JPS5979009A - Gas turbine blade - Google Patents

Abstract

PURPOSE:To increase the quantity of a cooling fluid flowing through small holes formed in the walls of the turbine blade having a cooling fluid passage therein by a method wherein the angle between the line of axis of each of the small holes and the opening surface of each of the holes is determined to be nearly 90 degrees. CONSTITUTION:The cooling fluid passage B is provided with the blade A and the cooling fluid is flowed out from the blade A through the small holes E and Ea formed in the walls D and G of the blade A. Further, the angle between the line of axis of each of the small holes E in the abdominal wall D of the blade A and the opening surface of the fluid inlet side of the small hole Ea is determined to be nearly 90 degrees. As a consequence, it is possible to reduce the loss of pressure of the cooling fluid flowing through the small holes Ea and to increase the quantity of the cooling fluid flowing through the above-mentioned small holes Ea.

Description

DETAILED DESCRIPTION OF THE INVENTION [Technical Field of the Invention] The present invention relates to blades of gas turbines, and particularly relates to a page break for 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 has a shaft 2 disposed within a cylindrical casing 1 via bearings, and a compressor 3 disposed between both ends of the shaft 2 and the casing 1. The high-pressure air compressed by the compressor 3 increases the pressure inside the combustor 5, and in this state fuel is injected and combusted, and the brown-temperature, high-pressure gas produced by this combustion is The rotational force of the shaft 2 is obtained from the K which is introduced into the power turbine 4 and expanded. and,
In the case shown in the figure, the compressor 3 is of an axial flow type with guide vanes 6 and rotor vanes 7 arranged in the axial direction, and the power turbine 4 has rotor blades 8 fixed to the shaft 2 and a casing 1. It is constructed by alternately arranging fixed stationary blades 9 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 control the combustion gas temperature at the inlet of the power turbine 4. When the combustion gas temperature is increased in this way, the gas flows at high speed through the rotary blades 9 and rotor blades 8, and the temperature of these blades also increases.
There is a risk that the allowable temperature of the metal materials that make up the blades will be exceeded. For this reason, in gas turbines in which the combustion gas temperature is 90υ℃ or higher, a cooling fluid passage is usually provided inside the blade body, and cooling fluid, such as air, is forced to flow through this passage. Trying to keep the temperature 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 directed into the gas flow from the outer surface of the blade body. A method of blowing out water is adopted.

[Problems with background technology]

As mentioned above, cooling fluid passages are provided in the rotor blades 8 and stationary blades 9 that constitute the 9-power turbine 4, and a portion of the air pressurized by the compressor 3 is guided into these passages to cool each member. In the method of blowing out the air after cooling into the gas flow from the external surface, the amount of cooling air flowing through the inside of the blade is greatly influenced by the position of the blade. For example, 9 in Figure 1
In the first-stage stator vane indicated by g, the difference between the outlet pressure of the compressor 3 and the rotary pressure of the stator vane 9a is small. Only a small amount of cooling air flows. That is, the pressure of the rotor of 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,
As a result, the above-mentioned differential pressure is small, and only a small amount of cooling air flows through the cooling fluid passage within the stator vane 9a. In this way, the static Jg 9 a is a wing 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 the m2 diagram are provided inside the stationary blades 9& to alleviate the decline in cooling performance due to insufficient amount of cooling air. ing. That is, a cavity B extending in the height direction of the blade body A is provided from the intermediate portion to the leading edge in the blade body A, and a multi-branch flow path C communicating with the cavity B is provided from the intermediate portion to the trailing edge. A part of the air guided into the cavity B is transferred to the wall 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 of the blade through a plurality of small holes E provided with the axis S inclined by θ and φ (0=φ), respectively, with respect to the outer surface of the blade. Injected toward the inner surface of the wall G existing 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 in B. I try to let them do it. 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 to the multi-branch flow path C described above.
The air is blown out from the rear end of the blade main body A to the outside of the blade through. In other words, this blade provides convection cooling when cooling air flows through the cavity B, impingement cooling when cooling air is injected into the pores B, and cooling air that flows through the multi-branched flow path C. Convection cooling is used to suppress the decline in cooling performance caused by a small amount of cooling air.

However, in the wing configured as above,
There were 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 or pure air, considering the shape of the blade body A, the leading edge and intermediate ventral sides of the blade body A are impinged from the inside and are cooled from the outside. It is desirable to use convective cooling with inclined pores in conjunction with film cooling. 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 oblique 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 mentioned above. Therefore, if both cooling methods are used together, a large cooling effect can be obtained, but when used together in this way, the pressure loss increases significantly, and as a result, for example, the first
When the difference between the cooling air supply pressure and the pressure at the blade rotor is originally small, as in the case of the first-stage stator vane 9a shown in the figure, the amount of cooling air decreases significantly, and this amount decreases. The negative aspects associated with this are revealed. Therefore, the pores E are simply provided with the axis S inclined with respect to the outer surface, as shown in Fig. 3.
The pressure loss in the pores E is relatively large, and 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 these circumstances, and its purpose is to sufficiently reduce the pressure loss of the cooling fluid passage provided inside, thereby reducing the cooling fluid supply pressure and the pressure around the blades. A sufficient amount of cooling fluid can be flowed even when the difference between , is characterized in that pores for convection cooling and film cooling are provided in 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 walls shown in FIG. Tilt by θ degrees with respect to the above axis? The angle φ between the fMS and the opening surface H of the cooling fluid inflow side of the pore Ea is approximately 90 degrees. Note that in FIG. 4, I indicates a convex portion provided to set the angle φ between the opening surface H and the axis S to around 90 degrees, and J similarly indicates a concave portion.

〔Effect of the invention〕

With the above configuration, the pore Ea can now be formed.

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 V between the axis of the branch pipe N and the inner surface of the wall of the container M is determined. When α is changed, the loss coefficient X of the branch tube N changes as shown in FIG. Expressing this in the formula, X = 0.5 + 0.3 co
sα+Q, 2coaα. 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 flow velocity in the branch pipe N is at its maximum, and the flow area is at its maximum. As will be seen, in the present invention,
Since the angle φ between the axial center line S of the pore Ea and the cooling fluid inflow side opening surface H of the pore Ea is large, the amount of the cooling fluid can be improved accordingly. Furthermore, since the pressure loss in the pore Ea can be sufficiently reduced,
As in the first stage stator vane 9a shown in FIG. 1, the difference between the cooling air supply pressure and the pressure of the blade rotation can contribute to improving the safety of the small 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 an example in which the present invention is applied to the first stage stationary blade 9a in FIG. 1, 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, a wall D located on the leading edge and ventral side of the wing body A [the pore shown in FIG. A plurality of small holes for impingement cooling are provided with a plurality of small holes Ea for convection cooling and film cooling, in which the angle φ between the core wire S and the opening surface H on the cooling fluid inflow side is set near 90 degrees ( (not shown) toward the wall and the inner surface of G.

In this way, since the pores Ea provided in the wall 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 example, even when used in combination with impingement cooling, the required amount of cooling air can be passed through the pores Ea.
As a result, the blade body A can be cooled well.

In addition, in the example, the shape of the pore E provided in the wall G is set in the same way as the conventional one.

[Brief explanation of drawings]

FIG. 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; Figure 3 is an enlarged view of the pores for convection cooling and film cooling provided in the blade body; The figure is a diagram for explaining the characteristic points of the blade of the present invention, Figures 5 and 6 are diagrams for explaining the basis of the effects of the present invention, and Figure 7 is a diagram for explaining the features of the blade of the present invention.
The figure is a cross-sectional view of a blade 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 flow path,
D. G...Wall, E r E a... Pores for convection cooling and film cooling. Applicant: Makoto Ishizaka, Director, Agency of Industrial Science and Technology - Figure 1 Figure 2 Figure 3 Figure 4 Figure 6 Branch angle (closing)

Claims (1)

[Claims] A cooling fluid passage is provided in the blade body, and all or part of the cooling fluid guided into the passage is directed to a wall existing between the passage and the outer surface of the blade body, the axis of which is inclined with respect to the outer surface. In a gas turbine blade in which air is blown out of the blade body through a plurality of pores, all or part of the pores are located between the axial center line and the opening surface located on the cooling fluid inflow side. A blade of a gas turbine characterized in that the angle of is set in the vicinity of 90 degrees.