JPH1061405A - Stator vane for axial flow turbomachine - Google Patents

Abstract

(57) [Problem] To provide a stationary blade of an axial flow type turbomachine capable of reducing fluid loss as a whole blade. In a stationary blade of an axial flow turbomachine arranged in a plurality of annular cascade flow paths, the stationary blade has a blade trailing edge viewed from the rotor axial direction and a projected shape thereof viewed from the same direction. It has a convex shape in the blade pressure surface direction, and at least a region from the root portion to a predetermined height in the radial direction of the blade from the end of the blade has a curved shape convex to the blade pressure surface direction side when viewed from the same direction. A curved area is formed, and a predetermined area including a radial center area of the stationary blade is formed as a linear area formed in a straight line, and a projected shape of a leading edge and a trailing edge of the stationary blade on a meridional plane is formed. Has a convex shape on the front side of the rotor blade, and a region from the root of the end of the blade to a predetermined height in the radial direction of the blade has a curved region convex from the end to the front side. A predetermined linear region including a position that is formed and that is a half of the blade length is formed as a linear linear region.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a stationary blade of an axial type turbo machine such as a steam turbine, a gas turbine, and an axial compressor.

[0002]

2. Description of the Related Art As a conventional example of a stationary blade structure of an axial flow turbomachine, there is known an investing of Leaned Nozzle Effects o.
n Low Pressure Steam Turbine Efficiencies, Proc. o
f theAdvances in Steam Turbine Technology for Powe
As described in r Generation PWR-Vol.10 ASME Power division, there is a structure in which a stationary blade is simply tilted in the rotating blade rotation direction. This is intended to suppress peeling of the root portion of a long stationary blade such as a steam turbine low pressure stage. Also, The Influence of Blade Dean o
n Turbine Losses, ASME Paper No. 90-GT-55
There is a structure having an arcuate shape in which the trailing edge line of the stationary blade is symmetrical in the height direction of the blade when viewed from the axial direction.
This is intended to suppress the development of secondary flow vortices generated in the vicinity of the inter-blade side wall of the stationary blade. In contrast,
There is a structure described in Japanese Patent Publication No. 6-92723 in which the trailing edge of the stationary blade is viewed from the axial direction and has an asymmetrical bow shape in the height direction of the blade. Further, there is a structure having an arcuate shape that is asymmetrical in the height direction of the blade when the trailing edge line of the stator blade is viewed from the axial direction or the meridional plane as described in JP-A-6-81603. .

[0003]

The prior art described above is aimed at suppressing the separation of the base portion of the stationary blade and the boundary layer or secondary flow vortex developed near the side wall between the blades. There is a limit to the reduction of fluid loss of the steel.

An object of the present invention is to provide a stationary blade of an axial flow type turbomachine capable of reducing fluid loss as a whole blade.

[0005]

A first feature of the present invention is as follows.
In the stationary blade of the axial-flow type turbomachine arranged in the annular cascade flow passage, the shape of the trailing edge of the stationary blade viewed from the rotor axis direction is convex toward the blade pressure surface side when viewed from the same direction. In the end portion of the blade, a region from at least the root portion to a predetermined height in the radial direction of the blade from the root portion is formed as a curved curved region convex toward the blade pressure surface direction side when viewed from the same direction, A straight linear region is formed in a predetermined region including the position of half the blade length of the stationary blade, and the shape viewed from the meridian plane of the leading edge and the trailing edge of the stationary blade is convex toward the front side of the moving blade. In the end portion of the wing, at least a region from the root portion to a predetermined height in the radial direction of the wing is formed with a curved region that is convex from the end portion to a front stage side, and the wing length is A straight line region is formed in a predetermined region including a half position. It is preferable that the curved region is formed together with a root portion and a tip portion of the blade length.

[0006] Thus, the development of the side wall boundary layer can be suppressed, the secondary flow vortex generated in the flow between the blades can be reduced, and the blade length can be reduced without largely changing the outflow angle and the like of the fluid flowing through the stator blades. Since the loss near the center can be reduced at the same time, the fluid loss of the entire blade can be reduced.
The upstream side is, for example, an upstream side of steam flowing through the blade or a supply port side that supplies steam to a steam turbine including the blade.

It is preferable that the curvature of the curved region is formed so as to gradually decrease as the distance from the end portion increases. Alternatively, it is preferable that the curvature of the curved region is formed so as to gradually increase as the distance from the end portion increases, and then to gradually decrease.

[0008] A second feature of the present invention is that the curved area is:
Let h be the blade length, and let b be the maximum circumferential displacement from the radial radiation passing through the blade root position to the blade trailing edge line in the projected shape of the blade trailing edge viewed from the rotor axis direction, and b be the projected shape on the meridional plane. Where f is the maximum displacement in the rotor axis direction from the blade root position of the leading edge or trailing edge to the leading edge or trailing edge of the blade, 0 <b / h <0.2, 0 < f /
h <0.2. By satisfying the above range, the loss can be further reduced.

[0009]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS A first embodiment of the present invention will be described below with reference to FIGS. FIG. 1 is a view of a trailing edge line of a stationary blade when the stationary blade of the present invention is viewed from a downstream axial direction. FIG.
FIG. 3 is a diagram of a leading edge line and a trailing edge line of the vane of the present invention when viewed from the meridian plane. 1 and 2 can be projected shapes when viewed from various directions and the like.

1 is an upper diaphragm, 2 is a lower diaphragm, 3 is a stationary blade, 4 is a rotating direction of a moving blade, 5 is a leading edge of a stationary blade,
Numeral 6 indicates the trailing edge of the stationary blade, and numeral 7 indicates the flow direction (of steam). Stationary wing 3
The place where the lower diaphragm 2 contacts with the base 2 of the stationary blade 3
1, where the stationary blade 3 and the upper diaphragm 1 are in contact with each other is shown as a tip portion 22 thereof. Also, 31 indicates the blade pressure surface direction, and 32 indicates the blade suction surface direction.

Here, the circumferential inclination angle θ is an angle formed by a tangent to the trailing edge of the blade at each radial position of the stationary blade and a line connecting the center of the rotation axis and the trailing edge of the blade. Θ can also be approximated by the angle between the tangent to the trailing edge of the blade at each radial position of the stationary blade and the extension of the straight line region of the blade.

The axial inclination angle δ is an angle between a tangent to the leading edge of the blade at each radial position of the stator blade and a line passing through the leading edge at the same position and perpendicular to the rotation axis. Alternatively, it can be approximated by an angle between a tangent to the trailing edge of the blade at each radial position of the stationary blade and an extension of the straight line region of the blade.

In the present embodiment, the shape of the trailing edge of the stationary blade viewed from the rotor axis direction has a convex shape on the blade pressure surface side as viewed from the same direction, and the root portion and the tip end of the blade end portion. In the region from the portion to the predetermined height in the radial direction of the blade, a curved curved region that is convex toward the blade pressure surface direction side when viewed from the same direction is formed, and includes a position of half the blade length of the stationary blade. In a predetermined region, a linear straight region is formed, and the shape of the stationary blade as viewed from the meridional plane of the leading edge and trailing edge has a convex shape on the front stage side of the moving blade, and among the ends of the blade, At least a region from the root portion to a predetermined height in the radial direction of the wing is formed in a predetermined curved region having a curved shape convex to the front stage from the end portion, and a predetermined region including a position that is half the wing length, A straight line region is formed.

For example, as shown in FIG. 1, the trailing edge line of the stationary blade 3 is directed to the rotating blade rotation direction 4 near the root and the tip so that the maximum inclination angle θ in the circumferential direction of the upper diaphragm 1 and the lower diaphragm 2 is increased. From the value, the radius is smoothly reduced so as to have an arcuate shape toward the center in the radial direction, and the curvature is reduced except at the vicinity of the root and the tip.

By adopting such a flow path shape, the flow is pressed toward the side wall near the root and the tip, and the development of the boundary layer and the secondary flow vortex developed on the side wall is suppressed. In the region where the secondary flow has little effect on the flow, it is possible to prevent the flow from being adversely affected, so that the loss of the entire blade due to the flow can be reduced.

Further, as shown in FIG. 2, the leading edge 5 of the stationary blade is directed toward the leading edge 5 of the stationary blade near the root and the tip, from the maximum value of the circumferential inclination angle θ at the upper diaphragm 1 and the lower diaphragm 2. The shape is smoothly reduced so as to form an arcuate shape toward the center in the radial direction, the curvature is reduced at portions other than the vicinity of the root and the tip, and the shape has a linear shape at least for one section.

By adopting such a flow path shape, the flow is pushed toward the side wall near the root and the tip, and the development of the boundary layer and the secondary flow vortex developed on the side wall is suppressed. Thus, it is possible to prevent the flow from being adversely affected in a region where the influence of the secondary flow is small, thereby reducing the loss due to the flow. In this case, no inclination is given in the circumferential direction, so that the outflow angle at the blade outlet is not affected.

As an example of a configuration for obtaining the above-mentioned effect, in the case of a blade having a blade length of 97.3 mm, the blade trailing edge shape viewed from the downstream side in the rotor axial direction, and the blade leading edge and trailing edge shape viewed from the meridian plane , Each curved area can be 32 mm. However, the length of the curved region is not limited to this, and can be appropriately selected as long as the above-mentioned effect can be obtained. For example, it can be selected from a range of about 25 to 40 mm.

FIG. 5 shows the relationship between the blade position and the blade length. In this embodiment, the distribution of the loss of the flow generated by the stationary blade is reduced over the entire area in the blade length direction as shown in FIG.

A is a conventional stationary vane which does not have an arcuate shape, b is a comparative stationary vane which is uniformly arcuate both circumferentially and axially, and c is a stationary vane of this embodiment. It is. As is clear from FIG. 5, the stator blade of this embodiment has the smallest loss over the entire region from the root to the tip. It should be noted that the stator vane b of the comparative example in which the conventional stator vane a has been improved has a lower loss near the side wall than the conventional stator vane a, but has a reverse effect near the center of the vane. It was found that the loss tended to increase. As a result, it was found that there was a limit in reducing the loss as a whole.

In the present embodiment c, since the loss due to the boundary layer at the end can be reduced and the loss near the center of the blade can be reduced, the loss can be reduced over the entire region in the blade length direction. For this reason, the loss as a whole can be further reduced.

FIG. 6 shows the distribution of the outflow angle at the outlet of the stationary blade in the blade length direction.

The symbols a, b, and c used in FIG.
It is the same as FIG. According to the vane c of the present invention, the outflow angle distribution of the vane b is equal to the root and the vicinity of the tip, and the outflow angle distribution of the conventional vane a is equal to the other region around the center of the vane. .

With the stationary blade having the structure of this embodiment, a deviation in the relative speed of steam or the like introduced to the moving blade located behind the stationary blade can be suppressed, so that the moving blade has a favorable performance over the entire blade length. Steam or the like can be introduced in the direction of relative speed.

Further, the blade having the shape of this embodiment can be applied to a blade having a large secondary flow. In a steam turbine facility including a plurality of turbines, for example, a high-pressure turbine or an intermediate-pressure turbine is used.

Further, for example, it is preferable that the length is such that a synergistic effect between the straight region and the curved region can be sufficiently exhibited. For example, it is preferable to adapt to a wing having an aspect ratio (wing height h / wing width w) of about 0.3 to 3.5.

For a high pressure turbine to which steam generated in a boiler or the like is supplied first, the same ratio is 0.3 to 2.5. For a reheat turbine to which steam output from the high pressure turbine is reheated and supplied, It is desirable that the ratio be applied to a wing having a ratio of about 0.5 to 3.5.

The curvature of the curved area is formed so as to gradually decrease as the distance from the end increases.

The curvature at the end can be a curvature of a circular arc having a radius approximately equal to the length of the curved area.

Alternatively, the curvature of the curved region is formed so as to gradually increase as the distance from the end portion increases, and then to gradually decrease. For example,
The position where the curvature is maximized is near a region where the other diagonal intersects with the curve when a quadrangle having an end and a point where the curved region ends as one diagonal is drawn.

Further, according to the present embodiment, the loss due to fluid can be reduced in the entire blade area as described above by changing the stacking manner of the blade sections constituting the blade without changing the blade shape itself. It is possible to provide a steam turbine blade having stable blade strength and a steam turbine including the blade.

Further, since an existing airfoil can be used, the existing turbine can be modified so as to be adapted to a required stage blade such as a steam turbine having a conventional vane.

FIGS. 7 and 8 show a second embodiment of the present invention. Basically, the configuration of the first embodiment can be applied. FIG.
FIG. 3 is a view of a trailing edge line of the stationary blade of the present invention when viewed from the downstream rotor axial direction, similarly to FIG. 1. FIG. 8 is a diagram of a leading edge line and a trailing edge line of the vane of the present invention when viewed from the meridian plane. In the present embodiment, in order to improve the flow at the blade root portion where the side wall boundary layer and the secondary flow strongly appear, the lower diaphragm 2 is arranged so that the trailing edge of the stationary blade 3 is directed to the blade rotation direction 4 only near the root. From the maximum value of the inclination angle θ in the circumferential direction, the shape is smoothly reduced so as to have an arcuate shape toward the center in the radial direction, the curvature is reduced except at the vicinity of the root, and at least one section has a linear shape. To Similarly, from the maximum value of the circumferential inclination angle θ of the lower diaphragm 2, the leading edge 5 of the stationary blade is formed in an arcuate shape toward the center in the radial direction so as to approach the leading edge 5 of the stationary blade near the root. The shape is smoothly reduced, the curvature is reduced at a portion other than the vicinity of the root, and the shape has a linear shape at least for one section.

By adopting such a flow path shape, the flow near the blade root where the secondary flow appears strongly is easily pushed in the side wall direction, and the boundary layer developed on the root side wall and the secondary flow vortex are suppressed. Since the flow loss near the center of the blade can be reduced, the loss can be reduced to some extent as a whole, though not as much as in the first embodiment. This is effective when the flow loss at the root is large.

When the boundary layer develops at the tip of the blade and the secondary flow is large, the arcuate portion, which is the curved region applied in the second embodiment, may be applied to the tip of the blade. In this case, a similar effect can be obtained.

A third embodiment will be described with reference to FIGS. In the third embodiment, basically, the configuration of the first embodiment can be applied.

9 and 10, the maximum displacement of the circumferential arcuate shape from the radial radiation passing through the blade root position X when viewed from the axial direction of the vane according to the present invention is b, and the blade root in the meridional plane is b. Is the maximum displacement of the axially arcuate shape from the axial position Y of f.

In the present embodiment, the curved area is defined as a blade length h, and a circumferential direction from a radial ray passing through a blade root position in a projected shape of the blade trailing edge viewed from the rotor axial direction to a blade trailing edge line. Is the maximum displacement in the rotor axial direction from the blade root position of the leading edge or trailing edge to the leading edge or trailing edge in the shape of projection on the meridian plane, and f is the maximum displacement of each. 0 <b / h <0.2, 0 <f / h <0.2.

FIG. 11 shows the relationship between b / h and f / h and loss when b and h are simultaneously changed by the same amount in the present invention. From the figure, b / h and f / h are each 0.
If <b / h <0.2, 0 <f / h <0.2, then B = f
The flow loss can be reduced as compared with the case where = 0. Thus, the range in which the effect of the present invention is further exhibited can be defined by the maximum displacement of each of 0 <b / h <0.2, 0 <f / h <0.2. From these findings, the present invention is not necessarily limited to b = f
It does not need to be. Further, when a value of b or f exceeding the above range is used, the three-dimensional channel shape cannot perform a predetermined blade function, and conversely, the flow loss increases as compared with the case of b = f = 0. You can see that

Therefore, this embodiment can also be applied to a wing having a wing length h of about 97.3 mm as described above.

A fourth embodiment will be described with reference to FIG.
This embodiment is basically applicable to the configuration of the first embodiment. FIG. 12 shows a stationary blade 3 and a moving blade 12 having a certain radial height.
FIG. In the figure, assuming that the chord length is c, the length of the camber line of the blade is 1 and the lift coefficient is CL, and the height in the radial direction of the secondary flow vortex from the side wall is s, a method of estimating steam turbine efficiency And as a result, Turbomachinery Vol. 11, No. 4,
S = 0.28c as described in (1983)
((l / c) √CL) ＾ 0.85. However,
The pitch (at the root of the blade) is t and the vane inlet angle is λ
₁ , the exit angle of the stationary blade is λ ₂ , λ∞ = tan ~ ¹ (2 / (cotλ ₁ + cot
λ ₂ )), CL = 2 (t / c) (cotλ ₂ −cotλ ₁ )
sinλ∞. Here, w is the wing span. From the above, s can be obtained if the geometric shape and arrangement of the stationary blade are determined. Therefore, when the radial height of the arcuate portion, which is the curved region, is r, the arcuate shape is within the range of 0 <r <s. Make up the part.

[0042] With this, emphasis is placed on reducing the loss due to the fluid in the central portion of the total loss, and the loss near the central portion excluding the region near the end portion is more reliably reduced while reducing the overall loss. Can be reduced.

For example, it is preferable to adapt to a case where a secondary flow occurs due to a boundary layer or the like at an end portion in a range smaller than the calculated s value. In this case, it is desirable to adapt to a wing having an aspect ratio larger than 2.0 in consideration of the influence of the boundary layer.

Further, by forming so that s ≦ r <1/2 · h, the secondary flow loss at the end portion of the total loss can be reduced at the radial height of the secondary flow blade. Is emphasized, and the loss in the same part can be more reliably reduced while reducing the loss as a whole.

Based on these values, a curved portion is appropriately formed in a region where the secondary flow occurs, and a linear portion is formed in a region where the influence of the secondary flow is small. The effect can be further exhibited. In this case, the aspect ratio is 2.
It is preferred to adapt to wings below zero.

The reference s for obtaining the height r is, for example, h = 97.3 mm, t = 38.8 mm, λ ₁ = 90 deg,
λ ₂ = 13.8 deg, l = 90.0 mm, c = 74.6 mm,
Accordingly, λ∞ = 26.2 deg and Cl = 1.87. Therefore, s is about 32 mm. Based on this, it is possible to set the height r in the radial direction of the arcuate portion, which is a curved region, and the like.

FIGS. 3 and 4 are cross-sectional views of the stationary blade of the comparative example, showing the case where the stationary blade moves to the pressure surface side and the upstream side as it leaves the side wall, respectively.

As shown in the comparative diagram of FIG. 3, in the case of a blade having a predetermined inclination in the circumferential direction by moving the stacking state of the blade cross section toward the pressure surface, only the circumferential inclination angle θ is increased. In this case, the outflow angle of the flow near the side wall is deflected in the axial direction from the design point. It is not preferable from the viewpoint of the outflow angle. Also, as in the comparative example of FIG. 4, even if the blade is simply inclined to the upstream side in the axial direction, the loss suppressing effect may be obtained at the end portion than the straight blade, but the loss as a whole blade is The reduction effect is limited.
Further, the loss may increase depending on the degree of δ.

When the stationary blade of the present embodiment is used, the radial distribution of the inclination angle θ in the circumferential direction is made to have an appropriate arcuate shape protruding in the direction of the blade pressure surface, and the distribution of the inclination angle δ in the radial direction is simultaneously adjusted. It can be distributed so that it has an appropriate bow shape that protrudes in the blade leading edge direction, and the stator blade shape is such that the radial distribution of the circumferential inclination angle θ is made into an arc shape that protrudes in the blade pressure surface direction. Compared with this, the flow is pressed against the side wall more, and the development of the side wall boundary layer can be suppressed and the secondary flow vortex generated in the inter-blade flow can be reduced without largely changing the outflow angle.

At the same time, since the secondary flow vortex develops around the side wall, especially in the case of a wing having a long blade length, by forming an arcuate shape only near the upper and lower side walls, the secondary flow vortex near the center of the blade is formed. The loss can be reduced without adversely affecting the flow field having relatively little influence.

In the flow path between blades of the axial-flow type turbomachine,
The bending of the flow path generates a centrifugal force, and an imbalance between this and a pressure gradient generates a secondary flow vortex. That is, in the boundary layer developed on the upper and lower side walls between the blades, the circumferential pressure gradient prevails the centrifugal force, and a velocity component is generated in the circumferential direction to form a pair of vortices. This vortex is called a secondary flow vortex and causes fluid loss. However, since the generation region develops around the side wall, the influence is small in other regions sandwiching the center of the blade. Therefore, in the case of a blade with a long blade, if the contour of the blade is formed in a curved shape with a large curvature everywhere in the axial direction or the radial direction as viewed from the meridian plane, it will be in the area where the influence of secondary flow vortices near the blade center is small. This may have an adverse effect on the blade, resulting in an increase in the loss of the conventional blade profile compared to a straight vane.

By applying the stationary blade of the above embodiment,
The loss can be suppressed.

[0053]

According to the present invention, it is possible to provide a stationary blade of an axial flow type turbomachine capable of reducing fluid loss as a whole blade.

[Brief description of the drawings]

FIG. 1 is a projection view showing an inclination of a trailing edge of a stationary blade when the stationary blade of one embodiment is viewed from a downstream axial direction.

FIG. 2 is a projection view of a leading edge and a trailing edge when the stationary blade of one embodiment is viewed from a meridian plane.

FIG. 3 is a projection view of a comparative example showing a blade surface inclination when the stationary blade moves toward the pressure surface side as moving away from the side wall.

FIG. 4 is a projection view of a comparative example showing the blade surface inclination when the stationary blade moves upstream as the distance from the side wall increases.

FIG. 5 is a comparison diagram of a loss distribution of a flow generated in a stationary blade.

FIG. 6 is a comparison diagram of a vane outflow angle distribution.

FIG. 7 is a projection view showing the inclination of the trailing edge of the stationary blade when the stationary blade of the embodiment is viewed from the axial direction on the downstream side.

FIG. 8 is a projection view of the vane of the embodiment as viewed from the meridian plane.

FIG. 9 is a front view of the wing of the embodiment.

FIG. 10 is a side view of the wing of one embodiment.

FIG. 11 is a diagram relating to a loss when b / n and f / h are simultaneously changed.

FIG. 12 is a sectional view of a stationary blade and a moving blade according to one embodiment.

[Explanation of symbols]

DESCRIPTION OF SYMBOLS 1 ... Upper diaphragm, 2 ... Lower diaphragm, 3 ... Stator blade, 4 ... Rotating blade rotation direction, 5 ... Stator blade leading edge, 6 ... Stator blade trailing edge,
7: flow direction, 8: stationary blade at side wall, 9: stationary blade at position away from side wall, 10: inflow direction of stationary blade, 11: outflow direction of stationary blade, 12: moving blade.

 ──────────────────────────────────────────────────の Continuing from the front page (72) Kuniyoshi Tsubouchi, Inventor 7-2-1, Omika-cho, Hitachi City, Ibaraki Pref. Hitachi, Ltd.

Claims (7)

[Claims] 1. A stationary blade of an axial flow turbomachine arranged in a plurality of annular cascade passages, wherein a shape of a trailing edge of the stationary blade viewed from a rotor axis direction is a blade pressure when viewed from the same direction. A surface region having a convex shape on the surface side, and at least a region from the root portion to a predetermined height in the radial direction of the blade from among the end portions of the blade is a curved curved region convex toward the blade pressure surface direction side when viewed from the same direction. Is formed in a predetermined area including a position of a half of the blade length of the stationary blade, a linear straight region is formed, and the shape of the stationary blade viewed from a meridian plane of a leading edge and a trailing edge is a moving blade. The front end of the wing has a convex shape, and at least a region from the root portion to a predetermined height in the radial direction of the wing from the end portion of the wing is formed with a curved region convexly convex from the end portion to the front side. An axial flow type turbine, wherein a linear region is formed in a predetermined region including a position where the blade length is half. Stationary wing of bo machine. 2. The axial-flow turbomachine according to claim 1, wherein the curvature of the curved region is formed so as to gradually decrease as the distance from the end portion increases. Wings. 3. A stationary blade of an axial-flow type turbomachine according to claim 1, wherein the curvature of the curved region is formed so as to gradually increase with distance from the end portion, and then gradually decreases. A stator vane for an axial flow turbomachine characterized by being formed. 4. The stationary blade of an axial-flow type turbomachine according to claim 1, wherein said curved region is a chord length of c, a length of a camber line of the blade is l, a lift coefficient is CL, and said linear region is Assuming that the height in the radial direction from the blade root side end of the starting stator blade is r, it is formed so as to satisfy the relationship of 0 <r <0.28c ((l / c) √CL) ＾ 0.85. A stationary vane of an axial flow turbomachine characterized by the following. 5. The stationary blade of an axial-flow turbomachine according to claim 1, wherein said curved region has a chord length of c, a length of a blade camber line of 1, and a lift coefficient of CL. If the height in the radial direction from the blade root end of the stationary blade at which the linear region starts is r, and the blade length is h, 0.28c ((l / c) √CL) ＾ 0.85 ≦ r < A stator vane for an axial-flow type turbomachine formed to satisfy a relationship of 1/2 · h. 6. A vane for an axial-flow turbomachine according to claim 1, wherein said curved region extends from a root portion and a tip portion of a blade end to a predetermined height in a radial direction of the blade. The region of (1), wherein the curved region is formed, is a vane of the axial-flow type turbomachine. 7. A vane for an axial flow turbomachine according to any one of claims 1 to 5, wherein said curved region has a blade length h and a blade in a projected shape of said blade trailing edge viewed from a rotor axial direction. Let b be the maximum circumferential displacement from the radial radiation passing through the root position to the wing trailing edge line, and the rotor from the wing root position of the leading or trailing edge in the projected shape to the meridian plane to the wing leading or trailing edge. When the maximum displacement in the axial direction is f, the respective maximum displacements are 0 <b / h <0.2, 0 <f /
A stationary blade for an axial-flow type turbomachine, wherein h <0.2.