JP5398515B2 - Radial turbine blades - Google Patents

Description

  The present invention is used for an exhaust turbocharger, a small gas turbine, an expansion turbine or the like of an internal combustion engine, and has a working gas flowed in a radial direction from a spiral scroll to a moving blade of a turbine rotor to act on the moving blade. The present invention relates to a rotor blade structure in a radial turbine configured to rotationally drive the turbine rotor by causing the turbine rotor to flow out in the rear axis direction.

  In a relatively small supercharger (exhaust turbocharger) used for an internal combustion engine for automobiles, etc., a turbine rotor blade located inside a scroll from a spiral scroll formed in a turbine casing. A radial turbine configured to rotationally drive the turbine rotor by causing the turbine rotor to flow in a radial direction and then acting on the rotor blades and then flowing out in the axial direction is often employed.

  FIG. 9 shows an example of a turbocharger using such a radial turbine 1. In the figure, a spiral scroll 3 is formed in the turbine casing 2 and a gas outlet passage 5 is formed on the inner peripheral side. In addition, a compressor 9 is provided in the compressor casing 7, and a bearing housing 11 that connects the turbine casing 2 and the compressor casing 7 is formed.

  A plurality of rotor blades 15 are fixed to the outer periphery of the turbine rotor 13 at equal intervals in the circumferential direction. A diffuser 17 is provided at the air outlet of the compressor 9. The turbine rotor 13 and the compressor 9 are connected by a rotor shaft 19. ing. A pair of bearings 21 that are attached to the bearing housing 11 and support the rotor shaft 19 are provided. The turbine rotor 13, the compressor 9 and the rotor shaft 19 rotate around a rotation center 23.

  In the turbocharger equipped with such a radial turbine 1, exhaust gas from an internal combustion engine (not shown) enters the scroll 3 and enters the outer periphery of the plurality of rotor blades 15 while circling along the spiral of the scroll 3. After flowing into the rotor blade 15 from the end face and flowing radially toward the center side of the turbine rotor 13 to perform expansion work on the turbine rotor 13, it flows out in the axial direction and is sent out of the machine from the gas outlet passage 5. It has become so.

The shape of the moving blade 15 is shown in FIGS. 8A shows the meridional shape of the moving blade 15, and FIG. 8B shows the blade cross-sectional shapes of the AA, BB, and CC lines in FIG. Show.
The AA line cross-sectional shape shows the shape near the hub side wall, the CC line cross-sectional shape shows the shape near the shroud side wall, and the BB line cross-sectional shape shows the height between the hub side and the shroud side. The shape of the center part of a direction (Z direction) is shown. In any of the cross-sectional positions of A, B, and C, it has a shape extending radially in the radial direction.

On the other hand, the gas inflow velocity of the gas flowing into the moving blade 15 while circling along the spiral of the turbine scroll 3 located on the upstream side of the moving blade 15 is different in the velocity distribution in the height direction (Z direction) of the moving blade 15. have. That is, as shown in FIG. 7, the gas inflow velocity C is formed in the vicinity of the inlet end face 25 (see FIG. 6) of the moving blade 15 and has a width of 10% to 20% of the height of the inlet end face 25 3. the dimensions boundary layer, is a circumferential component is circumferential velocity C _theta gas velocity C corner clogging shroud side 27 and hub side 29 central large ends of the inlet end face 25 is reduced.
The radial velocity C _R is the radial component, as shown in FIG. 7, the height direction distribution as corners, i.e. the shroud side 27 and hub side 29 of the central portion is smaller at both ends becomes larger of the inlet end face 25 It has become.

Further, the shape of the moving blade 15 is the same as the diameter of the inlet end face 25 over the entire height of the shroud side 27, the central portion, and the hub side 29 as shown in FIG. ₂ = the _{U 1.} Therefore, different gas flows relative angle beta in the height direction of the animal wing 15, when adjusted to optimize the gas inlet relative angle beta ₁ in the center portion shown in FIG. 6 (b), FIG. 6 (c) The gas inlet relative angle β ₂ on the wall side, that is, the hub side 29 and the shroud side 27, becomes larger than the gas inlet relative angle β _{1 at the} center due to flow distortion from the scroll 3. W ₁ and W ₂ are gas inflow relative velocities, and C ₁ and C ₂ are gas inflow absolute velocities.

  For this reason, in such a prior art, gas flows into the back side (negative pressure surface side) of the moving blade 15 with a collision angle (incidence angle) on the hub side 29 and the shroud side 27. In addition to causing a collision loss at the inlet, an increase in the collision angle on the hub side 29 and the shroud side 27 promotes an increase in the secondary flow loss inside the rotor blade 15, and has a problem of causing a decrease in turbine efficiency.

  Therefore, the present applicant configures the inflow relative angle at the moving blade inlet uniformly in the height direction of the moving blade, so that the gas collision loss caused by the variation in the gas inflow relative angle and the 2 There has been a proposal to suppress the next flow loss and increase the turbine efficiency (Patent Document 1).

As shown in FIG. 5, the technique of Patent Document 1 forms a central portion of the inlet end surface 25 into which the working gas flows in a flat shape, and also linearly cuts off corner portions on the shroud side 27 and the hub side 29. By forming the portion 31, the both end radii of the inlet end face 25 are formed to be smaller than the central portion.
Thus, by changing the cut-off amount of the cut-off portion 31, the both ends of the inlet end face of the moving blade 15, that is, the shroud side 27 and the hub side 29 are adjusted to the inner circumference according to the gas flow distribution at the moving blade 15 inlet. The relative inflow angle β of the gas flowing into the moving blade 15 can be adjusted so as to be an optimum angle in the height direction of the moving blade 15.

  Then, the collision angle of the gas at the moving blade inlet is made constant in the height direction of the moving blade, so that the collision loss at the moving blade inlet and the moving blade caused by the nonuniform gas relative inflow angle in the height direction of the moving blade The increase in secondary flow loss in the interior is avoided to prevent a decrease in turbine efficiency.

Japanese Patent No. 3534730

  However, in recent years, a radial turbine having a small diameter relative to the compressor diameter has been selected for the purpose of improving the response in the turbocharger of the radial turbine, and further, the exhaust gas temperature tends to be higher to improve the exhaust gas performance. . For example, in general, a gasoline engine rises from about 900 ° C. to about 1050 ° C., and a diesel engine also rises from about 750 ° C. to about 850 ° C.

For this reason, the value of the speed ratio (U / C ₀ ) of the turbine performance parameter tends to be small, and the efficiency characteristic in which the maximum efficiency point of the turbocharger exists in the speed ratio (U / C ₀ ) region smaller than the conventional one. The speed ratio (U / C ₀ ) is matched at a small operating point.

C ₀ of the speed ratio of the performance parameters (U / C ₀₎ indicates a theoretical gas velocity is represented by C 0 _{= f} (T, [pi) gas temperature T and the turbine pressure ratio [pi function as is It means the theoretical gas velocity obtained when a gas having pressure and temperature is expanded to a certain pressure and temperature. U represents the peripheral speed of the moving blade, and is expressed as a function of the rotational speed and the moving blade diameter D, such as U = f (N, D). Accordingly, U decreases as the blade diameter decreases, C ₀ increases as the exhaust gas temperature increases, and the parameter speed ratio (U / C ₀ ) tends to decrease.
Also, the velocity triangle at the moving blade inlet when the speed ratio (U / C ₀ ) of the performance parameter is small, as shown in FIG. Come to head.

Therefore, when the technique of Patent Document 1 described above is applied to a turbine in which the speed ratio (U / C ₀ ) of such performance parameters is small, that is, the gas inflow relative angle β at the moving blade inlet is set to the value of the gas moving blade. In order to make it constant in the height direction, the central part of the inlet end face of the rotor blade is formed in a flat shape, and when the corners are cut straight on the shroud side and the hub side, the gas relative to the shroud side and the hub side The inflow angle β is more inclined to the negative side, and the collision loss with respect to the moving blade inlet becomes larger, leading to lower turbine efficiency.
Particularly, since the gas flow rate of 50% or more flows into the rotor blade from the vicinity of both the shroud side and hub side walls, the influence of the collision loss in the vicinity of both the wall portions is large.

Therefore, in view of the problems of the prior art, the present invention adjusts the velocity triangle of the inlet of the turbine blade, particularly the blade inlet near the shroud side and hub side walls, or the shape of the inlet of the blade. Is adjusted so that the inlet shape of the moving blade along the flow direction of the gas relative inflow velocity component reduces the collision loss of the inflowing gas generated near both walls of the moving blade inlet and develops inside the moving blade An object of the present invention is to provide a moving blade of a radial turbine capable of suppressing the secondary flow and thereby improving the turbine efficiency of a turbine having a matching point (operating point) having a small speed ratio (U / C ₀ ). To do.

In order to solve the above problems, the present invention is configured to cause a working gas to flow radially from a spiral scroll formed in a turbine casing to a rotor blade of a turbine rotor positioned inside the scroll. In a moving blade of a radial turbine configured to rotationally drive the turbine rotor by causing the turbine rotor to flow out in the axial direction after acting on the blade,
The direction of the blade shape in the vicinity of both the shroud side and hub side walls forming the height direction of the moving blade inlet into which the working gas flows is determined by the inflow velocity (C) of the working gas flowing into the moving blade inlet and the moving blade It is characterized by being formed so as to coincide with the inflow direction of the gas relative inflow velocity component of the speed triangle formed by the circumferential rotational speed (U) and the gas relative inflow speed (W).

According to this invention, the shape of the moving blade inlet in the vicinity of both the shroud side and the hub side walls is defined by the inflow speed (C) of the working gas flowing into the moving blade inlet and the rotational speed (U) in the circumferential direction of the moving blade. In order to have a shape that matches the inflow direction of the gas relative inflow velocity component of the velocity triangle formed by the gas relative inflow velocity (W), the working gas smoothly flows into the rotor blades near both walls, and the motion near both walls The collision loss generated at the blade inlet can be reduced, the gas flow collision loss can be reduced, the secondary flow developed inside the rotor blade can be suppressed, and the turbine efficiency can be improved.
This effectively improves the turbine efficiency in a turbine having such a characteristic that the gas relative inflow angle β (−β _A , −β _C , see FIG. 1B) is inclined to the minus side toward the blade tip. Will be able to.

That is, collisions that occur at the blade inlets near both walls by adjusting the velocity triangles at the blade inlets near both walls so that the inlet shape of the blade matches the inflow direction of the gas relative inflow velocity (W) Loss can be reduced.
Further, since the gas flow rate of 50% or more flows into the rotor blade from the vicinity of both the shroud side and hub side walls, the effect of improving the turbine efficiency can be greatly obtained by improving the flow in the vicinity of the both wall portions. .

Preferably, only the tip portions of the blade inlet near the shroud-side and hub-side walls are inclined according to the flow direction of the gas relative inflow velocity component.
In this way, only the tip portions in the vicinity of both walls are inclined in accordance with the flow direction of the gas relative inflow velocity (W), thereby being indicated by the velocity triangle A and the velocity triangle C shown in FIG. It can be easily adjusted to the flow direction of the gas relative inflow velocity (W).

  Furthermore, when the inclination angle is the same on the shroud side and the hub side, it is possible to make the structure more easily matched to the flow direction of the gas relative inflow velocity (W). Since the inflow velocity (C) of the working gas flowing into the moving blades flows under substantially the same conditions in the wall portions on the shroud side and the hub side, if the diameters on the shroud side and the hub side are set to be the same, the moving blades The circumferential rotational speed (U) is the same on the shroud side and the hub side, and the inclination angle of the tip can be set the same on the shroud side and the hub side.

Preferably, the diameter of the shroud side and the hub side is larger than the central portion of the moving blade inlet in the height direction, and the blade shape of the shroud side and the hub side moving blade inlet is straightened in the radial direction. It is good to do.
Thus, by extending the shroud side and the hub side in the radial direction from the central portion, the rotational speed (U) in the circumferential direction of the rotor blade is increased as shown in FIG. The relative inflow angle β in the flow direction of velocity (W) can be made close to zero and along a moving blade that extends straight in the radial direction. The working gas flows smoothly into the moving blade, reducing the collision loss that occurs at the moving blade inlet. And collision loss of gas flow can be reduced.

  Furthermore, by setting the diameter on the shroud side and the diameter on the hub side to the same diameter, it is possible to easily make a structure that matches the flow direction of the gas relative inflow velocity (W).

According to the present invention, the relative inlet of the gas can be obtained by adjusting the velocity triangle of the inlet of the turbine blade, particularly the blade inlet near the shroud and hub walls, or adjusting the inlet shape of the blade. By adopting the shape of the moving blade inlet along the flow direction of velocity (W), the collision loss of the inflowing gas generated near both walls of the moving blade inlet is reduced, and the secondary flow developed inside the moving blade is suppressed. be able to. Thereby, the turbine efficiency of the turbine having a matching point (operating point) with a small speed ratio (U / C ₀ ) can be improved.

1 shows a first embodiment of the present invention, (a) is a partial cross-sectional view showing a meridional shape of a moving blade, (b) is an AA line, a BB line, a CC line of (a). It is each blade | wing sectional drawing. It is a perspective view of the moving blade part of 1st Embodiment. The second embodiment is shown, (a) is a partial cross-sectional view showing a meridional shape of a moving blade, (b) is a blade of each of the DD, EE, and FF lines of (a). It is sectional drawing. It is a perspective view of the moving blade part of 2nd Embodiment. It is explanatory drawing of a prior art and is a corresponding figure of Fig.1 (a). It is explanatory drawing of a prior art and is a corresponding figure of FIG. It shows the circumferential component C _theta and the radial velocity C _R of the gas inlet velocity. It is explanatory drawing of a prior art and is a corresponding figure of Fig.1 (a), (b). It is a block diagram of the whole supercharger using a radial turbine.

  Hereinafter, the present invention will be described in detail with reference to embodiments shown in the drawings. However, the dimensions, materials, shapes, relative arrangements, and the like of the component parts described in this embodiment are not intended to limit the scope of the present invention to that unless otherwise specified.

(First embodiment)
A first embodiment of the present invention will be described with reference to FIG. 1, FIG. 2, and FIG.
FIG. 9 shows an example of a turbocharger using the moving blades of the radial turbine of the present invention. In the figure, a spiral scroll 3 is formed in the turbine casing 2 and a gas outlet is formed on the inner peripheral side. A passage 5 is formed, a compressor 9 is provided in the compressor casing 7, and a bearing housing 11 that connects the turbine casing 2 and the compressor casing 7 is formed.

  A plurality of rotor blades 50 are fixed to the outer periphery of the turbine rotor 13 at equal intervals in the circumferential direction. A diffuser 17 is provided at the air outlet of the compressor 9. The turbine rotor 13 and the compressor 9 are connected by a rotor shaft 19. ing. A pair of bearings 21 that are attached to the bearing housing 11 and support the rotor shaft 19 are provided. The turbine rotor 13, the compressor 9 and the rotor shaft 19 rotate around a rotation center 23.

In the turbocharger equipped with such a radial turbine 1, exhaust gas from an internal combustion engine (not shown) enters the scroll 3 and enters the scroll 3 along the spirals of the scroll 3, and the inlets on the outer peripheral side of the plurality of rotor blades 50. After flowing into the rotor blade 50 from the end face 52 and flowing radially toward the center side of the turbine rotor 13 to perform expansion work on the turbine rotor 13, it flows out in the axial direction and is sent out from the gas outlet passage 5 to the outside of the machine. It has come to be.
The basic configuration of the turbocharger using the radial turbine 1 described above is the same as that of the conventional technology. The present invention relates to the improvement of the gas inlet portion of the turbine rotor blade 15.

As shown in FIGS. 1A, 1B, and 2, in the first embodiment of the present invention, the shroud side 27 and the hub that form the height direction of the inlet end face (moving blade inlet) 52 of the moving blade 50 are used. In the vicinity of both walls of the side 29, the blade tip portions L ₁ and L ₂ having a height direction (width direction) of about 20% are inclined in the inflow direction of the inflow gas. The central portion in the height direction has a blade cross-sectional shape that extends straight in the radial direction as in the prior art (FIG. 8).

That is, as shown in FIG. 1 (b), the flow direction of the inlet gas cross-sectional shape of the blade 50 in the A-A line cross-sectional position of the hub side 29, i.e., inclined in the W _A in the same direction as the direction of the velocity triangle A hub-side inclined portion 54 having a cross-sectional shape is formed. Similarly, the shroud side inclined section 56 inclined in the W _C in the same direction also velocity triangle cross-sectional shape of the blade 50 in the line C-C cross-sectional position of the shroud side 27 are formed.

As shown in FIG. 1B, the speed triangle of the present invention includes an inflow speed C of the working gas flowing into the moving blade 50, a rotational speed U in the circumferential direction of the moving blade 50, and a gas relative inflow speed W. The relative inflow angle β in the flow direction of the gas relative inflow velocity W is not a characteristic directed to the plus side (see FIG. 6 of β ₁ , β ₂ ) as described in the prior art, but the minus side (− β _A , −β _C (see FIG. 1).

  As described in the technical background section, for the purpose of improving the response in recent years, this corresponds to the reduction of the compressor diameter and the increase of the exhaust gas temperature. This is due to the increase in the inflow velocity C of the working gas that accompanies the conversion, and the fact that the length relationship between the two sides of the velocity triangle changes.

Further, as shown in FIG. 1, the position of the inlet end surface 52 of the rotor blade 50 is set to D ₁ having the same diameter both in the vicinity of both the shroud side 27 and hub side 29 walls and in the central portion. Therefore, the rotational speed U _A and U _C in the circumferential direction of each of the blade 50 on the hub side 29 and shroud side 27 becomes the same, the inflow velocity C of the working gas, both walls near the shroud side 27 and hub side 29 Are substantially the same, the working gas inflow velocities C _A and C _C in the vicinity of both the shroud side 27 and the hub side 29 walls are the same, and the relative inflow angle β is −β _A = −β _C. Thus, the inclination angle of the hub side inclined portion 54 and the shroud side inclined portion 56 is set to this angle.

  In this way, by making the inclination angle the same for both the hub-side inclined portion 54 and the shroud-side inclined portion 56, the moving blade in which the gas relative inflow velocity W flows more easily than the shape of the moving blade with different angles set. 50 can be manufactured.

As an example, it is preferable to set the inclination angle of the hub-side inclined portion 54 and the shroud-side inclined portion 56 by inclining about 20 °, but depending on the moving blade diameter (diameter) D ₁ in the vicinity of both walls of the moving blade 50. Since it is only necessary to set the inclination angles of the hub-side inclined portion 54 and the shroud-side inclined portion 56 so as to match the relative inflow angle β due to the formed speed triangle, the hub-side inclined portion 54 and the shroud-side inclined portion 56 are separately provided. Of course, the angle may be set.

In this manner, by tilting only the tip portions in the vicinity of both walls of the inlet of the moving blade 50 in accordance with the relative inflow angle β of the gas relative inflow velocity W, it easily matches the inflow direction of the gas relative inflow velocity W. Therefore, it is possible to reduce the collision loss to the moving blade 50 generated at the moving blade inlet near the shroud side 27 and the hub side 29 of the moving blade 50.
As a result, the speed ratio (U / C ₀ ) of the performance parameter is set such that the gas relative inflow angle β (−β _A , −β _C , FIG. 1B) is inclined to the minus side toward the blade tip. Turbine efficiency can be improved efficiently in a turbine having a small matching point (operating point).

Furthermore, since a gas flow rate of 50% or more flows into the rotor blade 50 from the vicinity of both walls on the shroud side 27 and the hub side 29, improving the flow in the vicinity of both walls can improve the turbine efficiency. Efficiently large.
Further, if the inlet end surface 52 is inclined at the full height (full width), the effect of improving the response due to the downsizing of the blade diameter is reduced.

(Second Embodiment)
Next, a second embodiment will be described with reference to FIGS. In addition, the same code | symbol is attached | subjected to the same thing as the structural member demonstrated in 1st Embodiment, and description is abbreviate | omitted.
The second embodiment, by adjusting the velocity triangle extending the respective Dotsubasa径D ₂ of the hub side 29 and shroud side 27 of the blade 70 to a diameter larger than Dotsubasa径D ₁ of the central portion, radial It is made to flow in along the shape of the moving blade 70 straightly extended.

As shown in FIGS. 3A and 3B, in the second embodiment, in the height direction in the vicinity of both walls of the shroud side 27 and the hub side 29 that form the height direction of the inlet end surface 72 of the moving blade 70. (width direction) wing tip of about 20% L _1, hub-side protruding the Dotsubasa径D ₂ of L ₂ is set larger than Dotsubasa径D ₁ of the height direction central portion of the tip to the triangular A protrusion 74 and a shroud side protrusion 76 are formed. Here Dotsubasa径D ₁ is a rotor blade diameter as a reference the same as Dotsubasa径D ₁ of the first embodiment.

As shown in FIG. 3B, the cross-sectional shape of the moving blade 70 at the cross-sectional position on the DD line on the hub side 29, the cross-sectional shape at the cross-sectional position on the FF line on the shroud side 27, and the EE line at the center portion The cross-sectional shapes at the cross-sectional positions are all straightly extending in the radial direction. In the second embodiment, since the inlet portion of the rotor blade 70 has a shape extending straight in the radial direction, the relative inflow angle β is substantially zero for both β _D = β _F (W _D , W _F , FIG. 3 (b)).

Then, the gas relative inflow velocity W _D direction of the velocity triangles in D-D line cross-sectional position of the hub side 29, straight so as to be the same direction as the direction of the blade 70, the circumferential speed U of the rotor blade 70 Adjust _D. That is, the inflow velocity C _D of the working gas is determined by the operating conditions of the engine, because it is a value that can not be adjusted by the blade shape, by changing the rotational speed U _D of the circumferential direction of the rotor blade 70, i.e. the blades 70 Adjust by changing the diameter.
Similarly, as the gas relative inflow velocity W _F direction of the velocity triangles in the cross-sectional shape of the blade 70 in the line F-F cross-sectional position of the shroud side 27 becomes the same direction and straight in the direction of the blade 70, adjusting the rotational speed U _F in the circumferential direction of the rotor blade 70.

Thus, in both the near wall of the shroud side 27 and hub side 29 of the inlet end face 72 of the blade 70, the hub-side projection 74 and the shroud-side projection 76 Dotsubasa径D ₂ and greater than D ₁ formed by the gas relative inflow velocity W _D of the velocity _triangle, the direction of the W _F, can be the direction of flow along the rotor blade 70 which is formed straight extending in a radial direction.
Note that by the both Dotsubasa径D ₂ the diameter of the hub-side projection 74 and the shroud-side projection 76, becomes easy manufacture, it may be set to different diameters as a matter of course.

Further, by adjusting the rotor blade inlet velocity triangles by Dotsubasa径D _2, and the gas relative inflow velocity W _D in both walls near the inlet of the moving blade _70, W _F of the relative inflow angle β substantially zero dynamic By adopting a direction that matches the shape of the blade 70 extending straight in the radial direction, the inflow direction of the gas relative inflow speed can be easily matched with the direction of the rotor blade 70, and the shroud side 27 and the hub of the rotor blade 70 can be matched. It is possible to reduce the collision loss to the moving blade 70 that occurs at the moving blade inlets near both walls of the side 29.
Thereby, turbine efficiency can be improved efficiently in a turbine having a matching point (operating point) having a small speed ratio (U / C ₀ ) of performance parameters.

  In addition, since a gas flow rate of 50% or more flows into the rotor blade 50 from the vicinity of both walls on the shroud side 27 and the hub side 29, improving the flow in the vicinity of both wall portions improves the turbine efficiency. Efficiently large.

According to the present invention, the relative inlet of the gas can be obtained by adjusting the velocity triangle of the inlet of the turbine blade, particularly the blade inlet near the shroud and hub walls, or adjusting the inlet shape of the blade. By adopting the shape of the moving blade inlet along the flow direction of velocity (W), the collision loss of the inflowing gas generated near both walls of the moving blade inlet is reduced, and the secondary flow developed inside the moving blade is suppressed. be able to. Thereby, since the turbine efficiency of the turbine having a matching point (operating point) with a small speed ratio (U / C ₀ ) can be improved, it is suitable for application to a moving blade of a radial turbine.

DESCRIPTION OF SYMBOLS 1 Radial turbine 2 Turbine casing 3 Scroll 13 Turbine rotor 15, 50, 70 Rotor blade 27 Shroud side 29 Hub side 52, 72 Inlet end face 54 Hub side inclined part 56 Shroud side inclined part 74 Hub side protruding part 76 Shroud side protruding part

Claims (5)

By causing the working gas to flow radially from the spiral scroll formed in the turbine casing to the rotor blades of the turbine rotor located inside the scroll, and to flow out in the rear axial direction after acting on the rotor blades. In a moving blade of a radial turbine configured to rotationally drive the turbine rotor,
The direction of the blade shape in the vicinity of both the shroud side and hub side walls forming the height direction of the moving blade inlet into which the working gas flows is determined by the inflow velocity (C) of the working gas flowing into the moving blade inlet and the moving blade A moving blade of a radial turbine characterized by being formed so as to coincide with an inflow direction of a gas relative inflow velocity component of a speed triangle formed by a circumferential rotational speed (U) and a gas relative inflow speed (W) .   2. The rotor blade for a radial turbine according to claim 1, wherein only the tip portions of the rotor blade inlet near the shroud side and hub side walls are inclined in accordance with the flow direction of the gas relative inflow velocity component.   The blade of a radial turbine according to claim 2, wherein the inclination angle is the same on the shroud side and the hub side.   The diameter of the shroud side and the hub side is formed larger than the central portion of the blade inlet in the height direction, and the blade shape of the blade inlet on the shroud side and the hub side extends straight in the radial direction. The rotor blade of the radial turbine according to claim 1.   The radial turbine rotor blade according to claim 4, wherein a diameter on the shroud side and a diameter on the hub side are the same.

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