A transonic stator vane applied to a centripetal turbine, centripetal turbine

Abstract

The present invention discloses a transonic stator and a centripetal turbine applied to a centripetal turbine. The transonic stator adopts a three-stage radial air flow passage and a "helmet-shaped" blade profile design. The width of the first-stage passage gradually decreases along the flow direction, greatly improving the air flow acceleration effect and significantly reducing the flow loss at the blade inlet. The width of the second-stage radial air flow passage remains unchanged, facilitating the inspection of the throat area of the stator. The change in the width of the third-stage radial air flow passage is coupled according to the air flow Mach number at the outlet of the stator blade, effectively reducing the flow loss in the vane-less area when the air flow Mach number at the outlet of the stator blade is in different situations, and further improving the acceleration performance of the stator. The "helmet-shaped" blade profile adopts a leading edge with a small curvature radius and a blade back with zero curvature design, which can effectively improve the leading edge load of the blade and the acceleration ability of the cascade passage, reduce the blade profile loss, and reduce the size and weight of the stator.

Description

Technical Field

[0001] The present invention relates to the technical field of centripetal turbine guide vane design. In particular, it relates to a transonic guide vane applied to a centripetal turbine. Additionally, it also relates to a centripetal turbine employing the above transonic guide vane. Background Art

[0002] One of the major development trends of current gas turbine engines is to gradually develop towards higher power-to-weight ratios and higher thermodynamic cycle parameters, with continuously increasing turbine inlet temperatures and turbine expansion ratios. This has led to the increasingly frequent adoption of higher single-stage expansion ratios and aerodynamic loads in the design of turbine components to achieve the goals of reducing the number of stages, weight, and cost. Due to a series of advantages such as strong single-stage work capacity, few components, and simple structure, centripetal turbines are widely used in gas turbine engine power plants such as auxiliary power units and small-power-class turboprop shafts. The main function of the turbine guide vane is to accelerate the airflow and impart pre-rotation. Due to the characteristic of radial airflow in the centripetal turbine guide vane, as the flow path radius decreases, the airflow passage area along the flow direction often decreases. For a turbine where the outlet velocity of the guide vane blades needs to reach the supercritical state, this makes it impossible to achieve the goal of continuous expansion and acceleration of the airflow after it flows out of the guide vane blades, thus affecting the work capacity of the downstream impeller. In addition, since the maximum outer contour size of the centripetal turbine is usually determined by the structural size of the guide vane, for an aviation gas turbine device pursuing lightweight and high structural compactness, it is required that the structural size of the turbine guide vane be as small as possible, which conflicts with the requirement of the transonic guide vane to increase the guide vane flow area to achieve supersonic flow, increasing the design difficulty of the transonic guide vane.

[0003] Currently, Chinese invention patent CN108533332A proposes a converging-diverging nozzle with an "inward concave" back surface. Through the design of the blade profile, the minimum width position (throat) of the cascade is located somewhere in the middle of the cascade channel formed by two adjacent blades. This enables the airflow to continue expanding and accelerating in the cascade channel after reaching the critical velocity at the minimum width position of the cascade, thereby achieving the purpose of transonic flow. However, this structure has high requirements for the back surface blade profile design and processing, and has poor control over the flow loss when the blade outlet Mach number is between 1.1 and 1.2. On the other hand, restricted by the contraction of the flow area in the blade-free region at the outlet of the guide vane blades, the transonic airflow at the blade outlet will inevitably slow down when it reaches the downstream rotor blades, and the expected high transonic flow purpose cannot be achieved.

[0004] Therefore, the existing conventional guide vanes and air flow channels are not coupledly designed according to the flow characteristics of a transonic centripetal turbine guide, resulting in poor acceleration ability of the guide, high flow loss, and difficulty in meeting the usage requirements of a centripetal turbine with a large expansion ratio. In addition, the load at the inlet section of the guide vane is relatively low and the acceleration performance is poor, resulting in a relatively long blade, a large size of the guide, and affecting the structural size and weight of the entire turbine component. Summary of the Invention

[0005] The present invention provides a transonic guide and a centripetal turbine applied to a centripetal turbine to solve the technical problems that the existing guide of the centripetal turbine cannot achieve both transonic flow, reduce flow loss, and lower the processing and manufacturing difficulty and the size of the guide vane.

[0006] According to one aspect of the present invention, there is provided a transonic guide applied to a centripetal turbine, including a guide outer ring, a guide inner ring, and a plurality of guide vanes. The guide outer ring and the guide inner ring are symmetrically arranged around the engine axis. An annular air flow channel is formed between the guide outer ring and the guide inner ring. The plurality of guide vanes are evenly arranged in the annular air flow channel at the same angular interval in the circumferential direction. The annular air flow channel includes an axial air flow channel and a radial air flow channel communicating therewith. The radial air flow channel is composed of three sections. Among them, the first section of the radial air flow channel extends from the inlet to a position 10% - 15% of the blade height from the leading edge of the blade. The second section of the radial air flow channel extends from the end point of the first section of the radial air flow channel to the throat of the guide vane. The third section of the radial air flow channel extends from the end point of the second section of the radial air flow channel to the end point of the guide outer ring. The channel width of the first section of the radial air flow channel gradually decreases in the flow direction. The channel width of the second section of the radial air flow channel remains unchanged in the flow direction. The channel width change of the third section of the radial air flow channel is coupledly designed according to the air flow Mach number at the outlet of the guide vane.

[0007] Further, in the first section of the radial air flow channel, the cross-sectional profile of the guide outer ring is a straight line, and the cross-sectional profile of the guide inner ring gradually expands from the end point of the first section of the radial air flow channel towards the inlet. In the second section of the radial air flow channel, the cross-sectional profiles of the guide outer ring and the guide inner ring are both straight lines.

[0008] Further, the ratio of the channel width at the starting point to the channel width at the end point of the first section of the radial air flow channel is between 1.15 and 1.25.

[0009] Furthermore, when the airflow velocity at the outlet of the guide vane is subsonic, the channel width of the third-stage radial airflow channel gradually decreases or remains unchanged along the flow direction, so that the airflow passage area of the third-stage radial airflow channel gradually decreases along the flow direction; when the airflow velocity at the outlet of the guide vane is transonic, the channel width of the third-stage radial airflow channel gradually increases along the flow direction, so that the airflow passage area of the third-stage radial airflow channel gradually increases along the flow direction.

[0010] Furthermore, the leading edge of the guide vane adopts a helmet-shaped design to reduce the blade profile curvature between the starting point and the throat point of the blade profile on the suction side.

[0011] Furthermore, the ratio of the maximum inscribed circle diameter between the suction side and the pressure side of the guide vane to the axial chord length of the blade profile is between 0.09 and 0.11.

[0012] Furthermore, the ratio of the maximum inscribed circle diameter between the suction side and the pressure side of the guide vane to the trailing edge diameter is between 5 and 6.

[0013] Furthermore, the ratio of the distance from the starting point of the blade profile on the suction side to the center of the maximum inscribed circle to the axial chord length of the blade profile is between 0.15 and 0.28, and the ratio of the distance from the ending point of the blade profile on the suction side to the center of the maximum inscribed circle to the axial chord length of the blade profile is between 0.9 and 1.

[0014] Furthermore, the angle between the line connecting the starting point of the blade profile on the suction side and the center of the maximum inscribed circle and the line connecting the ending point of the blade profile on the suction side and the center of the maximum inscribed circle is between 160° and 180°.

[0015] In addition, the present invention also provides a centripetal turbine using the transonic guide vane described above.

[0016] The present invention has the following effects:

[0017] The transonic guide vane applied to a centripetal turbine according to the present invention adopts a three-stage design for the radial air flow passage. The first-stage radial air flow passage extends from the inlet to a position 10% - 15% of the blade height from the leading edge of the blade. The second-stage radial air flow passage extends from the end point of the first-stage radial air flow passage to the throat of the guide vane. The third-stage radial air flow passage extends from the end point of the second-stage radial air flow passage to the end point of the outer ring of the guide vane. Among them, the channel width of the first-stage radial air flow passage gradually decreases along the flow direction. Coupled with the gradual decrease in the channel width of the cascade, the air flow passage area is further reduced along the air flow direction on the basis of the existing guide vane, greatly improving the air flow acceleration effect of the first-stage radial air flow passage. Moreover, it can also reduce the lap radius with the combustion chamber and the overall outline size. More importantly, it can reduce the flow velocity of the air flow in the axial air flow passage when entering the inlet of the radial air flow passage, greatly reducing the flow loss at the blade inlet. The channel width of the second-stage radial air flow passage remains unchanged. Due to the gradual decrease in the channel width of the cascade, the air flow can still be accelerated continuously, and it is beneficial to ensure that the throat of the guide vane is less affected by the machining quality of the outer ring of the guide vane and the radial position of the blade, facilitating the inspection of the throat area of the guide vane. The change in the channel width of the third-stage radial air flow passage is designed by coupling according to the air flow Mach number at the outlet of the guide vane, effectively reducing the flow loss in the non-blade area when the air flow Mach number at the outlet of the guide vane is in different situations. Whether the air flow Mach number at the outlet of the guide vane is subsonic or supersonic, the air flow at the outlet of the guide vane can be further accelerated, improving the acceleration performance of the guide vane.

[0018] In addition, the centripetal turbine of the present invention also has the above advantages.

[0019] In addition to the purposes, features and advantages described above, the present invention has other purposes, features and advantages. The present invention will be further described in detail below with reference to the drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] The drawings forming a part of this application are used to provide a further understanding of the present invention. The schematic embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation to the present invention. In the drawings:

[0021] Figure 1 is a schematic meridional flow passage diagram of a conventional centripetal turbine guide vane.

[0022] Figure 2 is a right view schematic diagram of a conventional centripetal turbine guide vane.

[0023] Figure 3 is a schematic flow passage design diagram of the radial air flow passage of the transonic guide vane applied to a centripetal turbine according to a preferred embodiment of the present invention.

[0024] Figure 4 It is another schematic diagram of the flow channel design of the radial air flow channel of the transonic guide vane applied to the centripetal turbine in the preferred embodiment of the present invention.

[0025] Figure 5 It is yet another schematic diagram of the flow channel design of the radial air flow channel of the transonic guide vane applied to the centripetal turbine in the preferred embodiment of the present invention.

[0026] Figure 6 It is a schematic diagram of the elemental airfoil of the conventional centripetal turbine guide vane.

[0027] Figure 7 It is a schematic diagram of the elemental airfoil of the guide vane and the cascade channel in the preferred embodiment of the present invention.

[0028] Figure 8 It is a schematic diagram of the maximum inscribed circle of the elemental airfoil of the guide vane in the preferred embodiment of the present invention.

[0029] Figure 9 It is a schematic diagram of the helmet-shaped leading edge of the guide vane in the preferred embodiment of the present invention.

[0030] Figure 10 It is a schematic diagram of the curvature of the suction side profile of the guide vane in the preferred embodiment of the present invention.

[0031] Figure 11 It is a schematic diagram of the comparison of the blade surface Mach numbers between the guide vane in the preferred embodiment of the present invention and the conventional guide vane. Detailed implementation manners

[0032] The following will describe the embodiments of the present invention in detail with reference to the accompanying drawings. However, the present invention can be implemented in many different ways defined and covered by the following.

[0033] Such as Figure 1 and Figure 2As shown in the figure, a conventional centripetal turbine guide vane includes a guide vane outer ring, a guide vane inner ring, and a plurality of guide vanes. The guide vane outer ring and the guide vane inner ring are symmetrically arranged around the engine axis. An annular air flow passage is formed between the guide vane outer ring and the guide vane inner ring. The plurality of guide vanes are evenly arranged in the annular air flow passage at the same angular intervals in the circumferential direction. The annular air flow passage includes an axial air flow passage and a radial air flow passage communicated therewith. Among them, the radial air flow passage is composed of the area where the guide vanes are located and the bladeless area. The profiles of the guide vane outer ring and the guide vane inner ring in the guide vane area are basically designed with straight lines, that is, the widths of the air flow passages in the guide vane area are basically the same. On the one hand, this causes a relatively large flow velocity when the air flow turns from the axial air flow passage into the inlet of the radial air flow passage, increasing the flow loss at the inlet of the guide vanes. On the other hand, since the air flow passage area of the guide vane is related to the width of the air flow passage and the radius of the air flow passage, and the radius of the radial air flow passage gradually decreases along the air flow direction, even if the width of the radial air flow passage remains unchanged, the air flow passage area of the radial air flow passage still decreases along the air flow direction. Therefore, when the air flow Mach number at the outlet of the guide vanes reaches the critical value, due to the reduction of the air flow passage area in the bladeless area, the air flow cannot continue to expand and accelerate after flowing out of the guide vanes, so that the velocity of the air flow when reaching the impeller blades will decrease and supersonic flow cannot be achieved. In addition, in the current design, the profile design of the guide vanes is not considered in coupling with the channel design. Only based on the acceleration ability of the cascade channel, there is a relatively complex wave system in the bladeless area, resulting in poor acceleration ability of the guide vane and large flow loss; moreover, it also greatly increases the complexity of blade manufacturing and inspection. In view of the above-mentioned disadvantages of the existing centripetal turbine guide vane channel and profile design, the present invention optimizes the design of the centripetal turbine guide vane, adopts a three-stage design structure and a coupling design with the guide vane profile to improve the acceleration performance of the guide vane, so as to achieve high supersonic flow of the air flow while reducing the flow loss and the manufacturing difficulty, and can well meet the requirements of a centripetal turbine with a large expansion ratio.

[0034] Specifically, as Figure 3 and Figure 4As shown in the figure, a preferred embodiment of the present invention provides a transonic guide vane applied to a centripetal turbine, which includes a guide vane outer ring, a guide vane inner ring and a plurality of guide vanes. The guide vane outer ring and the guide vane inner ring are symmetrically arranged around the engine axis. An annular air flow channel is formed between the guide vane outer ring and the guide vane inner ring. The annular air flow channel includes an axial air flow channel and a radial air flow channel communicating with it. The plurality of guide vanes are evenly arranged at the same angular interval in the circumferential direction in the radial air flow channel of the annular air flow channel. Among them, the radial air flow channel consists of three sections. Among them, the first section of the radial air flow channel extends from the inlet to a position 10% - 15% of the blade height from the leading edge of the blade (i.e., at point F), that is, the distance from the end of the first section of the radial air flow channel to the leading edge of the blade is 10% - 15% of the blade height. The second section of the radial air flow channel extends from the end of the first section of the radial air flow channel to the throat of the guide vane (i.e., at point T). The third section of the radial air flow channel extends from the end of the second section of the radial air flow channel to the end of the guide vane outer ring (i.e., at point E). The channel width of the first section of the radial air flow channel gradually decreases along the flow direction. The channel width of the second section of the radial air flow channel remains unchanged along the flow direction. The change of the channel width of the third section of the radial air flow channel is coupled and designed according to the air flow Mach number at the throat of the guide vane.

[0035] It can be understood that the channel width of the first section of the radial air flow channel gradually decreases along the flow direction. Coupled with the gradual decrease of the cascade channel width, the air flow passage area is further reduced along the air flow direction on the basis of the existing guide vane, greatly improving the air flow acceleration effect of the first section of the radial air flow channel. And it can also reduce the lap radius with the combustion chamber and reduce the overall outline size. More importantly, it can reduce the flow velocity of the air flow in the axial air flow channel when it enters the inlet of the radial air flow channel, greatly reducing the flow loss at the blade inlet. The channel width of the second section of the radial air flow channel remains unchanged. Due to the gradual decrease of the cascade channel width, the air flow can still be accelerated continuously, and it is beneficial to ensure that the influence of the guide vane outer ring processing quality and the radial position of the blade on the throat of the guide vane is small, which is convenient for inspecting the throat area of the guide vane. The change of the channel width of the third section of the radial air flow channel is coupled and designed according to the air flow Mach number at the outlet of the guide vane, effectively reducing the flow loss of the air flow at the outlet of the guide vane flowing through the vane-less area at different speeds. Whether the air flow Mach number at the outlet of the guide vane is subsonic or supersonic, the air flow at the outlet of the guide vane can be further accelerated, improving the acceleration performance of the guide vane.

[0036] Specifically, when the air flow velocity at the throat of the guide vane is subsonic, the channel width of the third-stage radial air flow channel gradually decreases or remains unchanged along the flow direction, so that the air flow passage area of the third-stage radial air flow channel gradually decreases along the flow direction, thereby continuously accelerating the air flow flowing out of the guide vane to achieve high subsonic flow; when the air flow at the throat of the guide vane reaches the critical point, the channel width of the third-stage radial air flow channel gradually increases along the flow direction, so that the air flow passage area of the third-stage radial air flow channel gradually increases along the flow direction, thereby, according to the principle of the Laval nozzle, continuously expanding and accelerating the air flow flowing out of the guide vane to achieve transonic flow.

[0037] It can be understood that in the first-stage radial air flow channel, the cross-sectional profile of the guide vane outer ring is a straight line, while the cross-sectional profile of the guide vane inner ring gradually expands from the end point of the first-stage radial air flow channel (i.e., point F) towards the inlet, so that the channel width of the first-stage radial air flow channel is always greater than the channel width b2 at point F along the air flow direction, and further the air flow passage area along the air flow direction gradually decreases. With this design, on the one hand, the lap radius with the combustion chamber can be reduced, and the overall outer contour size can be reduced; on the other hand, the flow velocity of the air flow in the axial air flow channel when entering the inlet of the radial air flow channel can be reduced, greatly reducing the flow loss at the blade inlet, and more importantly, greatly improving the acceleration effect of the air flow. Optionally, the ratio of the channel width b1 at the starting point of the first-stage radial air flow channel to the channel width b2 at the end point thereof is between 1.15 and 1.25, preferably 1.2.

[0038] In the second-stage radial air flow channel, the cross-sectional profiles of both the guide vane outer ring and the guide vane inner ring are straight lines, so that the channel widths are exactly the same, that is, b2 = b3, to ensure that the throat of the guide vane is less affected by the machining quality of the guide vane outer ring and the radial position of the blade, facilitating the control and inspection of the guide vane throat area. Although the channel width of the air flow channel remains unchanged here, as the cascade channel width gradually decreases from point F to point T, the air flow passage area of the second-stage radial air flow channel still gradually decreases, and the air flow can still be compressed and accelerated.

[0039] The third-stage radial air flow channel is the existing vane-less area channel, and the cross-sectional profiles of the guide vane inner ring and the guide vane outer ring here can adopt different change forms according to the Mach number of the air flow at the outlet of the guide vane. For example, Figure 3As shown, when the airflow Mach number at the outlet of the guide vane reaches or exceeds the critical Mach number, the profile of the outer ring section of the guide vane maintains the same straight shape as that of the second-stage radial airflow passage, while the inner ring of the guide vane expands outward starting from point T. The channel width of the third-stage radial airflow passage gradually increases along the flow direction, that is, b3 is less than b4, ensuring that the airflow passage area gradually increases along the airflow direction, that is, A3 < A4. According to calculations, when the airflow Mach number at the outlet of the guide vane is 1.1 and A4 / A3 = 1.12, the airflow Mach number at point E can reach 1.2. For example, as Figure 4 As shown, when the airflow Mach number at the outlet of the guide vane reaches or exceeds the critical Mach number, the profile of the inner ring section of the guide vane and the profile of the outer ring section of the guide vane are symmetrically distributed along the center line and expand outward along the airflow direction starting from point T respectively, so that the airflow passage area continuously increases along the airflow direction, that is, A3 < A4. For example, as Figure 5 As shown, when the airflow at the outlet of the guide vane is subsonic, that is, when the Mach number is lower than the critical Mach number, the profile of the outer ring section of the guide vane contracts inward starting from the corresponding position of point T, and the profile of the inner ring section of the guide vane expands outward starting from point T. However, the channel width of the third-stage radial airflow passage gradually decreases along the flow direction, that is, b3 > b4, and the airflow passage area of the third-stage radial airflow passage gradually decreases along the flow direction, so as to further accelerate the airflow flowing out of the guide vane, thereby achieving high subsonic flow. In addition, when the airflow Mach number at the outlet of the guide vane is subsonic, the profiles of the inner ring and the outer ring of the guide vane can both contract inward starting from point T, but the installation gap between the inner ring of the guide vane and the impeller needs to be ensured.

[0040] It can be understood that by including the airflow passage in the vane-less area within the radial airflow passage formed by the outer ring and the inner ring of the guide vane, the profiles of the outer ring and the inner ring of the guide vane corresponding to the vane-less area are coupled and designed according to the airflow Mach number at the outlet of the guide vane. Whether the airflow Mach number at the outlet of the guide vane is subsonic or supersonic, the airflow flowing out of the outlet of the guide vane can be further accelerated through the coupled design of the profiles of the outer ring and the inner ring of the guide vane, and the flow loss in the vane-less area is effectively reduced, improving the acceleration performance of the guide vane.

[0041] It can be understood that, as Figures 6 to 9As shown, the guide vane is stacked by no less than two two-dimensional elemental airfoils along the blade height direction according to a certain rule. For the elemental airfoil, it has the following characteristics: The two-dimensional elemental airfoil has a suction side, a pressure side, a leading edge, and a trailing edge that is approximately circular. The airfoil profile satisfies curvature continuity. The suction side, pressure side, and leading edge side profiles are generated by Bessel curves, or other forms of spline curves can also be used. A cascade channel is formed between the suction side and pressure side of the elemental airfoils of any two adjacent guide vanes at the same blade height. The width of the cascade channel gradually contracts along the airflow direction, that is, the airfoil throat is located at the trailing edge of the blade outlet. At the leading edge position on the pressure side of the blade, there is an incoming flow stagnation point P1. After the incoming flow impacts on P1, the airflow velocity stagnates to zero, and then the airflow is divided into two parts. One part of the airflow bypasses the leading edge on the pressure side and flows along the suction side towards the trailing edge of the suction side, and the other part of the airflow flows along the pressure side towards the blade outlet. Due to the reduction of the flow area, before reaching the throat position P2, the velocities of these two parts of the airflow will continuously increase. When the pressure ratio between the inlet and outlet of the guide vane reaches above the critical pressure ratio, when the airflow reaches P2, the airflow Mach number will reach 1.0. Since the area of the cascade channel expands after P2, therefore, the airflow will continue to accelerate to the supersonic state along the oblique cut section of the cascade channel and finally enter the above-mentioned vaneless area channel. According to the continuous expansion of the area of the vaneless area channel, further acceleration is achieved.

[0042] For a conventional centripetal turbine guide vane, to reduce the inlet airflow attack angle of the blade, the inlet construction angle β of the blade is selected to be the same as the incoming flow angle, which makes the bending degree of the suction side at the leading edge of the blade larger, and the suction side airfoil presents an "outward convex" style. Due to the tangential chord length b t of the airfoil and the axial chord length b x and the installation angle α approximately have the relationship: b t = b x *tanα. With the increase of the bending degree of the leading edge, on the premise that the axial chord length b x remains unchanged, it will inevitably lead to an increase in the installation angle α, thereby increasing the length of the tangential chord length b t , resulting in an increase in the overall size and weight of the guide vane. On the other hand, due to the increase in the bending degree of the curve between the starting point (q point) and the throat point (p2 point) of the suction side profile, the curvature is large, and the curve length is long, the acceleration of the fluid flowing around from point p1 is poor, resulting in a low load near the leading edge point of the conventional airfoil and failing to fully utilize the advantage of the longer chord length of the guide vane blade.

[0043] Therefore, the present invention adopts a leading edge design with a small radius of curvature, and the leading edge of the blade profile is generally in a "helmet shape", which reduces the bending degree of the profile line between point q and point p2, that is, reduces the blade profile curvature here, thereby improving the air flow acceleration at the leading edge. On the premise of keeping the chord length unchanged, the load near the leading edge is increased, and the tangential chord length of the blade back is effectively reduced, effectively reducing the size and weight of the guide vane. Specifically, the present invention redesigned the elemental blade profile, and the ratio of the maximum inscribed circle diameter between the blade back side and the blade basin side of the guide vane blade to the axial chord length b of the blade profile x is between 0.09 and 0.11. The ratio of the maximum inscribed circle diameter between the blade back side and the blade basin side of the guide vane blade to the trailing edge diameter is between 5 and 6. The distance L from the starting point q of the blade back side profile line of the guide vane blade to the center of the maximum inscribed circle qo and the axial chord length b of the blade profile x is between 0.15 and 0.28. The distance L from the end point of the blade back side profile line to the center of the maximum inscribed circle ho and the axial chord length b of the blade profile x is between 0.9 and 1. The included angle θ between the line connecting the starting point of the blade back side profile line and the center of the maximum inscribed circle and the line connecting the end point of the blade back side profile line and the center of the maximum inscribed circle is between 160° and 180°. The blade profile installation angle α is between 20° and 40°. And, the blade profile line behind the throat on the blade back side is nearly a straight line, and the blade profile curvature distribution on the blade back side is as Figure 10 shown. In the chord length region from 30% to 90%, the blade curvature is almost zero, that is, the blade profile is almost a straight line without any bending, so as to reduce the convexity of the blade profile as much as possible, thereby reducing the interference to the transonic flow behind the throat and reducing the flow loss. The comparison of the blade surface Mach numbers between the blade profile design of the present invention and the conventional blade profile design is as Figure 11 shown. The blade profile design of the present invention can accelerate the air flow passing through the guide vane to above the critical Mach number, and can make the air flow Mach number at the inlet of the rotor blade finally reach not less than 1.3.

[0044] It can be understood that the blade adopting the helmet-shaped leading edge design of the present invention can increase the load at the inlet section of the guide vane blade, greatly improve the air flow acceleration at the leading edge of the blade, further enhance the acceleration performance of the guide vane, and can also effectively reduce the tangential chord length of the blade back, which is beneficial to reducing the size and weight of the guide vane.

[0045] In addition, the present invention further provides a centripetal turbine, which adopts the transonic nozzle as described above. Among them, the centripetal turbine is preferably a centripetal turbine with a large expansion ratio. The high-temperature and high-pressure gas after combustion in the combustion chamber enters the air flow passage of the nozzle axially and then enters the nozzle blades radially. After three-stage acceleration in the radial air flow section, it enters the centripetal impeller to drive the impeller to do work, and completes the transformation from radial flow to axial flow, and finally flows out of the impeller.

[0046] The above are only the preferred embodiments of the present invention and are not intended to limit the present invention. For those skilled in the art, the present invention can have various changes and modifications. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A transonic guide vane applied to a centripetal turbine, characterized in that, It includes an outer guide vane ring, an inner guide vane ring and a plurality of guide vanes. The outer guide vane ring and the inner guide vane ring are symmetrically arranged around the engine axis. An annular air flow channel is formed between the outer guide vane ring and the inner guide vane ring. The plurality of guide vanes are evenly arranged in the annular air flow channel at the same angular interval in the circumferential direction. The annular air flow channel includes an axial air flow channel and a radial air flow channel communicating therewith. The radial air flow channel consists of three sections. Among them, the first section of the radial air flow channel extends from the inlet to a position 10% - 15% of the blade height from the leading edge of the blade. The second section of the radial air flow channel extends from the end point of the first section of the radial air flow channel to the throat of the guide vane. The third section of the radial air flow channel extends from the end point of the second section of the radial air flow channel to the end point of the outer guide vane ring. The channel width of the first section of the radial air flow channel gradually decreases in the flow direction. The channel width of the second section of the radial air flow channel remains unchanged in the flow direction. The channel width change of the third section of the radial air flow channel is coupledly designed according to the air flow Mach number at the outlet of the guide vane.

2. The transonic guide vane applied to a centripetal turbine according to claim 1, characterized in that, In the first section of the radial air flow channel, the cross-sectional profile of the outer guide vane ring is a straight line, and the cross-sectional profile of the inner guide vane ring gradually expands from the end point of the first section of the radial air flow channel towards the inlet. In the second section of the radial air flow channel, the cross-sectional profiles of the outer guide vane ring and the inner guide vane ring are both straight lines.

3. The transonic guide vane applied to a centripetal turbine according to claim 2, wherein The ratio of the channel width at the starting point of the first section of the radial air flow channel to the channel width at the end point thereof is between 1.15 and 1.

25.

4. The transonic guide vane applied to a centripetal turbine according to claim 1, wherein, When the air flow velocity at the outlet of the guide vane is subsonic, the channel width of the third section of the radial air flow channel gradually decreases or remains unchanged in the flow direction, so that the air flow passage area of the third section of the radial air flow channel gradually decreases in the flow direction; when the air flow velocity at the outlet of the guide vane is transonic, the channel width of the third section of the radial air flow channel gradually increases in the flow direction, so that the air flow passage area of the third section of the radial air flow channel gradually increases in the flow direction.

5. The transonic guide vane applied to a centripetal turbine according to claim 1, characterized in that The leading edge of the guide vane adopts a helmet-shaped design to reduce the blade profile curvature between the starting point and the throat point of the suction side profile.

6. The transonic guide vane applied to a centripetal turbine according to claim 5, wherein The ratio of the maximum inscribed circle diameter between the suction side and the pressure side of the guide vane to the axial chord length of the blade profile is between 0.09 and 0.

11.

7. The transonic guide vane applied to a centripetal turbine according to claim 6, characterized in that, The ratio of the maximum inscribed circle diameter between the suction side and the pressure side of the guide vane to the trailing edge diameter is between 5 and 6.

8. The transonic guide vane applied to a centripetal turbine according to claim 7, wherein, The ratio of the distance from the starting point of the suction side profile of the guide vane to the center of the maximum inscribed circle to the axial chord length of the blade profile is between 0.15 and 0.28, and the ratio of the distance from the ending point of the suction side profile to the center of the maximum inscribed circle to the axial chord length of the blade profile is between 0.9 and 1.

9. The transonic guide vane applied to a centripetal turbine according to claim 8, wherein The angle between the line connecting the starting point of the suction side profile and the center of the maximum inscribed circle and the line connecting the ending point of the suction side profile and the center of the maximum inscribed circle is between 160° and 180°.

10. A centripetal turbine, characterized in that, Adopt the transonic guide vane according to any one of claims 1 - 9.

Images