JPH11200802A - Blades for turbomachinery - Google Patents

Abstract

(57) [Summary] An object of the present invention is to provide a liquid crystal display device having a negative incident angle.
Suppress the increase in the thickness of the fluid boundary layer on the ventral side of the leading edge of the rotor blade, or suppress the increase in energy loss depending on the incident angle by suppressing the separation of the fluid. It is in. In order to solve the above problems, a turbomachine rotor blade according to the present invention uses a point A at which the x-coordinate becomes minimum in the profile of an airfoil cross section near the leading edge, and an airfoil near the leading edge. Y at the cross-sectional contour
Point B is the point where the coordinates are minimum, point C is the point where the y coordinate is equal to point A in the profile of the airfoil section near the leading edge, and points B and x are equal.
Assuming that a point having the same coordinates and located on the line segment AC is a point D, the length of the line segment DC is at least twice the length of the line segment DB, and an arc-shaped arc formed near the point B. The radius of curvature of the profile of the airfoil section is made larger than the radius of curvature of the profile of the arcuate airfoil section forming the leading edge.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a compressor for compressing a fluid (air or the like) and a rotor blade for a turbomachine used for a turbine driven by the fluid (combustion gas, steam, air or the like). The present invention relates to a rotor blade for a turbomachine used in a steam turbine of a thermal power plant and a nuclear power plant, a gas turbine of a thermal power plant and an aircraft, a supercharger of an automobile, and the like.

[0002]

2. Description of the Related Art A rotor blade for a turbomachine used in a compressor and a turbine (hereinafter, referred to as a "rotor blade") is provided on a plurality of cylindrical surfaces around a rotation axis or a flow obtained by analysis. Designing a two-dimensional airfoil on multiple conical surfaces corresponding to the surface, arranging the two-dimensional airfoil in the spanwise direction, and interpolating between the two-dimensional airfoils in the spanwise direction Design by In general, the shape of the leading edge of a two-dimensional airfoil (the most upstream part of the fluid flowing into the bucket) is determined by the radius of the arc forming the leading edge (hereinafter, referred to as
It is called "leading edge radius". ), The included angles of the tangents at both ends of the arc forming the leading edge (hereinafter, referred to as “leading edge wedge angle”), and the included angle between the direction opposite to the rotating direction of the moving blade and the leading edge direction (hereinafter, referred to as “the leading edge wedge angle”). "Entrance angle") is determined as a parameter.

[0003] The inlet angle is designed to oppose the expected inflow angle (the inflow angle of the fluid flowing into the blade) at any two-dimensional airfoil position of the blade. However, the flow actually flowing into the moving blade is caused by the loss caused by the stationary blade located on the upstream side of the moving blade and the side wall boundary layer of the fluid at the tip (blade top) and root (blade bottom) of the moving blade. As a result, the flow velocity of the fluid locally decreases, so that the inflow direction of the fluid may be deflected in a direction facing the rear side (convex side) of the bucket. The deviation between the predicted inflow angle (entrance angle) and the actual inflow angle is called the incident angle. And if the angle of incidence is zero or very small,
Low energy loss in flow. However, when the incident angle increases and exceeds a predetermined value, the energy loss of the fluid rapidly increases. This is because when the angle of incidence exceeds a certain value,
The thickness of the wing surface boundary layer increases, especially on the ventral side of the leading edge of the blade,
Alternatively, flow separation occurs.

Conventionally, in order to suppress such an increase in energy loss depending on the incident angle, the radius of the leading edge is simply increased or the wedge angle of the leading edge is increased.

[0005]

In particular, the entrance angle is small,
A moving blade having a small pressure drop in the moving blade has a large relative inflow velocity of fluid in the moving blade, and a flow path formed between the circumferentially adjacent moving blades (hereinafter, referred to as an “inter-blade flow path”). Since the pressure drop in the flow direction for stabilizing the flow in the above is also small, the rate of increase in energy loss when having an incident angle is large. However, the leading edge radius and the leading edge wedge angle have a large effect on the overall shape of the airfoil, and cannot be freely determined by considering only the incident angle characteristics as described above. In particular, in an airfoil with a small inlet angle and a small pressure drop at the moving blade, if the shape of the inter-blade flow path is appropriately designed to stabilize the flow, the leading edge radius can be increased or the leading edge wedge angle can be increased. Is difficult to reduce.

In addition, in the airfoil at the root of the impeller of the impulse stage of the steam turbine, which is generally considered to have a small inlet angle, if the leading edge radius is increased, the outer peripheral surface of the rotor develops and the side wall boundary which flows down. Due to the effects of layers and interference, the secondary flow, called a horseshoe-shaped vortex, increases and energy loss increases.

Therefore, in an airfoil having a small inlet angle and a small pressure drop at the moving blade, or an airfoil at the root of the moving blade in the impulse stage of the steam turbine, the leading edge radius is increased or the leading edge wedge angle is increased. And it is difficult to suppress an increase in energy loss depending on the incident angle.

It is an object of the present invention to provide a rotor blade for a turbomachine in which the shape of the leading edge is specified to suppress an increase in energy loss depending on the incident angle.

[0009]

In order to achieve the above object, a rotor blade for turbomachine of the present invention is fixed to a rotor, rotates together with the rotor with the inflow of fluid, and
The airfoil has an airfoil cross section having an inlet angle of 45 degrees or less, which is an included angle between the direction opposite to the rotation direction and the leading edge direction. Furthermore, on the cross section of the airfoil, an orthogonal coordinate system is defined in which the axial direction of the rotor and the direction in which the inflow direction of the fluid is positive is the x-axis and the rotational direction is the y-axis. A point at which the x coordinate of the orthogonal coordinate system is minimum in the profile of the airfoil section is point A, and a point B at which the y coordinate of the orthogonal coordinate system is minimum in the profile of the airfoil cross section near the leading edge. In the profile of the airfoil section near the edge, the point A
Point C and point B, where y and y coordinate of the rectangular coordinate system are equal.
And the point on the line segment AC where the x coordinate of the orthogonal coordinate system is equal is point D, the length of the line segment DC is at least twice the length of the line segment DB, and the vicinity of the point B Is larger than the radius of curvature of the profile of the arcuate airfoil section forming the leading edge.

[0010] Alternatively, in order to achieve the above object, a rotor blade for turbomachine of the present invention is fixed to a rotor, rotates together with the rotor with the inflow of fluid, and rotates in the opposite direction to the rotation direction. It has an airfoil section with an angle of 90 degrees or less obtained by adding an entrance angle indicating an included angle with the leading edge direction and an exit angle indicating an included angle between the direction opposite to the rotation direction and the trailing edge direction. Furthermore, on the cross section of the airfoil, an orthogonal coordinate system is defined in which the axial direction of the rotor and the direction in which the inflow direction of the fluid is positive is the x-axis and the rotational direction is the y-axis. A point at which the x coordinate of the rectangular coordinate system is minimum in the profile of the airfoil section is point A, and a point y in the rectangular coordinate system is the profile of the airfoil section near the leading edge.
Point B is the point where the coordinates are minimum, point C is the point on the contour of the airfoil section near the leading edge where the point A is equal to the y coordinate of the rectangular coordinate system, and point B is the x coordinate of the rectangular coordinate system. Are equal to each other and the point on the line segment AC is a point D, the length of the line segment DC is at least twice the length of the line segment DB and an arc-shaped wing formed near the point B. The radius of curvature of the profile of the mold section is greater than the radius of curvature of the profile of the arcuate airfoil section forming the leading edge.

Alternatively, in order to achieve the above object, a turbomachine rotor blade according to the present invention is a turbomachine rotor blade used in an impulse stage of a steam turbine, and has a shape similar to An x-axis is defined as the direction in which the fluid inflow direction is positive in the axial direction of the rotor of the steam turbine, and the orthogonal direction is defined as the rotation direction as the y-axis. The point at which the x coordinate of the rectangular coordinate system is minimum is point A, the point at which the y coordinate of the orthogonal coordinate system is minimum at the contour of the airfoil near the leading edge is point B, and the airfoil section near the leading edge. The point C is a point where the point A and the y coordinate of the Cartesian coordinate system are equal, and the point B is the point where the x coordinate of the Cartesian coordinate system is equal and the line segment AC
When the point at the top is point D, the shape of the airfoil cross section at the root is formed near line B, with the length of line segment DC at least twice the length of line segment DB. The radius of curvature of the profile of the arcuate airfoil section is greater than the radius of curvature of the profile of the arcuate airfoil section forming the leading edge.

[0012]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS An embodiment of a rotor blade for a turbo machine according to the present invention will be described below.

FIG. 1 is a cross-sectional view of a steam turbine to which a rotor blade for turbomachine of the present invention is applied. In FIG. 1, reference numeral 1 denotes steam from a steam generator (such as a boiler), 2 denotes a moving blade (blade) rotated by the steam 1, 3 denotes a shroud for suppressing vibration of the moving blade 1 due to rotation, and 4 denotes a moving blade 2. Vanes (nozzles) that guide the steam 1 to the rotor, 5 is a rotor of the steam turbine, 6 is a rotating shaft of the moving blades 2 and the rotor 5, 7 is an annular diaphragm outer ring,
Reference numeral 8 denotes an annular diaphragm inner ring, 9 denotes a steam turbine casing, 10 denotes a cylindrical surface centered on the rotating shaft 6, and 11 denotes a conical surface corresponding to a flow surface expected by analysis.

The rotor blade for a turbo machine according to the present embodiment is used as a rotor blade of a steam turbine of a thermal power plant, and is used particularly in an impulse stage thereof. The tip of the vane 4 (the outer peripheral side in the radial direction of the rotor 5) is fixed to the vehicle compartment 9 via the diaphragm outer ring 7. On the other hand, the root ends (the inner peripheral side of the rotor 5 in the radial direction) of the circumferentially adjacent stationary blades 4 are connected (fixed) using the diaphragm inner ring 8. A shroud 3 is provided at the tip of the moving blade 2. The shroud 3 is formed integrally with the bucket 2. Alternatively, it is fixed to the rotor blade 2 by Tenon caulking. Generally, the shroud 3 integrally formed with the rotor blade 2 is referred to as “integral shroud” or “integral cover” or “continuous shroud” or “continuous cover”. On the other hand, the root end of the rotor blade 2 is fixed to the outer periphery of the rotor 5.

The combination of the moving blades 2 and the stationary blades 4 is called a “paragraph”. Ten paragraphs are provided in the circumferential direction. Also, several paragraphs are provided in the axial direction of the rotor 5 (axial direction of the rotating shaft 6). As the flow of the steam 1 goes downstream, the blade lengths of the moving blades 2 and the stationary blades 4 increase.

Then, the steam 1 is accelerated in the moving blade 4 and the flow direction of the steam 1 is changed in the circumferential direction.
Lead to. When the steam 1 is guided to the moving blade 2, the steam 1 accelerates and expands in a flow path between the circumferentially adjacent moving blades 2, and the moving blade 2
Is fixed, and the rotor 5 rotates about the rotation axis. A generator connected to one end of the rotor 5 converts rotational energy (mechanical energy) into electrical energy to generate power.

The moving blade 2 and the stationary blade 4 have a plurality of cylindrical surfaces 10,
Alternatively, it is designed by designing a two-dimensional airfoil on a plurality of conical surfaces 11 and interpolating between the airfoils. Bucket 2
The cross section of the cylindrical surface 10 or the conical surface 11 is referred to as an airfoil cross section.

Hereinafter, terms used in the description of the present embodiment will be defined.

FIG. 2 shows an airfoil of an arbitrary cross section of the rotor blade for turbomachine of the present invention. In FIG. 2, 19 is the inflow direction of steam 1, 20 is a group of circles, 21 is a warp line (camber line), 22
Is the leading edge, 23 is the trailing edge, 24 is the leading edge circle, 24a is the leading edge arc, 25 is the trailing edge circle, 26 is the back face, 27 is the ventral surface, 28 is the chord line, 29 is the leading edge tangent, 30 is the trailing edge Edge tangent line, 31 is a circumferential direction (rotation direction of rotor blade 2 and rotor 5) line, αin is an entrance angle,
αaut indicates an exit angle, βin indicates an inflow angle, i indicates an incident angle, and WLE indicates a leading edge wedge angle.

The circle group 20 is a circle inscribed in the airfoil. The warp line 21 is a center line of an airfoil obtained by connecting the centers of the group of circles 20. The leading edge 22 refers to an intersection located on the upstream side of the airfoil among intersections of the warp line 21 and the airfoil surface (the profile of the airfoil). The trailing edge 23 is an intersection of the warp line 21 and the wing surface, which is located downstream of the airfoil. The leading edge circle 24 refers to a circle including the leading edge 22 in the circle group 20. The radius of the leading edge circle 24 is called the leading edge radius. The leading edge arc 24a is an arc that overlaps the leading edge circle 24 on the wing surface. Trailing edge circle 25 is circle group 2
0 means a circle including the trailing edge 23. The radius of the trailing edge circle 25 is called the trailing edge radius. The back surface 26 refers to a wing surface located on the convex side of the airfoil with respect to the warp line 21. The abdominal surface 27 refers to a wing surface located on the concave side of the airfoil with respect to the warp line 21. The chord line 28 is a line segment connecting the leading edge 22 and the trailing edge 23.
The length of the chord line 28 is referred to as a chord length c. Leading edge tangent 29
Denotes a tangent to the warp line 21 at the front edge 22. Of the leading edge tangent line 29, the direction toward the upstream of the steam 1 is called the leading edge direction. The trailing edge tangent 30 refers to the tangent of the warpage 21 at the trailing edge 23. The direction of the trailing edge tangent 30 toward the downstream of the steam 1 is referred to as the trailing edge direction. The entrance angle αin is equal to the leading edge tangent 29.
From the circumferential line 31 (angle between the leading edge direction and the circumferential line 31). The exit angle αout refers to the angle of the trailing edge tangent 30 from the circumferential line 31 (angle between the trailing edge direction and the circumferential line 31). The inflow angle βin refers to an angle from the circumferential line 31 in the inflow direction 19 (angle between the inflow direction 19 and the circumferential line 31).
The incident angle i is an angle obtained by subtracting the inflow angle βin from the entrance angle αin (i = αin−βin). The leading edge wedge angle WLE is
The angle formed by the tangents at both ends of the leading edge arc 24a. In addition, in the rotor blade 2 used in the impulse stage, the angle obtained by adding the inlet angle αin and the outlet angle αout of the airfoil at the root is generally 60 to 90 degrees, and is used in the reaction stage. Small compared to that of the moving blade. Further, in the rotor blade 2 used for the impulse stage, the inlet angle αin
Is 45 degrees or less, and from the root to the tip,
The entrance angle αin increases. That is, the airfoil is formed to be twisted from the root to the tip.

Hereinafter, the flow of the steam 1 between the moving blade 2 and the stationary blade 4 will be described.

FIGS. 3 and 4 show a general speed triangle between the moving blade and the stationary blade. 3 and 4, c1 is a velocity vector of the steam 1 flowing into the stationary blade 4, c2 and c2a are velocity vectors of the steam 1 flowing out of the stationary blade 4, c3 a velocity vector of the steam 1 flowing out of the moving blade 2, U is a velocity vector of the moving blade 2, w 2 and w 2 a are relative velocity vectors of the steam 1 flowing into the moving blade 2 with respect to the moving blade 2, and w 3 is outflow from the moving blade 2 with respect to the moving blade 2. 2 shows a relative velocity vector and a velocity vector of the steam 1. That is, the direction of the arrow indicates the direction of the speed, and the size of the arrow indicates the size of the speed. 3 and 4, reference numeral 32 denotes the rotation direction of the rotor blade 2 (circumferential line 31).
Is shown.

The flow of the steam 1 flowing into the stationary blade 4 (velocity vector c1) is accelerated in the stationary blade 4 and turned in the rotation direction 32 (velocity vector c2). Since the moving blade 2 is rotating at the speed vector U, the relative flow of the steam 1 in the rotating coordinate system of the outflow speed of the steam 1 from the stationary blade 4, that is, the inflow speed to the moving blade 2, is equal to the speed vector w 2. Become. The moving blade 2 is designed so that the entrance angle αin of the moving blade 2 matches the velocity vector w2.

However, as shown in FIG. 5, the outflow velocity of the steam 1 from the stationary blade 4 is reduced due to the stationary blade loss and the side wall boundary layer.
The speed is reduced from the speed vector c2 to the speed vector c2a. The deceleration region of the steam 1 due to the boundary layer on the blade surface of the stationary blade 4 and the shape of the trailing edge of the stationary blade 4 is referred to as “stationary vane wake”.
Due to the deceleration of the steam 1, the speed triangle between the stationary blade 4 and the moving blade 2 changes, and the speed vector w2 becomes the speed vector w2.
changes to a. The inflow angle of the velocity vector w2a has a negative incident angle i with respect to the inflow angle of the velocity vector w2, that is, the entrance angle αin of the moving blade 2. Then, when having a negative incident angle i, the abdominal surface 27 near the leading edge 22 of the rotor blade 2
The energy loss in the bucket 2 increases rapidly because the boundary layer of the flow of the steam 1 on the side becomes thicker or the steam 1 is easily separated from the ventral surface 27.

Hereinafter, the flow of the steam 1 near the leading edge 22 of the bucket 2 will be described.

FIG. 5 is a schematic view showing the vicinity of the leading edge when the incident angle i is zero in a conventional rotor blade for turbomachinery. FIG. 6 shows the relationship between the dimensionless static pressure p / P0 and the axial length when the incident angle i is zero in the conventional rotor blade for turbomachinery. FIG. 7 is a schematic view showing the vicinity of the leading edge of a conventional rotor blade for turbomachine when having a negative incident angle i. FIG.
2 shows the relationship between the dimensionless static pressure p / P0 and the axial length when the conventional blade for a turbomachine has a negative incident angle i. The dimensionless static pressure p / P0 is defined as the wing surface (the rear surface 26).
And the static pressure p on the abdominal surface 27) is made dimensionless by the inflow relative total pressure P0. Further, the axial length refers to the rotor 5
Of the moving blade 2 with respect to the axial direction of the moving blade 2 (the blade width of the moving blade 2). Further, the conventional blade for turbomachinery is a leading edge arc 2
4a, a front edge radius from the front edge 22 to the back surface 26 side;
This means that the radius of the front edge from the front edge 22 to the abdominal surface 27 is substantially equal. 5 to 8, reference numeral 40 denotes a stagnation point of the flow of the steam 1,
Reference numeral 41 denotes a streamline of the steam 1 leading to the front edge 22, 42 denotes a back streamline of the steam 1 flowing to the back surface 26 side, and 42a denotes a back stream line of the steam 26.
The upper back upstream line 42c is a back flow tube formed between any stream line 42 remote from the back surface 26 and the stream line 42a.
2d is the cross-sectional area of the back-side flow tube of the flow tube 42c, 42dmin is the minimum point of the cross-section of the back-side flow tube where the cross-sectional area 42d is minimum near the leading edge 22, and 43 is the ventral surface streamline of the steam 1 flowing to the ventral surface 27 side. , 43a are the abdominal surface upstream line on the abdominal surface 27, 43c are the abdominal surface 2
Ventral flow tube 43d formed between an arbitrary streamline 42 and streamline 43a away from 7; 43d, the cross sectional area of the ventral side flow tube 43c; 43dmin, the cross sectional area 43d near the leading edge 22; The minimum point on the cross section of the flow tube on the ventral side is shown.

As shown in FIG. 5, when the angle of incidence i is zero, it means that the inflow angle βin and the entrance angle αin match, that is, when the stagnation point 40 and the leading edge 22 match. Say. At this time, the stream line of the steam 1 is on the wing surface (the back surface 26 and the abdomen surface 27) like the back surface stream line 42 on the back surface 26 side and like the ventral surface stream line 43 on the ventral surface 27 side. It is smooth along. For this reason, the rear-side flow tube cross-sectional area 42d and the ventral-side flow tube cross-sectional area 43d do not have a portion that locally increases sharply in the flow direction (streamline direction) of the steam 1.

In FIG. 6, a point 40p where p / P0 = 1.0 corresponds to the stagnation point 40 in FIG.
The point 42p where the / P0 is constricted corresponds to the minimum point 42dmin on the back side flow tube cross section in FIG. 5, and the point 43p where the dimensionless static pressure p / P0 is minimum is the ventral side flow tube cross section in FIG. This corresponds to the minimum point 43dmin. Although the dimensionless static pressure p / P0 increases on the downstream side of the minimum point 43dmin on the abdominal side flow tube cross section, the amount of increase is small and does not cause an increase in the energy loss of the rotor blade 2. This corresponds to the fact that the above-described cross-sectional area 43d of the ventral side flow tube does not locally increase in the flow direction of the steam 1 (stream line direction).

In FIG. 7, the axis direction of the rotor 5 (the flow direction is positive) is the x-axis, and the rotation direction 32 of the rotor blade 2 is the y-axis.
Form an xy coordinate system. As shown in FIG. 7, the case where the stagnation point 40 has a negative incident angle i means the case where the stagnation point 40 is shifted to the vicinity of the point A where the x coordinate becomes minimum. Therefore, the ventral-side flow tube cross-sectional area 43d is minimized near the point B where the y coordinate is minimized near the leading edge 22, and rapidly increases downstream of the point B.

In FIG. 8 as in FIG. 6, p / P0 =
The point 40p at which 1.0 is obtained corresponds to the stagnation point 40 in FIG. 7, and the point 42p at which the non-dimensional static pressure p / P0 is narrowed corresponds to the minimum point 42dmin on the back side flow tube cross section in FIG. The point 43p at which the dimensionless static pressure p / P0 is a minimum corresponds to the minimum point 43dmin on the cross section of the ventral-side flow tube in FIG. Compared to FIG. 6, in FIG. 8, the minimum point 43 mm
Downstream of n, the dimensionless static pressure p / P0 rises sharply. The boundary layer of the rear upstream line 42a is rapidly thickened or peeled off due to the rapid pressure increase. Therefore, the mixing loss due to the mixing of the low-energy fluid in the boundary layer near the back surface 26 and the main flow having no velocity gradient increases, and the airfoil loss increases. Further, when the flow separates from the wing surface, the flow becomes unsteady, and the airfoil loss rapidly increases.

Such an increase in the airfoil loss due to the negative incident angle i results in a small reduction ratio of the flow path area formed between the circumferentially adjacent vanes 4 (the acceleration of the steam 1 in the vanes 4). Is small), and the velocity vector w2 of the steam 1 flowing into the bucket 2
Is large, and is remarkable in the rotor blade 2 whose entrance angle αin is 45 degrees or less.

In consideration of the above, in the present embodiment, the shape of the leading edge 22 of the moving blade 2 is specified, thereby suppressing an increase in energy loss depending on the negative incident angle i.

FIG. 9 is a schematic view of an airfoil of a rotor blade for turbomachine of the present invention. The inside of the circle is an enlarged view of the leading edge 22. As shown in FIG. 9, the inflow direction 19 of the steam 1 is set to the positive direction, the axial length is set to the x-axis, and the rotating direction 32 of the rotor blade 2 is set.
An xy coordinate system is set with the length of the y axis as the y axis. In FIG. 9, point A
Is the point where the x coordinate is minimum in the airfoil, point B is the point near the leading edge 22 of the airfoil where the y coordinate is minimal, and point C is the point A and y
A point D on the abdominal surface 27 having the same coordinates indicates a point on the line segment AC having the same x coordinate as the point B. In the present embodiment, the line segment DC
Is longer than twice the length of the line segment DB, and the radius of curvature at the point B is larger than the radius of curvature of the leading edge 22.

The cause of the increase in energy loss depending on the negative angle of incidence i is that the cross-sectional area 43 of the ventral flow tube near the leading edge 22 is large.
When d decreases rapidly near point B and increases rapidly from point B toward the downstream side of the abdominal surface 27, a strong reverse pressure gradient is generated. To suppress this strong reverse pressure gradient, the airfoil surface BC on the ventral surface 27 downstream of point B
It is necessary to increase the radius of curvature of. Therefore, first, the radius of curvature at the point B is made larger than the radius of curvature of the leading edge 22. Since the point B is a point where the x coordinate becomes a minimum near the leading edge 22 of the airfoil, the gradient (dy / dx) B becomes zero. Therefore, the average curvature of the airfoil surface BC on the abdominal surface 27 can be approximately determined by giving the gradient (DC / DB) of the line segment BC. In order to make the radius of curvature of the airfoil surface BC on the abdominal surface 27 large enough to suppress energy loss due to the negative incident angle i, the line segment BC
(DC / DB) must be greater than a predetermined threshold. According to experience, the predetermined threshold value is set to 2 in this embodiment. On the other hand, the gradient of the line segment BC (DC /
For example, in the design of the flow path formed between the rotor blades 2 adjacent to the circumferential line 31, (DC / D)
If B) is too large, the inlet flow path becomes narrow. Therefore, the flow path is determined by using a monotonous throttle flow path.

FIG. 10 is an enlarged schematic view showing the vicinity of the leading edge of the airfoil of the turbomachine rotor blade of the present invention. In FIG. 10, the xy coordinate system is defined as a point B as the origin, the length of the rotating blade 32 in the rotation direction 32 as the y-axis, and the length in the axial direction as dimensionless as the length of the line segment DB as the x-axis. Take. The airfoil surface BC can be approximated as a line segment BC. Therefore, assuming that the line segment BC is an arc having a radius r, the slope (dy / dx) B of the point B is zero, so that the airfoil surface BC is

[0036]

X ² + (y−r) ² = r ² (Equation 1) Further, if the coordinates of the point C are (xc, 1), the relationship between the radius r and xc, that is, the gradient (DC / DB) is as follows.

[0037]

(Equation 2)

Can be expressed by When the gradient (DC / DB) = xc = 2 as in the present embodiment, r = 2.5.
This indicates that the average curvature of the airfoil surface BC becomes about 2.5 times the length of the line segment DB. By having such a large radius of curvature on the airfoil surface BC, the steam 1
Of the leading edge 2 has a negative incident angle i.
The generation of a large reverse pressure gradient on the side of the abdominal surface 27 in the vicinity of 2 can be suppressed, and an increase in energy loss can be suppressed.

FIG. 11 shows the relationship between the dimensionless static pressure p / P0 and the axial length when the blade has a negative angle of incidence i in the rotor blade for turbomachine of the present invention. The rotor blade for a turbomachine (the rotor blade 2 of the present embodiment) of the present invention has a ventral side flow tube cross-sectional area 43d.
Can be suppressed from being locally reduced, so that the dimensionless static pressure p / P0 is located downstream of the minimal point 43dmin on the cross section of the flow tube on the ventral side.
The magnitude of the reverse pressure gradient generated on the downstream side of the point 43p at which the pressure becomes minimal can be reduced. Thereby, the moving blade 2
Has a negative angle of incidence i, the vapor 1
Energy can be prevented from suddenly becoming thicker, or the flow of the steam 1 can be prevented from separating from the abdominal surface 27, so that when the rotor blade 2 has a negative incident angle i, the energy loss sharply increases. This has the effect of suppressing the increase.

FIG. 12 shows the relationship between the inflow angle and the energy loss in the rotor blade for turbine machine of the present invention. FIG.
The deviation between the entrance angle αin and the inflow angle βin is the incident angle i, and has a negative incidence angle i when the value of the inflow angle βin is larger than the value of the entrance angle αin. Further, the minimum value of the energy loss is E (corresponding to the energy loss when βin = αin in the conventional type), and the energy loss is twice as large as E, that is, 2 times.
When the value of the larger inflow angle at E is γin, H
_Αin indicates an incident angle margin indicating a deviation between the inflow angle and γin when the energy loss is E.

According to FIG. 12, in the present invention, the energy loss for an inflow angle larger than the inflow angle at which the energy loss is minimized, that is, when there is a negative incident angle, is smaller than the conventional energy loss.

FIG. 13 shows the relationship between the gradient (DC / DB) and the incident angle margin in the rotor blade for turbine machine of the present invention. The incident angle margin H_αin is a gradient (DC / DB)
Rapidly increases at a position exceeding about 2. In addition, the stationary blade 4
Has a trailing edge thickness, and there is a minimum 45 at the negative angle of incidence caused by the trailing edge thickness. The incident angle margin H_αin with respect to the lower limit value 2 of the gradient (DC / DB) of the present invention is larger than the minimum value 45 of the negative incident angle as shown in FIG. When designing the rotor blade 2, it is necessary to set the gradient (DC / DB) to two or more predetermined values. Then, the predetermined value will be described with reference to FIG. 13 so that the incident angle margin H_αin is larger than the maximum incident angle caused by the stationary blade 4 and the side wall located upstream of the moving blade 2. It is determined. The predetermined value of the gradient (DC / DB) is preferably about 3 to 5 in design.

FIG. 14 shows an experimental result of the paragraph loss distribution in the rotor blade for turbine machine of the present invention. FIG. 15 shows the positional relationship between the measurement points of the stationary blade, the moving blade, and the paragraph loss. FIG.
5, 46 is a measurement point for measuring paragraph loss, 47 is a wake of the stationary blade (flow of steam 1 between the stationary blade 4 and the trailing edge and the leading edge 22 of the moving blade 2), and 48 is a wake of the moving blade. (Rearing edge 23 of moving blade 2 and measurement point 4
6). The paragraph loss refers to an energy loss from the leading edge of the stationary blade 4 to the trailing edge of the moving blade 2 (within the paragraph).

The paragraph loss is measured at a measurement point 46 located downstream of the trailing edge 23 of the bucket 2. The measurement point 46 is a position at which the moving blade wake 48 can be measured with respect to the stationary blade 4 when the moving blade wake 47 interferes with the leading edge 22 of the moving blade 2. Although the relative positional relationship between the stationary blade 4 and the measuring point 46 does not change, the relative positional relationship between the moving blade 2 and the measuring point 46 because the rotating blade 2 rotates around the rotation axis 6. Changes in the circumferential direction. Then, at the measurement point 46, a temporal change in paragraph loss due to the rotation of the bucket 2 is obtained. In FIG. 14, the time axis is replaced with the circumferential direction of the coordinate system fixed to the moving blade 2. In FIG. 14, the place where the paragraph loss is low, that is, the place where the distribution of the paragraph loss is constant, corresponds to the rotor blade wake 48. Further, where the paragraph loss is high, that is, the distribution of the paragraph loss is mountain-like, a negative angle of incidence is generated on the moving blade 2 due to the relative deflection of the stator blade wake 47 with respect to the leading edge 22, and the energy loss increases. To do so. According to FIG. 14, the present invention has a smaller energy loss due to a negative incident angle than the conventional type. That is, according to the present invention, the effect of reducing the energy loss due to the negative incident angle, thereby reducing the paragraph loss, improving the paragraph performance, and improving the efficiency of a turbo machine such as a steam turbine or a gas turbine. To play.

According to the present embodiment, the physical mechanism of the increase in the energy loss due to the negative incident angle is larger than the increase in the leading edge radius or the leading edge wedge angle of the conventional moving blade. Therefore, an effect of effectively suppressing an increase in energy loss due to a negative incident angle is obtained as compared with the related art. Further, since the shape feature is given in a dimensionless form, the influence on the entire airfoil is small. Therefore, the present embodiment can be applied to the vicinity of the root of the rotor blade 2 where the leading edge radius could not be increased, and the effect of suppressing an increase in energy loss due to a negative incident angle can be obtained. Play.

FIG. 16 shows a change in the gradient (dy / dx) of the surface near the leading edge of the rotor blade for turbomachine of the present invention. FIG.
FIG. 7 shows a schematic view of the vicinity of the leading edge of the rotor blade for turbomachine of the present invention. The curve 50 in FIG. 16 and FIG. 17A show the case where the airfoil surface BC has no inflection point. A curve 51 in FIG. 16 and FIG. 17B show a case where the airfoil surface BC has an inflection point E. Note that the x-axis in FIG. 16 corresponds to the x-axis in FIG.

Since the gradient (dy / dx) C of the point C is limited by the shape of the flow path formed between the adjacent moving blades 2,
The gradient (dy / dx) C at the point C is constant. Further, as described above, the gradient (dy / dx) B of the point B is zero by definition. When the airfoil surface BC having the gradients at both ends fixed as described above has the inflection point E, the gradient (dy) of the airfoil surface BC is smaller than when the airfoil surface BC does not have the inflection point E. / Dx) temporarily increases. That is, as shown in FIG. 16, in the curve 50 where the airfoil surface BC does not have the inflection point E, the gradient (dy / d) of the airfoil surface BC
While the maximum value of x) is the gradient (dy / dx) C of the point C, the curve 51 in which the airfoil surface BC has the inflection point E is represented by the gradient (inflection point E of the inflection point E on the airfoil surface BC ( dy / dx) E is the point C
Is larger than the gradient (dy / dx) C of the curve 51, and becomes the maximum value in the curve 51. Further, as shown in FIG. 17, when the airfoil surface BC has an inflection point E, the gradient (dy / dx) E becomes large near the inflection point E, so that the cross-sectional area 43d of the ventral surface side flow tube is reduced. The spread increases and the reverse pressure gradient increases. As described above, when the reverse gradient increases near the inflection point E, the boundary layer of the airfoil surface BC becomes thicker or peels off, so that the energy loss of the moving blade 2 increases.
Also, the flow is more stable on a convex surface than on a concave surface. In consideration of the above, as the airfoil shape in the vicinity of the leading edge 22 of the rotor blade 2, the length of the line segment DC is twice or more the length of the line segment DB, and the radius of curvature at the point B is In addition to making the radius of curvature larger than the radius of the edge 22, by making the airfoil surface BC not to have the inflection point E, the steam 1 flows into the rotor blade 2 at a negative incident angle. In such a case, it becomes more effective to suppress the boundary layer from becoming thicker or coming off near the leading edge 22 of the bucket 2, and the paragraph loss can be reduced.

In the above embodiment, a two-dimensional airfoil has been described, but the length of the line segment DC is at least twice the length of the line segment DB, and the radius of curvature at the point B is the radius of curvature of the leading edge. The same effect can be obtained by applying the criterion of larger size to all the design sectional airfoils of the three-dimensional moving blade.

[0049]

According to the present invention, an orthogonal coordinate system is defined on an airfoil cross-section, in which the axial direction of the rotor and the positive direction of the fluid inflow direction are the x-axis and the rotational direction is the y-axis. A point at which the x-coordinate of the orthogonal coordinate system is minimum in the profile of the airfoil section near the leading edge is point A, and the y coordinate of the rectangular coordinate system is minimum at the profile of the airfoil surface near the leading edge. A point B, a point C in the contour of the airfoil cross section near the leading edge where the point A and the y coordinate of the rectangular coordinate system are equal, a point C, and the point B and the x coordinate of the rectangular coordinate system are equal and a line When the point on the minute AC is point D,
The length of the line segment DC is at least twice the length of the line segment DB, and
By making the radius of curvature of the profile of the arc-shaped airfoil section formed near point B larger than the radius of curvature of the profile of the arc-shaped airfoil cross-section forming the leading edge, a negative angle of incidence is obtained. In this case, on the ventral side of the leading edge of the rotor blade, the thickness of the fluid boundary layer is suppressed from increasing, or the separation of the fluid is suppressed, thereby reducing the energy loss depending on the incident angle. This has the effect of suppressing the increase.

[Brief description of the drawings]

FIG. 1 is a cross-sectional view of a steam turbine to which a rotor blade for turbomachine of the present invention is applied.

FIG. 2 is a view showing an airfoil of an arbitrary cross section of a rotor blade for turbomachine of the present invention.

FIG. 3 is a diagram showing a general speed triangle between a moving blade and a stationary blade.

FIG. 4 is a diagram showing a general speed triangle between a moving blade and a stationary blade.

FIG. 5 is a schematic view of the vicinity of a leading edge when an incident angle i is zero in a conventional rotor blade for turbomachinery.

FIG. 6 is a diagram showing the relationship between the dimensionless static pressure p / P0 and the axial length when the incident angle i is zero in a conventional rotor blade for turbomachinery.

FIG. 7 is a schematic view of the vicinity of a leading edge of a conventional rotor blade for turbomachine when having a negative incident angle i.

FIG. 8 is a diagram showing the relationship between the dimensionless static pressure p / P0 and the axial length when the blade has a negative angle of incidence i in a conventional rotor blade for turbomachinery.

FIG. 9 is a schematic view of an airfoil of a rotor blade for turbomachine of the present invention.

FIG. 10 is an enlarged schematic view of the vicinity of the leading edge of the airfoil of the rotor blade for turbomachine of the present invention.

FIG. 11 shows the relationship between the dimensionless static pressure p / P0 and the axial length when the blade for a turbomachine of the present invention has a negative angle of incidence i.

FIG. 12 is a diagram showing a relationship between an inflow angle and an energy loss in a turbine machine rotor blade of the present invention.

FIG. 13 is a view showing a relationship between a gradient (DC / DB) and an incident angle margin in the rotor blade for turbine machine of the present invention.

FIG. 14 shows an experimental result of a paragraph loss distribution in the rotor blade for turbine machine of the present invention.

FIG. 15 is a diagram illustrating a positional relationship between measurement points of a stationary blade, a moving blade, and a paragraph loss.

FIG. 16 is a diagram showing a change in the gradient (dy / dx) of the surface near the leading edge of the rotor blade for turbomachine of the present invention.

FIG. 17 is a schematic view of the vicinity of a leading edge of a rotor blade for turbomachine of the present invention.

[Explanation of symbols]

1 steam, 2 blades, 3 shrouds, 4 vanes, 5 blades
Rotor, 6: rotating shaft, 7: diaphragm outer ring, 8: diaphragm inner ring, 9: vehicle compartment, 10: cylindrical surface, 11: conical surface, 19: inflow direction, 20: circle group, 21: warping line, 22
... leading edge, 23 ... trailing edge, 24 ... leading edge circle, 24a ... leading edge arc, 25 ... trailing edge circle, 26 ... back face, 27 ... ventral surface, 28 ... chord line, 29 ... leading edge tangent, 30 ... rear Edge tangent line, 31 circumferential line, 40 stagnation point of flow, 41 streamline, 42 back streamline, 42a back upstream line, 42c back flow tube, 42d
... cross-sectional area of the back flow tube, 42dmin ... minimum point of the cross-section of the back flow tube, 43 ... ventral streamline, 43a ... ventral upstream line, 43c ...
Ventral flow tube, 43d ... Ventral flow tube cross-sectional area of flow tube 43c,
43dmin: minimum point on cross-section of the flow tube on the ventral side, 45: minimum value of negative incident angle, 46: measurement point for measuring paragraph loss, 47: wake of stationary blade, 48: wake of blade, αin: inlet angle, αout ... exit angle, βin ... inflow angle, i ... incidence angle, WLE ... leading edge wedge angle, H_αin ... incidence angle margin.

Claims (3)

[Claims] 1. An airfoil fixed to a rotor and rotated together with the rotor with the inflow of fluid, and having an inlet angle of 45 degrees or less, which indicates an included angle between a direction opposite to the rotation direction and a leading edge direction. In a turbomachine rotor blade having a cross-section, on the airfoil cross-section, an orthogonal coordinate system in which the axial direction of the rotor and the direction in which the fluid inflow direction is positive is the x-axis, and the rotation direction is the y-axis. A point at which the x coordinate of the orthogonal coordinate system is minimum in the profile of the airfoil section near the leading edge is defined as a point A, and the y coordinate of the orthogonal coordinate system is minimum at the profile of the airfoil section near the leading edge. A point B is defined as a point B, and the point A is defined by the outline of the airfoil section near the leading edge and the point y in the rectangular coordinate system.
A point having the same coordinates is defined as a point C, the point B and x in the rectangular coordinate system.
When a point having the same coordinates and located on the line segment AC is point D, the length of the line segment DC is at least twice the length of the line segment DB, and
A rotor blade for a turbomachine, wherein a radius of curvature of an arcuate airfoil section formed near point B is larger than a radius of curvature of an arcuate airfoil section forming a leading edge. . 2. An inlet angle which is fixed to a rotor, rotates together with the rotor with the inflow of fluid, and indicates an included angle between a direction opposite to the rotation direction and a direction of a leading edge thereof; An angle obtained by adding the exit angle indicating the included angle between the direction and the trailing edge direction has an airfoil section having an airfoil cross section of 90 degrees or less, on the airfoil cross section, in the axial direction of the rotor and An orthogonal coordinate system is defined in which the direction in which the fluid inflow direction is positive is the x axis and the rotational direction is the y axis, and the x coordinate of the orthogonal coordinate system is minimized in the profile of the airfoil section near the leading edge. A point is point A, a point at which the y coordinate of the orthogonal coordinate system is minimum in the profile of the airfoil section near the leading edge is point B, and the point A and the rectangular coordinates are the point of the profile of the airfoil section near the leading edge. System y
A point having the same coordinates is defined as a point C, the point B and x in the rectangular coordinate system.
When a point having the same coordinates and located on the line segment AC is point D, the length of the line segment DC is at least twice the length of the line segment DB, and
A turbomachine operating system characterized in that the radius of curvature of the profile of the arcuate airfoil section formed near point B is larger than the radius of curvature of the profile of the arcuate airfoil section forming the leading edge. Wings. 3. A rotor blade for a turbomachine used in an impulse stage of a steam turbine, wherein, on an airfoil cross-section, the axial direction of the rotor of the steam turbine and the direction in which the inflow direction of the fluid is positive are x. Axis and the rotation direction are defined as the y-axis. A point at which the x-coordinate of the rectangular coordinate system is minimized in the profile of the airfoil section near the leading edge is defined as a point A, and the blade near the leading edge is defined as a point. The point at which the y coordinate of the rectangular coordinate system is the minimum in the contour of the mold section is point B, and the point A and y in the rectangular coordinate system are the contour of the airfoil section near the leading edge.
A point having the same coordinates is defined as a point C, the point B and x in the rectangular coordinate system.
When a point having the same coordinates and located on the line segment AC is designated as a point D, the shape of the airfoil cross section at the root is such that the length of the line segment DC is twice or more the length of the line segment DB, and A radius of curvature of a profile of an arcuate airfoil section formed near point B is larger than a radius of curvature of a profile of an arcuate airfoil section forming a leading edge. .