JPS6327524B2 - - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION [Technical Field of the Invention] The present invention relates to gas turbine blades, and particularly to improvements in blades employing a fluid cooling system.

[Background technology of the invention]

In general, gas turbines have many advantages over reciprocating engines, such as being smaller, lighter, and capable of producing greater horsepower. As shown in FIG. 1, such a gas turbine usually includes a shaft 2 within a cylindrical casing 1.
A compressor 3 and a power turbine 4 are constructed between both ends of the shaft 2 and the casing 1, respectively, and the pressure inside the combustor 5 is controlled by the high pressure air compressed by the compressor 3. In this state, the fuel is injected and combusted, and the high-temperature, high-pressure gas generated by this combustion is guided to the power turbine 4 and expanded, thereby obtaining the rotational force of the shaft 2. In the case shown in the figure, the compressor 3 has guide vanes 6 and rotor vanes 7 arranged in the axial direction to form an axial flow type, and the power turbine 4 has rotor blades 8 fixed to the shaft 2 and a casing. 1 and fixed vanes 9 are arranged alternately in the axial direction.

By the way, in the gas turbine configured as described above, in order to increase the output efficiency, it is necessary to increase the combustion gas temperature at the inlet of the power turbine 4. When the combustion gas temperature is increased in this way, this gas flows around the stationary blades 9 and moving blades 8 at high speed, which causes the temperature of these blades to rise as well, raising the allowable temperature of the metal materials that make up the blades. There is a risk that it will be exceeded. For this reason, in gas turbines with combustion gas temperatures of 900°C or higher, cooling fluid passages are usually provided inside the blade body, and cooling fluid, such as air, is forced to flow through this passage. temperature is maintained at a safe value. Recently, there has been a trend to increase the combustion gas temperature in order to improve output efficiency, and in order to cope with this trend, the cooling fluid passages provided within the blade body have also become more complex.

Therefore, as described above, a blade in which a cooling fluid passage is provided in the blade body to cool the blade body always requires a cooling fluid supply source. Providing a separate supply source as this cooling fluid supply source inevitably increases the cost of the plant. For this reason, normally a part of the air that has been made highly pressurized by the compressor 3 is guided and used as a cooling fluid, and the air that has passed through the cooling fluid passage in the blade body is introduced into the gas stream from the outer surface of the blade body. A method is adopted in which the air is blown out towards the target.

[Problems with background technology]

As described above, cooling fluid passages are provided in the rotor blades 8 and stationary blades 9 that constitute the power turbine 4, and a part of the air pressurized by the compressor 3 is guided into these passages to cool each blade. In the case of a method in which the air that has been subjected to cooling is blown out from the outer surface of the blade into the gas flow, the amount of cooling air flowing through the blade is greatly influenced by the position of the blade. For example, in the first stage stator vane indicated by 9a in FIG. 1, the difference between the outlet pressure of the compressor 3 and the pressure around this stator vane 9a is small, so normally the cooling inside the stator vane 9a is Only a small amount of cooling air flows through the fluid passages. That is, the pressure around the stationary blade 9a is the value obtained by subtracting the pressure loss within the combustor 5 from the compressor outlet pressure, and since the pressure loss is usually small, the above-mentioned difference is The pressure is low, and only a small amount of cooling air flows through the cooling fluid passage within the stationary vane 9a. In this way, the stationary blade 9a is a blade that is affected by its position, and it is difficult to secure the required amount of cooling air. Therefore, in conventional gas turbines, cooling fluid passages as shown in FIG. 2 are provided in the stationary blades 9a to alleviate the decrease in cooling performance due to insufficient amount of cooling air. I have to. That is, a cavity B extending in the height direction of the wing body A is provided from the middle part to the leading edge part in the wing body A, and a multi-branch flow path C is provided from the middle part to the trailing edge end leading to the cavity B. A part of the air guided into the cavity B is transferred to the wall D existing between the cavity B and the leading edge outer surface and ventral outer surface of the wing body A, as shown in an enlarged view in FIG. The remaining cooling air is blown out from the blade through a plurality of pores E provided with the axis S inclined at θ and φ (where θ=φ) with respect to the outer surface of D, and the remaining cooling air is blown into the cavity B. The liquid is injected toward the inner surface of the wall G that exists between the cavity B and the dorsal outer surface of the wing body A through a plurality of small holes (not shown) provided in the partition plate F inserted therein. That's what I do. Then, a part of the cooling air blown out from the small holes in the partition plate F is blown out of the blade through the small holes E provided in the wall G in the same manner as described above, and the remaining air is sent through the multi-branched flow path C mentioned above. The air is blown out from the rear end of the wing body A to the outside of the wing. In other words, this blade has convection cooling when the cooling air flows through the cavity B, convection cooling when the cooling air flows through the pores E, and pores E because the pores E are provided at an angle. Cooling air blown from the partition plate F flows along the outer surface of the blade, resulting in so-called film cooling. Cooling air is blown out from small holes provided in the partition plate F onto the inner surface of the wall G, resulting in impingement cooling and multi-branch flow path C. The convection cooling caused by the flow of air through the inside suppresses the deterioration in cooling performance due to a small amount of cooling air.

However, the blade configured as described above has the following problems. That is, in a gas turbine blade, the parts that are heated to the highest temperature are the leading edge outer surface and the ventral outer surface. Therefore, it is necessary to properly cool these. When this type of blade is cooled using fluid, that is, air, considering the shape of the blade body A, there is impingement cooling from the inside and sloping It is desirable to use convection cooling through pores and film cooling in combination. That is, in impingement cooling, a high-speed air flow is injected toward the surface to be cooled, so the cooling effect is large. Further, in the case of inclined pores, a large convection cooling effect can be obtained due to the increase in heat exchange area due to the inclination, and a film cooling effect can also be obtained as described above. Therefore, if both cooling methods are used together, a large cooling effect can be obtained, but when used in combination in this way, the pressure loss increases significantly, and as a result, for example, the first stage stator blade 9a shown in FIG. Originally, the difference between the cooling air supply pressure and the pressure around the blade is
If it is small, the amount of cooling air will be significantly reduced, and the downside of this reduction in amount will appear. Therefore, conventionally, for blades located in the above-mentioned positions, only the slanted pores E are used to cool the leading edge and the ventral side of the intermediate portion of the blade body A, as shown in Fig. 2. However, since the pore E is simply provided with the axis S inclined with respect to the outer surface as shown in Fig. 3, the pressure loss in the pore E is relatively small. As a result, it is difficult to obtain the required amount of cooling air, and cracks often occur in the blade body A due to insufficient cooling.

[Purpose of the invention]

The present invention was developed in view of the above circumstances, and its purpose is to sufficiently reduce the pressure loss of the cooling fluid passage provided inside, and to reduce the pressure loss of the cooling fluid supply pressure and the air flow around the blades. Even when the difference in pressure is small, a sufficient amount of cooling fluid can flow, thereby demonstrating good cooling characteristics.
Moreover, it is an object of the present invention to provide a gas turbine blade that can be used in combination with an effective impingement cooling method.

[Summary of the invention]

The gas turbine blade according to the present invention is characterized in that pores for convection cooling and film cooling are provided on the outer surface of the blade body so that one end thereof is open, as shown in FIG. That is, D and G in the figure are walls similar to the wall shown in FIG. 3, and this wall D has small holes for convection cooling and film cooling.
Ea, whose axis S is inclined by θ degrees with respect to the outer surface of the wall D in the cross section of the blade, and the axis S and the pore
The angle φ between the cooling fluid inflow side opening of Ea and the plane H is approximately 90 degrees. In addition, in FIG. 4, a convex portion provided to set the angle φ between the plane H in which the opening opens and the axis S to be around 90 degrees is shown, and J similarly represents a concave portion.

〔Effect of the invention〕

With the above configuration, the diameter of the pore Ea,
Assuming that the inclination angle θ, the effective length on the axial center line, and the pressure difference between the inside and outside across the wall D are the same in the conventional blade, the heat exchange internal area of the pores Ea in the blade of the present invention is equal to that of the conventional blade. The amount of cooling fluid flowing through the pores Ea can be significantly increased under the same conditions, and film cooling can also be significantly improved. That is, as shown in FIG. 5, a straight branch pipe N is connected to the wall of the cooling fluid container M, and the angle α between the axis of the branch pipe N and the inner surface of the wall of the container M is When changing, the loss coefficient X of the branch pipe N changes as shown in FIG. Expressing this as a formula, X=0.5+
It becomes 0.3cosα+0.2cos ² α. As is clear from FIG. 6, when the branch angle α is 90 degrees, the loss coefficient is the smallest. Therefore, at this time, the branch pipe N
The flow velocity within is maximum, and the flow rate is maximum. As will be seen, in the present invention, the axis S of the pore Ea
Since the angle φ between the hole Ea and the plane H where the cooling fluid inflow side opening of the pore Ea opens is approximately 90 degrees, the conditions are set to allow the maximum amount of cooling fluid to flow. Therefore, heat can be effectively removed from the inner surface of the pore Ea, and the pore
Since the amount of cooling fluid blown out from Ea is large, the film cooling effect can be improved accordingly. Furthermore, since the pressure loss in the pores Ea can be sufficiently reduced, impingement cooling is possible even when the difference between the cooling air supply pressure and the pressure around the blade is small, such as the first stage stationary blade 9a shown in Fig. 1. It is possible to flow the required amount of cooling air even when both are used together. Therefore, even if the blade is a first-stage stationary blade, the blade main body can be cooled well, and this can ultimately contribute to improving the safety of the blade and stabilizing the operation of the gas turbine.

[Embodiments of the invention]

An embodiment of the present invention will be described below with reference to the drawings.

FIG. 7 shows the first stage stationary blade 9 in FIG. 1 according to the present invention.
This figure shows an example applied to a, and the same parts as in FIG. 2 are designated by the same reference numerals. Therefore, the explanation of the overlapping parts will be omitted.

In this embodiment, the leading edge of the wing body A and the wall D located on the ventral side have the pores shown in FIG. A plurality of small holes Ea for convection cooling and film cooling are provided in which the angle φ between the core wire S and the plane H where the cooling fluid inflow side opening opens is set to be around 90 degrees, and a partition tube P is provided in the cavity B. The cooling air guided into the partition tube P is injected toward the inner surfaces of the walls D and G from a plurality of small holes (not shown) provided in the partition tube P for impingement cooling. ing.

In this way, since the pores Ea provided in the wall D are set to the above-mentioned conditions, the pressure loss in these pores Ea can be reduced for the reason mentioned above. Therefore, as shown in the embodiment, even if impingement cooling is used in combination, the required amount of cooling air can be passed through the pores Ea, and as a result, the blade body A can be cooled well.

In the embodiment, the shape of the pore E provided in the wall G is set to be the same as that of the conventional one, but it may be set to the same shape as that provided in the wall D. Moreover, the present invention is of course applicable not only to the first-stage stator blades but also to the stator blades and rotor blades located at subsequent stages.

[Brief explanation of drawings]

Figure 1 is a diagram for explaining the schematic configuration of a gas turbine, Figure 2 is a cross-sectional view of the blade body of the first stage stationary blade in a conventional gas turbine, and Figure 3 is a diagram showing the convection cooling and film provided on the blade body. FIG. 4 is an enlarged view showing the cooling pores; FIG. 4 is a diagram for explaining the features of the blade of the present invention; FIGS. 5 and 6 are diagrams for explaining the basis of the effects of the present invention. , FIG. 7 is a cross-sectional view of a blade main body of an embodiment in which the present invention is applied to a first stage stationary blade of a gas turbine. A...Blade body, B...Cavity, C...Multi-branched channels, D, G...Wall, E, Ea...Small holes for convection cooling and film cooling.

Claims (1)

[Claims] 1 A cooling fluid passage having a cross section in the cross-sectional shape of the blade is provided in the blade body, and the cooling fluid guided into the passage flows in the cross-sectional direction of the blade, all or in part, between the passage and the blade body. The airflow is caused to flow out of the blade body through a plurality of pores provided in a thin wall existing between the blade and the outer surface with the axis line inclined with respect to the outer surface within the cross section of the blade. A gas turbine blade characterized in that, in whole or in part, the angle between its axis and the plane in which the cooling fluid inflow side opening opens is approximately 90 degrees.