the present invention, the cooling air for film cooling the main body is blown out from the cooling air passage in the dividing plane between the head and the main body, and covers the side of the main body from its front end. There will be no edge balloons, and there will be balloons from the sides.

When performing film cooling from the leading edge of the blade, the dynamic pressure of the mainstream gas is applied to the leading edge of the blade, so the pressure of the cooling air needs to be higher than this.In order to maintain this pressure difference, the pressure of the mainstream gas system must be increased. Sometimes this is intentionally lowered, but when blowing out from the trailing edge of the blade, the mainstream gas accelerates and the pressure decreases, so the pressure difference between the mainstream gas and the cooling air is maintained, and the mainstream gas system There is no need to lower the pressure, which increases the efficiency of the gas turbine.

In addition, in the present invention, the angle at which the dividing line between the head and the main body touches the outer surface of the blade is set at a large angle, so that the cooling air blown out of the blade through the cooling air passage in the dividing surface is directed toward the rear of the blade at a small angle. It will be blown out.

Therefore, even if the pressure of the cooling air becomes higher than the pressure of the mainstream gas and is blown out vigorously, a cooling air layer is formed along the blade surface, and the cooling performance and aerodynamic performance are not impaired.

Further, in the present invention, cooling air is no longer blown out from the leading edge of the blade, but instead is blown out from the side surface of the blade and the trailing edge of the blade. The amount of cooling air blown out of the blade is determined by the total cross-sectional area of the cooling air outlet, depending on the pressure difference between the cooling air and the mainstream gas. There is a pressure distribution along the blade surface, and the blade structure is complex in order to blow out a predetermined amount of cooling air at each position. When the cooling air is no longer blown out and the mainstream gas accelerates and is blown out from the lower pressure blade side and trailing edge, a predetermined amount of cooling air is blown out according to the pressure distribution of the mainstream gas along the blade surface. The wing structure for ejection is simple.

In addition, in the present invention, when dividing the wing into the head and the main body, the wings are separated so that the main body side is convex.
Even if the direction of the aerodynamic force acting on the head changes, this force can be effectively supported by the main body.

Further, the combination of the head and the main body is concave and convex, and it is possible to prevent the head from shifting from the main body, creating a step, causing separation of the mainstream gas flowing on the wing surface, and reducing aerodynamic performance.

In addition, even if the ceramic head is damaged for some reason, the main body has an airfoil shape with a convex tip, so a certain level of aerodynamic performance will be maintained, and even if the head is damaged, Can be easily replaced.

[Brief explanation of the drawing]

Figure 1-A, Figure 1-B, Figure 1-C, Figure 1-
Figure D is a plan sectional view of different conventional cooling type stator vanes, Figures 1-E are side sectional views of other conventional cooling type stator blades, Figure 2 is a system diagram of a gas turbine, and Figure 3 is a diagram of a gas turbine.
The figure is a sectional view of the stator blade of a gas turbine according to an embodiment of the present invention, FIG. 4 is a sectional view of the stator blade along the camber line of FIG. 3, and FIG. FIG. 6 is a side view of the main part of FIG. 4. 1...Blade, 10...Gas turbine turbine section,
11... Generator, 12... Head, 13... Main body, 18... Platform, 19... Shroud, 20... Cap, 21... Hole, 22...
Groove, 23...ceramic member, 24...cooling air outlet.

Claims (1)

[Claims] 1. Separate the head of the stator blade of a gas turbine from other stator blade structures and form it with ceramic, and provide holes or grooves in the platform and shroud of the stator blade, and insert holes or grooves into these holes or grooves. A stationary blade for a gas turbine, characterized in that both upper and lower ends of the head are attached, and further, the part from the front end of the mainstream gas passage surface of the platform to the rear end of the head is formed of ceramic.