JP5946707B2 - Axial turbine blade - Google Patents

Description

  The present invention relates to an axial flow turbine, and more particularly to a turbine rotor blade used in a gas turbine or a steam turbine of a power plant.

  In recent years, further improvement in power generation efficiency of power plants has been demanded from the viewpoint of reducing environmental load, and further improvement in performance of turbines has become an important issue. The dominant factors of turbine performance include paragraph loss, exhaust loss, mechanical loss, etc. In particular, it is considered that reducing paragraph loss is most effective for improving performance. There are various types of paragraph loss. (1) Airfoil loss caused by the blade shape itself, (2) Secondary flow loss caused by the flow crossing the flow path between blades, (3) Operation There is a leakage loss caused by fluid leaking out of the flow path between blades. Of these, the airfoil loss of (1) is a loss caused by separation between the blade surface and the working fluid, or due to a mismatch between the geometric inlet angle of the blade and the working fluid inflow angle flowing into the blade. By simply considering the outflow angle of the turbine stationary blade provided upstream of the turbine rotor blade, the radius of the inflow angle near the end wall surface due to the effect of the viscosity of the end wall surface or the leakage from the blade tip to the outside The change in direction could not be taken into account, and the airfoil loss could not be reduced effectively.

  With respect to such a problem, Patent Document 1 aims to provide an axial turbine that prevents a mismatch between the blade geometric inlet angle of the turbine blade and the working fluid inflow angle. Techniques have been proposed to reduce the wing geometric inlet angle toward the center. With the above configuration, the blade geometric inlet angle in the vicinity of the end wall can be made to coincide with the direction of the working fluid inflow angle, and the airfoil loss can be reduced.

JP 2007-9761 A

  By the way, the mismatch between the blade geometric inlet angle and the working fluid inflow angle also contributes to the interference between the blade and the end wall boundary layer, which increases the secondary flow loss as well as the airfoil loss. It is also. Therefore, in order to effectively improve the stage performance of the turbine, it is important to simultaneously reduce not only the airfoil loss but also the secondary flow loss.

  However, in the prior art, there is no description about blade warpage (flow turning amount) which is deeply related to the secondary flow loss, and there is a possibility that the paragraph performance cannot be sufficiently obtained.

  SUMMARY OF THE INVENTION An object of the present invention is to provide an axial flow turbine blade that can simultaneously reduce the blade loss and the secondary flow loss and improve the stage performance of the turbine.

In order to achieve the above object, the turbine rotor blade of the axial turbine according to the present invention includes a half straight line extending from the leading edge to the upstream side of the blade camber line of the turbine rotor blade, and a blade back side angle formed by the turbine circumferential direction. Is defined as the inlet angle, the half straight line extending the blade camber line of the turbine blade from the trailing edge to the downstream side, and the blade back side angle formed by the turbine circumferential direction is defined as the outlet angle, and the inlet angle and the outlet When the turning amount of the blade defined by the angle is defined as the warp angle , the inlet angle distribution defined by the inlet angle at each blade height has a maximum value in the blade root region, and in the blade tip region. The rate of change of the inlet angle is larger than the rate of change of the inlet angle in the blade midspan region, and the outlet angle distribution defined by the outlet angle at each blade height has a maximum value in the blade root region. The change rate of the exit angle in the tip region is the wing mid spa Becomes larger than the change rate of the outlet angle in the area, tilt angle distribution defined by the warp angle at each blade height has a maximum value at the wing root region, the rate of change of the warp angle at the blade tip region Is characterized in that the airfoil is formed to be larger than the rate of change of the warp angle in the blade midspan region .

  According to the present invention, considering the radial change of the inflow angle in the vicinity of the end wall portion, matching the blade geometric inlet angle and the direction of the working fluid inflow angle reduces the airfoil loss and at the same time the end wall portion. Secondary flow loss can be reduced by reducing the warp angle of the nearby airfoil and reducing the blade load near the end wall.

1 is a perspective view of a steam turbine rotor blade according to a first embodiment of the present invention. FIG. 2 is an explanatory diagram in which blade section sections obtained by cutting the steam turbine rotor blade shown in FIG. 1 along a section of a blade root section and a section in the vicinity of the blade root section are superimposed. It is a graph showing the inlet angle distribution of the blade height direction of the turbine rotor blade which concerns on 1st Example of this invention. It is explanatory drawing which represents typically radial direction distribution of the speed | velocity | rate of the steam which flows in into a turbine rotor cascade from the turbine rotor cascade upstream. It is explanatory drawing which represents typically the blade element cross section of the stationary blade row in a blade root part, and a speed triangle. It is explanatory drawing which represents typically radial direction distribution of the speed | velocity | rate of the steam which flows in into a turbine rotor cascade from upstream of a turbine rotor cascade, considering the influence of an end wall and leakage. It is explanatory drawing which represents typically the speed triangle which considered the blade element cross section of the stationary blade row | line | column in a blade root | route part, the influence of an end wall, and leakage. It is a graph showing the inlet angle of the blade height direction of the turbine rotor blade which concerns on 1st Example of this invention, and the inflow angle distribution of steam. It is a graph showing the exit angle distribution of the blade height direction of the turbine rotor blade which concerns on 1st Example of this invention. It is a graph showing the curvature angle distribution of the blade height direction of the turbine rotor blade which concerns on 1st Example of this invention. It is explanatory drawing which represents typically the secondary flow which generate | occur | produces in a general turbine rotor cascade. It is explanatory drawing showing the definition of an entrance angle, an exit angle, and a curvature angle. It is a graph showing the efficiency distribution of the blade height direction of the turbine rotor blade which concerns on 1st Example of this invention. It is explanatory drawing which piled up the blade | wing element cross section cut | disconnected by the cross section of the blade root part of the turbine rotor blade which concerns on 2nd Example of this invention, and the cross section of a blade root part vicinity. It is explanatory drawing showing the definition of a front edge shape. It is a graph showing the dimensionless leading edge radius distribution of the blade height direction of the turbine rotor blade which concerns on 2nd Example of this invention. It is explanatory drawing which piled up the blade | wing element cross section cut | disconnected by the cross section of the blade root part of the turbine rotor blade which concerns on 3rd Example of this invention, and the cross section of a blade root part vicinity. It is meridional sectional drawing of the turbine rotor blade which concerns on 3rd Example of this invention. It is a graph showing the axial code length distribution of the blade height direction of the turbine rotor blade which concerns on 3rd Example of this invention.

  Hereinafter, embodiments of the present invention will be described in detail with reference to the drawings as appropriate. Throughout the drawings, the same components are denoted by the same reference numerals.

  Each example described below is an example in which the present invention is applied to a steam turbine rotor blade. The operation principle of the present invention is the same even in the case of gas turbines having different working media, and this patent can be applied to general axial turbines. In each embodiment described below, the blade root side is mainly described, and the blade tip side is partially omitted.

  A first embodiment of the present invention will be described. FIG. 1 is a perspective view of a steam turbine rotor blade according to this embodiment. FIG. 1 shows an excerpt of a part of a blade row structure provided in the turbine circumferential direction. The turbine stage of the steam turbine according to the present embodiment includes a plurality of turbine blades 1 installed in the circumferential direction of a turbine rotor (not shown), and a diaphragm (not shown) on the upstream side in the steam flow direction of the turbine blade 1. 2) turbine stationary blades (not shown) installed between the inner and outer rings in the circumferential direction of the turbine. The turbine rotor blade 1 is attached to a rotor (not shown) via a platform 3 on the inner peripheral side in the turbine radial direction, and a shroud 4 is provided at the tip on the outer peripheral side in the radial direction of the turbine. In addition, a seal structure (not shown) is provided between the shroud 4 provided on the front end side of the turbine rotor blade 1 and a stationary body facing the shroud 4.

  The main steam, which is a working fluid, flows out from a turbine stationary blade provided upstream of the turbine rotor blade 1 in the steam flow direction, and then flows into the turbine rotor blade 1. The steam turbine rotates the turbine rotor together with the turbine rotor blade 1 by causing the main steam flow that flows out from the turbine stationary blade provided on the upstream side of the turbine rotor blade 1 to flow into the turbine rotor blade 1. Electric power is generated by converting rotational energy and electric energy through a generator (not shown) connected to.

  FIG. 2 is a conceptual diagram in which the blade element cross sections obtained by cutting the blade root section 6 of the turbine rotor blade 1 shown in FIG. 1 at the blade root section 6 and the blade section 5 near the blade root section are overlapped. The blade element cross section is a cross section obtained by cutting a wing along a plane having a constant wing height, and is a wing shape that does not include a fillet. The fillet is a member that smoothly connects the blade surface and the end wall surface provided on the outer periphery of the airfoil at the blade root portion and the blade tip portion when the blade is joined to the platform and the shroud. In this embodiment, as in FIG. 1, the entrance angle and the exit angle of the blade root part are smaller than those of the airfoil in the vicinity of the blade root part. It is characteristic that the edge and the rear edge have a shape protruding to the ventral side.

  Details of the configuration and operation of this embodiment will be described.

  FIG. 3 shows a dimensionless blade height obtained by dividing the radial difference y from the blade root portion on the vertical axis to each blade element cross section by the blade length H, where the radial difference from the blade root portion to the blade tip portion is the blade length H. The inlet angle α is plotted against the depth y / H. As shown in FIG. 12, the inlet angle α is an angle formed by a half line 21 that extends the blade camber line 25 of the turbine blade from the front edge to the upstream side and the circumferential direction 20 in the blade element cross section. This is the angle on the side (negative pressure surface 16 side). For convenience of explanation, a region where the dimensionless blade height y / H is 0.0 <y / H ≦ 0.3 is a blade root region, and the dimensionless blade height y / H is 0.3 <y / H <. The region of 0.7 is defined as the blade midspan region, and the region where the dimensionless blade height y / H is 0.7 ≦ y / H <1.0 is defined as the blade tip region. In this embodiment, as shown in FIG. 3, the inlet angle α has a maximum value in the blade root region, and the rate of change of the inlet angle α in the blade tip region is higher than the rate of change in the blade midspan region. It is characteristic that it is getting bigger. Here, the rate of change is a value obtained by differentiating the dimensionless blade height y / H once.

  In the distribution curve of the inlet angle α of the first embodiment, the inlet angle α increases from the blade root portion toward the maximum value in the blade root region, and decreases from the blade root portion side toward the blade tip portion in the blade tip region. In addition, it is distributed in a convex shape between the blade root portion and the blade tip portion.

  FIG. 4 is a schematic diagram of the radial distribution of the velocity of the steam flowing into the turbine rotor cascade from the upstream of the turbine rotor cascade. In FIG. 4, the arrow marked 7 indicates the absolute speed of the steam flowing into the turbine blade, the arrow marked 9 indicates the peripheral speed of the turbine blade, and the arrow marked 8 indicates the peripheral speed of the turbine blade. Represents the relative velocity of the steam flowing into the rotor blade. FIG. 5 is an explanatory diagram schematically showing a blade element cross section of a stationary blade row and a velocity triangle in the blade root portion. FIG. 6 is a schematic diagram of the radial distribution of the velocity of the steam flowing into the turbine rotor cascade from the upstream of the turbine rotor cascade in consideration of the influence of the end wall and leakage. In FIG. 6, the broken line with the reference numeral 10 indicates the absolute velocity that flows into the turbine rotor blade that does not consider the viscous effect, and the arrow with the reference numeral 11 indicates the absolute velocity of the steam that flows into the turbine rotor blade that considers the viscous effect. The arrow with the speed and the sign 12 represents the relative speed of the steam flowing into the turbine rotor blade considering the viscous effect, and the solid line with the sign 13 represents the absolute speed distribution of the steam considering the viscous effect. FIG. 7 is an explanatory diagram schematically showing the blade element cross section of the stationary blade row and the velocity triangle in the blade root portion in consideration of the influence of the end wall and leakage.

  As shown in FIGS. 4 and 5, in the turbine rotor cascade, the outflow angle 32 from the stationary cascade provided upstream of the turbine rotor cascade, and the absolute velocity 7 are assumed to be uniform in the radial direction. The radial distribution of the relative inflow angles 33 flowing into the turbine rotor cascade is different from the inflow angles 33 at each blade height position due to the effect of the rotation of the turbine rotor cascade. Become. Furthermore, in the actual flow, as shown in FIGS. 6 and 7, the flow is affected by the viscosity from the wall surface in the vicinity of the end wall, and a slow speed region (boundary layer) develops. When the working fluid leaks from the blade tip, the flow rate on the blade tip side decreases (the axial velocity is lost), so the inflow angle 33 near the end wall changes greatly. Therefore, when the design is performed under the assumption that the inflow angle 33 and the absolute velocity 7 are uniform in the radial direction, a mismatch between the blade geometric inlet angle and the working fluid inflow angle occurs as shown in FIG. Cause airfoil loss.

  On the other hand, in the case of the turbine rotor blade of the present embodiment, as shown in FIG. 3, the inlet angle in the vicinity of the end wall is set large so as to coincide with the direction of the boundary layer flow. The mismatch between the blade geometric inlet angle and the actual working fluid inflow angle is alleviated, and the airfoil loss can be reduced.

  FIG. 9 is a plot of the blade exit angle β against the dimensionless blade height y / H on the vertical axis. As shown in FIG. 12, the exit angle β is an angle formed by a half line 22 that extends the blade camber line 25 of the turbine blade from the trailing edge to the downstream side and the circumferential direction 20 in the blade element cross section. The angle on the 17th side. In this embodiment, as shown in FIG. 9, the exit angle β has a maximum value in the blade root region, and the change rate of the exit angle β in the blade tip region is higher than the change rate in the blade midspan region. It is characteristic that it is getting bigger.

  More specifically, in the distribution curve of the exit angle β of the first embodiment, the exit angle β increases from the blade root portion toward the maximum value in the blade root region, and the blade tip from the blade root portion side in the blade tip region. It decreases toward the part and is distributed in a convex shape between the blade root part and the blade tip part.

Further, FIG. 10 is a plot of the warp angle γ representing the turning amount of the blade against the dimensionless blade height y / H on the vertical axis. The warp angle γ is expressed by the following equation.
γ = α + β−180 ° (Formula 1)

  Here, α is an inflow angle and β is an outflow angle. In this embodiment, as shown in FIG. 10, the warp angle γ has a maximum value in the blade root region, and the rate of change of the warp angle γ in the blade tip region is higher than the rate of change in the blade midspan region. It is characteristic that it is getting bigger. The warp angle γ increases from the blade root portion to the maximum value, decreases from the maximum value to the blade tip portion, and is distributed in a convex shape between the blade root portion and the blade tip portion.

  FIG. 11 is a schematic diagram of the secondary flow 18 generated in a general turbine rotor cascade. For simplicity, the platform and shroud are not shown. As shown in FIG. 11, the turbine rotor blade has a pressure surface 17 formed on the blade side and a suction surface 16 formed on the blade back side. The main steam flows from the blade leading edge 14 of the turbine blade, passes through the inter-blade passage formed between the pressure surface 17 and the suction surface 16, and then flows out to the blade trailing edge 15 of the turbine blade. To do. When the steam main stream passes through the flow path between the blades, a pressure gradient is created that balances the centrifugal force caused by the bending of the flow path (flow diversion) between the blades. At that time, in the vicinity of the end wall surface, the flow (wall boundary layer) that lost momentum due to the viscosity from the wall surface cannot flow against the pressure gradient generated between the blades and moves to the suction surface side. . As a result, the end wall boundary layer interferes with the suction surface 16 and rolls up as a flow path vortex 19 at each of the blade root portion and blade tip portion of the suction surface 16 into the inter-blade flow path, causing a large loss flow. Form.

  In particular, in a cascade structure composed of low aspect ratio blades, flow path vortices 19 generated on the blade tip side and blade root side interfere with each other to form a lossy flow. The flow path vortex 19 causes a reduction in work efficiency that the turbine rotor blade should originally perform, and causes a secondary flow loss.

  Furthermore, the generation of the secondary flow 18 in the inter-blade flow path becomes a factor that causes a flow that does not follow the airfoil shape at the end wall portion. Since the flow in the vicinity of the end wall portion does not follow the airfoil, a deviation (deviation angle) occurs between the blade geometric outlet angle and the working fluid outflow angle. Therefore, the secondary flow 18 is generated downstream of the turbine blade. This makes it difficult to predict the working fluid inflow angle to the turbine vane of the next stage and increases the airfoil loss of the turbine vane of the next stage.

  In the case of the present invention, as shown in FIG. 3 and FIG. 9, the inlet angle α near the end wall portion is decreased to make the inlet angle α coincide with the flow direction in the end wall boundary layer, and the outlet angle β Is also made small, the warp angle γ in the vicinity of the end wall portion is made small, and the turning of the flow is made small. Therefore, the pressure gradient generated between the blades is reduced, and the secondary flow 18 can be reduced. As a result, the interference between the blade and the end wall boundary layer is reduced, the growth of the flow path vortex 19 can be suppressed, and the secondary flow loss is reduced.

  Further, the small pressure gradient generated between the blades means that the blade load (pressure difference between the pressure surface and the suction surface) is small. This blade load is proportional to the work the turbine is getting from the fluid, but it can also be understood as the amount of resistance to blade flow. In the present invention, since an airfoil with small warpage is adopted in the vicinity of the end wall portion, it can be said that the blade load in the vicinity of the end wall portion, that is, resistance to flow is small, and it is easier to flow than the main flow portion. As a result, the flow distribution of the flow in the turbine rotor blade is biased toward the end wall, so that the development of the boundary layer can be suppressed, and the secondary flow loss is reduced.

  Further, by reducing the secondary flow amount, the deviation angle in the vicinity of the end wall portion can be reduced, which leads to improvement of the blade loss of the turbine stationary blade at the next stage.

  On the other hand, when the warp angle γ is reduced, work can hardly be obtained from the fluid in that region. Therefore, if the range for reducing the warp angle γ is increased, the performance may deteriorate. . However, it is speculated that the amount of performance improvement due to the secondary flow 18 being suppressed and the secondary flow loss being reduced is greater than the amount of performance reduction due to the drop in blade load, and the present invention is applied. Can be said to improve the overall paragraph performance.

  Therefore, according to the present embodiment, the airfoil loss can be reduced and the secondary flow loss can be reduced at the same time, and the efficiency in the vicinity of the end wall surface can be improved as shown in FIG. As a result, turbine stage performance can be improved. In FIG. 13, the solid line represents the present invention, and the broken line represents the conventional efficiency.

  The range in which the inlet angle α is adjusted in the radial direction depends on the thickness of the inflowing boundary layer, that is, the flow state. Further, as described above, if the range in which the warp angle γ is reduced is made too large, the performance may be deteriorated. Considering the above, the dimensionless blade height y / H is in the range of 0.00 <y / H ≦ 0.25, and the dimensionless blade height y / H is 0.75 ≦ y / The range of H <1.00 is the blade tip region, and the inlet angle α and the outlet angle β have maximum values in the blade root region, and the change rates of the inlet angle α and the outlet angle β in the blade tip region are It can be said that it is effective to configure it to be larger than the rate of change.

  Another embodiment of the present invention is shown in FIG. The description will focus on the differences from the first embodiment.

  The present embodiment is an example in which the leading edge shape of the blade section in the vicinity of the end wall portion of the first embodiment is blunt. Here, as shown in FIG. 15, the blunt leading edge is a dimensionless value obtained by dividing the radius 26 of the inscribed circle at the leading edge of the blade element cross section by the maximum radius 27 of the inscribed circle of the blade element cross section. When the leading edge radius Rn is used, it means that the dimensionless leading edge radius Rn is a leading edge of 0.5 or more.

  As described in the first embodiment, the radial distribution of the inflow angle α flowing into the turbine rotor cascade is a distribution that is largely twisted in the radial direction in the vicinity of the end wall, and the blade geometric inlet angle and the working fluid In order to prevent mismatch with the inflow angle, it is desirable to make the inlet angle α near the end wall portion large so that the inlet angle α of the blade matches the direction of the boundary layer flow. However, the appropriate value of the entrance angle α at the end wall portion depends on the speed deficit degree due to the influence of the wall surface, that is, the degree of development of the boundary layer. Therefore, it is difficult to completely match the direction of flow in the end wall boundary layer with the blade inlet angle α.

  In the present embodiment, in addition to the features of the first embodiment, as shown in FIG. 14, the dimensionless leading edge radius in the blade section 24 of the blade root portion is the dimensionless front in the blade section 23 in the vicinity of the blade root portion. It is configured to be larger than the edge radius, that is, the change rate of the dimensionless leading edge radius Rn is negative in the blade root portion, and the change of the dimensionless leading edge radius Rn in the blade height direction is monotonously decreased. It is characteristic that Due to the blunt leading edge, even if there is a deviation (incidence) between the blade geometric inlet angle and the working fluid inflow angle, the flow flowing into the blade leading edge does not separate from the blade along the blade. It begins to flow. That is, the airfoil incident characteristics can be improved.

  Therefore, according to the present embodiment, the incidence characteristic of the airfoil in the vicinity of the end wall can be improved, and the airfoil loss can be further effectively reduced.

  With this configuration, it is possible to improve the incident characteristics in the blade root part. However, the deviation between the blade geometric inlet angle and the working fluid inflow angle occurs not only at the blade root portion but also at the blade tip portion. Therefore, it is desirable to take the same configuration at the blade tip. That is, as shown in FIG. 16, the rate of change of the dimensionless leading edge radius Rn is negative at the blade root portion and positive at the blade tip portion, and is dimensionless in the blade height direction in the vicinity of the blade root portion in the blade root region. It can be said that it is effective that the change in the leading edge radius Rn is monotonously decreased and the change in the dimensionless leading edge radius Rn in the blade height direction is monotonously increased in the vicinity of the blade tip in the blade tip region. In the example of Example 2, the dimensionless leading edge radius distribution is distributed in a concave shape between the blade root portion and the blade tip portion. Further, the vicinity of the blade root portion may be within the range where the dimensionless blade height y / H is within 0.25 within the blade root region. Further, the vicinity of the blade tip may be within the range of the dimensionless blade height y / H of 0.75 or more within the blade tip region.

Making the leading edge shape somewhat blunt will certainly lead to improved incident characteristics, but making the leading edge shape blunt increases the contact area between the working fluid and the blade surface. Therefore, the blade surface friction loss is also increased. Therefore, in order to increase the reduction amount of the airfoil loss more than the increase amount of the blade surface friction loss, it is necessary to increase the leading edge shape to some extent.
Considering the above, it can be said that it is effective to configure the dimensionless leading edge radius Rn at the blade root portion and the blade tip portion to satisfy 0.6 ≦ Rn.

  Another embodiment of the present invention is shown in FIG. The description will focus on the differences from the first embodiment.

  In the present embodiment, the axial cord length of the blade section in the vicinity of the end wall portion of the first embodiment is increased. The axial cord length is the axial length of the wing. As described in the first embodiment, the secondary flow 18 generated between the turbine rotor cascades induces the generation of the flow path vortex 19 and causes the secondary flow loss. Therefore, it is desirable to reduce the secondary flow amount by reducing the pressure difference between the blades, that is, by reducing the blade load, by taking measures such as reducing the warpage in the blade shape near the end wall.

  In the present embodiment, in addition to the features of the first embodiment, as shown in FIG. 17, the axial cord length Cx in the blade section 29 of the blade root portion is equal to the axial cord length C in the blade section 28 near the blade root portion. It is characteristic that the rate of change in the axial cord length is negative in the blade root part and the change in the axial cord length in the blade height direction in the blade root region is monotonously decreased. It is. Increasing the shaft cord length is equivalent to lowering the load per unit length acting on the blade cross section, so increasing the shaft cord length suppresses the secondary flow and reduces the secondary flow loss. can do. Furthermore, increasing the axial cord length of the blade root part can increase the blade cross-sectional area at the blade root part and reduce the local centrifugal stress caused by the rotation of the turbine rotor blade. The problem is also improved.

  By the said structure, the problem of the intensity | strength by the loss by the secondary flow in a blade | wing root part and the local centrifugal stress can be improved. However, the problem of strength due to loss due to secondary flow and local centrifugal stress occurs not only at the blade root but also at the blade tip. Therefore, it is desirable to take the same configuration at the blade tip. That is, as shown in FIG. 19, the change rate of the axial cord length is negative at the blade root portion and positive at the blade tip portion, and the change in the axial cord length in the blade height direction in the vicinity of the blade root portion in the blade root region. It can be said that it is effective to configure so that the change in the axial cord length in the blade height direction in the vicinity of the blade tip in the blade tip region is monotonously increased. In the example of the third embodiment, the axial code length distribution is distributed in a concave shape from the blade root portion side toward the blade tip portion side.

  On the other hand, when the shaft cord length is increased, the contact area between the working fluid and the blade surface is increased in that region, so that the blade surface friction loss increases and the performance may deteriorate. However, it is speculated that the amount of performance improvement due to the secondary flow 18 being suppressed and the secondary flow loss being reduced is greater than the amount of performance reduction due to the drop in blade load, and the present invention is applied. Can be said to improve the overall paragraph performance.

  Therefore, according to the present embodiment, the blade load of the airfoil near the end wall can be further reduced, and the secondary flow loss can be further effectively reduced by reducing the secondary flow amount.

  In this embodiment, as shown in FIG. 18, a typical example in which an airfoil having a long axial cord length in the vicinity of the end wall portion is configured such that the trailing edge position protrudes downstream from the conventional trailing edge position 31. However, it is not always necessary to project the trailing edge position to the downstream side, and the attachment position in the flow direction of the airfoil having a long axial cord length in the vicinity of the end wall portion is not a problem.

  Further, as described above, by increasing the shaft cord length, the blade surface friction loss increases, and the performance may deteriorate. Considering the above, if the value obtained by dividing the maximum value of the axial code length within an arbitrary range of the dimensionless blade height y / H by the minimum value of the axial code length is the local axial code length ratio Θ, The dimensionless blade height y / H is within the range of 0.00 ≦ y / H ≦ 0.25, and the local axis code length ratio Θ is within the range of 1.0 ≦ Θ ≦ 1.5, and the dimensionless blade height. It is effective to configure so that the local axis code length ratio θ within the range of 0.75 ≦ y / H ≦ 1.00 is within the range of 1.0 ≦ Θ ≦ 1.5. I can say that.

DESCRIPTION OF SYMBOLS 1 Turbine rotor blade 2 Turbine stationary blade 3 Platform 4 Shroud 5, 23, 28 Blade element cross section 6, 24, 29 Blade element cross section near blade root part 14 Blade leading edge 15 Blade trailing edge 16 Suction surface 17 Positive Pressure surface 18 Secondary flow 19 Channel vortex 20 Circumferential direction 21, 22 Semi straight line 25 Blade camber line 26 Inscribed circle radius 27 at leading edge Inscribed circle maximum radius 32 Outflow angle 33 Inflow angle

Claims (10)

A turbine blade of an axial turbine,
A half straight line extending the blade camber line of the turbine blade from the leading edge to the upstream side and a blade back side angle formed by the turbine circumferential direction are defined as an inlet angle, and the blade camber line of the turbine blade from the trailing edge When defining a half line extending downstream and a blade back side angle formed by the turbine circumferential direction as an exit angle, and defining a turning amount of the blade defined by the inlet angle and the outlet angle as a warp angle ,
The inlet angle distribution defined by the inlet angle at each blade height has a maximum value in the blade root region, and the rate of change of the inlet angle in the blade tip region is the ratio of the inlet angle in the blade midspan region. Greater than the rate of change,
The outlet angle distribution defined by the outlet angle at each blade height has a maximum value in the blade root region, and the rate of change of the outlet angle in the blade tip region is the ratio of the outlet angle in the blade midspan region. Ri name greater than the rate of change,
The warp angle distribution defined by the warp angle at each blade height has a maximum value in the blade root region, and the rate of change of the warp angle in the blade tip region is the warp angle in the blade midspan region. A turbine rotor blade characterized in that an airfoil is formed to be larger than a rate of change . The turbine rotor blade according to claim 1,
In the blade root region, the inlet angle distribution increases from the blade root portion to the maximum value, and decreases in the blade tip region from the blade root portion side toward the blade tip portion. Distributed in a convex shape between the tip ,
The exit angle distribution increases from the blade root portion to the maximum value in the blade root region, and decreases from the blade root portion side toward the blade tip portion in the blade tip region. While distributed in a convex shape between the tip ,
In the blade root region, the warp angle distribution increases from the blade root portion to the maximum value, decreases from the maximum value to the blade tip portion, and is distributed in a convex shape between the blade root portion and the blade tip portion. A turbine rotor blade characterized by forming an airfoil as described above. The turbine rotor blade according to claim 1 or 2,
A turbine rotor blade characterized in that a leading edge shape of a blade element cross section of at least one of a blade root part and a blade tip part is a blunt type. The turbine rotor blade according to claim 3, wherein
When the inscribed circle radius at the leading edge of the blade section is divided by the maximum inscribed circle radius of the blade section, the dimensionless leading edge radius is defined.
A turbine rotor blade having a dimensionless leading edge radius of the blade root portion or blade tip portion having the blunt type leading edge shape is 0.6 or more. The turbine rotor blade according to any one of claims 1 to 4, wherein
When the inscribed circle radius at the leading edge of the blade section is divided by the maximum radius of the inscribed circle of the blade section, the dimensionless leading edge radius is defined.
The dimensionless leading edge radius distribution defined by the dimensionless leading edge radius at each blade height is:
The rate of change of the dimensionless leading edge radius in the blade root is negative, and decreases monotonously from the blade root to the blade tip in the vicinity of the blade root in the blade root region.
The airfoil is formed so that the rate of change of the dimensionless leading edge radius at the blade tip is positive and increases monotonously from the blade root side toward the blade tip near the blade tip in the blade tip region. Turbine blades. The turbine rotor blade according to claim 5,
The turbine blade according to claim 1, wherein the dimensionless leading edge radius distribution is formed in an airfoil shape so as to be distributed in a concave shape between a blade root portion and a blade tip portion. The turbine rotor blade according to claim 1 or 2,
The shaft code length distribution defined by the shaft code length at each blade height is
An airfoil shape that monotonously decreases from the blade root part toward the blade tip side near the blade root part in the blade root area, and monotonously increases from the blade root part side toward the blade tip part near the blade tip part in the blade tip area. A turbine rotor blade characterized by forming The turbine rotor blade according to claim 1 or 2,
The shaft code length distribution defined by the shaft code length at each blade height is
The rate of change of the axial cord length in the blade root part is negative, and monotonously decreases from the blade root part toward the blade tip side in the vicinity of the blade root part in the blade root region,
The airfoil is formed such that the rate of change of the axial cord length at the blade tip is positive and increases monotonously from the blade root side toward the blade tip near the blade tip in the blade tip region. Turbine blade. The turbine rotor blade according to claim 8, wherein
The turbine rotor blade according to claim 1, wherein the axial code length distribution is formed in an airfoil shape so as to be distributed in a concave shape between the blade root portion side and the blade tip portion side. The turbine rotor blade according to any one of claims 7 to 9,
The length y from the blade root part to each blade element section divided by the blade length H is defined as the dimensionless blade height y / H,
When a value obtained by dividing the maximum value of the axial code length within an arbitrary range of the dimensionless blade height y / H by the minimum value of the axial code length is defined as a local axial code length ratio Θ,
The dimensionless blade height y / H is in the range of 0.00 ≦ y / H ≦ 0.25, and the local axis code length ratio Θ is in the range of 1.0 ≦ Θ ≦ 1.5,
Dimensionless blade height y / H is in the range of 0.75 ≦ y / H ≦ 1.00, and the local axis code length ratio Θ is in the range of 1.0 ≦ Θ ≦ 1.5. Turbine blade.