Large-expansion-ratio guider and centripetal turbine

Abstract

The invention relates to the technical field of aero-engines, and discloses a large-expansion-ratio guider and a centripetal turbine. The large-expansion-ratio guider comprises an outer ring end wall; the inner ring end wall and the outer ring end wall are coaxially arranged, and a flow channel is formed between the inner ring end wall and the outer ring end wall; the multiple guide vanes are arranged in the flow channel in the circumferential direction of the outer ring end wall at intervals, the two ends of each guide vane are connected with the outer ring end wall and the inner ring end wall correspondingly, and every two adjacent guide vanes, the wall face of the outer ring end wall and the wall face of the inner ring end wall define a vane grid channel; a plurality of convex parts and a plurality of groove parts are arranged on the outer ring end wall and / or the inner ring end wall; the protruding parts are arranged on the pressure face side of the guide vane one by one, and the groove parts are arranged on the suction face side of the guide vane one by one. By the adoption of the non-axisymmetric end wall structure, formation and development of main loss vortex systems such as channel vortexes and angle vortexes can be restrained from the source, mutual interference between the main loss vortex systems and shock waves in a cascade channel is remarkably reduced, and therefore flow loss in the guider is greatly reduced under the working condition of large expansion ratio and high load.

Description

Technical Field

[0001] This invention relates to the field of aero-engine technology, specifically to a high expansion ratio guide vane and a radial turbine. Background Technology

[0002] Centripetal turbines, with their compact structure and high single-stage expansion ratio, are widely used in small gas turbines, auxiliary power units, and turbochargers. As a key component of centripetal turbines, the guide vane plays a crucial role in effectively converting the thermal and pressure energy of high-temperature, high-pressure gas into kinetic energy and providing downstream impellers with inlet conditions that meet aerodynamic design requirements. With the continuous improvement of turbine equivalent power and performance requirements in aero-engine systems, centripetal turbines are rapidly developing towards higher expansion ratios. To adapt to this trend and achieve unconventional expansion, modern high-expansion-ratio centripetal turbine guide vanes generally adopt high-load designs and introduce a double-throat design consisting of a "geometric throat" and an "aerodynamic throat." The geometric throat is the physically smallest cross-section of the converging channel formed between the two guide vanes in the guide vane; the gas flow velocity increases and pressure decreases here, and it can be accelerated to the speed of sound. The aerodynamic throat, located downstream of the geometric throat, is a narrower channel formed between the guide vanes and the impeller blades. Through these two convergence channels, the airflow can continue to expand to supersonic speeds, thereby extracting more energy and increasing the turbine's equivalent power.

[0003] In traditional guide vanes, the endwalls of the flow channels typically employ axisymmetric cylindrical or conical structures. However, this traditional design reveals serious flaws under high expansion ratio conditions. The strong lateral pressure gradient generated by high-load operation drives low-energy fluid near the endwalls to migrate from the blade pressure surface to the suction surface, forming larger and stronger channel vortices. Simultaneously, the horseshoe vortex formed by the impact of the incoming flow at the blade leading edge has its pressure-side branch entrained into the channel vortex under the influence of the lateral pressure gradient, further intensifying the secondary flow in the end region. These complex secondary flow structures not only consume energy themselves, leading to total pressure loss, but also violently mix with the suction surface boundary layer, reducing the efficiency of the mainstream flow. Under high expansion ratio conditions, the resulting aerodynamic losses can account for more than 30% of the total guide vane losses, becoming a major bottleneck restricting turbine efficiency improvement. Furthermore, under high expansion ratio and high-load conditions, a supersonic flow field forms early within the guide vane channel, reaching supersonic speeds approximately 30% of the way along the flow direction. The expansion shock wave generated within the channel interacts strongly with the secondary vortex in the end region, further increasing aerodynamic losses. Even more serious is the fact that the powerful vortex structure leads to extremely uneven distribution of the airflow angle at the guide vane outlet, severely affecting the aerodynamic performance and operational stability of the downstream rotor blades. Furthermore, the unique radial inflow conditions of the radial turbine are also prone to inducing flow separation under high loads, further deteriorating the flow field quality. Therefore, traditional axisymmetric endwall designs are no longer sufficient to meet the urgent requirements of advanced high-expansion-ratio radial turbines for high efficiency and high stability. Summary of the Invention

[0004] To address the problems of existing technologies failing to suppress the severe secondary flow driven by strong pressure gradients, the surge in aerodynamic losses caused by the mixing of secondary flow and boundary layer, the interference of shock waves and secondary flow, and the severe deterioration of the outlet flow field under special operating conditions with high expansion ratio and high load, this invention provides a high expansion ratio guide and centripetal turbine. It adopts a non-axisymmetric structure, which can control the formation and development of secondary flow in the end region, effectively suppress the intensity and scale of major loss vortex systems such as channel vortices and angular vortices, significantly reduce their mutual interference with shock waves in the channel, and improve the uniformity of the airflow angle and velocity distribution at the guide outlet.

[0005] In a first aspect, the present invention provides a large expansion ratio guide, comprising an outer ring end wall; an inner ring end wall coaxially disposed with the outer ring end wall and forming a flow channel between them; a plurality of guide vanes spaced circumferentially within the flow channel, wherein the two ends of the guide vanes are respectively connected to the outer ring end wall and the inner ring end wall, and adjacent two guide vanes, the wall surface of the outer ring end wall, and the wall surface of the inner ring end wall enclose a vane cascade channel; the outer ring end wall and / or the inner ring end wall are provided with a plurality of protrusions and a plurality of grooves, wherein the plurality of protrusions and grooves are all disposed on one side where the guide vanes are located, the plurality of protrusions are disposed one by one on the pressure surface side of the guide vanes, and the plurality of grooves are disposed one by one on the suction surface side of the guide vanes.

[0006] The high expansion ratio guide vane provided by this invention forms a non-axisymmetric endwall structure by constructing protrusions and grooves on the outer and / or inner ring endwalls corresponding to the pressure and suction sides of the guide vane. This structure can actively intervene in and reshape the migration path of low-energy fluid in the end region. The protrusions effectively block and weaken the intensity of secondary flow migrating from the pressure to the suction side by creating local lateral pressure barriers. The grooves, by providing controllable low-energy fluid collection and guidance channels, prevent disordered accumulation and the formation of strong vortices in the suction corner region. This synergistic effect suppresses the formation and development of major loss vortex systems such as channel vortices and corner vortices at the source, and significantly reduces their mutual interference with shock waves within the blade passage. This results in a substantial reduction in internal flow losses under high expansion ratio and high load conditions. Simultaneously, the optimized end region flow field makes the outlet airflow angle and velocity distribution of the guide vane more uniform, providing downstream impellers with excellent aerodynamic quality, ultimately effectively improving the overall turbine stage efficiency and operational stability.

[0007] Preferably, the protrusion extends from the leading edge to the trailing edge of the guide vane along the intersection line of the pressure surface of the guide vane and the outer or inner annular end wall. Wherein, if the protrusion and the groove are located on the outer annular end wall, the intersection line described therein is the intersection line of the pressure surface of the guide vane and the outer annular end wall; if the protrusion and the groove are located on the inner annular end wall, the intersection line described therein is the intersection line of the pressure surface of the guide vane and the inner annular end wall. This arrangement allows the protrusion to continuously apply a stable lateral blocking effect to the low-energy fluid in the end region along the entire pressure surface side channel of the guide vane, thereby effectively suppressing the generation and migration of secondary vortex systems such as channel vortices from the source, and helping to maintain the stability of the pressure distribution from the geometric throat to the aerodynamic throat region.

[0008] Preferably, the maximum arch height of the protrusion is located within the range of 10% to 40% of the total flow length from the inlet of the blade channel. This can suppress the flow over-expansion caused by the bending of the end walls between the two throats, thereby avoiding the resulting shock wave enhancement and flow channel separation.

[0009] Preferably, the maximum arch height of the protrusion does not exceed 10% of the axial thickness of the guide vane. This height limitation ensures that the mainstream flow path is not excessively distorted, resulting in additional flow losses such as bending losses, while still generating sufficient end-zone secondary vortex lateral flow blocking effect.

[0010] Preferably, a geometric throat is formed between two adjacent guide vanes. The groove extends from the leading edge of the guide vane to the corresponding geometric throat along the intersection of the suction surface of the guide vane and the outer or inner annular end wall. This positioning avoids the formation of a vortex backflow zone between the two throats. It is understood that, in the channel between two adjacent guide vanes, the section containing the shortest straight segment connecting the pressure surface of one guide vane to the suction surface of the adjacent guide vane, and perpendicular to both profiles, is the geometric throat. Downstream of the geometric throat is the aerodynamic throat, and the aerodynamic throat and the geometric throat together form a double throat.

[0011] Preferably, the maximum recess depth of the groove is located within the range of 5% to 15% of the total flow length from the inlet of the blade cascade channel. Setting the maximum recess depth here allows for intervention and control during the generation stage of major secondary vortex systems such as horseshoe vortices and channel vortices, effectively guiding the low-energy fluid towards the mainstream region of the blade cascade channel for pre-dissipation, thereby weakening the intensity and scale of the subsequently developing vortex systems from the source.

[0012] Preferably, the maximum recess depth of the groove does not exceed 10% of the axial thickness of the guide vane. This depth can prevent the recess from being too deep, thus avoiding the formation of low-energy fluid accumulation and rotating corner regions, while still generating sufficient end-zone secondary vortex lateral flow blocking effect.

[0013] Preferably, the transition region wall between the edge of the protrusion and the outer or inner ring end wall, and the transition region wall between the edge of the groove and the outer or inner ring end wall, both have continuous curvature. This eliminates geometric discontinuities in the transition region, avoids local flow separation caused by steps or sharp corners, and also helps improve the stress distribution of the structure, avoiding stress concentration, thus achieving a high degree of balance between aerodynamic performance and structural integrity.

[0014] Preferably, the wall surfaces corresponding to the protrusions and the grooves have continuous curvature. Continuous curvature in the protrusions and grooves avoids sudden changes in flow direction and pressure, ensuring that the airflow maintains a smooth boundary layer state as it passes through them, thus preventing additional aerodynamic losses due to flow separation.

[0015] Secondly, the present invention also provides a centripetal turbine equipped with the aforementioned high expansion ratio guide. Apart from the aforementioned high expansion ratio guide, the remaining structure of the centripetal turbine of the present invention is that of a high expansion ratio centripetal turbine with a double throat, as is common in the prior art. Its specific structural composition and working principle are also prior art, and therefore will not be described in detail.

[0016] The beneficial effects of this invention are as follows: By constructing protrusions and grooves on the outer and / or inner ring endwalls corresponding to the pressure and suction sides of the guide vanes, a non-axisymmetric endwall structure is formed, which can actively intervene in and reshape the migration path of low-energy fluid in the end region. The protrusions, by generating local lateral pressure barriers, can effectively block and weaken the intensity of secondary flow migrating from the pressure side to the suction side; the grooves, by providing controllable low-energy fluid collection and guiding channels, can prevent its disordered accumulation and formation of strong vortices in the suction side corner region. This synergistic effect can suppress the formation and development of major loss vortex systems such as channel vortices and corner vortices from the source, and significantly reduce their mutual interference with shock waves within the blade channel, thereby achieving a significant reduction in internal flow losses under high expansion ratio and high load conditions. Simultaneously, the optimized end-region flow field makes the outlet airflow angle and velocity distribution of the guide vane more uniform, providing downstream impellers with excellent aerodynamic quality inflow, ultimately effectively improving the overall turbine stage's efficiency and operational stability. Attached Figure Description

[0017] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the specific embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0018] Figure 1 This is a partial cross-sectional view of a high expansion ratio guide according to an embodiment of the present invention; Figure 2 This is a schematic diagram of the partial structure of the inner ring end wall from one perspective; Figure 3 This is a schematic diagram of the partial structure of the inner ring end wall from another perspective; Figure 4 A schematic diagram of the double larynx; Figure 5 This is a schematic diagram of angle θ; Figure 6 Comparison diagram of the distribution of secondary vortex cores before and after setting the protrusion and groove sections; Figure 7 Mach number distribution cloud map of the 0.9 blade height section before setting the protrusion and groove; Figure 8 Mach number distribution cloud map of the 0.9 blade height section after setting the protrusion and groove.

[0019] Explanation of reference numerals in the attached figures: 1. Outer ring end wall; 101. Flow channel; 102. Blade passage; 1021. Geometric throat; 2. Inner ring end wall; 201. Protrusion; 202. Groove; 3. Guide vane; 4. Impeller; 401. Aerodynamic throat. Detailed Implementation

[0020] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0021] The following is combined Figures 1 to 8 The following describes embodiments of the present invention.

[0022] According to an embodiment of the present invention, in one aspect, a guide with a large expansion ratio is provided, combined with Figures 1 to 4As shown, it includes an outer ring end wall 1; an inner ring end wall 2, which is coaxially arranged with the outer ring end wall 1 and forms a flow channel 101 between them; a plurality of guide vanes 3, which are arranged circumferentially in the flow channel 101 along the outer ring end wall 1. The two ends of the guide vanes 3 are respectively connected to the outer ring end wall 1 and the inner ring end wall 2. Two adjacent guide vanes 3, the wall surface of the outer ring end wall 1, and the wall surface of the inner ring end wall 2 enclose a blade channel 102. The inner ring end wall 2 is provided with a plurality of protrusions 201 and a plurality of grooves 202. The plurality of protrusions 201 and grooves 202 are all provided on the side where the guide vanes 3 are located. The plurality of protrusions 201 are provided one by one on the pressure surface side of the guide vane 3, and the plurality of grooves 202 are provided one by one on the suction surface side of the guide vane 3. That is, each guide vane 3 corresponds to one protrusion 201 and one groove 202. Specifically, the protrusion 201 is formed by arching up a portion of the inner ring end wall 2 to the side where the guide vane 3 is located, and the groove 202 is formed by recessing a portion of the inner ring end wall 2 to the opposite side.

[0023] It is understandable that each guide vane 3 has a pressure surface and a suction surface. The pressure surface is a blade basin-shaped surface facing the incoming flow and bearing a higher static pressure of gas, while the suction surface is a blade back-shaped surface facing away from the incoming flow and generating a lower static pressure of gas due to the acceleration of the airflow. Together, they constitute the lateral boundary of the blade passage 102. The pressure difference between them is the fundamental reason for driving the deflection and acceleration of the airflow, and it is also the main driving force for inducing lateral secondary flow in the endwall region.

[0024] The high expansion ratio guide vane provided in this embodiment forms a non-axisymmetric endwall structure by constructing protrusions 201 and grooves 202 on the inner ring endwall 2 corresponding to the pressure and suction sides of the guide vane 3. This structure can actively intervene in and reshape the migration path of low-energy fluid in the end region. The protrusions 201 effectively block and weaken the intensity of secondary flow migrating from the pressure side to the suction side by generating local lateral pressure barriers. The grooves 202 prevent the disordered accumulation of low-energy fluid in the suction corner region, thus avoiding the formation of strong vortices. This synergistic effect can suppress the formation and development of major loss vortex systems such as channel vortices and corner vortices from the source, and significantly reduce their mutual interference with shock waves in the blade channel 102. This results in a significant reduction in internal flow losses under high expansion ratio and high load conditions. At the same time, the optimized end region flow field makes the outlet airflow angle and velocity distribution of the guide vane more uniform, providing aerodynamically superior incoming flow to the downstream impeller 4, ultimately effectively improving the working efficiency and operational stability of the entire turbine stage.

[0025] Furthermore, the protrusion 201 extends from the leading edge of the guide vane 3 to the trailing edge of the guide vane 3 along the intersection of the pressure surface of the guide vane 3 and the inner ring end wall 2. This arrangement allows the protrusion 201 to continuously apply a stable lateral blocking effect to the low-energy fluid in the end region along the entire pressure surface side channel of the guide vane 3, thereby effectively suppressing the generation and migration of secondary vortex systems such as channel vortices from the source, and helping to maintain the stability of the pressure distribution in the region from the geometric throat 1021 to the aerodynamic throat 401.

[0026] Furthermore, the maximum arch height of the protrusion 201 is located within a range of 10% to 40% of the total flow length from the inlet of the blade channel 102. This can suppress the flow over-expansion caused by the bending of the end walls between the two throats, thereby avoiding the resulting shock wave enhancement and flow channel separation.

[0027] Furthermore, the maximum arch height of the protrusion 201 does not exceed 10% of the axial thickness of the guide vane 3. This height limit ensures that the mainstream flow path does not undergo excessive distortion, resulting in additional flow losses such as bending losses, while still generating sufficient end-zone secondary vortex lateral flow blocking effect.

[0028] Further, see Figure 4 A geometric throat 1021 is formed between two adjacent guide vanes 3. A groove 202 extends from the leading edge of the guide vane 3 to the corresponding geometric throat 1021 along the intersection of the suction surface of the guide vane 3 and the inner ring end wall 2. The geometric throat 1021 is not significantly shaped to ensure that the gas flow capacity is the same as the design. This location avoids the formation of a vortex backflow zone before the double throats. It can be understood that the geometric throat 1021 is the section containing the shortest straight segment among the straight segments connecting the pressure surface of one guide vane 3 and the suction surface of the adjacent guide vane 3, and perpendicular to both profiles, in the channel between two adjacent guide vanes 3. Downstream of the geometric throat 1021 is the aerodynamic throat 401, which is a narrower channel formed between the guide vane 3 and the downstream impeller 4 blades. The aerodynamic throat 401 and the geometric throat 1021 form a double throat.

[0029] Furthermore, the maximum recess depth of the groove 202 is located within the range of 5% to 15% of the total flow length from the inlet of the blade channel 102. In this embodiment, the maximum recess of the groove 202 is located near 10% of the total flow length from the inlet of the blade channel 102. Setting the maximum recess depth here allows for intervention and control during the generation stage of major secondary vortex systems such as horseshoe vortices and channel vortices, effectively guiding the low-energy fluid towards the main flow region of the blade channel 102 for pre-dissipation, thereby weakening the intensity and scale of the subsequently developing vortex systems from the source.

[0030] Furthermore, the maximum depression depth of the groove portion 202 does not exceed 10% of the axial thickness of the guide vane 3. This depth can avoid the formation of low-energy fluid accumulation and rotating corner regions due to excessive depression while generating sufficient transverse flow blocking effects of the end region secondary vortices.

[0031] Furthermore, the wall surfaces of the transition regions between the edge of the protrusion portion 201 and the inner ring end wall 2 and between the edge of the groove portion 202 and the inner ring end wall 2 are both curvature continuous. This can eliminate the geometric discontinuity in the transition region, avoid local flow separation caused by steps or sharp corners, and is also beneficial to improving the stress distribution of the structure and avoiding stress concentration, thereby achieving an efficient balance between aerodynamic performance and structural integrity.

[0032] Furthermore, the wall surfaces corresponding to the protrusion portion 201 and the groove portion 202 are both curvature continuous. The curvature continuous protrusion portion 201 and groove portion 202 can avoid sudden changes in the flow direction and pressure, ensure that the airflow can maintain a smooth boundary layer state when flowing through the protrusion portion 201 and the groove portion 202, and avoid additional aerodynamic losses caused by flow separation.

[0033] Optionally, in order to refine the positions of the protrusion portion 201 and the groove portion 202, refer to Figure 5 , before shaping the inner ring end wall 2 (i.e., setting the protrusion portion 201 and the groove portion 202), first evenly divide the cascade passage 102 into n intervals along the flow direction, that is, divide it with n - 1 circumferential control lines; each circumferential control line intersects with the center line of the cascade passage 102 at a point, and control the angle θ between the normal line at this point and the corresponding circumferential control line within the range of 20 to 65 degrees; then simulate the inner ring end wall 2 to determine the supersonic shock wave region in the cascade passage 102. When shaping the inner ring end wall 2, both the protrusion portion 201 and the groove portion 202 avoid these supersonic shock wave regions. Here, n is set according to the length of the cascade passage 102, and preferably 6 < n < 10 to ensure the precise coordinated control of the "wave system" and "vortex", and at the same time not generate large control variables resulting in a long optimization cycle. It should be noted that the circumferential control line is a reference line set along the circumferential direction (circumferential) in the annular cascade passage of rotating machinery (such as turbines, compressors) for constraining or evaluating parameter distributions, which is a prior art, so no further description will be made here.

[0034] As Figure 6 can be seen, after adopting the non-axisymmetric end wall design (i.e., setting the protrusion portion 201 and the groove portion 202), due to the reduction of the strength and size of the outlet passage vortex, the degree of deflection of the outlet airflow decreases, the relative inlet airflow angle of the moving blade decreases, and the inlet attack angle decreases. Comparing Figure 7 and Figure 8It is evident that by adopting the non-axisymmetric endwall design, the flow separation bubble area near the leading edge of the suction surface of impeller 4 is reduced, effectively improving the inlet conditions of impeller 4.

[0035] Alternatively, the same protrusion 201 and groove 202 can be provided only on the outer ring end wall 1, or the protrusion 201 and groove 202 can be provided on both the outer ring end wall 1 and the inner ring end wall 2, so as to achieve coordinated three-dimensional control of the gas flow of the entire blade channel 102.

[0036] According to an embodiment of the present invention, another aspect provides a centripetal turbine, which is equipped with the large expansion ratio guide of the aforementioned embodiment. Apart from the aforementioned large expansion ratio guide, the rest of the structure of the centripetal turbine in this embodiment is a large expansion ratio centripetal turbine with a double throat in the prior art. Its specific structural composition and working principle are all prior art, and therefore will not be described in detail.

[0037] Although embodiments of the invention have been described in conjunction with the accompanying drawings, those skilled in the art can make various modifications and variations without departing from the spirit and scope of the invention, and such modifications and variations all fall within the scope defined by the appended claims.

Claims

1. A guide with a large expansion ratio, characterized in that, include: Outer ring end wall (1); The inner ring end wall (2) is coaxially arranged with the outer ring end wall (1) and a flow channel (101) is formed between them; Multiple guide vanes (3) are arranged circumferentially within the flow channel (101) along the outer ring end wall (1). The two ends of the guide vanes (3) are respectively connected to the outer ring end wall (1) and the inner ring end wall (2). Two adjacent guide vanes (3), the wall surface of the outer ring end wall (1), and the wall surface of the inner ring end wall (2) enclose a blade cascade channel (102). The outer ring end wall (1) and / or the inner ring end wall (2) are provided with a plurality of protrusions (201) and a plurality of grooves (202). The plurality of protrusions (201) and grooves (202) are all located on the side where the guide vane (3) is located. The plurality of protrusions (201) are located on the pressure surface side of the guide vane (3), and the plurality of grooves (202) are located on the suction surface side of the guide vane (3).

2. The guide with a large expansion ratio according to claim 1, characterized in that, The protrusion (201) extends from the leading edge of the guide vane (3) to the trailing edge of the guide vane (3) along the intersection line of the pressure surface of the guide vane (3) and the outer ring end wall (1) or the inner ring end wall (2).

3. A guide with a large expansion ratio according to claim 1, characterized in that, The maximum arch height of the protrusion (201) is located within the range of 10% to 40% of the total flow length from the inlet of the blade channel (102).

4. A guide with a large expansion ratio according to claim 1, characterized in that, The maximum arch height of the protrusion (201) does not exceed 10% of the axial thickness of the guide vane (3).

5. A guide with a large expansion ratio according to claim 1, characterized in that, A geometric throat (1021) is formed between two adjacent guide vanes (3). The groove (202) extends from the leading edge of the guide vane (3) to the front of the corresponding geometric throat (1021) along the intersection of the suction surface of the guide vane (3) and the outer ring end wall (1) or the inner ring end wall (2).

6. A guide with a large expansion ratio according to claim 1, characterized in that, The maximum recess depth of the groove (202) is located within the range of 5% to 15% of the total flow length from the inlet of the blade channel (102).

7. A guide with a large expansion ratio according to claim 1, characterized in that, The maximum recess depth of the groove (202) does not exceed 10% of the axial thickness of the guide vane (3).

8. A guide with a large expansion ratio according to claim 1, characterized in that, The transition area wall between the edge of the protrusion (201) and the outer ring end wall (1) or the inner ring end wall (2), and the transition area wall between the edge of the groove (202) and the outer ring end wall (1) or the inner ring end wall (2) are all curved continuously.

9. A guide with a large expansion ratio according to claim 1, characterized in that, The wall surfaces corresponding to the protrusion (201) and the groove (202) have continuous curvature.

10. A centripetal turbine, characterized in that, The device is equipped with a large expansion ratio guide as described in any one of claims 1 to 9.

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