JPS6360205B2 - - 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 performance and increase output, and as a result, gas turbine blades are exposed to high temperatures. Since there is no material available, the method used is to cool the wings.

Stator blades (hereinafter referred to as blades for convenience in this explanation) used in conventional gas turbines are shown in Figure 1-A,
As shown in the examples in Figure 1-B, Figure 1-C, and Figure 1-D, the blade 1 is formed hollow, cooling air is introduced here, and the inside is cooled by convection. What is shown is that a core 4 is provided inside a hollow blade 1, and cooling air is introduced into the core 4, and the air is blown out toward the inner surface of the blade through a number of pores 5 at the tip of the core 4, and is locally cooled. In addition, cooling air is introduced into the hollow blade 1 and blown out of the blade through numerous pores 6 at the leading edge of the blade, as shown in Figure 1-B, which increases heat transfer and is forcedly cooled.
There are blades 1 covered with a cooling air layer to block heat from high-temperature combustion gas and film-cooled, as shown in Figure 1-C. This leads to wing 1 in Figure 1-D.

Note that in the above-mentioned Figures 1-A to 1-D, the same parts are indicated by the same part numbers.

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 flow back into the blade, which will affect the cooling performance. On the other hand, if the pressure difference is too large, the cooling air will be blown out forcefully, and if the blowing angle to the blade surface is large, it will be difficult to form a cooling air layer along the blade 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. , is becoming more complex.

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 forms the head and main body of a stator vane of a gas turbine separately, and provides holes and grooves having dimensions slightly larger than the head in the platform and shroud of the stator vane. , by fitting both upper and lower ends of the head into the holes and grooves.

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 bracket form, 19 is a shroud, and 20 is a cap.

Next, in this wing 1, the head 12 and the main body 13 are formed separately, and the head 12 is formed from the main body 1.
3. The platform 18 and the shroud 19 can be made of the same heat-resistant alloy, or the main body 13
It may be formed from a ceramic material different from the above.

The range of the head 12 is the range where the mainstream gas is blocked 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 one body.

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.

The stator vane of the present invention is configured as described above, and the cooling air is guided to the tip portion 14 and the trailing edge portion 2 of the main body portion 13, and the cooling air guided to the hollow portion of the tip portion 14 is directed to the main body portion 13. 13 Tip cooling air outlet 1
7 into the cooling air passage 16 between the head 12 and the main body 13, and is blown out of the blade through the cooling air passage 16, covering the main body 13 with a layer of cooling air and cooling 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 even if the hole 21 is provided in the shroud 19 and the groove 22 is provided in the platform 18,
Alternatively, holes may be provided on both sides.

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. does not need to be considered.

For this reason, it is also possible to use ceramic for the head 12, which could not conventionally be used to construct the wing 1 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.

Therefore, in the present invention, the mainstream gas is dammed,
The head of the leading edge, which is the hottest part of the wing, is formed separately from the main body, so even if the head expands due to high temperature, it will not affect the main body.
The structural strength of the wing as a whole can be maintained sufficiently.

In addition, in the present invention, a cooling air passage is provided in the dividing plane between the head and the main body, and cooling air is blown out from the cooling air passage to the side of the main body, so that the main body can be film-cooled, and the blade as a whole can be cooled. For example, there will be no leading edge airflow, and air will be emitted from the side.

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 above wing, the angle at which the parting line between the head and the main body touches the outer surface of the wing can be set large, so that the cooling air blown out of the wing through the cooling air passage in the parting surface is It will blow out at a small angle backwards.

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 air 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 from the inside of the blade to the outside 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, so when blowing from the leading edge of the blade and the side of the blade, etc. , the mainstream gas has a pressure distribution along the blade surface, and the blade structure is complicated to make the amount of cooling air blown out at each position a predetermined amount. If the cooling air is no longer blown out from the leading edge, the mainstream gas is accelerated, and the air is blown out from the lower pressure side of the blade and the trailing edge of the blade, the cooling air will be distributed according to the pressure distribution of the mainstream gas along the blade surface. The blade structure for blowing out a predetermined amount becomes 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 separately formed head is damaged for some reason, the main body has an airfoil shape with a convex tip, so a certain level of aerodynamic performance is maintained, and even if the head is damaged, can also be easily replaced.

[Brief explanation of the drawing]

Figure 1-A, Figure 1-B, Figure 1-C, and Figure 1
-D is a sectional view of different conventional cooling type stator blades, FIG. 2 is a system diagram of a gas turbine, and FIG. 3 is a sectional view of a stator blade of a gas turbine according to an embodiment of the present invention. There is a
FIG. 3 is a cross-sectional view of the wing head shown in the figure. 1... Blade, 10... Turbine part of gas turbine, 1
1... Generator, 12... Head, 13... Main body, 18...
Platform, 19...shroud, 20...cap, 21...hole, 22...groove.

Claims (1)

[Claims] 1. The head and main body of a stator blade of a gas turbine are formed separately, and holes and grooves having dimensions slightly larger than the head are provided in the platform and shroud of the stator blade, and the holes and grooves are formed in the platform and shroud of the stator blade. A stationary blade for a gas turbine, characterized in that both upper and lower ends of the head are mounted in a groove.