Turbo nozzle, expander and air cycle machine

Abstract

The present application provides a turbine nozzle, an expander and an air cycle machine. The turbine nozzle includes a wheel disc (1), one end of the wheel disc (1) is provided with a convex ring (2), and nozzle vanes (3) are circumferentially arranged on the end face of the convex ring (2). In the airfoil structure of the nozzle vanes (3), the camber line curvature of the first 50% of the meridional chord length near the leading edge is less than the camber line curvature of the last 50% of the meridional chord length far from the leading edge of the nozzle vanes (3), and the camber line curvature change rate of the first 50% of the meridional chord length near the leading edge is less than the camber line curvature change rate of the last 50% of the meridional chord length far from the leading edge of the nozzle vanes (3). The camber line curvature of the last 50% of the meridional chord length far from the leading edge of the nozzle vanes (3) gradually increases. According to the turbine nozzle of the present application, the shock boundary layer interference loss can be reduced, the efficiency of the turbine nozzle can be improved, the enthalpy difference of the refrigeration system can be increased, and the system entropy increase can be effectively reduced.

Description

Technical Field

[0001] This application relates to the technical field of aircraft air conditioners, and particularly to a turbine nozzle, an expander and an air cycle machine. Background Art

[0002] Currently, the power of the mainstream aircraft cabin air conditioner comes from the bleed air of the aircraft engine. The high-temperature and high-pressure gas led out by the engine enters the refrigeration pack unit. After driving the expansion wheel, it drives the compression wheel and the ram fan to operate. The expansion wheel is mainly used to provide power and generate a low-temperature air flow. The compression wheel is mainly used to increase the pressure of the low-pressure air outside the aircraft at high altitude, and at the same time, there is a temperature increase effect. The main purpose of the ram fan is to perform heat exchange for the heat exchanger.

[0003] The centripetal turbine nozzle in the aircraft air cycle machine is used to accelerate the air flow at the turbine inlet to reach the transonic state, so as to drive the turbine rotor to rotate, and then output power to drive the compressor and the fan to rotate. Since the air flow on the blade surface of the transonic nozzle needs to be accelerated to supersonic speed, shock waves are easily generated and shock wave boundary layer interference occurs, which leads to local flow separation, increases the flow loss, causes a large increase in the entropy of the aircraft environmental control system, weakens the refrigeration effect, and at the same time increases the fuel compensation loss of the aircraft. Summary of the Invention

[0004] Therefore, the technical problem to be solved by this application is to provide a turbine nozzle, an expander and an air cycle machine, which can reduce the shock wave boundary layer interference loss, improve the efficiency of the turbine nozzle, increase the enthalpy difference of the refrigeration system, and effectively reduce the system entropy increase.

[0005] To solve the above problems, this application provides a turbine nozzle, including a wheel disc. A convex ring is provided at one end of the wheel disc. Nozzle blades are circumferentially arranged on the end surface of the convex ring. In the airfoil structure of the nozzle blades, the curvature of the mean camber line of the first 50% of the meridional chord length near the leading edge is less than the curvature of the mean camber line of the last 50% of the meridional chord length far from the leading edge of the nozzle blade, and the curvature change rate of the mean camber line of the first 50% of the meridional chord length near the leading edge is less than the curvature change rate of the mean camber line of the last 50% of the meridional chord length far from the leading edge of the nozzle blade. The curvature of the mean camber line of the last 50% of the meridional chord length far from the leading edge of the nozzle blade gradually increases.

[0006] Preferably, the curvature C of the mean camber line of the first 50% of the meridional chord length near the leading edge v1 ≤0, and the curvature C of the mean camber line of the last 50% of the meridional chord length far from the leading edge 0≤C v2 ≤3.8462m -1 .

[0007] Preferably, along the direction from the leading edge to the trailing edge, the mean camber line of the nozzle blade is concave towards the center of the wheel disc.

[0008] Preferably, the outer ring radius of the convex ring is R 1, the inner ring radius is R 2 , R 1 / R 2 = 1.1394 to 1.5416.

[0009] Preferably, the roulette is further provided with a relief hole, the relief hole is located on the inner circumferential side of the convex ring, an installation ring is formed between the relief hole and the convex ring, and the radius of the relief hole is R 3 , the inner ring radius of the convex ring is R 2 , R 3 / R 2 = 0.7393 to 0.9035.

[0010] Preferably, the axial height of the nozzle vane is H, and the inner ring radius of the convex ring is R 2 , H / R 2 = 0.058 to 0.066.

[0011] Preferably, taking the intersection point of the leading edge and the mean camber line as the origin, the meridian direction as the x-axis, and the circumferential direction as the y-axis to establish a coordinate system, the meridian direction is the radial direction passing through the center O of the roulette, and the circumferential direction is the tangential direction around the center O of the roulette. The non-dimensional distribution function of the mean camber line in the meridian and circumferential coordinate systems is:

[0012] θ / C m = -1.14682×10 -4 + 0.01282(m / C m ) - 1.10957×10 -4 (m / C m ) 2 + 1.24871×10 -6 (m / C m ) 3 + 2.20646×10 -9 (m / C m ) 4

[0013] where θ is the tangential direction coordinate around the O axis, C m is the component of the chord length of the nozzle vane in the meridian direction, and m is the meridian direction coordinate.

[0014] Preferably, taking the intersection point of the leading edge and the mean camber line as the origin, the meridian direction as the x-axis, and the thickness of the nozzle vane as the y-axis to establish a coordinate system, the meridian direction is the radial direction passing through the center O of the roulette, and the non-dimensional thickness distribution function of the nozzle vane is:

[0015] t / C = 2.0×[0.04281 + 0.00347(m / C m ) - 6.28517×10 -5 (m / C m )2 +2.43263×10 -7 (m / C m ) 3

[0016] where t is the nozzle thickness, C is the chord length of the nozzle vane, C m is the component of the chord length of the nozzle vane in the meridian direction, and m is the coordinate in the meridian direction.

[0017] Preferably, the maximum thickness TH of the nozzle vane is 16% - 20% of the chord length of the nozzle vane, the position TP where the maximum thickness of the nozzle vane is located is at the position of 30% - 40% of the chord length from the intersection of the leading edge and the mean camber line, and the maximum thickness LeH of the leading edge is 8% - 10% of the chord length of the nozzle vane.

[0018] According to another aspect of the present application, a turbine expander is provided, including a turbine nozzle, and the turbine nozzle is the above-mentioned turbine nozzle.

[0019] Preferably, the turbine expander further includes a sealing ring and a bearing housing. The sealing ring is installed on the bearing housing, and the turbine nozzle is fixedly installed on the sealing ring.

[0020] According to another aspect of the present application, an air cycle machine is provided, including the above-mentioned turbine nozzle or the above-mentioned turbine expander.

[0021] The turbine nozzle provided by the present application includes a wheel disc. A convex ring is provided at one end of the wheel disc. Nozzle vanes are circumferentially arranged on the end surface of the convex ring. In the airfoil structure of the nozzle vanes, the curvature of the mean camber line of the first 50% of the meridian chord length near the leading edge is less than the curvature of the mean camber line of the last 50% of the meridian chord length far from the leading edge of the nozzle vanes, the rate of change of the curvature of the mean camber line of the first 50% of the meridian chord length near the leading edge is less than the rate of change of the curvature of the mean camber line of the last 50% of the meridian chord length far from the leading edge of the nozzle vanes, and the curvature of the mean camber line of the last 50% of the meridian chord length far from the leading edge of the nozzle vanes gradually increases. By optimizing the structure of the mean camber line, the turbine nozzle can make the nozzle vanes form a post-loading characteristic, with basically no loading or even negative loading before 50% of the chord length of the nozzle, and gradually increasing deflection after 50% of the chord length, so that the first half of the nozzle vanes with a smaller rate of change of curvature and a smaller curvature effectively control the acceleration process of the airflow in the nozzle, while the second half of the nozzle vanes with a larger rate of change of curvature and a larger curvature inhibit the sudden deceleration of the high-speed airflow, thereby eliminating the flow field structure with shock wave organization that is easily generated by conventional nozzles, reducing the shock wave boundary layer interference loss, improving the efficiency of the turbine nozzle, increasing the enthalpy difference of the refrigeration system, and effectively reducing the entropy increase of the system. Description of the Drawings

[0022] Figure 1 is a three-dimensional structural schematic diagram of the turbine nozzle according to an embodiment of the present application;

[0023] Figure 2 Structural schematic diagram of a turbine nozzle according to an embodiment of the present application;

[0024] Figure 3 Cross-sectional structural schematic diagram of the turbine nozzle in the A-A direction according to an embodiment of the present application;

[0025] Figure 4 Coordinate structural schematic diagram of the turbine nozzle according to an embodiment of the present application;

[0026] Figure 5 Structural diagram of the mid-arc distribution of the nozzle blades of the turbine nozzle according to an embodiment of the present application;

[0027] Figure 6 Structural diagram of the thickness distribution of the nozzle blades of the turbine nozzle according to an embodiment of the present application;

[0028] Figure 7 Three-dimensional structural schematic diagram of an air circulation machine according to an embodiment of the present application;

[0029] Figure 8 Partial cross-sectional structural schematic diagram of the air circulation machine according to an embodiment of the present application;

[0030] Figure 9 Enthalpy-entropy diagram of a shock nozzle and a non-shock nozzle;

[0031] Figure 10 Shock wave organizational structure of the nozzle flow field in the related art;

[0032] Figure 11 Shock wave organizational structure of the nozzle flow field according to an embodiment of the present application;

[0033] Figure 12 Structural diagram of the Mach number distribution of the nozzle elementary flow surface according to an embodiment of the present application;

[0034] Figure 13 Structural diagram of the entropy distribution of the nozzle spanwise flow surface according to an embodiment of the present application.

[0035] The reference numerals are represented as:

[0036] 1. Turbine disk; 2. Convex ring; 3. Nozzle blade; 4. Relief hole; 5. Sealing ring; 6. Bearing housing; 7. Fan; 8. Compressor; 9. Expander. Detailed implementation manners

[0037] Refer to in combination Figures 1 to 6As shown, according to an embodiment of the present application, the turbine nozzle includes a wheel disc 1. A convex ring 2 is provided at one end of the wheel disc 1. Nozzle vanes 3 are circumferentially arranged on the end face of the convex ring 2. In the airfoil structure of the nozzle vanes 3, the curvature of the mean camber line of the first 50% of the meridional chord length near the leading edge is less than the curvature of the mean camber line of the last 50% of the meridional chord length away from the leading edge of the nozzle vanes 3. The rate of change of the curvature of the mean camber line of the first 50% of the meridional chord length near the leading edge is less than the rate of change of the curvature of the mean camber line of the last 50% of the meridional chord length away from the leading edge of the nozzle vanes 3. The curvature of the mean camber line of the last 50% of the meridional chord length away from the leading edge of the nozzle vanes 3 gradually increases.

[0038] By optimizing the structure of the mean camber line, the turbine nozzle can enable the nozzle vanes 3 to form a post-loading characteristic. There is basically no loading or even negative loading before 50% of the chord length, and the turning gradually increases after 50% of the chord length. The first half of the nozzle vanes 3 with a relatively small rate of change of curvature and a relatively small curvature of the mean camber line effectively controls the acceleration process of the air flow in the nozzle, while the second half of the nozzle vanes 3 with a relatively large rate of change of curvature and a relatively large curvature of the mean camber line suppresses the sudden deceleration of the high-speed air flow, thereby eliminating the flow field structure with shock waves that is prone to occur in conventional nozzles, reducing the shock wave boundary layer interference loss, improving the efficiency of the turbine nozzle, increasing the enthalpy difference of the refrigeration system, and effectively reducing the system entropy increase.

[0039] In one embodiment, the curvature C of the mean camber line of the first 50% of the meridional chord length near the leading edge v1 ≤0, and the curvature 0 ≤ C of the mean camber line of the last 50% of the meridional chord length away from the leading edge v2 ≤3.8462 m -1 . By limiting the curvature of the mean camber line of the first 50% of the meridional chord length near the leading edge and the curvature of the mean camber line of the last 50% of the meridional chord length away from the leading edge, a structure can be formed in which the curvature of the first half of the mean camber line of the nozzle vanes is close to zero and the curvature of the second half gradually increases, thereby using this mean camber line structure to optimize and limit the airfoil structure, reducing or avoiding the generation of shock waves, and improving the efficiency of the turbine nozzle.

[0040] In one embodiment, along the direction from the leading edge to the trailing edge, the mean camber line of the nozzle vanes 3 is concave towards the center of the wheel disc 1, which can define the turning structure of the nozzle vanes, so that the turning of the nozzle vanes mainly occurs in the area where shock waves are likely to form, effectively eliminating the shock waves on the blade surface. The air flow decelerates from supersonic to subsonic gradually without generating shock waves, thereby constructing a shock-wave-free flow field organizational structure, reducing the nozzle loss, and improving the efficiency of the environmental control system.

[0041] In one embodiment, the outer ring radius of the convex ring 2 is R 1 , and the inner ring radius is R 2 , R 1 / R 2 = 1.1394 - 1.5416. Preferably, R 1 / R 2 The optimal value of it is 1.3405. This value range can ensure the rationality of the turbine nozzle assembly structure and is conducive to the mating assembly with the turbine volute. When the optimal value is selected, it can guide the airflow to accelerate smoothly in the fluid channel formed by the nozzle vane 3, reduce the supersonic region, and eliminate the possible shock wave loss.

[0042] In one embodiment, the disk 1 is further provided with a relief hole 4. The relief hole 4 is located on the inner circumferential side of the convex ring 2. An installation ring is formed between the relief hole 4 and the convex ring 2. The radius of the relief hole 4 is R 3 , and the inner ring radius of the convex ring 2 is R 2 , R 3 / R 2 = 0.7393 - 0.9035, and the optimal value is 0.8214. This value range can ensure the rationality of the turbine nozzle assembly structure and provide space for the installation of the downstream turbine rotor. When the optimal value is selected, it can ensure that the clearance between the static and moving components is minimized without geometric interference between the nozzle vane 3 and the turbine rotor, thereby reducing leakage.

[0043] In one embodiment, the axial height of the nozzle vane 3 is H, and the inner ring radius of the convex ring 2 is R 2 , H / R 2 = 0.058 - 0.066, and the optimal value is 0.0619. The above ratio limits the construction of the three-dimensional flow channel of the nozzle and can ensure that the nozzle design meets the predetermined flow rate requirements.

[0044] In one embodiment, taking the intersection point of the leading edge and the mean camber line as the origin, taking the meridian direction as the x-axis, and taking the circumferential direction as the y-axis to establish a coordinate system. The meridian direction is the radial direction passing through the center O of the disk 1, and the circumferential direction is the tangential direction around the center O of the disk 1. The dimensionless distribution function of the mean camber line in the meridian and circumferential coordinate systems is:

[0045] θ / C m = -1.14682×10 -4 + 0.01282(m / C m ) - 1.10957×10 -4 (m / C m ) 2 + 1.24871×10 -6 (m / C m ) 3 + 2.20646×10 -9 (m / C m ) 4

[0046] where θ is the tangential direction coordinate around the O axis, C m is the component of the chord length of the nozzle vane 3 in the meridian direction, and m is the meridian direction coordinate.

[0047] In one embodiment, taking the intersection point of the leading edge and the mean camber line as the origin, the meridian direction as the x-axis, and the thickness of the nozzle vane 3 as the y-axis to establish a coordinate system, the meridian direction is the radial direction passing through the center O of the disk 1, and the dimensionless thickness distribution function of the nozzle vane 3 is as follows:

[0048] t / C = 2.0×[0.04281 + 0.00347(m / C m ) - 6.28517×10 -5 (m / C m ) 2 + 2.43263×10 -7 (m / C m ) 3

[0049] where t is the nozzle thickness, C is the chord length of the nozzle vane 3, C m is the component of the chord length of the nozzle vane 3 in the meridian direction, and m is the coordinate in the meridian direction.

[0050] The nozzle vane shaping method provided in this application is to first determine the coordinates of the mean camber line CL, and then symmetrically superimpose the thickness t in the normal direction of the mean camber line to form the upper surface U and the lower surface L of the nozzle vane. In the axial direction perpendicular to the circular plane passing through the center O, the nozzle vane 3 is obtained by stretching the height. Other nozzle vanes are generated by performing a rotational array around O. In one embodiment, there are a total of 19 vanes in the circumferential direction.

[0051] The turbine nozzle of this application is generated by superimposing the mean camber line distribution of the nozzle vanes in Appendix Figure 5 and the nozzle vane thickness distribution in Appendix Figure 6 .

[0052] The dimensionless distribution function of the above-mentioned mean camber line in the meridian and circumferential coordinate systems and the dimensionless thickness distribution function of the nozzle vane 3 can control the air flow acceleration and deceleration process in adjacent vane channels, thereby eliminating shock waves and reducing flow losses.

[0053] In one embodiment, the maximum thickness TH of the nozzle vane 3 is 16% - 20% of the chord length of the nozzle vane 3, the position TP where the maximum thickness of the nozzle vane 3 is located is at the position of 30% - 40% of the chord length from the intersection point of the leading edge and the mean camber line, and the maximum thickness LeH of the leading edge is 8% - 10% of the chord length of the nozzle vane 3. This distribution is to form a "fat head type" nozzle vane 3 to adapt to a wider range of incoming air flow angle conditions. On the other hand, this thickness distribution is also to form a specified throat area under the specified number of vanes, so as to accelerate the air flow to the required speed.

[0054] For the nozzle vane 3 provided in this application, the mid-arc turning is mainly concentrated in the middle and rear sections, so its aerodynamic load is also concentrated in this area. Due to the rear-loading characteristic of the nozzle vane 3 and the fact that the mid-arc of the nozzle vane 3 is concave inward towards the center of the disk 1 in the circumferential direction, the turning of the nozzle vane 3 mainly occurs after 50% of the chord length. This form of load distribution has the effect of reducing or even eliminating the shock wave at the nozzle outlet, reducing aerodynamic losses, and improving turbine efficiency.

[0055] According to Figure 9 For the nozzle flow process 1 to 2sw with shock waves and the nozzle flow process 1 to 2fs without shock waves in [reference], it can be deduced that the efficiency of the nozzle without shock waves is higher than that of the nozzle with shock waves, that is, Hfs / Hs > Hsw / Hs. Obviously, the entropy increase of the nozzle without shock waves is also less than that of the nozzle with shock waves, that is, ΔSfs < ΔSsw. Therefore, the nozzle without shock waves has the beneficial effects of improving the air cycle efficiency, increasing the nozzle enthalpy difference, reducing the system entropy increase, improving the refrigeration capacity of the environmental control system, and reducing the fuel compensation loss.

[0056] According to Figure 10 The nozzle in the related technology and Figure 11 From the comparison diagram of the shock wave organizational structure of the flow field of the nozzle in the embodiment of this application, it can be seen that the shock-wave-free nozzle provided in this application has a rear-loading characteristic and is concave inward in the circumferential direction. The turning of the nozzle vane 3 mainly occurs after 50% of the chord length. This structure is more likely to induce a shock-wave-free flow field structure.

[0057] Figure 12 and Figure 13 The flow field numerical simulation results of [reference] show that the shock-wave-free nozzle in the embodiment of this application effectively eliminates the shock wave on the blade surface, and the air flow gradually decelerates from supersonic to subsonic without generating a shock wave, thereby constructing a shock-wave-free flow field organizational structure, reducing the nozzle loss, and improving the efficiency of the environmental control system.

[0058] Combined with reference to Figure 7 and Figure 8 As shown, according to the embodiment of this application, the expander includes a turbine nozzle, and the turbine nozzle is the above-mentioned turbine nozzle.

[0059] The expander further includes a sealing ring 5 and a bearing seat 6. The sealing ring 5 is installed on the bearing seat 6, and the turbine nozzle is fixedly installed on the sealing ring 5. In this embodiment, the turbine nozzle is fixedly connected to the sealing ring 5 through eight circumferential fixing holes, axially positioned with the sealing ring 5 through the first axial positioning surface, and axially positioned with the bearing seat 6 through the second axial positioning surface. The first axial positioning surface and the second axial positioning surface are not in the same plane, and the first axial positioning surface protrudes from the second axial positioning surface. The turbine nozzle also has a radial positioning surface, and the turbine nozzle is radially positioned and matched with the bearing seat 6 through the radial positioning surface.

[0060] Combined with reference to Figure 7and Figure 8 As shown, according to an embodiment of the present application, the air circulation machine includes the above-mentioned turbine nozzle or the above-mentioned expander 9.

[0061] The air circulation machine further includes a fan 7 and a compressor 8. The fan 7, the compressor 8 and the expander 9 are located on the same rotating shaft. The airflow does work on the expander 9, causing the expander 9 to drive the rotating shaft to rotate, and then driving the fan 7 and the compressor 8 to work through the rotating shaft.

[0062] The turbine inlet is connected to the high-pressure airflow. After flowing through the nozzle blades 3 under the guiding action of the turbine volute and accelerating to the transonic state, it drives the turbine rotor to do work. The temperature of the airflow drops, and after passing through the turbine outlet and accessing the mixing chamber for temperature regulation, it is sent into the cabin.

[0063] It is easy for those skilled in the art to understand that, on the premise of no conflict, the above-mentioned various advantageous manners can be freely combined and superimposed.

[0064] The above are only the preferred embodiments of the present application and are not intended to limit the present application. Any modifications, equivalent replacements, and improvements made within the spirit and principle of the present application shall be included in the protection scope of the present application. The above is only the preferred implementation manner of the present application. It should be noted that for those of ordinary skill in the art, without departing from the technical principle of the present application, several improvements and variations can be made, and these improvements and variations should also be regarded as the protection scope of the present application.

Claims

1. A turbine nozzle, characterized in that, it includes a wheel disc (1), one end of the wheel disc (1) is provided with a convex ring (2), nozzle vanes (3) are circumferentially arranged on the end face of the convex ring (2), in the airfoil structure of the nozzle vanes (3), the curvature of the mean camber line of the first 50% of the meridional chord length close to the leading edge is less than the curvature of the mean camber line of the last 50% of the meridional chord length far from the leading edge of the nozzle vanes (3), the rate of change of the curvature of the mean camber line of the first 50% of the meridional chord length close to the leading edge is less than the rate of change of the curvature of the mean camber line of the last 50% of the meridional chord length far from the leading edge of the nozzle vanes (3), and the curvature of the mean camber line of the last 50% of the meridional chord length far from the leading edge of the nozzle vanes (3) gradually increases.

2. The turbine nozzle according to claim 1, characterized in that, The camber line curvature C of the first 50% of the meridional chord length near the leading edge v1 ≤0, and the camber line curvature of the last 50% of the meridional chord length away from the leading edge satisfies 0 ≤ C v2 ≤3.8462 m -1 .

3. The turbine nozzle according to claim 1, characterized in that, Along the direction from the leading edge to the trailing edge, the mean camber line of the nozzle vanes (3) is concave towards the center of the wheel disc (1).

4. The turbine nozzle according to claim 1, characterized in that, The outer ring radius of the convex ring (2) is R 1 , and the inner ring radius is R 2 , R 1 / R 2 = 1.1394 to 1.5416 5. The turbine nozzle according to claim 1, characterized in that, The roulette wheel (1) is further provided with a relief hole (4), the relief hole (4) is located on the inner circumferential side of the convex ring (2), an installation ring is formed between the relief hole (4) and the convex ring (2), and the radius of the relief hole (4) is R 3 , the inner ring radius of the convex ring (2) is R 2 , R 3 / R 2 = 0.7393 to 0.9035.

6. The turbine nozzle according to claim 1, characterized in that, The axial height of the nozzle vane (3) is H, and the inner ring radius of the convex ring (2) is R 2 , H / R 2 = 0.058 to 0.

066.

7. The turbine nozzle according to claim 1, characterized in that, Taking the intersection point of the leading edge and the mean camber line as the origin, taking the meridional direction as the x-axis, and taking the circumferential direction as the y-axis to establish a coordinate system, the meridional direction is the radial direction passing through the center O of the wheel disc (1), and the circumferential direction is the tangential direction around the center O of the wheel disc (1). The dimensionless distribution function of the mean camber line in the meridional and circumferential coordinate systems is: θ / C m = -1.14682×10 -4 +0.01282(m / C m ) - 1.10957×10 -4 (m / C m ) 2 +1.24871×10 -6 (m / C m ) 3 +2.20646×10 -9 (m / C m ) 4 where θ is the tangential direction coordinate about the O axis, and C m is the component of the chord length of the nozzle vane (3) in the meridional direction, and m is the meridional direction coordinate.

8. The turbine nozzle according to claim 1, characterized in that, Taking the intersection point of the leading edge and the mean camber line as the origin, taking the meridional direction as the x-axis, and taking the thickness of the nozzle vanes (3) as the y-axis to establish a coordinate system, the meridional direction is the radial direction passing through the center O of the wheel disc (1). The dimensionless thickness distribution function of the nozzle vanes (3) is: t / C = 2.0×[0.04281 + 0.00347(m / C m ) - 6.28517×10 -5 (m / C m ) 2 + 2.43263×10 -7 (m / C m ) 3 where t is the nozzle thickness, C is the chord length of the nozzle vane (3), and C m is the component of the chord length of the nozzle vane (3) in the meridian direction, and m is the meridian direction coordinate.

9. The turbine nozzle according to claim 1, characterized in that, The maximum thickness TH of the nozzle vanes (3) is 16% - 20% of the chord length of the nozzle vanes (3), the position TP where the maximum thickness of the nozzle vanes (3) is located is at a position 30% - 40% of the chord length away from the intersection point of the leading edge and the mean camber line, and the maximum thickness LeH of the leading edge is 8% - 10% of the chord length of the nozzle vanes (3).

10. An expander, including a turbine nozzle, characterized in that, the turbine nozzle is the turbine nozzle according to any one of claims 1 to 9.

11. The expander according to claim 10, characterized in that, the expander further includes a sealing ring (5) and a bearing seat (6), the sealing ring (5) is installed on the bearing seat (6), and the turbine nozzle is fixedly installed on the sealing ring (5).

12. An air cycle machine, characterized in that, it includes the turbine nozzle according to any one of claims 1 to 9 or the expander according to any one of claims 10 to 11.

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