JP2000204903A - Axial turbine - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a highly efficient axial turbine by reducing shock wave loss and incident angle loss near a low pressure last stage. SOLUTION: In an axial turbine, blades displaced in a circumferential direction on a blade pressure surface side against a straight line radially extending a back edge line 5 of the blade from a root back edge part of the blade in the radial direction are juxtaposed in plural lines on diaphragms 3, 4 in the circumferential direction, and the diaphragms 3, 4 having the blades are juxtaposed in plural stages in an axial direction. In the axial turbine, provided that a root pitch of the blade is (t) and maximum displacement for the back edge line 5 of the blade against the straight line radially extending from the root back edge part of the blade in the radial direction is δ, δ/t=C(constant) value ranges from 0.6 to 1.5 in the blade near a low pressure last stage of the turbine.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to an axial turbine and, more particularly, to a plurality of blades having a shape displaced to the blade pressure surface side are arranged in a circumferential direction on a diaphragm, and the diaphragm having the blades is arranged in the axial direction. The present invention relates to an axial flow turbine which is provided in a plurality of stages in parallel.

[0002]

2. Description of the Related Art Among the wings generally used in the prior art, the simplest technique for improving the flow inside the paragraph by changing the shape or structure of the wing is to improve the degree of reaction of the wing at the root. Is simply inclined toward the circumferential blade pressure surface side.

[0003] This includes, for example, a paper preprint collection Pro
c. of the Advances in Steam Turbine Technology for
Power Generation PWR-Vol.10 The title of the paper published in ASME Power Division is `` An Investigation of Leaned N
ozzle Effects on Low Pressure Steam Turbine "and is called Tangential Lean Tsubasa.

Further, as a technique for further improving the internal flow of the paragraph by designing the wing structure three-dimensionally, a bow in which the trailing edge of the wing is symmetrical at the center in the span direction when viewed from the axial direction. Some are shaped (curved). This includes, for example, ASMEPaper No. 90-GT-55,
Thesis title `` The Influence of Blade Lean on Turbine Los
ses "and the like, and is intended to reduce the loss due to the secondary flow vortex generated near the side wall of the stationary blade.

Further, as disclosed in Japanese Patent Application Laid-Open No. Sho 62-170707, the trailing edge of the blade is projected in the circumferential direction, and the amount of projection is maximum at the end of the trailing edge of the blade on the upper wall side of the diaphragm. Some have a wing structure. In addition, as shown in JP-A-3-189304, a stationary blade having an asymmetrical bow shape when the trailing edge of the stator blade is viewed from the axial direction, and in JP-A-6-081603. As described above, there are also vanes having an arcuate shape that are asymmetric in the span direction when viewed from the axial direction or the rear surface of the child.

[0006] The above-mentioned blade having a three-dimensionally designed blade structure with a curved trailing edge line of the blade is generally called a "curved stacking blade". It is a means of suppressing boundary layers and secondary flow vortices. However, with these blades, the amount of protrusion of the blade pressure surface changes the degree of reaction of the blade and the angle of incidence of the flow to the downstream moving blade. It becomes difficult. To solve this, Japanese Patent Laid-Open No. 10-13 / 1998
Japanese Patent No. 1707 proposes an optimum value for reducing the side wall loss based on the ratio of the protrusion amount of the blade pressure surface of the Curved Stacking blade to the blade root pitch.

[0007]

The above-mentioned blade is used in which the trailing edge line of the blade is displaced to the circumferential blade pressure surface side with respect to the straight line radially extending from the blade root trailing edge in the radial direction. The technique of suppressing the separation of the stator blade root and the boundary layer or secondary flow vortex that develops near the side wall has the effect of reducing loss when the flow inside the turbine stage is a subsonic flow. On the other hand, when the flow inside the turbine stage is transonic or supersonic flow, as in the last stage of the low pressure of the turbine and the stage before it,
Reducing the shock wave loss by reducing the Mach number of the flow is necessary for improving the efficiency of the turbine. For that purpose, the above publication (Japanese Unexamined Patent Application Publication No. 10-13170)
No. 7) is not enough to control the generation of a shock wave.

That is, in the inter-blade flow near the low-pressure final stage of the axial flow turbine, that is, in the inter-blade flow in the low-pressure last stage of the axial flow turbine and one stage before it, the shock wave is generated as the flow becomes transonic or supersonic. appear. Since this causes fluid loss, it is necessary to suppress the occurrence. On the other hand, in each stage of the axial flow type turbine composed of multiple stages, the flow pattern is designed so that the degree of reaction of the turbine and the angle of incidence of the flow on the rotor blades become necessary values.

However, when the blade length is long, such as the blade used in the low-pressure final stage and the stage preceding the low-pressure stage, the blade root reaction rate decreases, and the outflow angle from the stationary blade deviates from the inflow angle into the moving blade. Furthermore, in a blade in which the blade pressure surface is displaced in the circumferential direction, a blade load change in the span direction occurs, so that it is difficult to obtain a flow pattern as designed. This makes it difficult to match the angle of incidence of the flow to the moving blade with the mechanical angle of the moving blade, causing loss of the incident angle.

The present invention has been made in view of the foregoing, and has as its object the object of the present invention is in the vicinity of the low pressure final stage of an axial turbine,
For example, it is an object of the present invention to provide a highly efficient axial flow turbine by reducing shock wave loss and incident angle loss in a low pressure final stage of an axial flow turbine and a stage preceding the low pressure stage.

[0011]

That is, the present invention relates to a blade having a shape in which a trailing edge line of a blade is displaced in a circumferential direction toward a blade pressure surface side with respect to a straight line radially extending from a blade root trailing edge in a radial direction. In an axial flow turbine in which a plurality of diaphragms having the blades are arranged in the circumferential direction on the diaphragm and a plurality of diaphragms having the blades are arranged in the axial direction in a plurality of stages, the blade root pitch is t, and the radius from the blade root trailing edge is radius. Δ / t = C, where δ is the maximum displacement of the trailing edge line with respect to a straight line extending radially in the direction.
The value of (constant) is formed so as to be in the range of 0.6 <C <1.5 for blades near the low pressure final stage of the turbine so as to achieve the intended purpose.

Further, according to the present invention, a blade having a shape displaced in the circumferential direction toward the blade pressure surface side with respect to a straight line obtained by radially extending the trailing edge line of the blade from the trailing edge portion of the blade root is formed on the diaphragm. In the axial flow turbine in which a plurality of diaphragms having the blades are juxtaposed in the axial direction, a plurality of diaphragms having the blades are juxtaposed in the axial direction, and the blade root pitch is radially extended from the blade root rear edge in the radial direction. Assuming that the maximum displacement of the trailing edge line with respect to the straight line is δ, the value of δ / t = C (constant) is 1.0 <C <1.5 at the low pressure final stage of the turbine, and before the low pressure final stage. In one stage, it is formed so as to be in the range of 0.6 <C <0.9.

[0013] Further, the wing trailing edge line is curved and displaced so as to protrude toward the wing pressure surface side with respect to a straight line obtained by radially extending the trailing edge line of the wing from the trailing edge of the blade root, or the wing trailing edge line is straight. A plurality of blades having a shape inclined toward the circumferential blade pressure surface side while being maintained in a row are arranged on the diaphragm in the circumferential direction, and the diaphragm having the blades is arranged in a plurality of stages in the axial direction. In the type turbine, when the blade root pitch is t and the maximum bending displacement of the blade trailing edge line with respect to a straight line radially extending from the blade root trailing edge in the radial direction is δ,
The value of / t = C (constant) is in the range of 1.0 <C <1.5 in the low pressure final stage of the turbine, and 0.6 <C <0.9 in the first stage before the low pressure final stage. It was formed so that it might become.

[0014]

DESCRIPTION OF THE PREFERRED EMBODIMENTS The present invention will be described below in detail with reference to the illustrated embodiments. First, FIG. 1 shows a meridional shape of a general steam turbine. The stationary blade 1 is fixed to the upper wall 4 and the lower wall 3 of the diaphragm, and controls the flow direction of the steam flow 6 into the moving blade 2. A plurality of the stationary blades 1 are arranged on the diaphragm in the circumferential direction, and a plurality of diaphragms having the stationary blades are arranged in the axial direction.

FIGS. 2 and 3 show shapes of the trailing edge line 5 of the stationary blade 1 as viewed from the downstream side. FIG. 2 shows a Curved Stacking blade in which the trailing edge of the blade has a curved shape and protrudes toward the circumferential pressure surface, and FIG. 3 shows a straightened trailing edge of the blade. This is a tangential lean blade that is inclined toward the circumferential pressure surface.

The stationary blade of the present invention is particularly formed as follows. That is, assuming that the blade root pitch is t and the maximum displacement of the blade trailing edge line with respect to a straight line radially extending from the blade root trailing edge in the radial direction is δ, the value of δ / t = C (constant) is Is formed in the range of 0.6 <C <1.5 in the blade near the low pressure final stage.

That is, when the blade of the present invention is applied to the stationary blade 1 with the blade displaced toward the circumferential blade pressure surface side with respect to a straight line radially extending from the blade root rear edge in the radial direction from the blade root trailing edge. ,
The trailing edge line 5 of the blade is displaced toward the blade pressure surface in the circumferential direction with respect to a straight line 7 extending radially in the radial direction from the trailing edge of the blade root. The blade length is defined by H, the maximum protrusion amount is defined by δ, and the blade root pitch is defined by t.

The blade whose trailing edge line as shown in FIGS. 2 and 3 is displaced toward the circumferential blade pressure surface side with respect to a straight line radially extending from the blade root trailing edge in the radial direction is described above. In addition to the suppression of secondary flow vortices as described above, there is an effect of reducing the shock wave loss due to the reduction of the root Mach number and the rotor blade incident angle loss as the outflow angle increases.

By presenting the similar parameters of the blade flow, the degree of reaction of the turbine at the low pressure final stage of the axial flow type turbine composed of multiple paragraphs and the stage immediately preceding the low pressure stage and the position of the rotor blade located downstream are determined. By controlling the angle of incidence and further limiting the range of similarity parameter values, it is possible to reduce losses in each paragraph and improve efficiency. In particular, by displacing the blade pressure surface based on the similar parameters presented in the present invention, the same loss reduction effect can be obtained even if the number of blades, blade length, and plant are changed.

In order to present similar parameters of blades displaced toward the blade pressure surface in the circumferential direction with respect to a straight line obtained by radially extending the trailing edge line of the blade radially from the trailing edge of the blade root, the blade discharge angle is the same. The condition which becomes becomes. The outflow angle is a representative quantity representing the effect of the blade which is displaced toward the blade pressure surface in the circumferential direction with respect to a straight line radially extending the trailing edge line of the blade radially from the blade root trailing edge. In the above-mentioned blade, the blade load changes in the span direction, and accordingly, the outflow angle changes in the span direction. Therefore, by examining the amount of change in the outflow angle, the effect of the blade displaced to the circumferential blade pressure surface side with respect to the straight line radially extending the trailing edge line of the blade from the blade root trailing edge is evaluated. be able to.

In order to simulate the flow around such a wing, a plurality of symmetric Curved S at the center in the span direction between parallel walls.
The flow in a cascade test with tacking wings was simulated using three-dimensional turbulence analysis. As a result of performing a parameter survey using the above turbulence analysis, the ratio δ / t between the maximum protrusion amount δ and the blade root pitch t shown in FIGS.
Was found to be a similar parameter for obtaining a similar outflow angle distribution. The calculation results are shown below.

FIG. 4 shows the outflow angle distribution in the span direction when the blade length H is 120 mm, the blade root pitch is fixed at 36.1 mm, and the protrusion amount δ is changed. FIG.
The horizontal axis shows the position in the span direction in a dimensionless manner by the blade length H. From FIG. 4, it can be seen that as the protrusion amount δ is increased, the change in the outflow angle increases with the change in the blade load in the span direction. In particular, the outflow angle near the side wall is large, which means that the degree of reaction at the root portion of the turbine stage increases, and it is possible to reduce the outflow Mach number at the base of the stationary blade.

It is possible to control the value of the protrusion amount δ by finding a similar parameter for controlling such a change in the outflow angle, and further determining the range of the similar parameter so as to optimize the efficiency of the turbine stage. This is the main feature of the present invention.

FIG. 5 shows the outflow angle distribution in the span direction when the blade length H is changed while the value of the ratio δ / t of the protrusion amount δ and the blade root pitch t is kept constant. From FIG. 5, it can be seen that the outflow angle distribution in the span direction is substantially similar by keeping the value of δ / t constant even when the value of the blade length H changes. FIG. 6 shows a case where the value of the ratio δ / t between the protrusion amount δ and the blade root pitch t is kept constant while the blade length H is kept constant and the blade root pitch t is enlarged by expanding the blades in a similar manner. The outflow angle distribution in the span direction is shown.

FIG. 6 shows that even if the value of the blade root pitch t changes due to the similar enlargement of the blade, the value of the ratio δ / t of the protrusion amount δ to the blade root pitch t is kept constant, so that the span direction is maintained. It can be seen that the outflow angle distributions are almost similar.

From the results of the three-dimensional turbulence analysis described above, it can be seen that in the blade in which the trailing edge of the blade is displaced toward the circumferential blade pressure surface side with respect to a straight line radially extending from the blade root trailing edge in the radial direction, It was shown that the ratio δ / t between the protrusion amount δ and the blade root pitch t is a similar parameter that determines the blade outflow angle distribution in the span direction. This is applied to the stationary blade of the axial-flow turbine, and is shown in FIGS. 2 and 3, which are views of the stationary blade viewed from the downstream side.

Generally, a turbine is composed of a plurality of paragraphs, and the blade length and the number of blades, that is, the blade root pitch are different in each paragraph. When applying a blade displaced to the circumferential blade pressure surface side with respect to a straight line radially extending in the radial direction from the portion, when forming a cascade, the ratio δ / t of the protrusion amount δ and the blade root pitch t is determined. A configuration using similar parameters is performed. By implementing such a cascade configuration, it is possible to control the degree of reaction of the turbine stage and the outflow angle from the stationary blade, that is, the angle of incidence on the rotor blade located on the downstream side.

In the following, a case is used in which the trailing edge of the blade is displaced to the circumferential blade pressure surface side with respect to a straight line radially extending from the blade root trailing edge in the radial direction to the stationary blade of the axial flow turbine. of,
The effect of the change in the value of the similarity parameter δ / t on the flow around the wing will be described. First, as the first effect,
A description will be given of a change in the inflow angle at the blade root portion by changing δ / t of the stationary blade.

If the angle of inflow to the moving blade does not match the mechanical angle of the moving blade, a loss of the inflow angle occurs, leading to a reduction in efficiency. Therefore, it is necessary to construct a blade that matches these angles. FIG. 7 shows the change in the distribution of the blade inflow angle in the span direction when the value of δ / t of the stationary blade is changed. Thereby, δ / t
It can be understood that the blade root inflow angle increases with an increase in the value of.

The solid line in FIG. 7 shows the distribution of the mechanical angles of the moving blades in the span direction. When not using a blade that is displaced to the circumferential blade pressure surface side with respect to a straight line that radially extends the trailing edge line of the vane radially from the blade root trailing edge to the stator vane, or δ
When a blade having a small value of / t is used, the blade inflow angle near the blade root is smaller than the mechanical angle of the blade. If the value of δ / t is too large, the blade inflow angle near the blade root becomes larger than the mechanical angle of the blade. Therefore, by setting the value of δ / t in an appropriate range, the inflow angle at the blade root and the mechanical angle can be matched, and the inflow angle loss can be reduced.

Next, as a second effect, a description will be given of a change in the exit Mach number at the base portion of the stationary blade due to a change in δ / t of the stationary blade. FIG. 8 shows the spanwise distribution of the exit Mach number of the stationary blade when the value of δ / t is changed. Thus, it is understood that the root Mach number of the blade is reduced as the value of δ / t is increased. Therefore, δ / t
Is larger, the shock wave loss at the root of the blade can be reduced. However, in order to increase the value of δ / t, the value of the protrusion amount δ must be increased or the blade root pitch t
Needs to be reduced. If the value of the protrusion amount δ is increased, the blade area increases, and if the value of the blade root pitch t is decreased, the number of blades increases, so that both increase the contact area between the fluid and the wall in the turbine stage, and the friction loss is reduced. Increase. Therefore, by limiting the range of the value of the similarity parameter δ / t, it is possible to maximize the effect of reducing the loss.

In one stage before the low pressure final stage of the steam turbine, the trailing edge of the vane is displaced toward the circumferential blade pressure surface side with respect to a straight line radially extending from the trailing edge of the blade root in the radial direction. Even when the blade is not used, the Mach number flowing into the rotor blade is 0.9 or less, the flow near the rotor blade inlet is a subsonic flow, and no shock wave is generated. Therefore, in the first stage before the final stage of low pressure, use a vane in which the trailing edge line of the vane is displaced to the circumferential blade pressure surface side with respect to a straight line radially extending from the blade root trailing edge in the radial direction. Contributes mainly to the effect of reducing the shock wave loss in the stationary blade.

On the other hand, in the low pressure final stage, a vane is used in which the trailing edge line of the vane is displaced to the circumferential blade pressure surface side with respect to a straight line radially extending from the trailing edge of the blade root in the radial direction. If not, the Mach number flowing into the moving blade exceeds 0.9, and the flow field around the moving blade becomes transonic flow because it is accelerated on the suction side of the moving blade. Occurs. Therefore, in the low pressure final stage, the vane is used in which the trailing edge line of the vane is displaced to the circumferential blade pressure surface side with respect to the straight line radially extending from the trailing edge of the blade root in the radial direction. The effect of reducing the shock wave loss can be obtained in both the blade and the moving blade.

Hereinafter, the first embodiment of the present invention will be described with reference to FIGS.
0 will be described. 9 and 10 show the ratio δ between the protrusion amount δ and the blade root pitch t in three different plants.
/ T is plotted on the abscissa, and the ordinate is a blade whose trailing edge is displaced toward the circumferential blade pressure surface side with respect to a straight line radially extending from the blade root trailing edge in the radial direction. The turbine stage efficiency in the case where the blade trailing edge line is radially extended from the blade root trailing edge in a radial direction to the straight line radially extending from the blade efficiency The paragraph efficiency improvement amount Δη (%) is shown.

FIG. 9 shows the result when the Curved Stacking blade is used, and FIG. 10 shows the result when the tangential lean blade is used. 9 and 10, L-0 represents the last stage of the low-pressure stage of the steam turbine, and L-1 represents the paragraph efficiency improvement amount Δη (%) in the stage before L-0. These results were obtained by performing a three-dimensional turbulence analysis on the flow between the vanes and the moving vanes simultaneously.

9 and 10, first, in the low pressure final stage, the ratio δ / of the protrusion amount δ and the blade root pitch t at which the value of the paragraph efficiency improvement amount Δη (%) becomes the maximum value depends on the plant.
It can be seen that the range of the value of t is 0.6 <δ / t <0.9. In the first stage before the low-pressure final stage, the range of the ratio δ / t of the protrusion amount δ and the blade root pitch t at which the value of the paragraph efficiency improvement amount Δη (%) becomes the maximum value is 1.
It can be seen that 0 <δ / t <1.5.

As described above, the range of the ratio δ / t of the protrusion amount δ and the blade root pitch t at which the value of the paragraph efficiency improvement amount Δη (%) becomes the maximum value is different between the low pressure final stage and the preceding stage. That is, as described above, in the first stage before the low-pressure final stage, the blade in which the trailing edge line of the blade is displaced to the circumferential blade pressure surface side with respect to a straight line radially extending radially from the blade root trailing edge portion. This is because the effect of reducing the shock wave loss due to the use mainly contributes to the stationary blade, and contributes to the reduction of the shock wave loss in both the stationary blade and the moving blade in the low pressure final stage.

As described above, the efficiency of an axial flow turbine using a blade displaced toward the circumferential blade pressure surface side with respect to a straight line radially extending the trailing edge line of the blade radially from the blade root trailing edge portion. In order to improve the protrusion, the protrusion amount δ and the blade root pitch t
Of the ratio δ / t, the blade length and the number of blades, and even if the plant changes, in the low pressure final stage, 1.0 <δ / t <1.1.
5. 0.6 <δ / t <0.1 in the first stage before the low pressure final stage.
By selecting an appropriate value from the range of 9, efficiency can be improved.

Next, a second embodiment of the present invention will be described with reference to FIG. FIG. 9 shows three different plants:
The horizontal axis represents the ratio δ / t of the protrusion amount δ and the blade root pitch t, and the vertical axis represents the trailing edge line of the blade in the turbine stage efficiency when using the above Curved Stacking blade in the radial direction from the blade root trailing edge. FIG. 7 shows the stage efficiency improvement amount Δη (%) from the turbine efficiency when the blade displaced to the circumferential blade pressure surface side with respect to the radially extended straight line is not used. From this, in order to improve the efficiency of the axial flow turbine using the Curved Stacing blades, the value of the ratio δ / t of the protrusion amount δ and the blade root pitch t is changed by changing the blade length, the number of blades, and even the plant. , At the low pressure final stage, 1.0 <δ /
t <1.5, 0.6 <δ / in one stage before the low pressure final stage
Efficiency improvement can be achieved by selecting an appropriate value from the range of t <0.9.

A third embodiment of the present invention will be described with reference to FIG. FIG. 10 shows the ratio δ / t of the protrusion amount δ and the blade root pitch t in three different plants on the horizontal axis.
On the vertical axis, the turbine stage efficiency when using the tangential lean blade described above is displaced to the circumferential blade pressure surface side with respect to a straight line that extends the trailing edge of the blade radially from the trailing edge of the blade root in the radial direction. The graph shows the amount of increase in the stage efficiency Δη (%) from the turbine efficiency when the blade is not used.

Thus, in order to improve the efficiency of the axial turbine using the tangential lean blade, the value of the ratio δ / t between the protrusion amount δ and the blade root pitch t is determined by changing the blade length, the number of blades, and the plant. , An appropriate value is selected from the range of 1.0 <δ / t <1.5 in the low pressure final stage and 0.6 <δ / t <0.9 in the first stage before the low pressure final stage. This makes it possible to achieve an improvement in efficiency.

[0042]

As described above, according to the present invention, the shock wave loss and the incident angle loss in the vicinity of the low pressure final stage of the axial flow turbine, for example, in the low pressure final stage of the axial flow turbine and one stage before it, are reduced. An efficient axial turbine can be obtained.

[Brief description of the drawings]

FIG. 1 is a vertical sectional side view showing one embodiment of a main part of an axial flow turbine of the present invention.

FIG. 2 shows a curved stack of the stationary blade of the axial flow turbine of the present invention.
FIG. 3 is a vertical sectional front view of a turbine stage using ing blades.

FIG. 3 is a longitudinal sectional front view of a turbine stage using a tangential lean blade as a stationary blade of the axial flow turbine of the present invention.

FIG. 4 shows a change in the outflow angle in the span direction with respect to a protrusion δ of a blade displaced toward a circumferential blade pressure surface side with respect to a straight line obtained by radially extending a trailing edge line of a blade radially from a blade root trailing edge. FIG.

FIG. 5: The blade length was changed while the δ of the blade was displaced toward the circumferential blade pressure surface side with respect to a straight line obtained by radially extending the trailing edge line of the blade from the blade root trailing edge in the radial direction, and changing the blade length. FIG. 10 is a characteristic diagram showing a change in the outflow angle in the span direction in the case.

FIG. 6 shows that the wing of the wing is displaced toward the wing pressure surface in the circumferential direction with respect to a straight line obtained by radially extending the trailing edge line of the wing radially from the trailing edge of the wing root, and the wing size and the wing root FIG. 9 is a characteristic diagram illustrating a change in the outflow angle in the span direction when the pitch is changed.

FIG. 7 shows the dynamics when the δ / t of a blade displaced toward the blade pressure surface in the circumferential direction is changed with respect to a straight line obtained by radially extending the trailing edge line of the blade radially from the blade root trailing edge. It is a characteristic view which shows the distribution of the blade inflow angle in the span direction.

FIG. 8 is a graph showing a static state when δ / t of a blade displaced toward a circumferential blade pressure surface side is changed with respect to a straight line obtained by radially extending a trailing edge line of a blade radially from a blade root trailing edge. FIG. 5 is a characteristic diagram showing a spanwise distribution of an exit Mach number of a blade.

FIG. 9 is a characteristic diagram showing the relationship between the magnitude of δ / t and the amount of improvement in paragraph efficiency when Curved Stacking blades are used.

FIG. 10 shows δ when a tangential lean wing is used
FIG. 9 is a characteristic diagram showing a relationship between the magnitude of / t and the paragraph efficiency improvement amount.

[Explanation of symbols]

Reference numeral 1 represents a stationary blade, 3 represents a lower wall of a diaphragm, 4 represents an upper wall of a diaphragm, 5 represents a trailing edge line of a stationary blade, 6 represents a flow of steam.

 ────────────────────────────────────────────────── ─── Continued from the front page (72) Inventor Shigeki Senoo 7-2-1, Omika-cho, Hitachi City, Ibaraki Prefecture

Claims (4)

[Claims] 1. A plurality of blades having a shape displaced in a circumferential direction toward a blade pressure surface side with respect to a straight line obtained by radially extending a trailing edge line of a blade from a blade root trailing edge portion in a radial direction on a diaphragm. In an axial flow turbine in which a plurality of diaphragms having the blades are juxtaposed in the axial direction in a plurality of stages, the blade root pitch is t, and the blades with respect to a straight line radially extending radially from the blade root rear edge in the radial direction. The maximum displacement of the trailing edge line is δ
Where the value of δ / t = C (constant) is formed in the range of 0.6 <C <1.5 in the blades near the low-pressure final stage of the turbine. Turbine. 2. A plurality of blades having a shape displaced in the circumferential direction toward the blade pressure surface side with respect to a straight line obtained by radially extending the trailing edge line of the blade from the trailing edge of the blade root toward the blade pressure surface. In an axial flow turbine in which a plurality of diaphragms having the blades are juxtaposed in the axial direction in a plurality of stages, the blade root pitch is t, and the blades with respect to a straight line radially extending radially from the blade root rear edge in the radial direction. The maximum displacement of the trailing edge line is δ
Where δ / t = C (constant) is 1.0 <C <1.5 at the low pressure final stage, and 0.6 <C <0.5 at the first stage before the low pressure final stage. An axial-flow turbine formed in the range of 9. 3. A wing having a shape that is curved and displaced so as to protrude toward the wing pressure surface side with respect to a straight line obtained by radially extending a trailing edge line of the wing from the trailing edge of the blade root to a wing pressure surface side. In an axial flow turbine in which a plurality of diaphragms having the blades are arranged in parallel in a plurality of stages in the axial direction, the blade root pitch is t, with respect to a straight line radially extending in the radial direction from the blade root trailing edge. Assuming that the maximum bending displacement of the blade trailing edge line is δ, the value of δ / t = C (constant) is 1.0 <C <1.5 in the low pressure final stage of the turbine, and 1 before the low pressure final stage. An axial flow turbine, wherein the stages are formed in a range of 0.6 <C <0.9. 4. A plurality of blades having a shape inclined toward the circumferential blade pressure surface side while keeping the trailing edge line of the blade straight, are arranged side by side in the circumferential direction on the diaphragm, and the diaphragm having the blade is axially mounted. In the axial-flow turbine having a plurality of stages arranged side by side in the direction, the blade root pitch is t, and the maximum displacement of the blade trailing edge line with respect to a straight line radially extending from the blade root trailing edge in the radial direction is δ.
Where δ / t = C (constant) is 1.0 <C <1.5 at the low pressure final stage, and 0.6 <C <0.5 at the first stage before the low pressure final stage. An axial-flow turbine formed in the range of 9.