• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The turbine nozzle is arranged circumferentially in an annular passage 4 formed between the inner and outer rings of the diaphragm and comprises a series of nozzle vanes 1 fixed to the inner and outer rings of the turbine diaphragm. The flow passage is formed between the pressure surface (F) and the suction surface (B) of the adjacent blades of the nozzle blades, the cross section of the flow passage is wing height (h) from the inner diameter and outer diameter surface (hub and tip end wall) ) And a range that extends into a curved line along with another range formed by a nearly straight line.
    公开号:KR20010052802A
    申请号:KR1020007014115
    申请日:1999-06-10
    公开日:2001-06-25
    发明作者:와타나베히로요시；하라다히데오미
    申请人:마에다 시게루；가부시키 가이샤 에바라 세이사꾸쇼；
    IPC主号:
    专利说明:

    Turbine Nozzle Vane {TURBINE NOZZLE VANE}
    In recent years, it has been recognized that improving turbine performance is important in order to improve energy consumption in the utility or machine operation of power generation in power plants.
    In order to improve the performance of the turbine, it is necessary to reduce the internal losses of each turbine stage. Internal losses at each of these turbine stages include blade contour losses, secondary flow losses, leak losses, and the like.
    Turbine stages with small aspect ratios (wing height / wing string) and small wing heights have a high rate of secondary flow losses. Therefore, it is possible to effectively reduce the secondary flow loss to improve the performance of the turbine.
    The secondary flow generating mechanism is described below.
    As shown in FIG. 15, the flow G flowing between the nozzle vanes 1 is generated by a pressure gradient from the pressure plane F to the suction plane B of each nozzle vane 1. You will receive strength. In the main flow away from the turbine end wall, the force due to the pressure gradient and the centrifugal force due to the rotation of the flow are in equilibrium. However, the flow in the boundary layer near the turbine end wall has a small kinetic energy and flows as the arrow J from the pressure face F to the suction face B by the force of the pressure gradient. In the latter part of the flow passage, the flow impinges on the suction surface B and rotates to form the flow passage vortex W. This flow passage vortex W generates a non-uniform energy distribution downstream of the nozzle blades by accumulating low energy fluid in the end wall boundary layer. Even if the non-uniform energy distribution becomes uniform downstream of the nozzle blade, a large energy loss occurs during the homogenization. In Fig. 15, E denotes a radial line and L denotes a hub end wall.
    Hereinafter, various examples which can suppress the secondary flow will be described.
    For example, as shown in FIG. 16, the blade 1 is inclined at an angle Θ with respect to the radial line E, thereby reducing the wing-to-wing pressure gradient near the hub end wall of the blade. In Fig. 16, reference numeral 2 denotes an outer ring, and reference numeral 3 denotes an inner ring. Further, as shown in Figs. 17 and 18, the nozzle blades 1 are bent at their opposite ends, so that the pressure surface F is directed toward the end wall. In FIG. 17, U represents an outer diameter surface. In FIG. 18, Θt represents the angle between the tangent line to the blade stacking line 1 at the tip end wall and the radius line E, and Θr is the tangent line and radius to the wing stacking line 1 at the hub end wall. The angle between the line E is shown, and h represents the wing height. According to the prior art, while using the same wing geometry, the secondary stacking line is tilted or bent in a direction to reduce the wing-to-wing pressure gradient near the end wall, thereby controlling the secondary flow to reduce losses.
    Another prior art controlled secondary flow by including a curved or inclined surface provided across the entire height of the nozzle blade, as disclosed in Japanese Patent Laid-Open No. 10-77801.
    In the conventional configuration, in order to control the pressure gradient, the nozzle blades need to be greatly inclined or bent, and many efforts to satisfy the above requirements cause problems in the mechanical strength or manufacturing process of the nozzle blades.
    In addition, in the bent or inclined blade, the flow distribution at the exit of the blade may be significantly different from the flow distribution on the blade is not bent or inclined.
    For example, FIG. 19 shows the abscissa indicating the position along the wing height, expressed as a dimensionless ratio with respect to the height h, and the absolute velocity (V = (Vt 2 + Vm 2 ) 0.5 ). A graph having a vertical axis representing the circumferential velocity Vt and the meridional velocity Vm, expressed as a specific dimensional ratio, is shown. This graph (FIG. 19) shows that the flow velocity distribution (indicated by the solid line) of a typical wing and the flow velocity distribution (indicated by a dashed line) of a bent wing are different at opposite ends of the wing.
    If the nozzle vanes are curved and combined with conventional rotary vanes located downstream of the nozzle vanes, the flow from the nozzle vanes will not match the rotary vanes, so the curved nozzle vanes may not be effective. In this case, a new rotary vane is required that can match the flow from the exit of the curved nozzle vane, so this configuration cannot satisfy a wide range of applications.
    The present invention relates to a turbine nozzle, and more particularly, to a turbine nozzle having a series of nozzle vanes disposed circumferentially in an annular passage formed between an inner ring and an outer ring of a diaphragm and fixed to the inner and outer rings of the diaphragm. will be.
    1 is a perspective view of a turbine nozzle according to an embodiment of the present invention.
    2 is a flow passage cross-sectional view of the turbine nozzle of FIG.
    Figure 3 is a diagram showing the meridional distribution of the distance (Sh, St) of the nozzle blades according to the present invention.
    4A to 4D are diagrams showing changes in flow path cross section in the meridion direction with respect to the nozzle blades of a conventional turbine nozzle.
    5A to 5D are diagrams showing changes in flow path cross section in the meridion direction with respect to the nozzle blades of a turbine nozzle according to an embodiment of the invention.
    6 is a graph showing the relationship between the distances (Sh, St) and losses at x / Cx = 0.3.
    Fig. 7 is a graph showing the relationship between height Lh and Lt and loss.
    8 is a diagram illustrating meridional distribution of distances Sh and St of nozzle vanes according to an embodiment of the present invention.
    9 is a graph showing the relationship between the distance (Sh, St) and the loss at the leading edge.
    10 is a graph showing a comparison of the loss distribution at the exit of the wing according to the present invention and the conventional wing.
    FIG. 11 is a graph showing the static pressure distribution on the wing surface in the midspan of the wing. FIG.
    12 is a graph showing the static pressure distribution on the blade surface of the hub end wall of the turbine diaphragm.
    13 is a graph showing the velocity distribution at the wing exit.
    14A and 14B are diagrams showing a comparison of contour distributions for static pressure in the cross section of a flow path on a nozzle blade according to the present invention and a conventional nozzle blade, respectively.
    15 is a partial perspective view of flow in a conventional turbine nozzle.
    16 is a partial front view of a conventional nozzle with inclined vanes to reduce secondary flow loss.
    17 is a partial perspective view of a conventional nozzle with bent vanes to reduce secondary flow loss.
    FIG. 18 is a partial front view of the nozzle of FIG. 17; FIG.
    19 is a graph illustrating a comparison of flow velocity distributions of a conventional wing and a curved wing.
    It is therefore an object of the present invention to provide a turbine nozzle which can reduce the secondary flow loss while producing the same outlet flow as in a conventional blade and does not adversely affect the rotary vane downstream of the turbine nozzle.
    The turbine nozzle of one embodiment of the present invention is arranged in a circumferential direction in an annular passage 4 formed between an inner ring and an outer ring of the diaphragm, and a series of nozzle vanes 1 fixed to the inner and outer rings of the diaphragm. Wow; A flow passage formed between the pressure surface (F) and the suction surface (B) of a neighboring wing of the nozzle blades, the cross section of the flow passage extending from the inner and outer diameter surfaces (hub and tip end walls) along the wing height; And certain ranges formed by curved lines and other ranges formed by almost straight lines.
    Since the cross section of the flow passage within a predetermined range of the pressure surface and the suction surface includes a curved portion and a portion formed almost linearly, the turbine nozzle according to the present invention has a nozzle blade disclosed in Japanese Patent Laid-Open No. 10-77801. Is completely different from its structure.
    Turbine nozzles according to another aspect of the present invention are arranged in a circumferential direction in an annular passage 4 formed between an inner ring and an outer ring of a diaphragm, and a series of nozzle vanes 1 fixed to the inner and outer rings of the diaphragm 1 )Wow; Hub end of the turbine diaphragm facing the tip end wall of the turbine diaphragm in a predetermined range in the meridion direction of the nozzle wing and in the predetermined range between the tip end wall and the midspan of the wing. A pressure surface F of each nozzle blade facing the subwall; The tip end wall of the diaphragm facing the hub end wall of the turbine diaphragm in a predetermined range in the meridion direction of the nozzle wing and in the predetermined range between the tip end wall and the midspan of the wing and within the predetermined range between the hub end wall and the midspan of the wing. And a suction face (B) of each nozzle vane facing.
    Here, the predetermined range may include a range corresponding to at least 30% of the meridion width Cx of the nozzle blade from the leading edge 1f of the nozzle blade to the meridion direction x. The predetermined range is 20 to 40% of the blade height h from the hub end wall L of the nozzle blade 1 and the blade height h from the tip end wall U of the nozzle blade 1. It may also include a range corresponding to 20 to 40% of the).
    In the predetermined range, the pressure surface F of the nozzle blade 1 is disposed toward the tip end wall on the tip end wall side, that is, is bent to face the tip end wall, and is disposed toward the hub end wall on the hub end wall side. That is, it is bent to face the hub end wall, the suction surface (B) of the nozzle blade 1 is disposed toward the hub end wall of the tip end wall side, that is, bent to face the hub end wall, It is arranged toward the tip end wall, that is, it is bent to face the tip end wall.
    The line 1p on the pressure surface along the height of the nozzle blade 1 and the line 1s on the suction surface correspond to 20 to 40% from the hub end wall L along the height h of the nozzle blade 1. A central portion S formed in a substantially straight line is preferably except for a range C2 corresponding to 20 to 40% from the tip end wall U along the range C1 and the height h of the nozzle vane 1. Has In particular, the line and suction surface on the pressure surface F in the cross section of the flow path at any meridional position in the range of at least 30% from the leading edge 1f of the nozzle blade along the meridional width Cx of the nozzle vane. The line on (B) is the tip end along the height h of the nozzle blade 1 and the range C1 corresponding to 20-40% from the hub end wall L along the height h of the nozzle blade 1. Except for the range C2 corresponding to 20 to 40% from the subwall U, it preferably has a central portion S formed in a substantially straight line.
    The cross section of the flow passage is a line on the pressure surface F at the meridional position in the range of at least 30% from the leading edge 1f of the nozzle vane 1 along the meridional width Cx of the nozzle vane 1. And a line on the suction surface B, each of which comprises an almost straight line at the center of the nozzle blade.
    Extension line SE1 and hub end wall L of the pressure surface or central portion S on the suction surface formed almost linearly from the intersection point Pt1 between the line C1 on the pressure surface or the suction surface and the hub end wall L. The extension line SE2 and the tip end wall U of the central portion S from the distance Sh to the intersection point Pc1 therebetween and the intersection point Pt2 between the line C2 and the tip end wall U on the pressure or suction surface. The distance St to the intersection point Pc2 between) has a maximum value at the leading edge 1f of the nozzle blade, and at least of the blade height h at a position that is 30% of the meridian width from the leading edge of the nozzle blade. Has 4%.
    The maximum value of the distances Sh and St at the leading edge 1f of the nozzle blade 1 is preferably in the range of 5 to 15% of the blade height h.
    Sh or St is the distance between the intersections from the leading edge 1f of the nozzle blades to the point of 55 to 65% of the meridional width, the nozzle height h, and the blade width Cx from the leading edge 1f of the nozzle blades. If the meridian distance ratio to is expressed as ∧, the following equation is established.
    St / h, Sh / h = ΣAn and ∧ n
    Where An is a coefficient and n is an integer of 0 or more.
    In the above formula, the higher order term near zero is ignored. That is, n is an integer greater than or equal to zero, containing all higher-order terms that are not small enough to be ignored.
    These and other objects, features, and advantages of the present invention will become apparent from the following description taken in conjunction with the accompanying drawings, which illustrate preferred embodiments of the present invention.
    A turbine nozzle according to an embodiment of the present invention will be described with reference to the drawings.
    As shown in FIG. 1, a turbine nozzle according to the present invention comprises a series of nozzle vanes arranged in a circumferential direction y in an annular passage 4 formed between an inner ring 3 and an outer ring 2 of a diaphragm. It includes (1). This nozzle vane 1 has a hub and tip end walls L, U on opposing ends fixed to the outer diameter surface (tip end wall) of the inner ring 3 and the inner diameter surface (hub end wall) of the outer ring 2, respectively. ) The turbine nozzle of FIG. 1 is a perspective view seen from an upstream point. Each nozzle vane 1 has a wing profile or airfoil, and has a pressure face F and a suction face B.
    The flow passage formed between the pressure surface F and the suction surface B of the blade blades adjacent to each other among the nozzle blades has a cross section 4a at an arbitrary meridional position. This end surface 4a has the side edge formed by the line 1p on the pressure surface F, and the opposite side edge formed by the line 1s on the suction wing surface B. As shown in FIG. Each nozzle vane 1 has a width Cx in its meridion direction x. In Figure 1 z denotes the radial direction.
    On each nozzle blade 1, the line 1p on the pressure surface F and the line 1s on the suction surface B forming the end surface 4a are in the meridian direction x from the leading edge 1f. A range up to a point at least 30% of the width Cx and a range Lh, Lt corresponding to 20 to 40% of the wing height h inwardly from the hub and tip end walls L, U (see FIG. 2). (Ie, from the hub end wall L to the tip end wall U, and from the tip end wall U to the hub end wall L), the hub end wall L and the tip end wall U Are composed of straight or curved lines C1 and C2, respectively. The other portions of the lines 1p and 1s other than the ranges Lh and Lt, that is, the center portions of the lines 1p and 1s are constituted by the straight lines S.
    Thus, as shown in FIG. 2, in the flow passage 4a between the pressure surface F and the suction surface B of the adjacent nozzle blade 1, inward from the hub and tip end walls L and U. The range Lh, Lt corresponding to 20 to 40% of the blade height h is a straight or curved line C inclined from the pressure surface F to the suction surface B in the ends L and U directions. (C1, C2: parabola in the illustrated embodiment).
    Displacement from straight portion S on hub and tip end walls L, U, i.e. extension line SE1 of straight portion S at the intersection point Pt1 between the inclined line C1 and hub end wall L; A straight line S at the distance Sh to the intersection point Pc1 between the hub end wall L and the hub end wall L, and the intersection point Pt2 between the inclined line C2 and the tip end wall U in FIG. The distance St between the extension line SE2 (indicated by a dotted line in FIG. 2) and the outer diameter surface U of) is the maximum at the leading edge 1f of the nozzle blade, It gradually decreases toward the trailing edge of.
    The effect of the meridional range with the inclined portions C1 and C2 added is described below.
    In FIG. 3, various examples according to changes in the distances St and Sh in the meridian direction x are shown as characteristic curves (a) to (f). In FIG. 3, the horizontal axis is represented by x / Cx, and the vertical axis is represented by Sh / h and St / h. Here, x / Cx is defined as the meridional distance from the non-dimensionalized leading edge by the wing meridional width Cx. In the examples represented by these characteristic curves (a) to (f), the ratio of the distance Sh (= St) to the blade height h at the leading edge 1f is shown by the characteristic curve (a). Except that Sh / h = 0.09. The ratio of the ranges Lh, Lt to the wing height h is chosen such that Lh / h = Lt / h = 0.25.
    In the characteristic curve (a), since the distances Sh and St are Sh = 0 and St = 0 in the whole nozzle blade, this represents the conventional nozzle blade contour.
    4A to 4D show the change of the cross section of the flow path in the meridion direction with respect to the conventional nozzle blade (characteristic curve (a)). 5A to 5D show the change of the flow path cross section in the meridion direction with respect to the nozzle blade (characteristic curve (e)) according to the present invention.
    For comparison, FIG. 6 shows the total pressure loss of the nozzle blades represented by the characteristic curves (a) to (f) for the distance Sh at the meridional distance x / Cx = 0.3 in the viscous flow analysis. It is shown by calculation.
    Referring to Figure 6, as the distance Sh at x / Cx = 0.3 increases, the loss decreases to Sh / h = 0.046, and at Sh / h> 0.046 the characteristic curves (d), (e), (f) ) Remains almost unchanged.
    Considering simplicity or ease in manufacturing, as shown in FIG. 3, nozzle vanes (characteristic curves (d) and (e)) having a distance Sh that decreases to almost zero at x / Cx = 0.6 The photographic portions C1 and C2 span the entire meridional width, and the distance Sh is preferred for a constant nozzle blade (characteristic curve f) over the entire longitudinal width. Because the flow passage is a simpler form.
    The effects on the ranges Lh and Lt in the wing height to which the inclined portions C1 and C2 are added are described below.
    FIG. 7 shows the nozzle blades whose distance (Sh, St) distribution decreases to almost zero at x / Cx = 0.6, and Sh / h at the leading edge is 0.09 (characteristic curves (b), (c), The effect of the loss on the range Lh and Lt in the wing height to which the inclined portions C1 and C2 are added to (d) and (e)) is shown.
    According to FIG. 7, the nozzle blades according to the invention suffer less losses than conventional nozzle blades, regardless of the size of the ranges Lh and Lt, in particular the losses are minimal in the range 0.2 < Lh / h and Lt / h < 0.4. Becomes
    The effect on the distance (Sh, St) at the leading edge of the nozzle blade is described below.
    FIG. 8 shows nozzle blades with different distances Sh and St at the leading edges as characteristic curves (a) to (e), and FIG. 9 shows the total pressure loss of the nozzle blades, calculated by viscous flow analysis. The horizontal axis in FIG. 9 indicates Sh / h (= St / h) at the inlet of the nozzle blade.
    Referring to Fig. 8, the distribution of the distance Sh, St in the meridion direction of each nozzle blade represented by the characteristic curves (b) to (e) decreases to almost zero at x / Cx = 0.6.
    9, it can be seen that the nozzle blades represented by the characteristic curves (b) to (e) in which Sh / h reaches about 0.16 at the leading edge of the nozzle blade suffer less loss than the conventional nozzle blade. This nozzle blade (characteristic curves (b) to (d)) is preferable because the loss is particularly minimum in the range of 0.05 < Sh / h < 0.15.
    10-13 show detailed results analytically calculated for the nozzle blades according to the invention and conventional nozzle nozzles in the prior art.
    Figure 10 shows the comparison of the loss distribution calculated by the viscous flow analysis in the cross section of the nozzle blade according to the present invention and the blade exit of the conventional nozzle blade, where Sh / h = 0.09, St / h = at the leading edge 0.106, Lh / h = Lt / h = 0.25, and the distribution for the distances Sh and St in the meridion direction decreases to almost zero at x / Cx = 0.6. In Fig. 10, the horizontal axis represents z / h and the vertical axis represents total pressure loss.
    Referring to FIG. 10, in a typical nozzle blade (indicated by a solid line), loss peaks due to secondary flow appear near the hub and tip end walls, and non-uniform flow occurs when they are mixed and diffused downstream of the blade. It can be seen that in the nozzle blades (indicated by dashed lines) according to the present invention causing a large loss, the loss peak generated by the secondary flow near the hub end wall is about 30% lower than that of a conventional nozzle blade.
    Fig. 11 shows the static pressure distribution on the blade surface in the midspan of the blade, and Fig. 12 shows the static pressure distribution on the blade surface on the hub end wall of the turbine diaphragm. 11 and 12, the horizontal axis represents x / Cx and the vertical axis represents P / PsO (non-dimensionalized surface pressure due to static pressure at the nozzle inlet). 11 and 12, the static pressure on the wing (indicated by the dashed line) and the normal wing (indicated by the solid line) according to the present invention is the same as in the midspan of the wing, but on the hub end wall of the wing according to the present invention. It can be seen that the rod (pressure difference between the pressure side and the suction side) is smaller than the wing inlet side.
    This change in the loading distribution of the wing, ie the fact that the wing rod according to the present invention is smaller at the wing inlet side than the conventional wing rod, is based on the change in the static pressure distribution in the flow passage cross section 4a in the nozzle. This will be described below.
    14A and 14B show the static pressure contour lines of the flow passage cross section 4a in the conventional nozzle blade and the nozzle blade according to the present invention. In the conventional nozzle blade, the static pressure contour is distributed substantially parallel to the line 1p on the pressure surface F and the line 1s on the suction surface B. FIG. In the vicinity of the line 1s on the suction surface B, the static pressure on the hub and tip end walls L and U is substantially the same as the static pressure at the center of the blade height.
    In the nozzle vane according to the present invention, the static pressure distribution across the vane height near the line 1s on the suction surface B is the center of the vane height near the hub end wall L and the tip end wall U (Fig. 2). Is larger by Sh and St than in the straight portion (S) range of. Therefore, since the static pressure near the line 1s on the suction surface B increases in the vicinity of the hub end wall L and the tip end wall U, the vane rod decreases.
    14A and 14B, dotted arrows SF1 and SF2 are located near both end walls from the line 1p on the pressure surface F to the line 1s on the suction surface B in the cross section 4a of the flow passage. Represents the secondary flow of.
    These secondary flows SF1 and SF2 are generated by the pressure difference (wing rod) between the pressure surface F and the suction surface B near the hub end wall L and the tip end wall U, and the secondary flow The strength of (SF1, SF2) is proportional to the size of the vane rod. Therefore, the nozzle blade according to the present invention can make a blade rod smaller than the conventional nozzle blade in the vicinity of the hub end wall L and the tip end wall U, and the secondary flow is further suppressed on the conventional nozzle blade. Therefore, the loss due to the secondary flow can be reduced.
    In addition, in the conventional secondary flow control nozzle shown in Figs. 15 to 18, the velocity distribution at the nozzle outlet varies greatly as shown in Fig. 19.
    However, in the nozzle vane according to the present invention, the velocity distribution at the blade exit (circumferential velocity Vt and meridional velocity Vm is expressed as a specific dimension relative to absolute velocity V = (Vt 2 + Vm 2 ) 0.5 ). Is maintained at about the same speed distribution of a conventional nozzle blade as shown in FIG.
    In conclusion, even if the nozzle blades of the conventional turbine stage are replaced by the nozzle blades according to the present invention, the turbine nozzle does not act against the rotating blade located downstream of the turbine stage.
    As described above, the turbine nozzle according to the present invention can suppress the secondary flow at the end of the nozzle blade, thereby reducing the loss due to the secondary flow. In addition, the turbine nozzle according to the present invention provides a velocity distribution at the same nozzle exit as in a conventional nozzle blade, and therefore does not act against the rotating blade located downstream of the turbine nozzle.
    Although certain preferred embodiments have been shown and described in detail in the present invention, it is evident that various modifications and embodiments are possible without departing from the scope of the appended claims.
    The present invention is suitable for turbines used to drive various machines, such as generators in power plants.
    权利要求:
    Claims (14)
    [1" claim-type="Currently amended] As a turbine nozzle,
    A series of nozzle vanes disposed circumferentially in an annular passage formed between the inner and outer rings of the diaphragm and fixed to the inner and outer surfaces (hub and tip end walls) of the diaphragm; And
    A flow passage formed between a pressure surface and a suction surface of a neighboring wing of the nozzle blades, the cross section of the flow passage extending from the hub and tip end walls along the wing height and substantially in a predetermined range formed by a curved line; A turbine nozzle comprising another range formed by straight lines.
    [2" claim-type="Currently amended] The method of claim 1,
    And the predetermined range includes a range corresponding to 20 to 40% of the height of the vanes from the hub and tip end walls.
    [3" claim-type="Currently amended] The method of claim 1,
    And the predetermined range comprises a range corresponding to at least 30% of the meridional width of the nozzle blade from the leading edge of the nozzle blade.
    [4" claim-type="Currently amended] The method of claim 1,
    The cross section of the flow passage is formed of a line on the pressure surface and a line on the suction surface at a meridian position within a range corresponding to at least 30% of the meridian width of the nozzle blade from a leading edge of the nozzle blade. And each line comprises a substantially straight line in the center portion of the nozzle vane that does not include a range corresponding to 20-40% of the vane height from the hub and tip end walls.
    [5" claim-type="Currently amended] The method of claim 4, wherein
    The distance from the point of intersection between the line on the pressure side or suction side and the hub end wall to the point of intersection between the extension line of the substantially straight line and the hub end wall and the point of intersection at the point of intersection between the line on the pressure side or suction side and the tip end wall. And the distance to the intersection between the extension line of the straight line and the tip end wall has a maximum value at the leading edge of the nozzle vane.
    [6" claim-type="Currently amended] The method of claim 5,
    The maximum value is a turbine nozzle, characterized in that 5 to 15% of the wing height.
    [7" claim-type="Currently amended] The method of claim 5,
    Wherein the distances at the leading edge of the nozzle vane are in the range of 5-15% of the vane height and are at least 5% of the vane height within a position of 30% of the nozzle vane meridional width. .
    [8" claim-type="Currently amended] As a turbine nozzle,
    A series of nozzle vanes disposed circumferentially in an annular passage formed between the inner and outer rings of the diaphragm and fixed to the inner and outer rings of the turbine diaphragm;
    The hub end wall as a pressure surface of the nozzle blades facing the tip end wall of the diaphragm in a predetermined range between the tip end wall and the midspan of the vane in a predetermined range in the meridion direction of the nozzle vane. The pressure surface facing the hub end wall of the diaphragm in a predetermined range between the midspans of the vanes;
    The hub end wall as a suction surface of each nozzle blade toward the hub end wall of a diaphragm in a predetermined range between the tip end wall and the midspan of the vane in a predetermined range in the meridion direction of the nozzle vane. And the suction surface facing the tip end wall of the diaphragm in a predetermined range between the blade and the midspan of the vane.
    [9" claim-type="Currently amended] The method of claim 8,
    And the predetermined range includes a range corresponding to 20 to 40% of the height of the vanes from the hub and tip end walls.
    [10" claim-type="Currently amended] The method of claim 8,
    And the predetermined range comprises a range corresponding to at least 30% of the meridional width of the nozzle blade from the leading edge of the nozzle blade.
    [11" claim-type="Currently amended] The method of claim 8,
    The cross section of the flow passage is formed of a line on the pressure surface and a line on the suction surface at a meridion position within a range corresponding to at least 30% of the longitudinal width of the nozzle blade from the leading edge of the nozzle blade. And each line comprises a substantially straight line at a center portion of the nozzle vane that does not include a range corresponding to 20-40% of the vane height from the hub and tip end walls.
    [12" claim-type="Currently amended] The method of claim 11,
    The distance from the point of intersection between the line on the pressure side or suction side and the hub end wall to the point of intersection between the extension line of the substantially straight line and the hub end wall and the point of intersection at the point of intersection between the line on the pressure side or suction side and the tip end wall. And the distance to the intersection between the extension line of the straight line and the tip end wall has a maximum value at the leading edge of the nozzle vane.
    [13" claim-type="Currently amended] The method of claim 12,
    The maximum value is a turbine nozzle, characterized in that 5 to 15% of the wing height.
    [14" claim-type="Currently amended] The method of claim 12,
    Wherein the distances at the leading edge of the nozzle vane are in the range of 5-15% of the vane height and are at least 4% of the vane height within a position of 30% of the nozzle vane meridional width. .
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    US6491493B1|2002-12-10|
    CN1163662C|2004-08-25|
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    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    法律状态:
    1998-06-12|Priority to JP10-164833
    1998-06-12|Priority to JP16483398
    1999-06-10|Application filed by 마에다 시게루, 가부시키 가이샤 에바라 세이사꾸쇼
    2001-06-25|Publication of KR20010052802A
    2006-03-31|Application granted
    2006-03-31|Publication of KR100566759B1
    优先权:
    申请号 | 申请日 | 专利标题
    JP10-164833|1998-06-12|
    JP16483398|1998-06-12|
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