Turbine rotor, expander and air cycle machine

Abstract

The present application provides a turbine rotor, an expander and an air cycle machine. The turbine rotor includes a disk (1) and blades arranged circumferentially along the disk (1). The flow attack angle at the leading edge of the blade is 0°. Wherein, the inlet flow angle α of the turbine rotor is the included angle between the flow direction FD and the axial direction of the turbine rotor, and the leading edge blade angle β of the turbine rotor is the included angle between the tangent line of the leading edge of the mean camber line of the blade profile and the axial direction of the turbine rotor. The flow attack angle i = α - β. According to the turbine rotor of the present application, the problem of large separated flow generated by the impact of the flow on the leading edge of the rotor can be solved, thereby reducing the flow loss of the flow and improving the efficiency of the turbine rotor.

Description

Technical Field

[0001] This application relates to the technical field of aircraft air conditioners, and particularly to a turbine rotor, 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 from 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 rotor in the aircraft air cycle machine is used to receive the transonic air flow from the nozzle. By deflecting the air flow in the radial direction into an axial subsonic air flow, the gas pressure and temperature are rapidly reduced, and at the same time, power is output to drive the compressor and the fan to rotate. The leading edge of the centripetal turbine rotor is often machined for easy processing, and it is prone to leading-edge flow separation under the impact of the air flow at a certain angle of attack, and even expands into a large separation structure in the channel, thus significantly increasing the flow loss. Summary of the Invention

[0004] Therefore, the technical problem to be solved by this application is to provide a turbine rotor, an expander, and an air cycle machine, which can solve the problem of large separation flow generated by the air flow impact on the rotor leading edge, thereby reducing the air flow loss and improving the efficiency of the turbine rotor.

[0005] To solve the above problems, this application provides a turbine rotor, including a wheel disc and blades arranged circumferentially along the wheel disc. The angle of attack of the air flow at the leading edge of the blade is 0°. Wherein, the inlet air flow angle α of the turbine rotor is the included angle between the air flow direction FD and the axis of the turbine rotor, and the leading edge blade angle β of the turbine rotor is the included angle between the tangent of the leading edge of the mean camber line of the blade profile and the axis of the turbine rotor. The angle of attack of the air flow i = α - β.

[0006] Preferably, by rotating the meridian streamline around the central axis O of the turbine rotor, the blade intersection line JL is formed by intersecting with the blade. The airfoil of the blade is obtained by unfolding the intersection line JL along the meridian coordinate m and the circumferential coordinate θ. Taking the meridian coordinate m as the x-axis, the circumferential coordinate θ as the y-axis, and the intersection point of the connection line of the two end points of the leading edge of the airfoil and the mean camber line of the airfoil as the center, the airfoil coordinate system of the blade is established.

[0007] Preferably, the blade includes a first blade and a second blade. The blade heights of the first blade and the second blade are the same. The axial length of the first blade is greater than the axial length of the second blade. The first blade and the second blade are alternately arranged circumferentially along the wheel disc.

[0008] Preferably, the leading edges of the first blade and the second blade are aligned.

[0009] Preferably, the first blade and the second blade are integrally machined on the same blank.

[0010] Preferably, in the airfoil coordinate system of the blade, the mean camber line of the airfoil has a distribution function in the meridional direction and the circumferential coordinate system that is a Boltzmann function. The intersection surface of the first blade and the disk is HU1, the tip surface of the first blade is SH1, the intersection surface of the second blade and the disk is HU2, and the tip surface of the second blade is SH2. Among them,

[0011] The distribution function of the mean camber line of HU1 in the meridional direction and the circumferential coordinate system is:

[0012] θ = 1.08076 - 1.08736 / (1 + e (m-2.99171) / 0.42001 )

[0013] The distribution function of the mean camber line of SH1 in the meridional direction and the circumferential coordinate system is:

[0014] θ = 4.11889 - 4.12590 / (1 + e (m-1.29564) / 0.18631 )

[0015] The distribution function of the mean camber line of HU2 in the meridional direction and the circumferential coordinate system is:

[0016] θ = 0.43835 - 0.09317 / (1 + e (m-1.76698) / 0.15371 )

[0017] The distribution function of the mean camber line of SH2 in the meridional direction and the circumferential coordinate system is:

[0018] θ = 0.88796 - 0.54092 / (1 + e (m-0.89473) / 0.13378 )

[0019] Preferably, in the airfoil coordinate system of the blade, the airfoil thickness of the first blade satisfies:

[0020] t = 2×[0.67564 + 0.01378(m / C m ) - 2.18765×10 -4 (m / C m ) 2 + 6.78833×10 -7 (m / C m ) 3 )

[0021] The airfoil thickness of the second blade satisfies:

[0022] t = 2×[0.68432 - 0.00101(m / C m) - 2.71351×10 -4 (m / C m ) 2 - 2.75724×10 -7 (m / C m ) 3 )

[0023] Where Cm is the component of the blade chord length in the meridional direction.

[0024] Preferably, in the airfoil coordinate system of the blade, the maximum thickness TH of the first blade is 1.6 - 2.0 mm, and the maximum thickness position TP is located at the 40% - 50% chord length position. The leading edge thickness t of the first blade LE is 0.6 - 0.7 mm, and the trailing edge thickness t TE is 0.55 - 0.65 mm.

[0025] Preferably, the flow passage of the turbine rotor is composed of a disk profile and a tip profile. The disk profile is formed by rotating the intersection surface HU1 of the first blade and the disk around the O axis and mapping it to the meridional plane. The disk profile includes a straight line segment HL1 and an arc segment HL2. The diameter of the disk is DI, and the radius of the arc segment HL2 is R HL2 , R HL2 / DI = 0.3089 - 0.4089; The tip profile is formed by rotating the tip surface of the first blade around the O axis and mapping it to the meridional plane. The tip profile includes an arc segment SL1 and a straight line segment SL2. The diameter of the disk is DI, and the radius of the arc segment SL1 is R SL1 , R SL1 / DI = 0.1078 - 0.2078.

[0026] Preferably, the arc segment is a quarter - arc segment.

[0027] Preferably, the diameter of the disk is DI, and the radial intake width of the turbine rotor is HI, HI / DI = 0.0635 - 0.0777.

[0028] Preferably, the diameter of the disk is DI, and the axial outlet blade height of the turbine rotor is HE, HE / DI = 0.20 - 0.30.

[0029] Preferably, the diameter of the disk is DI, and the hub diameter at the axial outlet of the turbine rotor is DE, DE / DI = 0.1779 - 0.1899.

[0030] According to another aspect of the present application, there is provided an expander including a turbine rotor, and the turbine rotor is the above - mentioned turbine rotor.

[0031] Preferably, the expander further includes a turbine inlet, a turbine volute, a turbine nozzle, and a turbine outlet. The air flow sequentially passes through the turbine inlet, the turbine volute, the turbine nozzle, and the turbine rotor and then flows out from the turbine outlet.

[0032] According to another aspect of the present application, there is provided an air cycle machine including the above-mentioned turbine rotor or the above-mentioned expander.

[0033] The turbine rotor provided by the present application includes a disk and blades arranged circumferentially along the disk. The angle of attack of the air flow at the leading edge of the blade is 0°. Wherein, the inlet air flow angle α of the turbine rotor is the included angle between the air flow direction FD and the axis of the turbine rotor, and the leading edge blade angle β of the turbine rotor is the included angle between the tangent line of the leading edge of the mean camber line of the blade profile and the axis of the turbine rotor. The angle of attack of the air flow i = α - β. The turbine rotor has redesigned and optimized the blades of the turbine rotor according to the inlet air flow angle, so that the angle of attack of the air flow at the leading edge position of the blades of the turbine rotor is 0°, thereby forming a zero angle of attack turbine rotor blade structure, enabling the blade angle to match the air flow angle, and using the zero angle of attack characteristic to solve the problem of large separated flow generated by the air flow impact at the leading edge of the rotor, thereby reducing the air flow loss, improving the working efficiency of the turbine rotor, and increasing the refrigeration capacity of the refrigeration system. Description of the Drawings

[0034] Figure 1 is a three-dimensional structural schematic diagram of a turbine rotor according to an embodiment of the present application;

[0035] Figure 2 is a structural dimension schematic diagram of a turbine rotor according to an embodiment of the present application;

[0036] Figure 3 is a schematic diagram of the meridional flow path profile of the disk and the wheel cover of a turbine rotor according to an embodiment of the present application;

[0037] Figure 4 is a schematic diagram of the definition of the blade coordinate system of a turbine rotor according to an embodiment of the present application;

[0038] Figure 5 is a coordinate diagram of the geometric definition of the blades of a turbine rotor according to an embodiment of the present application;

[0039] Figure 6 is a structural diagram of the distribution of the mean camber line of a turbine rotor according to an embodiment of the present application;

[0040] Figure 7 is a structural diagram of the thickness distribution of a turbine rotor according to an embodiment of the present application;

[0041] Figure 8 is a three-dimensional structural schematic diagram of an air cycle machine according to an embodiment of the present application;

[0042] Figure 9Schematic diagram of the turbine local structure of an air circulation machine according to an embodiment of the present application;

[0043] Figure 10 Velocity vector of the turbine elementary flow surface at a large angle of attack in the related art;

[0044] Figure 11 Velocity vector of the turbine elementary flow surface at zero angle of attack according to an embodiment of the present application.

[0045] The reference numerals are shown as:

[0046] 1. Disk; 2. First blade; 3. Second blade; 4. Turbine inlet; 5. Turbine volute; 6. Turbine nozzle; 7. Turbine outlet; 8. Turbine rotor. Specific embodiments

[0047] Referring to Figures 1 to 7 As shown, according to the embodiment of the present application, the turbine rotor includes a disk 1 and blades arranged circumferentially along the disk 1. The angle of attack of the air flow at the leading edge of the blade is 0°. Wherein, the inlet air flow angle α of the turbine rotor is the included angle between the air flow direction FD and the axial direction of the turbine rotor, and the leading edge blade angle β of the turbine rotor is the included angle between the tangent line of the leading edge of the mean camber line of the blade and the axial direction of the turbine rotor. The angle of attack of the air flow i = α - β.

[0048] The turbine rotor has redesigned and optimized the blades of the turbine rotor according to the inlet air flow angle, so that the angle of attack of the air flow at the leading edge position of the blades of the turbine rotor is 0°, thus forming a blade structure of the turbine rotor with zero angle of attack, making the blade angle match the air flow angle, and using the zero angle of attack characteristic to solve the problem of large separation flow generated by the air flow impact at the leading edge of the rotor, thereby reducing the air flow loss, improving the working efficiency of the turbine rotor, and increasing the refrigeration capacity of the refrigeration system.

[0049] In one embodiment, α is jointly determined by the upstream turbine nozzle 6 and the flow channel, while the leading edge blade angle β of the turbine rotor is comprehensively determined by the mean camber line CL, thickness t, leading edge thickness tLE, and trailing edge thickness tTE of the turbine rotor. Therefore, it can be seen from the formula i = α - β that by using a reasonable flow channel design and turbine rotor blade design, the inlet air flow angle α and the blade angle β are respectively controlled to ensure i = 0, thereby a turbine rotor with zero angle of attack can be constructed.

[0050] In one embodiment, the airfoil coordinate system of the blade is constructed in the following manner: by rotating the meridian streamline around the central axis O of the turbine rotor, the intersection line JL with the blade is formed, and the intersection line JL is flattened and expanded along the meridian coordinate m and the circumferential coordinate θ to obtain the airfoil of the blade. Taking the meridian coordinate m as the x-axis, the circumferential coordinate θ as the y-axis, and the intersection point of the connection line of the two endpoints of the leading edge of the airfoil and the mean camber line of the airfoil as the center, the airfoil coordinate system of the blade is established.

[0051] The above-mentioned meridian direction coordinate m is the coordinate of the meridian streamline direction mapped by the actual streamline from the inlet HI to the outlet HE of the turbine rotor on the meridian plane. The circumferential coordinate θ is the coordinate of the tangential direction around the O axis. The meridian plane is the plane generated by the intersection of the plane passing through the rotation axis O and the actual flow passage, as shown in the appendix. Figure 3 As shown. The circumferential projection of the actual three-dimensional flow passage on the meridian plane is the meridian flow passage, and the circumferential projection of the actual streamline on the meridian plane is the meridian streamline.

[0052] In one embodiment, the blades include a first blade 2 and a second blade 3. The blade heights of the first blade 2 and the second blade 3 are the same. The axial length of the first blade 2 is greater than that of the second blade 3. The first blade 2 and the second blade 3 are arranged alternately along the circumferential direction of the disk 1.

[0053] Since along the airflow direction, the flow passage width between two adjacent first blades 2 decreases rapidly, when the airflow flows along the guiding direction of the blades, the problem of a sharp decrease in the flow area will occur, resulting in easy local separation of the airflow and affecting the airflow flow efficiency. To avoid this problem, the blades are arranged in an alternating form of the first blade 2 and the second blade 3. The first blade 2 is used as a large blade, and the second blade 3 is used as a small blade. Small blades with the same blade height as the large blades are arranged between two adjacent large blades to divide the flow passage in the inlet area between the two large blades. At the same time, since the axial length of the small blade is less than that of the large blade, the small blade can be used to reduce the change rate of the flow passage width at the airflow inlet, so that the flow area at the inlet of the turbine rotor is basically the same as that at the outlet, thereby improving the stability of the airflow flow, increasing the airflow flow efficiency, and reducing the airflow flow loss.

[0054] The large blades and the small blades adopt a full blade height structure, which can avoid air leakage at the top of the small blades, further improve the aerodynamic performance of the turbine rotor, and reduce the airflow flow loss.

[0055] In one embodiment, the leading edges of the first blade 2 and the second blade 3 are aligned, so as to facilitate using the second blade 3 to divide the flow passage between adjacent blades in the inlet area of the first blade 2 and control the change rate of the flow passage width.

[0056] In one embodiment, the first blade 2 and the second blade 3 are integrally processed on the same blank. In this embodiment, based on the aerodynamic layout structure of large and small blades with full blade height and aligned leading edges, the large blades and the small blades are processed directly on the turbine rotor disk 1 of the cylindrical blank by five-axis side milling. The blade top is an open structure without a shroud, which avoids the problem of deformation of small-sized turbine blades caused by welding the shroud, reduces the processing difficulty, and improves the processing accuracy. The processing of the large and small blades can also be carried out by other methods, such as wire cutting, powder metallurgy, or 3D printing, etc.

[0057] The small blades are located in the middle positions between two adjacent large blades. In one embodiment, the numbers of both the large blades and the small blades are 9. The intersecting surface between the disk 1 and the blades is HU, and the tip surface of the blades is SH. Specifically, the intersecting surface between the disk 1 and the large blades is HU1, and the intersecting surface between the disk 1 and the small blades is HU2. The tip surface of the large blades is SH1, and the tip surface of the small blades is SH2.

[0058] In one embodiment, the flow passage of the turbine rotor is composed of the profile line of the disk 1 and the tip profile line. Among them, the profile line of the disk 1 is formed by rotating the intersecting surface HU1 between the first blade 2 and the disk 1 around the O axis and mapping it to the meridian plane. The profile line of the disk 1 includes a straight line segment HL1 and an arc segment HL2. The diameter of the disk 1 is DI, and the radius of the arc segment HL2 is R HL2 , R HL2 / DI = 0.3089 - 0.4089, preferably 0.3589.

[0059] The tip profile line is formed by rotating the tip surface of the first blade 2 around the O axis and mapping it to the meridian plane. The tip profile line includes an arc segment SL1 and a straight line segment SL2. The diameter of the disk 1 is DI, and the radius of the arc segment SL1 is R SL1 , R SL1 / DI = 0.1078 - 0.2078, preferably 0.1578.

[0060] By defining the disk profile line and the tip profile line, the root structure and the tip structure of the blades can be defined, so as to precisely design the flow passage profile line, effectively control the value that the inlet air flow angle α of the turbine rotor needs to satisfy, and facilitate the construction of a zero - attack - angle turbine rotor.

[0061] In one embodiment, the arc segment is a quarter - arc segment. As a preferred embodiment, both the arc segment HL2 and the arc segment SL1 are quarter - arc segments. Since the air flow flows into the turbine rotor radially and then flows out axially, which is equivalent to a 90 - degree turn of the air flow during the process of flowing through the turbine rotor. Therefore, setting both the root arc segment and the tip arc segment as quarter - arc segments makes the included angle between the guiding directions of the blades at the inlet position and the outlet position also 90°, which can match the flow turning of the air flow, improve the guiding effect on the air flow, reduce the flow loss of the air flow, and improve the air flow efficiency.

[0062] The blade modeling method of the zero - attack - angle turbine rotor provided by 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 large blades and small blades of the rotor. The blade angle β of the turbine rotor can be controlled through the design of the turbine rotor blades, and finally, the attack angle i of the rotor blades is controlled by calculating i = α - β.

[0063] In one embodiment, the diameter of the roulette 1 is DI, the radial intake width of the turbine rotor is HI, and HI / DI = 0.0635 - 0.0777, which can define a reasonable size of the turbine rotor inlet, thereby ensuring the size of the turbine inlet flow rate. Preferably, HI / DI = 0.0706.

[0064] In one embodiment, the diameter of the roulette 1 is DI, the axial outlet blade height of the turbine rotor is HE, and HE / DI = 0.20 - 0.30. Preferably, HE / DI = 0.25. By defining the ratio range of HE / DI, the outlet size of the turbine rotor can be defined, thereby ensuring the conservation of the outlet flow rate and the inlet flow rate and reducing the airflow loss.

[0065] In one embodiment, the diameter of the roulette 1 is DI, the hub diameter of the axial outlet of the turbine rotor is DE, and DE / DI = 0.1779 - 0.1899. Preferably, DE / DI = 0.1849. By defining the ratio range of DE / DI, it can be ensured that the roulette 1 has sufficient size for machining assembly holes, and on the basis of ensuring the structural strength of the roulette 1, the machining requirements of the assembly holes can be met.

[0066] In one embodiment, within the airfoil coordinate system of the blade, the distribution function of the mean camber line of the airfoil in the meridian direction and the circumferential coordinate system is the Boltzmann function. The intersection surface of the first blade 2 and the roulette 1 is HU1, the tip surface of the first blade 2 is SH1, the intersection surface of the second blade 3 and the roulette 1 is HU2, and the tip surface of the second blade 3 is SH2, where

[0067] The distribution function of the mean camber line of HU1 in the meridian direction and the circumferential coordinate system is:

[0068] θ = 1.08076 - 1.08736 / (1 + e (m-2.99171) / 0.42001 )

[0069] The distribution function of the mean camber line of SH1 in the meridian direction and the circumferential coordinate system is:

[0070] θ = 4.11889 - 4.12590 / (1 + e (m-1.29564) / 0.18631 )

[0071] The distribution function of the mean camber line of HU2 in the meridian direction and the circumferential coordinate system is:

[0072] θ = 0.43835 - 0.09317 / (1 + e (m-1.76698) / 0.15371 )

[0073] The distribution function of the mean camber line of SH2 in the meridian direction and the circumferential coordinate system is:

[0074] θ = 0.88796 - 0.54092 / (1 + e(m-0.89473) / 0.13378 )。

[0075] The turbine rotor blade provided by this application is a ruled surface, so only the distribution functions of the hub intersection surface HU and the tip intersection surface SH are given here.

[0076] In the airfoil coordinate system of the blade, the airfoil thickness of the first blade 2 satisfies:

[0077] t = 2 × [0.67564 + 0.01378(m / C m ) - 2.18765×10 -4 (m / C m ) 2 + 6.78833×10 -7 (m / C m ) 3 )

[0078] The airfoil thickness of the second blade 3 satisfies:

[0079] t = 2 × [0.68432 - 0.00101(m / C m ) - 2.71351×10 -4 (m / C m ) 2 - 2.75724×10 -7 (m / C m ) 3 )

[0080] Where Cm is the component of the blade chord length in the meridional direction, and the blade thickness is constant along the span direction in this coordinate system. Only the thickness distributions of the large and small blades under the two cross-sections of the hub intersection surface HU and the tip intersection surface SH are given here. For the blades between the hub and the tip, the thickness distribution is the same as this, that is, the blade thickness is constant along the span direction.

[0081] The mid-arc line of the subsonic zero-angle-of-attack rotor provided by this application is based on the aerodynamic layout structure of the large and small blades, constructs a zero-angle-of-attack turbine rotor, optimizes the Boltzmann function distribution of the mid-arc lines of the large and small blades by defining the distribution functions of the mid-arc lines of the large and small blades in the meridional direction and the circumferential coordinate system, and can optimize the geometric deflection angle distribution of the mid-arc lines of the large and small blades under subsonic flow conditions, reduce the curvature near the leading edge and the trailing edge, increase the curvature near the blade body part, and control the blade angle. As Figure 6 shown, this mid-arc line distribution form has the function of controlling the leading-edge angle of attack, reducing aerodynamic losses, and improving the efficiency of the turbine rotor.

[0082] In one embodiment, in the airfoil coordinate system of the blade, the maximum thickness TH of the first blade 2 is 1.6 - 2.0 mm, and the maximum thickness position TP is located at the 40% - 50% chord length position. The leading-edge thickness t of the first blade 2LE is 0.6 to 0.7 mm, and the trailing edge thickness t TE is 0.55 to 0.65 mm. Refer to Figure 7 shown. By optimizing the blade thickness and its position distribution, it is ensured that the turbine blade has sufficient strength to resist centrifugal stress, while taking into account the flow efficiency of the air flow in the turbine.

[0083] Refer to in combination Figure 10 and Figure 11 shown. It can be seen from the comparison between the two that Figure 11 the zero angle of attack turbine in Figure 10 effectively controls the angle of attack of the leading edge of the rotor blade, suppressing the

[0084] Table 1 Comparison of Performance Indexes between Zero Angle of Attack and Large Angle of Attack Rotors

[0085]

[0086]

[0087] It can be seen from the data comparison in Table 1 that under the same inlet and outlet boundary conditions, that is, the same inlet total pressure, inlet total temperature, and outlet pressure. Using the zero angle of attack rotor of the present application compared with the conventional large angle of attack rotor, the outlet temperature drops by 14 K, the refrigeration power increases by 42.3%, the isentropic efficiency of the turbine increases by 16.3 percentage points, from 70.1% to 86.4%, and the overall performance of the turbine is greatly improved.

[0088] Refer to in combination Figure 8 and Figure 9 shown. According to the embodiment of the present application, the expander includes a turbine rotor 8, and the turbine rotor 8 is the above-mentioned turbine rotor.

[0089] The expander further includes a turbine inlet 4, a turbine volute 5, a turbine nozzle 6, and a turbine outlet 7. The air flow flows through the turbine inlet 4, the turbine volute 5, the turbine nozzle 6, and the turbine rotor 8 in sequence and then flows out from the turbine outlet 7.

[0090] The turbine inlet 4 accesses high-pressure air flow. After being guided by the turbine volute 5, it flows through the turbine nozzle 6 and is accelerated to transonic state to drive the turbine rotor 8, and quickly reduces the gas pressure and temperature. At the same time, it outputs power to drive the compressor and the fan to rotate. The air flow is connected to the mixing chamber for temperature adjustment through the turbine outlet 7 and then sent into the cabin.

[0091] According to the embodiment of the present application, the air cycle machine includes the above-mentioned turbine rotor or the above-mentioned expander.

[0092] Those skilled in the art can easily understand that, on the premise of no conflict, the above advantageous ways can be freely combined and superimposed.

[0093] The above are only the preferred embodiments of the present application and are not intended to limit the present application. Any modifications, equivalent replacements, improvements, etc. 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 in the technical field, several improvements and variations can be made without departing from the technical principle of the present application, and these improvements and variations should also be regarded as the protection scope of the present application.

Claims

1. A turbine rotor, characterized in that, It includes a wheel disc (1) and blades arranged circumferentially along the wheel disc (1). The blades have blade leading edges, and the airflow attack angle of the blade leading edges is 0°. Wherein the inlet airflow angle α of the turbine rotor is the included angle between the airflow direction FD and the axis of the turbine rotor, and the leading edge blade angle β of the turbine rotor is the included angle between the tangent line of the leading edge of the mean camber line of the blade airfoil and the axis of the turbine rotor. The airflow attack angle i = α - β; The blade includes a first blade (2). The flow passage of the turbine rotor is composed of a disk profile and a tip profile. The disk profile is formed by rotating the intersection surface HU1 of the first blade (2) and the disk (1) around the central axis O of the turbine rotor and mapping it to the meridian plane. The disk profile includes a straight line segment HL1 and an arc segment HL2. The diameter of the disk (1) is DI, and the radius of the arc segment HL2 is R HL2 , R HL2 / DI = 0.3089 - 0.4089; The tip profile is formed by rotating the tip surface of the first blade (2) around the O axis and mapping it to the meridian plane. The tip profile includes an arc segment SL1 and a straight line segment SL2. The diameter of the disk (1) is DI, and the radius of the arc segment SL1 is R SL1 , R SL1 / DI = 0.1078 - 0.2078.

2. The turbine rotor according to claim 1, characterized in that, By rotating the meridian streamline around the central axis O of the turbine rotor, it intersects with the blade to form a blade intersection line JL. The intersection line JL is linearly developed along the meridian coordinate m and the circumferential coordinate θ to obtain the airfoil of the blade. Taking the meridian coordinate m as the x-axis, the circumferential coordinate θ as the y-axis, and the intersection point of the connection line of the two endpoints of the leading edge of the airfoil and the mean camber line of the airfoil as the center, the airfoil coordinate system of the blade is established.

3. The turbine rotor according to claim 2, characterized in that, The blade further includes a second blade (3). The first blade (2) and the second blade (3) have the same blade height. The axial length of the first blade (2) is greater than the axial length of the second blade (3). The first blade (2) and the second blade (3) are alternately arranged along the circumference of the wheel disc (1).

4. The turbine rotor according to claim 3, characterized in that, The leading edges of the first blade (2) and the second blade (3) are aligned.

5. The turbine rotor according to claim 3, characterized in that, The first blade (2) and the second blade (3) are integrally processed on the same blank.

6. The turbine rotor according to claim 3, characterized in that, In the airfoil coordinate system of the blade, the distribution function of the mean camber line of the airfoil in the meridian direction and the circumferential coordinate system is the Boltzmann function. The intersection surface of the first blade (2) and the wheel disc (1) is HU1, the tip surface of the first blade (2) is SH1, the intersection surface of the second blade (3) and the wheel disc (1) is HU2, and the tip surface of the second blade (3) is SH2. Wherein, The distribution function of the mean camber line of HU1 in the meridian direction and the circumferential coordinate system is: θ = 1.08076 - 1.08736 / (1 + e (m-2.99171) / 0.42001 ) The distribution function of the mean camber line of SH1 in the meridian direction and the circumferential coordinate system is: θ = 4.11889 - 4.12590 / (1 + e (m-1.29564) / 0.18631 ) The distribution function of the mean camber line of HU2 in the meridian direction and the circumferential coordinate system is: θ = 0.43835 - 0.09317 / (1 + e (m-1.76698) / 0.15371 ) The distribution function of the mean camber line of SH2 in the meridian direction and the circumferential coordinate system is: θ = 0.88796 - 0.54092 / (1 + e (m-0.89473) / 0.13378 )。 7. The turbine rotor according to claim 3, characterized in that, In the airfoil coordinate system of the blade, the airfoil thickness of the first blade (2) satisfies: t = 2 × [0.67564 + 0.01378(m / C m ) - 2.18765 × 10 -4 (m / C m ) 2 + 6.78833 × 10 -7 (m / C m ) 3 ​ The airfoil thickness of the second blade (3) satisfies: t = 2 × [0.68432 - 0.00101(m / C m ) - 2.71351 × 10 -4 (m / C m ) 2 - 2.75724 × 10 -7 (m / C m ) 3 ​ Where Cm is the component of the blade chord length in the meridian direction, and t is the airfoil thickness.

8. The turbine rotor according to claim 3, characterized in that, In the airfoil coordinate system of the blade, the maximum thickness TH of the first blade (2) is 1.6 to 2.0 mm, and the maximum thickness position TP is located at the 40% to 50% chord length position. The leading edge thickness t of the first blade (2) LE is 0.6 to 0.7 mm, and the trailing edge thickness t TE is 0.55 to 0.65 mm.

9. The turbine rotor according to claim 1, characterized in that, The arc segment is a quarter arc segment.

10. The turbine rotor according to any one of claims 1 to 9, characterized in that, The diameter of the wheel disc (1) is DI, and the radial intake width of the turbine rotor is HI. HI / DI = 0.0635 - 0.0777.

11. The turbine rotor according to any one of claims 1 to 9, characterized in that, The diameter of the wheel disc (1) is DI, and the axial outlet blade height of the turbine rotor is HE. HE / DI = 0.20 - 0.

30.

12. The turbine rotor according to any one of claims 1 to 9, characterized in that, The diameter of the wheel disc (1) is DI, and the hub diameter at the axial outlet of the turbine rotor is DE. DE / DI = 0.1779 - 0.1899.

13. An expander, comprising a turbine rotor (8), characterized in that, The turbine rotor is the turbine rotor according to any one of claims 1 to 12.

14. The expander according to claim 13, wherein, The expander further includes a turbine inlet (4), a turbine volute (5), a turbine nozzle (6) and a turbine outlet (7). The air flow sequentially flows through the turbine inlet (4), the turbine volute (5), the turbine nozzle (6) and the turbine rotor (8) and then flows out from the turbine outlet (7).

15. An air circulation machine, characterized in that, Comprising the turbine rotor according to any one of claims 1 to 12 or the expander according to any one of claims 13 to 14.

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