JPS6224606B2 - - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION [Technical Field of the Invention] The present invention relates to a gas turbine blade, and more particularly to an improvement in a blade provided with a fluid cooling structure.

[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. Such gas turbines typically
A compressor and a power turbine are connected to one shaft,
The pressure inside the combustor is increased using high-pressure air compressed by a compressor, and in this state fuel is injected into the combustor and combusted, and the high-temperature, high-pressure gas generated by this combustion is guided to the power turbine. It is configured to obtain rotational power by expanding it. A compressor is usually configured as an axial flow type in which guide vanes and rotor blades are arranged in the axial direction, and a power turbine is also configured in such a manner that moving blades and stationary blades are alternately arranged in the axial direction.

By the way, in the above-described gas turbine, it is said that the most effective way to increase the output efficiency is to increase the combustion gas temperature at the inlet of the power turbine. However, if the inlet gas temperature of the power turbine is increased, the blade temperature will increase due to the high temperature combustion gas. Current heat-resistant metals that make up the blades cannot be operated for long periods of time if the temperature exceeds 900°C. Therefore, the only way to extend the operational life of the blade is to lower the blade temperature by some means.

[Problems with background technology]

For the reasons mentioned above, various types of gas turbine blades equipped with cooling structures have been proposed. Figures 1 and 2 show the internal structure of a typical wing. That is, 1 in the figure is a wing consisting of a wing body 2 and a wing root part 3 integrally connected to this wing body 2, and a leading edge part F and a middle part N in this wing 1.
and cooling systems 11 and 12 at the trailing edge R, respectively.
There are 13.

The cooling system 11 includes a flow path 16 formed by a leading edge wall 14 and a partition wall 15 to extend in the height direction from the blade root 3 to near the tip of the blade body 2, and a flow path 16 that penetrates the leading edge wall 14. and a plurality of small holes 17 for cooling the film provided in the height direction. Therefore, this cooling system 11 has a convection cooling effect due to the cooling fluid flowing through the flow path 16 as shown by the arrow in the figure, and a convection cooling effect due to the cooling fluid flowing through each small hole 17. In addition, the cooling fluid blown out from each small hole 17 flows along the outer surface of the leading edge wall 14, thereby cooling the leading edge F of the blade body 2 by a film cooling effect. In the figure, reference numeral 18 denotes a wall constituting the flow path 16, and a plurality of walls projecting in the height direction from the inner surface located on the ventral and dorsal sides of the blade body 2 are used to actively stir the cooling fluid flowing through the blade body 2. The viewing promoter is shown.

Therefore, the cooling system 12 has partition walls 19 and 20.
It extends in the height direction from the blade root 3 to the vicinity of the tip of the blade body 2, and near the tip, the direction is changed 180 degrees around the leading edge side by the partition walls 19 and 21, and the blade body 2, and then around the front edge side by partition walls 21 and 15.
A bending channel 22 is formed so as to change direction by 180 degrees and extend again to the vicinity of the tip of the wing body 2, and a portion of this bending channel 22 between the partition walls 19 and 21 forms the belly of the wing body 2. Side wall and partition wall 2
1 and 15, penetrating the walls located on the ventral side and the dorsal side of the wing body 2, respectively, and consisting of a plurality of small holes 23 for film cooling formed in the height direction. There is. Therefore, this cooling system 12 also has the same convection cooling effect as the cooling fluid flows through the bent channels 22 and the small holes 23 as shown by the arrows in the figure, and the cooling fluid blown out from each small hole. The intermediate portion N of the wing body 2 is cooled by the film cooling effect caused by the film flowing along the ventral outer surface and the dorsal outer surface of the wing body 2. Note that 24 in the figure indicates a turbulence promoter.

On the other hand, the cooling system 13 runs from the blade root 3 to the blade body 2.
A flow path 25 extends in the height direction so that a portion of the blade body 2 is divided into two parts up to the vicinity of the tip of the blade body 2.
and pin fins 26 that connect the inner surface of the wall located on the ventral side of the wing body 2 and the inner surface of the wall located on the dorsal side at multiple locations, as shown by the arrows in the figure. The trailing edge R is cooled by bringing the guided cooling fluid into contact with the pin fins 26 and the like.

However, in the conventional blade configured as described above, there was a problem in that the cooling performance of the ventral side was particularly low at the intermediate portion N of the blade body 2 for the following reasons. That is, in the conventional blade, the middle part N
, the convection cooling effect caused by the cooling fluid flowing through the bent channel 22 and each small hole 23 and each small hole 23
Cooling is achieved by the film cooling effect caused by the cooling fluid blown out from the blades flowing along the outer surface of the blade body 2. Of these, the film cooling effect is achieved by the state in which the cooling fluid is blown out from each small hole 23. They differ significantly. Letting the film cooling effect be η, this η is expressed by the following equation.

η=T _g -T _f /T _g -T _c ...(1) However, in equation (1), T _g is the mainstream gas temperature, and T _c
is the outlet temperature of the cooling fluid, and T _f is the film temperature.

Moreover, η has the following relationship.

η∝f(M, x) ...(2) M=(ρ _c・u _c )/(ρ _g・u _g ) Here, the above variable is
is the mass flow rate ratio, x is the distance from the outlet position, ρ _g
is the density of the mainstream gas, u _g is the velocity of the mainstream gas, ρ _c is the cooling fluid density at the outlet position, and u _c is the cooling fluid blowing speed.

As can be seen from the above equation, the film cooling effect is M
and are closely related to x. Figure 4 shows an example of the cooling effect η when film cooling is performed on the ventral side of the wing body, with M as a parameter.In order to obtain the best cooling effect, an appropriate mass flow rate ratio must be You need to choose M. In particular, since the distance that the film lasts on the ventral side of the wing body is shorter than that on the dorsal side of the wing body 2, it is necessary to provide multiple rows of small holes 23 for cooling the film.

However, in the case of a conventional blade, the bending channel 22
Since the cooling fluid led to the film is directly led to the small hole 23 for cooling the film, it is not possible to set an appropriate mass flow rate ratio M, and as a result, the cooling of the ventral side in the middle part N in particular is It was not possible to improve performance. That is, the bending channel 22
Assuming that the respective parts of the blade are A, B, and C as shown in Figure 1, the pressure at these positions A, B, and C drops linearly as shown in Figure 5, and the As shown in FIG. 6, the pressure P ₁ on the ventral outer surface is much higher than the pressure P ₂ on the dorsal outer surface.
Under such conditions, it is essentially difficult to set an appropriate mass flow rate ratio M using the conventional structure, resulting in the drawback that the ventral outer surface of the intermediate portion cannot be cooled well.

[Purpose of the invention]

The present invention has been made in view of the above circumstances, and its purpose is to provide a cooling system using a cooling fluid, in particular, a bending channel in the middle part of the blade body.
In a system that is cooled by a so-called return flow method and a film cooling method, the ventral outer surface of the intermediate portion can be cooled well by the film cooling method, thereby further improving the efficiency of the gas turbine. Our goal is to provide turbine blades.

[Summary of the invention]

The gas turbine blade according to the present invention has a fluid flow path dedicated to film cooling, which guides the cooling fluid only to a small hole for film cooling opened in the ventral outer surface of the intermediate portion of the blade body, as described above. It is characterized by being provided independently from the flow path.

〔Effect of the invention〕

With the above configuration, the blowing pressure, flow rate, etc. of the cooling fluid blowing out from the small film cooling hole opened on the ventral outer surface of the intermediate portion of the wing body can be freely set without being influenced by other factors. Therefore, it is possible to select the optimum mass flow rate ratio M according to the position of each small hole, so the ventral outer surface of the intermediate part can be cooled better than in the conventional case, and as a result, the gas turbine You can obtain something that can contribute to improving efficiency. In addition, since the cooling fluid is supplied to the small holes for film cooling only from the fluid channel dedicated to film cooling, calculations for determining the flow rate distribution and mass flow ratio M between each fluid channel are extremely simple. ,
It is possible to exhibit cooling characteristics as designed.

[Embodiments of the invention]

FIG. 7 is a view showing a blade of a gas turbine according to an embodiment of the present invention, corresponding to FIG. 2, and the same parts are designated by the same reference numerals. Therefore, the explanation of the overlapping parts will be omitted.

In this embodiment, a plurality of small holes 31 for film cooling are provided in the height direction of the blade body 2, with one end opening on the outer surface of the blade body 2 located on the ventral side X at the intermediate portion N. In addition, four channels 32a, 32b, 32c, and 32d are provided in the wing body 2 in a four-row configuration in the chord direction of the wing body 2, and the other end side of each row of small holes 31 communicates with each row. The passages 32a, 32b, 32c, and 32d extend to the blade root and are connected to cooling fluid supply passages (not shown), respectively.

With such a configuration, the small holes 31 located in each row are independent of the bending channel 22 and are connected to dedicated flow channels 32a, 32b, 32c, and 32d provided for each row.
Cooling fluid will be supplied from Therefore, the mass flow rate ratio M can be freely set for each row by individually adjusting the pressure in the flow channels 32a, 32b, 32c, and 32d, so that the mass flow rate ratio M can be set freely compared to conventional blades. The side outer surface can be cooled well, and the above-mentioned effects can be obtained after all.

In addition, in the embodiment described above, each flow path 32
Although the means for distributing the cooling fluid to the channels a, 32b, 32c, and 32d is not shown, it can be easily set by providing an orifice plate or the like at the cooling fluid inlet of each flow path. Furthermore, by changing the flow cross-sectional area of each flow path 32a, 32b, 32c, and 32d in the height direction, the flow velocity of the cooling fluid blown out from each row of small holes 31 can be set as desired. can. Further, the number of columns is not limited to the number in the example. Furthermore, it goes without saying that the present invention can be applied to both rotor and stator blades.

[Brief explanation of the drawing]

FIG. 1 is a longitudinal sectional view showing a conventional gas turbine blade equipped with a cooling mechanism cut along a camber line;
Fig. 2 is a view of the wing cut along the Z-Z line in Fig. 1 and viewed in the direction of the arrow, Figs. 3 to 6 are illustrations for explaining problems with conventional wings, and Fig. 7 is FIG. 2 is a cross-sectional view of a wing according to an embodiment of the present invention, corresponding to FIG. 2; 2...Blade body, 22...Bending channel, 17, 23, 3
1...Small holes for film cooling, 32a, 32b, 3
2c, 32d...flow path.

Claims (1)

[Claims] 1. A convection cooling effect by causing a part of the cooling fluid guided into the blade body of the gas turbine to flow through a bent passage formed in the blade body, and a convection cooling effect on the ventral and dorsal outer surfaces of the blade body. In a gas turbine blade in which the intermediate portion of the blade body is cooled by a film cooling effect by blowing air from an open small hole, cooling is provided only in the small hole opened on the ventral outer surface within the blade body. A blade for a gas turbine, characterized in that a fluid passage dedicated to film cooling that guides fluid is provided independently of the bent passage.