Inlet flow channel structure and method of a super-transonic compressor plane cascade test piece

Abstract

This application belongs to the field of supersonic compressor planar blade cascade testing technology, and specifically relates to an inlet flow channel structure and method for a supersonic compressor planar blade cascade test piece, including: a pressure stabilizing section, a converging section, a supersonic nozzle, a suction device, a straight section, and a test piece; wherein, the inlet of the pressure stabilizing section is supplied with a gas source, the outlet of the pressure stabilizing section is connected to the inlet of the converging section, the outlet of the converging section is detachably connected to the inlet of the straight section, and the test piece is arranged at the outlet of the straight section; the straight section includes two side walls and an upper suction orifice plate and a lower suction orifice plate, both of which have openings, and suction devices are installed on both the upper and lower suction orifice plates; the test piece includes two end walls that are parallel to and connected to the two side walls of the straight section, and a blade row located between the two end walls, wherein the blades of the blade row are gradually inclined backward in an array from top to bottom, realizing the parametric design of the inlet flow channel of the supersonic planar blade cascade test piece based on an open type, and providing effective guidance for the design of other similar test equipment.

Description

Technical Field

[0001] This invention relates to the field of supersonic compressor planar blade cascade testing technology, and particularly to an inlet flow channel structure and method for a supersonic compressor planar blade cascade test piece. Background Technology

[0002] With the development of supersonic fan / compressor blade design, experimental research based on supersonic compressor blade design has also emerged. Supersonic transonic compressor planar cascade testing is a fundamental test in the design process of high-load, high-thrust-weight aero-engine fans / compressors. It allows for the study of flow mechanisms using relatively simple yet fully functional fluid testing equipment across a wide range of flight conditions, quickly providing fundamental information about the blades, such as loads, losses, and airflow deflection characteristics. Compared to actual rotor testing, supersonic transonic planar cascade testing can obtain similar information with significantly less time and cost. Moreover, when these crucial aerodynamic boundary conditions are accurately understood and controlled, experimental results can be clearly compared to numerical calculations.

[0003] The inlet flow channel in the supersonic planar vane test provides a uniform and stable flow field environment for the supersonic planar vane test specimen, enabling the effective flow of compressed gas through it. Furthermore, during the supersonic planar vane test, the inlet Mach number is a critical state control parameter. The accuracy of the inlet total pressure measurement and the rationality of the supersonic nozzle curve design directly determine the accuracy of the inlet Mach number. Ensuring that the total pressure testing instrument section in the pressure stabilization zone and the flow state monitoring of the supersonic nozzle meet the test requirements for flow field quality are extremely important.

[0004] The disadvantages of existing technical solutions are as follows:

[0005] ① The diameter of the steady flow section is not suitable. The diameter of the steady flow section is related to the airflow velocity in the steady flow section and the contraction ratio of the convergence section. If the diameter of the steady flow section is too small, the airflow velocity in the steady flow section will be too large, which will affect the flow field quality. If the diameter of the steady flow section is too large, it will increase the space occupied, weight and processing cost of the air intake structure of the test piece.

[0006] ② The convergent section is located between the steady flow section and the supersonic nozzle section. Its function is to smoothly transition the cross-sectional dimensions of the steady flow section to the inlet cross-sectional dimensions of the supersonic nozzle section, while accelerating the airflow to the inlet cross-sectional airflow parameters of the nozzle section without separation. If the profile of the convergent section is not suitable, it will cause airflow separation in the convergent section.

[0007] ③ The function of the supersonic nozzle section is to accelerate the airflow. If the profile curve design is not appropriate, it will cause airflow separation, uneven airflow at the throat, and failure to meet the inlet flow field requirements of the test piece.

[0008] ④ The pressure gradient in the straight section requires a suction device to adjust the airflow. An improperly designed suction device can cause shock waves in the straight section, reducing the inlet velocity of the test piece. Summary of the Invention

[0009] To address the aforementioned problems, this application provides an inlet flow channel structure for a supersonic compressor planar blade cascade test specimen, comprising:

[0010] Pressure stabilizing section, convergent section, supersonic nozzle, suction device, straight section, test specimen;

[0011] The pressure stabilizing section is supplied with an air source at its inlet, the outlet of the pressure stabilizing section is connected to the inlet of the converging section, the outlet of the converging section is detachably connected to the inlet of the straight section with a supersonic nozzle, and the outlet of the straight section is equipped with a test piece.

[0012] The straight section includes two side walls and an upper suction perforated plate and a lower suction perforated plate, both of which have openings. Both the upper suction perforated plate and the lower suction perforated plate are equipped with suction devices.

[0013] The test specimen includes two end walls that are parallel to the two side walls of the straight section, and a blade row located between the two end walls, wherein the blades of the blade row are distributed in an array that gradually tilts backward from top to bottom.

[0014] Preferably, the suction device of the lower suction opening plate includes a lower front suction device located at the front and a lower rear suction device located below the projection of the blade row.

[0015] A design method for the inlet flow channel structure of a supersonic compressor planar blade cascade test piece is provided. The method is used to design the inlet flow channel structure of the supersonic compressor planar blade cascade test piece. Step 1: Based on the set physical flow rate, pressure, inlet velocity and gas temperature of the gas passing through the test piece, the relationship between the pressure, velocity and cross-sectional area of ​​the gas passing through the pressure stabilization section is established, and then the diameter d of the pressure stabilization section is obtained.

[0016] Step 2: Select multiple curves as the convergence curves of the convergence segment. Optimize the multiple convergence curves with the shortest convergence segment length L and the goal of preventing separation on the tunnel wall when the airflow accelerates along the convergence segment. Select one convergence curve or fit multiple convergence curves as the final convergence curve.

[0017] Step 3: Use the Vitósinski curve as the profile curve of the subsonic contraction section of the supersonic nozzle, and design the profile curve of the contraction section of the supersonic nozzle so that no separation occurs on the tunnel wall when the airflow flows along the contraction section and the length of the contraction section is minimized.

[0018] Step 4: Set the opening diameter of the straight section according to the height of the straight section and the thickness of the straight section wall panel, and set the height of the suction device according to the height of the straight section.

[0019] Preferably, the various curves mentioned in step two include: the Wittsinski curve, the bicubic curve, or the quintic curve.

[0020] Preferably, in step four, the aperture is designed to be 1 / 80 to 1 / 100 of the height (H) of the straight section, the ratio of the thickness of the straight section wall plate to the aperture is 1 to 2, and the height of the suction device is 40% to 50% of the height of the straight section.

[0021] Preferably, the supersonic nozzle includes a subsonic contraction section and a supersonic diffusion section;

[0022] The profile curve of the subsonic contraction section adopts the Wittsinski curve;

[0023] The profile of the supersonic diffuser section includes, in sequence along the airflow direction, a throat arc segment with its center located outside the nozzle, an expansion straight line segment tangent to the throat arc segment, and an expansion curve segment with its inner diameter gradually increasing along the airflow direction.

[0024] Preferably, the nozzle length l is calculated based on the given nozzle length-to-half-height ratio Blh and the half-height h of the test section.

[0025] The boundary layer displacement thickness at the nozzle exit is calculated based on the nozzle length l and the given boundary layer thickness correction angle q.

[0026] The correction amount for the viscosity curve at the nozzle exit is calculated based on the boundary layer displacement thickness at the nozzle exit.

[0027] The ratio of nozzle exit area F1 to throat area F* is calculated based on the nozzle exit Mach number M. The throat height y* is then calculated based on the ratio, the correction amount of the viscosity curve at the nozzle exit, and the geometric relationship of the nozzle exit.

[0028] Calculate the radius of the arc segment of the throat based on the throat height y*;

[0029] Calculate Prandtl-Mayer angle v1 based on the nozzle exit Mach number M;

[0030] The maximum expansion angle βB is calculated based on Prandtl-Mayer angle v1;

[0031] The Mach number Mb at the end point B of the expanding straight line segment is calculated based on the maximum expansion angle βB.

[0032] The sonic radius r0 of the spring is calculated based on the Mach number Mb at the terminal point B and the throat height y*.

[0033] The radius rB of the final point B is calculated based on the radius r0 of the sound velocity of the spring, and the position of the final point B is obtained based on the radius rB of the final point B.

[0034] The expansion curve segment that gradually reduces the gas turning angle to 0 is calculated based on the nozzle exit Mach number M, the terminal point B Mach number Mb, and the maximum expansion angle βB.

[0035] Preferably, the exit size of the supersonic nozzle is R* = 0.5H, where H is the inlet size of the straight section.

[0036] Preferably, the supersonic nozzle includes several models with Mach numbers between 0.8 and 2.0.

[0037] The advantages of this application include:

[0038] ① Design the diameter of the pressure stabilizing section reasonably to reduce the space occupied and processing costs while ensuring the uniformity of the flow velocity and flow field in the pressure stabilizing section.

[0039] ②Develop a parameterized design method for the flow channel profile of the convergent section to reduce the design difficulty of the flow channel profile of the convergent section.

[0040] ③ Develop a parametric supersonic nozzle flow channel profile design method to reduce the design difficulty of supersonic nozzle flow channel profiles and realize the parametric design capability of supersonic nozzles in different Mach number ranges.

[0041] ④ Develop parameterized design methods for straight sections and suction devices to adapt to the suction requirements of different test inlet Mach numbers, unify design criteria, and avoid structural errors caused by different design methods that lead to differences in the flow field of the test specimen. Attached Figure Description

[0042] Figure 1 This is a schematic diagram of the Vickers curve of the convergence segment;

[0043] Figure 2 These are schematic diagrams of bicubic and quintic curves with convergence segments;

[0044] Figure 3 These are schematic diagrams of different types of convergence curves;

[0045] Figure 4 It is a fountain stream in a nozzle;

[0046] Figure 5 This is a schematic diagram of the nozzle curve;

[0047] Figure 6 Schematic diagram of the opening structure in the straight section;

[0048] Figure 7 Schematic diagram of the inlet flow channel structure of a supersonic compressor planar blade test piece. Detailed Implementation

[0049] To make the technical solution and advantages of this application clearer, the technical solution of this application will be described in a clearer and more complete manner below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of this application, and are only used to explain this application, not to limit this application. It should be noted that, for ease of description, only the parts related to this application are shown in the accompanying drawings. Other related parts can be referred to the general design. In the absence of conflict, the embodiments and technical features in the embodiments of this application can be combined with each other to obtain new embodiments.

[0050] Furthermore, it should be noted that, unless otherwise explicitly specified and limited, terms such as “installation,” “connection,” and “linkage” used in the description of this application should be interpreted broadly. For example, a connection can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium; or it can be a connection within two components. Those skilled in the art can understand its specific meaning in this application according to the specific circumstances.

[0051] This invention provides a design method for the inlet flow channel of a supersonic transonic compressor planar blade test piece and the overall structure of the inlet flow channel of the test piece obtained based on this design method.

[0052] On the one hand, a parameterized design method for the inlet flow channel of a supersonic transonic compressor planar blade cascade test piece is provided:

[0053] Step 1: Determine the diameter D of the voltage stabilizing section

[0054] The steady-flow section acts as a flow straightener, ensuring flow field quality. The diameter of the steady-flow section directly affects the wind tunnel's contraction ratio and also influences the structural dimensions of the diffuser and convergent sections. The airflow velocity in the steady-flow section should not be too high, otherwise it will affect the flow field quality; the velocity should be controlled at 10–15 m / s, and the pressure within the steady-flow section should be controlled at 2–3 atmospheres. Based on parameters such as the physical flow rate, pressure, inlet velocity, and airflow temperature range of the test specimen, the relationship between pressure, velocity, and cross-sectional area within the steady-flow section is derived using W = ρVA and p / ρ = RT, thus obtaining the diameter d of the steady-flow section.

[0055] Step 2: Determine the length of the convergence segment and the convergence curve

[0056] The convergence section is located between the steady-flow section and the supersonic nozzle section. Its function is to smoothly transition the cross-sectional dimensions of the steady-flow section to the inlet cross-sectional dimensions of the supersonic nozzle section, while accelerating the airflow to the inlet cross-sectional airflow parameters of the supersonic nozzle section without separation. The following points should be considered when designing the convergence section:

[0057] ① When the airflow accelerates along the converging section, no separation occurs on the tunnel wall;

[0058] ② The airflow at the outlet section of the converging section is uniform, parallel, and stable;

[0059] ③ The convergence section should not be too long or too short. The length of the convergence section should generally not be too long, mainly due to the cost of the equipment; the length of the convergence section should also not be too short, so as to avoid uneven airflow or even separation.

[0060] The aspect ratio (length of convergence section / diameter of convergence section inlet) of the convergence section in a transsupersonic wind tunnel is typically 0.5 to 1.3. The length L of the convergence section is determined based on the outlet diameter d of the steady flow section, as well as the structure and dimensions of other parts.

[0061] The convergence profile of the wind tunnel's convergence section is designed as a smoothly transitioning curved surface. Since the exit of the convergence section is rectangular, convergence curves are drawn at the midpoints of both the length and width, with the remaining sides smoothly transitioning. Commonly chosen convergence curves include the Vitósinski curve, bicubic curve, or quintic curve. Figures 1-3 As shown.

[0062] The Vickers curve formula is expressed as follows:

[0063]

[0064] In the formula, R1 is the radius of the inlet section of the convergence section (i.e., 0.5d, where d is the outlet diameter of the steady flow section);

[0065] R2—Radius of the exit section of the convergence segment;

[0066] R—radius of the cross section at an axial distance of x;

[0067] L — Length of the convergence segment.

[0068] The inner wall curve of the convergence segment can be calculated based on the above data.

[0069] When the shrinkage ratio (C=(R1 / R2)) 2 When the radius is relatively large (generally greater than 4), it can be achieved by shifting the axis. Let R′1 and R′2 represent the inlet and outlet radii of the actual convergence section, respectively. h If the axis shift is the radius, then the shrinkage ratio is C = (R′1 / R′2). 2 .

[0070] Let R1 = R′1 + R h R2 = R'2 + R h R1 = 2R2

[0071] Get R h =R′1-2R′2

[0072] Substituting into the original equation yields the Vickers curve of the inner wall of the convergent segment.

[0073] The formula for a hypercubic curve is expressed as follows:

[0074]

[0075] In the formula x m —The point connecting the two curves;

[0076] D2—Diameter of the exit section of the convergence segment (m);

[0077] D1—Diameter of the inlet section of the convergence segment (m);

[0078] D – Axial distance x is the diameter of the cross-section (m).

[0079] The formula for a fifth-power curve is expressed as follows:

[0080]

[0081] Step 3: Supersonic Nozzle Profile Curve

[0082] The supersonic nozzle (hereinafter referred to as the nozzle) is an important component for ensuring a uniform airflow at the design Mach number Ma in the test section. Its function is to accelerate the airflow isentropically. The nozzle consists of a subsonic contraction section and a supersonic diffusion section.

[0083] The subsonic contraction phase uses the Vitósinski curve. Its formula is:

[0084]

[0085] In the formula, R0 is the radius of the inlet section of the contraction section;

[0086] R * The radius of the exit section of the contraction section;

[0087] x is a relative coordinate;

[0088] R is the radius of the cross section at any x.

[0089] The diffusion section consists of an initial expansion section AB and a parallel section BC, with B being the inflection point. The initial expansion section is designed using the Crown method, the purpose of which is to transform the sonic flow from the throat into a spring-like flow at the inflection point, such as... Figure 4 As shown. The parallel section BC transforms this supersonic flow into a uniform flow parallel to the axis, and the nozzle shape is as follows. Figure 5 As shown.

[0090] The function of the converging section of the nozzle is to accelerate the airflow, and the converging section should meet three requirements:

[0091] ① When the airflow flows along the contraction section, no separation occurs on the tunnel wall. For a given contraction ratio,

[0092] An excessively short contraction section can lead to airflow separation on the wall and uneven airflow in the throat.

[0093] ②The airflow at the outlet of the contraction section is uniform, straight, and stable.

[0094] ③ The contraction section should not be too long, as an excessively long section will increase construction costs and energy loss. The profile of the supersonic diffuser section, along the airflow direction, includes, in sequence, a throat arc section centered on the outside of the nozzle, a straight expansion section tangent to the throat arc section, and an expansion curve section whose inner diameter gradually increases along the airflow direction.

[0095] The nozzle length l is calculated based on the given nozzle length-to-half-height ratio Blh and the half-height h of the test section.

[0096] The boundary layer displacement thickness at the nozzle exit is calculated based on the nozzle length l and the given boundary layer thickness correction angle q.

[0097] The correction amount for the viscosity curve at the nozzle exit is calculated based on the boundary layer displacement thickness at the nozzle exit.

[0098] The ratio of nozzle exit area F1 to throat area F* is calculated based on the nozzle exit Mach number M. The throat height y* is then calculated based on the ratio, the correction amount of the viscosity curve at the nozzle exit, and the geometric relationship of the nozzle exit.

[0099] Calculate the radius of the arc segment of the throat based on the throat height y*;

[0100] Calculate Prandtl-Mayer angle v1 based on the nozzle exit Mach number M;

[0101] The maximum expansion angle βB is calculated based on Prandtl-Mayer angle v1;

[0102] The Mach number Mb at the end point B of the expanding straight line segment is calculated based on the maximum expansion angle βB.

[0103] The sonic radius r0 of the stream is calculated based on the Mach number Mb of the terminal point B and the throat height y*; the radius rB of the terminal point B is calculated based on the sonic radius r0 of the stream; and the position of the terminal point B is obtained based on the radius rB of the terminal point B.

[0104] The expansion curve segment that gradually reduces the gas turning angle to 0 is calculated based on the nozzle exit Mach number M, the terminal point B Mach number Mb, and the maximum expansion angle βB.

[0105] Step 4: Straight Section and Suction Device

[0106] The straight section has a suction function and is generally of the slotted or perforated type. This invention uses an oblique hole perpendicular to the wall surface at a 30° angle. Figure 6As shown, the orifice diameter is designed to be 1 / 80 to 1 / 100 of the height (H) of the straight section, the ratio of the wall thickness of the straight section to the orifice diameter is 1 to 2, and the height of the suction device is 40% to 50% of the height of the straight section. The orifice diameter, the height of the straight section, and the height of the suction device will be designed in more detail according to the requirements of the experimental flow field.

[0107] On the other hand, an inlet flow channel structure for a supersonic planar blade cascade test specimen, obtained based on the inlet flow channel design method of the present invention, is provided. A schematic diagram of the inlet flow channel structure is shown below. Figure 7 As shown, it mainly consists of a pressure stabilizing section 1, a convergence section 2, a supersonic nozzle 3, a straight section 4, an upper suction device 5, a lower rear suction device 6, and a lower rear suction device 7. The parts are connected by bolts. The features are as follows: the inlet of the pressure stabilizing section 1 is connected to the intake pipe, the outlet diameter of the intake pipe is d, the outlet of the pressure stabilizing box 1 is connected to the inlet of the converging section 2, the inlet size of the converging section is R1=0.5d; the outlet of the converging section 2 is connected to the inlet of the supersonic nozzle 3, the length L of the converging section is determined according to the contraction ratio of the converging section; the outlet size R2 of the converging section 2 is the same as the inlet size of the supersonic nozzle 3; the outlet of the supersonic nozzle 3 is connected to the straight section 4, the inlet size of the straight section 4 is H, the outlet size of the supersonic nozzle 3 is R*=0.5H, and the length LP of the supersonic nozzle 3 is determined in advance according to the overall design requirements; the upper suction device 5, the lower rear suction device 6 and the lower rear suction device 7 are respectively connected to the straight section 4, the height of the suction device is h=(0.4~0.5)H; the outlet of the straight section 4 is connected to the test piece. The supersonic nozzle 3 has a fixed curve profile. The supersonic nozzle 3 with the corresponding exit speed is replaced according to the requirements of different inlet speeds of the test piece. The curve profiles of the supersonic nozzle 3 with different exit speeds can be output by the "Transonic Nozzle Profile Curve Design Program" in "Step 3".

[0108] This program can meet the design requirements of supersonic nozzle profile curves with an exit Mach number range of 0.8 to 2.0; the orifice plate with suction function in the straight section 4 can be replaced with an orifice plate of the appropriate orifice ratio and orifice form according to the test requirements; all other components are general-purpose components and do not need to be replaced; the total pressure test section on the pressure stabilizing section 1 is used to calculate the Mach number along the wall and the exit Mach number of the supersonic nozzle 3; the static pressure on the end wall of the supersonic nozzle 3 is used to calculate the Mach number along the wall of the supersonic nozzle 3, and its function is to monitor the operating status of the supersonic nozzle 3 during the test and whether the required flow field state has been reached.

[0109] After completing the design based on the invention, a numerical simulation calculation with a model ratio of 1:1 was performed to verify the rationality and correctness of the invention method.

[0110] Compared with existing technical solutions, the advantages of this invention are:

[0111] ① The parametric design of the inlet flow channel of the open-type supersonic planar blade cascade test specimen was realized, providing effective guidance for the design of other similar test equipment;

[0112] ②Based on the experimental requirements and the aerodynamic principles of supersonic nozzles, a supersonic nozzle profile curve design program was designed and developed, forming a series of supersonic nozzle profile curve design capabilities, and improving the accuracy and design efficiency of supersonic nozzle profile curves.

[0113] ③ The inlet flow channel structure of the test piece obtained based on this design method has a relatively simple processing technology, reduces the processing cost of the test piece, and achieves lightweighting, thus having a certain degree of versatility.

[0114] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A method for designing the inlet flow channel structure of a supersonic compressor planar blade cascade test specimen, wherein the inlet flow channel structure of the supersonic compressor planar blade cascade test specimen includes: Stabilizing section (1), converging section (2), supersonic nozzle (3), suction device, straight section (4), test specimen; Among them, the inlet of the pressure stabilizing section (1) is supplied with a gas source, the outlet of the pressure stabilizing section (1) is connected to the inlet of the converging section (2), the outlet of the converging section (2) is detachably connected to the inlet of the straight section (4) with a supersonic nozzle (3), and the outlet of the straight section (4) is provided with a test piece. The straight section (4) includes two side walls and an upper suction opening plate and a lower suction opening plate, both of which have openings. Both the upper suction opening plate and the lower suction opening plate are equipped with suction devices. The method includes the following steps: Step 1: Based on the set physical flow rate, pressure, inlet velocity and gas temperature of the gas passing through the test piece, establish the relationship between the pressure, flow rate and cross-sectional area of ​​the gas passing through the pressure stabilizing section (1), and then obtain the diameter d of the pressure stabilizing section (1); Step 2: Select multiple curves as the convergence curves of the convergence segment. Optimize the multiple convergence curves with the shortest convergence segment length L and the goal of preventing separation on the tunnel wall when the airflow accelerates along the convergence segment. Select one convergence curve or fit multiple convergence curves as the final convergence curve. Step 3: The Vitósinski curve is used as the profile curve of the subsonic contraction section of the supersonic nozzle (3). The profile curve of the contraction section of the supersonic nozzle (3) is designed so that there is no separation on the tunnel wall when the airflow flows along the contraction section and the contraction section length is the shortest. Step 4: Set the opening diameter of the straight section according to the height of the straight section and the thickness of the straight section wall panel, and set the height of the suction device according to the height of the straight section. The supersonic nozzle includes a subsonic contraction section and a supersonic diffusion section; The profile of the supersonic diffuser section includes, in sequence along the airflow direction, a throat arc segment with its center located outside the nozzle, an expansion straight line segment tangent to the throat arc segment, and an expansion curve segment with its inner diameter gradually increasing along the airflow direction. The nozzle length l is calculated based on the given ratio Blh of nozzle length l to half-height h of test section and half-height h of test section. The boundary layer displacement thickness at the nozzle exit is calculated based on the nozzle length l and the given boundary layer thickness correction angle q. The correction amount for the viscosity curve at the nozzle exit is calculated based on the boundary layer displacement thickness at the nozzle exit. The ratio of nozzle exit area F1 to throat area F* is calculated based on the nozzle exit Mach number M. The throat height y* is then calculated based on the ratio, the correction amount of the viscosity curve at the nozzle exit, and the geometric relationship of the nozzle exit. Calculate the radius of the arc segment of the throat based on the throat height y*; Calculate Prandtl-Mayer angle v1 based on the nozzle exit Mach number M; The maximum expansion angle βB is calculated based on Prandtl-Mayer angle v1; The Mach number Mb at the end point B of the expanding straight line segment is calculated based on the maximum expansion angle βB. The sonic radius r0 of the spring is calculated based on the Mach number Mb at the terminal point B and the throat height y*. The radius rB of the final point B is calculated based on the radius r0 of the sound velocity of the spring, and the position of the final point B is obtained based on the radius rB of the final point B. The expansion curve segment that gradually reduces the gas turning angle to 0 is calculated based on the nozzle exit Mach number M, the terminal point B Mach number Mb, and the maximum expansion angle βB.

2. The inlet flow channel structure design method for the supersonic compressor planar blade cascade test specimen as described in claim 1, characterized in that, The test piece includes two end walls that are parallel to the two side walls of the straight section (4) and a blade row located between the two end walls, wherein the blades of the blade row are distributed in an array that gradually tilts backward from top to bottom.

3. The method of designing an inlet flow passage configuration for a super-transonic compressor cascade test piece according to claim 1, characterized in that, The suction device of the lower suction opening plate includes a lower front suction device (7) located at the front and a lower rear suction device (6) located below the projection of the blade row.

4. The method of designing an inlet flow passage configuration for a super-transonic compressor cascade test piece according to claim 1, characterized in that, The various curves mentioned in step two include: the Wittsinski curve, the bicubic curve, or the quintic curve.

5. The method of designing an inlet flow passage configuration for a super-transonic compressor cascade test piece according to claim 1, wherein In step four, the aperture is designed to be 1 / 80 to 1 / 100 of the height H of the straight section, the ratio of the thickness of the straight section wall plate to the aperture is 1 to 2, and the height of the suction device is 40% to 50% of the height of the straight section.

6. The inlet flow channel structure design method for the supersonic compressor planar blade cascade test specimen as described in claim 1, characterized in that, The exit size of the supersonic nozzle (3) is R*=0.5H, where H is the inlet size of the straight section (4).

7. The method of designing an inlet flow passage configuration for a super-transonic compressor cascade test piece according to claim 1, wherein The supersonic nozzle (3) includes several models with Mach numbers between 0.8 and 2.0.

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