Flow matching design method for supersound serial blades

Abstract

The invention relates to a flow matching design method of a supersound serial blade, which comprises the following steps: based on geometric profile coordinate point parameters of the serial blade, establishing a blade grid channel flow area change curve through a geometric throat analysis program to realize geometric flow capacity evaluation; rANS simulation is adopted to extract flow field characteristic parameters, the area correction coefficient of the pneumatic throat is solved, and pneumatic flow capacity evaluation is achieved; determining the nominal attack angle range of the rear-row blades of the tandem blades under the single working condition by judging the flow capacity under different pneumatic throat areas; establishing a relation between the flow area of the pneumatic throat of the tandem blade and a shock wave system structure under different incoming flow working conditions; and finally, obtaining a blade profile optimization result of the supersound serial blade meeting the matching relationship. According to the method, the serial blade with the good matching relation is formed, the performance index of the serial blade is improved, the design efficiency of the serial blade can be effectively improved, and technical support can be provided for serial blade matching design and engineering application.

Description

Technical Field

[0001] The present invention relates to the technical field of turbomachinery, and particularly relates to a flow matching design method for supersonic tandem blades. Background Art

[0002] The tandem blade design is a design technique that divides the original conventional blade into front and rear row blades. By regenerating the momentum layer in the rear row to reduce the boundary layer separation loss, it has the performance characteristics of high load and low loss. Compared with conventional blades, the same performance can be achieved with fewer stages, and the weight and size of the compressor can be reduced.

[0003] In engine design, increasing the flow rate can bring higher thrust, so large flow rate has become the development direction of the next-generation compression system. In blade design, the flow capacity of the blade is not only determined by the one-dimensional flow area, i.e., the geometric profile, but also the correction of aerodynamic blockage (including boundary layer and wake blockage of the front row) should be considered. The configuration of conventional blades is relatively simple and less sensitive to the boundary layer separation of the blade and the aerodynamic blockage caused by the wake. However, due to the relatively complex coupling relationship between the front row blade and the rear row blade of the tandem blade, the front row blade is affected by the geometric throat and the boundary layer aerodynamics, and the rear row blade is affected by the geometric profile, the boundary layer and the wake aerodynamics. Therefore, the tandem blade will be affected by both the geometric throat and the aerodynamic throat. There is a lack of quantitative criteria and optimization methods for the aerodynamic blockage of tandem blades (as Figure 3 shown) in the prior art.

[0004] In the case where the oncoming flow is large and the flow capacity of the blade is insufficient, the blade will reach the upper limit of the flow capacity, and the excess air flow will cause the blade to work in an overflow state, generating a strong positive shock wave at the inlet, resulting in a mismatch of the shock wave system structure of the blade, an increase in flow reduction loss, and this phenomenon is called the blade unstart phenomenon. When the aerodynamic throat of the rear row of tandem blades is blocked, the front row approximately works in an outlet throttling state, and the oncoming gas overflows at the inlet of the front row blade, causing the front row blade to be in an unstarted state, resulting in a large total pressure loss and a reduction in the flow capacity.

[0005] In view of the above technical problems, the present invention proposes a flow matching design method for supersonic tandem blades. Summary of the Invention

[0006] The object of the present invention is to overcome the deficiencies of the prior art and provide a flow matching design method for supersonic tandem blades. Based on the geometric profile coordinate point parameters of the tandem blades, a curve of the flow area change of the cascade channel is established through a geometric throat analysis program to realize the evaluation of the geometric flow capacity; the RANS simulation is used to extract the flow field characteristic parameters, and the area correction coefficient of the aerodynamic throat is solved to realize the evaluation of the aerodynamic flow capacity; by judging the flow capacity under different aerodynamic throat areas, the nominal angle of attack range of the rear blades of the tandem blades under a single working condition is determined; the relationship between the flow area of the aerodynamic throat of the tandem blades and the shock wave system structure under different incoming flow conditions is established; finally, the blade profile optimization result of the supersonic tandem blades that meets the matching relationship is obtained.

[0007] The present invention solves its technical problems through the following technical solutions:

[0008] A flow matching design method for supersonic tandem blades, comprising the following steps:

[0009] Step 1: Select the tandem blades, use the geometric throat analysis program, input the modeling parameters of the tandem blades, generate the geometric profile coordinate point parameters, and obtain the flow area of the cascade channel of the tandem blades;

[0010] Step 2: According to the flow area of the cascade channel of the tandem blades obtained in Step 1, judge whether the geometric throat of the tandem blades meets the flow requirements;

[0011] 2.1 If it meets the flow requirements, use the RANS simulation to calculate the tandem blades to obtain the area correction coefficient of the aerodynamic throat of the tandem blades;

[0012] 2.2 If it does not meet the flow requirements, change the flow area of the cascade channel of the tandem blades;

[0013] Step 3: According to the area correction coefficient of the aerodynamic throat of the tandem blades obtained in 2.1 of Step 2, calculate the flow area of the aerodynamic throat of the tandem blades after aerodynamic correction;

[0014] Step 4: According to the flow area of the aerodynamic throat of the tandem blades obtained in Step 3, judge whether the aerodynamic throat of the tandem blades meets the flow requirements;

[0015] 4.1 If it meets the flow requirements, determine the nominal angle of attack range of the rear blades of the tandem blades under a single working condition;

[0016] 4.2 If it does not meet the flow requirements, change the flow area of the cascade channel of the tandem blades;

[0017] Step 5: According to the nominal angle of attack range of the rear blades of the tandem blades under a single working condition determined in 4.1 of Step 4, adjust the incoming flow condition of the tandem blades to determine the relationship between the flow area of the aerodynamic throat of the tandem blades and the shock wave system structure;

[0018] Step 6: Output the optimized blade profile of the supersonic tandem blade according to the relationship between the flow area of the aerodynamic throat of the tandem blade determined in Step 5 and the shock wave system structure;

[0019] Step 7: According to the flow area of the cascade passage of the tandem blade changed in 2.2 in Step 2, repeat Steps 2 to 6 to output the optimized blade profile of the supersonic tandem blade;

[0020] Step 8: According to the flow area of the cascade passage of the tandem blade changed in 4.2 in Step 4, repeat Steps 2 to 6 to output the optimized blade profile of the supersonic tandem blade.

[0021] Furthermore, the implementation method of the geometric throat analysis program in Step 1 is as follows: By importing the data points on the blade heights of the front row blade and the rear row blade of the tandem blade respectively, calculate the distance between a point on the suction surface of the front row blade and the rear row blade of the tandem blade and the intersection point of the normal line of this point and the tangent line of the adjacent pressure surface respectively, so as to form the flow area of the cascade passage of the tandem blade.

[0022] Furthermore, the method for obtaining the area correction coefficient of the aerodynamic throat of the tandem blade in 2.1 of Step 2 is as follows: Use RANS simulation to calculate the tandem blade, and quantify the aerodynamic factors such as the development of the boundary layer and the diffusion of the wake through the flow rate and flow field parameters of the tandem blade, and express them with the boundary layer correction coefficient σ1 and the wake correction coefficient σ2.

[0023] Furthermore, the implementation method for determining the relationship between the flow area of the aerodynamic throat of the tandem blade and the shock wave system structure in Step 5 is as follows: Based on the flow area of the aerodynamic throat of the tandem blade obtained in Step 3, establish a mapping change relationship between the flow area of the aerodynamic throat of the tandem blade under different oncoming flow conditions and the shock wave system structure of the tandem blade, so as to determine the nominal angle of attack change range of the rear row blade of the tandem blade under all working conditions.

[0024] Furthermore, the implementation method for changing the flow area of the cascade passage of the tandem blade in 2.2 of Step 2 is as follows: Adjust the flow area of the cascade passage of the rear row blade of the tandem blade, and adopt the method of changing the nominal angle of attack of the rear row blade of the tandem blade to define the throat area of the rear row blade of the tandem blade. The nominal angle of attack of the rear row blade of the tandem blade is i AB ; Its formula is i AB =β AB -α AB ; In the formula, β AB is the inlet geometric angle of the rear row blade of the tandem blade, and α AB is the outlet geometric angle of the front row blade of the tandem blade.

[0025] Further, in step 4, the implementation method of changing the flow area of the cascade channel of the tandem blades is as follows: adjust the flow area of the cascade channel of the rear blades of the tandem blades, and define the throat area of the rear blades of the tandem blades by changing the nominal angle of attack of the rear blades of the tandem blades. The nominal angle of attack of the rear blades of the tandem blades is i AB ; its formula is i AB =β AB -α AB ; in the formula, β AB is the inlet geometric angle of the rear blades of the tandem blades, and α AB is the outlet geometric angle of the front blades of the tandem blades.

[0026] The advantages and positive effects of the present invention are:

[0027] 1. The flow matching design method of the supersonic tandem blades of the present invention includes analyzing the flow area of the cascade channel of the tandem blades by using a geometric throat analysis program, and further calculating the aerodynamic throat area of the tandem blades based on the computational fluid dynamics simulation results. Compared with the traditional method that only considers the one-dimensional flow area of the cascade channel, the aerodynamic throat area can reflect the influence of aerodynamic blockage factors such as boundary layer effect and wake diffusion on the tandem blades. Designing with the comprehensive aerodynamic correction coefficient can improve the design accuracy and efficiency.

[0028] 2. The flow matching design method of the supersonic tandem blades of the present invention is based on the flow area of the aerodynamic throat of the tandem blades, and establishes the relationship between the flow area of the aerodynamic throat of the tandem blades and the shock wave system structure. The shock wave system structure can be controlled under different oncoming flow conditions, and an optimized result of the front and rear blades with a good matching relationship can be formed. Brief Description of the Drawings

[0029] Figure 1 is the flow chart of the flow matching design method of the supersonic tandem blades of the present invention;

[0030] Figure 2 is the schematic diagram of the geometric model of the tandem blades of the present invention;

[0031] Figure 3 is the schematic diagram of the aerodynamic blockage model to be solved by the present invention;

[0032] Figure 4 is the flow chart of the geometric profile curvature analysis method of the tandem blades of the present invention;

[0033] Figure 5 is the schematic diagram of the calculation result of the flow area of the cascade channel of the tandem blades of the present invention;

[0034] Figure 6 is the schematic diagram of the definition of the nominal angle of attack of the rear blades of the tandem blades of the present invention;

[0035] Figure 7 This is the calculation result of the geometric throat area of the rear row of the tandem blades of the present invention at different nominal angles of attack of the rear row;

[0036] Figure 8 This is the flow rate - total pressure ratio characteristic diagram of the tandem blades of the present invention at different nominal angles of attack of the rear row;

[0037] Figure 9 This is the schematic diagram of the front and rear flow fields of the optimized tandem blades of the present invention. Detailed implementation manners

[0038] The present invention will be further described in detail below through specific embodiments. The following embodiments are only descriptive and not restrictive, and the protection scope of the present invention cannot be limited thereby.

[0039] A flow matching design method for supersonic tandem blades includes the following steps:

[0040] Step 1: Select tandem blades, use a geometric throat analysis program, input the modeling parameters of the tandem blades, generate geometric profile coordinate point parameters, and obtain the flow - through area of the cascade channel of the tandem blades;

[0041] Among them, the tandem blades are a planar cascade layout composed of a front - row blade and a rear - row blade. The flow - through area of the cascade channel is the size of the gap between the blades in the direction perpendicular to the on - coming flow, as Figure 2 shown;

[0042] The implementation manner of the geometric throat analysis program is as follows: By importing the data points on the blade heights of the front - row blade and the rear - row blade of the tandem blades respectively, the blade heights of the front - row blade and the rear - row blade of the tandem blades are the blade parameters under the basic element level. Calculate the distance between a point on the suction surface of the front - row blade and the rear - row blade of the tandem blades and the intersection point of the normal line of this point and the tangent line of the adjacent pressure surface as the flow - through area. Traverse all the suction - surface data points to form the flow - through area of the cascade channel of the tandem blades, form the flow - through areas at the overlapping positions of the front - row blade, the rear - row blade, and the front - and - rear row blades, and select some of the calculation results of the flow - through areas as Figure 5 shown.

[0043] Step 2: According to the flow - through area of the cascade channel of the tandem blades obtained in Step 1, judge whether the geometric throat of the tandem blades meets the flow requirements;

[0044] 2.1 If it meets the flow requirements, use RANS simulation to calculate the tandem blades to obtain the area correction coefficient of the aerodynamic throat of the tandem blades;

[0045] The specific method is as follows: As Figure 3As shown in the figure, considering the complex flow environment of the tandem blades, the aerodynamic blockage effect of the tandem blades is analyzed by dividing it into four characteristic cross-sections of Area1 - 4. Among them, the Area1 cross-section is the incoming flow area of the tandem blades and is not affected by aerodynamic blockage; Area2 is the area affected by the front-row blades and is affected by boundary layer blockage; Area3 is the coupling area between the front-row blades and the rear-row blades, and Area4 is the area affected by the rear-row blades, which is affected by the wake blockage of the front-row blades and the boundary layer blockage of the rear-row blades. The RANS simulation is used to calculate the tandem blades. Through the flow rate and flow field parameters of the tandem blades, the aerodynamic factors such as the development of the boundary layer and the diffusion of the wake are quantified and expressed by the boundary layer correction coefficient σ1 and the wake correction coefficient σ2. The boundary layer correction coefficient σ1 and the wake correction coefficient σ2 are defined as the normalized ratios of the aerodynamic blockage effect to the flow area of the cascade channel of the tandem blades obtained in Step 1, and are used to characterize the dynamic attenuation characteristics of the flow capacity under actual working conditions. The boundary layer correction coefficient σ1 refers to the dynamic attenuation characteristics of the flow capacity caused by the boundary layer blockage of the front-row blades and the rear-row blades of the tandem blades, and the wake correction coefficient σ2 refers to the dynamic attenuation characteristics of the flow capacity caused by the wake separation of the front-row blades of the tandem blades.

[0046] 2.2. If the flow demand is not met, change the flow area of the cascade channel of the tandem blades;

[0047] The implementation method is: adjust the flow area of the cascade channel of the rear-row blades of the tandem blades, and define the throat area of the rear-row blades of the tandem blades by changing the nominal angle of attack of the rear-row blades of the tandem blades. The nominal angle of attack of the rear-row blades of the tandem blades is i AB ; its formula is i AB = β AB - α AB ; in the formula, β AB is the inlet geometric angle of the rear-row blades of the tandem blades, and α AB is the outlet geometric angle of the front-row blades of the tandem blades.

[0048] Step 3. According to the area correction coefficient of the aerodynamic throat of the tandem blades obtained in 2.1 of Step 2, calculate the flow area of the aerodynamic throat of the aerodynamically corrected tandem blades;

[0049] Step 4. According to the flow area of the aerodynamic throat of the tandem blades obtained in Step 3, judge whether the aerodynamic throat of the tandem blades meets the flow demand;

[0050] 4.1. If the flow demand is met, determine the range of the nominal angle of attack of the rear-row blades of the tandem blades under a single working condition;

[0051] 4.2. If the flow demand is not met, change the flow area of the cascade channel of the tandem blades;

[0052] The implementation method is as follows: adjust the flow area of the cascade passage of the rear-stage blades of the tandem blades, and define the throat area of the rear-stage blades of the tandem blades by changing the nominal angle of attack of the rear-stage blades of the tandem blades. The nominal angle of attack of the rear-stage blades of the tandem blades is i AB ; its formula is i AB =β AB -α AB ; in the formula, β AB is the inlet geometric angle of the rear-stage blades of the tandem blades, and α AB is the outlet geometric angle of the front-stage blades of the tandem blades.

[0053] Step 5: According to the nominal angle of attack range of the rear-stage blades of the tandem blades determined in 4.1 in Step 4, adjust the oncoming flow conditions of the tandem blades, and determine the relationship between the flow area of the aerodynamic throat of the tandem blades and the shock wave system structure;

[0054] The implementation method is as follows: based on the flow area of the aerodynamic throat of the tandem blades obtained in Step 3, by establishing a mapping relationship between the flow area of the aerodynamic throat of the tandem blades under different oncoming flow conditions and the shock wave system structure of the tandem blades, determine the change range of the nominal angle of attack of the rear-stage blades of the tandem blades under all working conditions.

[0055] Step 6: According to the relationship between the flow area of the aerodynamic throat of the tandem blades and the shock wave system structure determined in Step 5, output the optimized blade profile results of the supersonic tandem blades;

[0056] Step 7: According to the flow area of the cascade passage of the tandem blades changed in 2.2 in Step 2, repeat Steps 2 to 6 to output the optimized blade profile results of the supersonic tandem blades;

[0057] Step 8: According to the flow area of the cascade passage of the tandem blades changed in 4.2 in Step 4, repeat Steps 2 to 6 to output the optimized blade profile results of the supersonic tandem blades.

[0058] As Figure 1 shown, this embodiment takes the flow matching design process of the front and rear rows at the outlet stage of the fan with a tandem rotor / stator aerodynamic layout in an axial compressor as an example to illustrate the specific implementation of the present invention, including the following steps:

[0059] Step 1: First, use a geometric throat analysis program to analyze the flow area of the cascade passage of the tandem blades, combined with Figure 4 For Figure 1A further description is given of the flow area of the cascade passage of the tandem blades obtained using the geometric throat analysis program. First, a geometric surface coordinate point file containing blade height, geometric surface coordinate point parameters, and the number of blades is generated from the blade modeling file. The coordinate points of the tandem blade data are divided into the coordinate points at each blade height of the tandem blades, that is, the coordinate points at the elementary stage, and the front-row blades and the rear-row blades are distinguished. According to the geometric surface coordinate point parameters, the number of elementary stages is set as i, and the pitch is calculated through the number of blades of the tandem blades and the blade radii of the front-row blades and the rear-row blades to obtain the coordinate points of adjacent blades. The data coordinate points are divided into the pressure surface and the suction surface. The number of suction surface coordinate points is set as j, and the j coordinate points are traversed in sequence. The intersection point of the normal line at this point on the suction surface and the tangent line of the adjacent pressure surface is calculated, and the distance from a point on the suction surface to the above intersection point is the flow area of the cascade passage at this point. The flow areas of the cascade passages between the front-row blades and the front-row blades, between the front-row blades and the rear-row blades, and between the rear-row blades and the rear-row blades are calculated in sequence, and all the flow area parameters of the cascade passages are output. This process is repeated until all i elementary stages are traversed to obtain the flow area of the cascade passage of the blade profile at each elementary stage, forming the flow area of the cascade passage of the tandem blades. The flow capacity of this blade profile is detected through the aerodynamic function to meet the flow requirements.

[0060] As Figure 6 shown, by modifying the nominal angle of attack of the rear-row blades to adjust the flow area of the cascade passage of the rear-row blades, the smaller the nominal angle of attack, the smaller the flow area of the rear-row cascade passage. The nominal angles of attack of the rear-row blades of the tandem blades are respectively set in the range of -4.5° to 2°, where Baseline represents the prototype scheme without modifying the nominal angle of attack. As Figure 7 shown, for the minimum values of the flow areas of the cascade passages at 10%, 50%, and 90% of the blade height after changing the nominal angle of attack of the rear-row blades, that is, the throat areas, based on the geometric throat area at a nominal angle of attack of 0°, the ratio of the throat area of the rear-row blades after adjusting the nominal angle of attack to the throat area at an angle of attack of 0° is obtained.

[0061] Step 2: According to the flow area of the cascade passage of the tandem blades obtained in Step 1, determine whether the geometric throat of the tandem blades meets the flow requirements;

[0062] 2.1. If it meets the flow requirements, use RANS simulation to calculate the tandem blades to obtain the area correction coefficient of the aerodynamic throat of the tandem blades;

[0063] The generated blade geometric parameters are input into the AutoGrid5 module of IGG to automatically generate a three-dimensional calculation model. Boundary conditions such as rotational speed, back pressure, and pre-whirl angle are set through the CFX commercial computational fluid dynamics software for RANS numerical simulation. Flow field parameters are extracted based on the simulation results, including the boundary layer velocity distribution, total pressure loss coefficient in the wake region, and shock wave position. By quantifying the correlation function between boundary layer blockage and the wake blockage of the front row, the boundary layer correction coefficients σ1 of the front and rear blades of the tandem blades and the wake correction coefficient σ2 of the tandem blades are dynamically calculated; the boundary layer correction coefficient σ1 refers to the dynamic attenuation characteristic of the flow capacity caused by the boundary layer blockage of the front and rear blades of the tandem blades, and the wake correction coefficient σ2 refers to the dynamic attenuation characteristic of the flow capacity caused by the wake separation of the front blade of the tandem blades.

[0064] 2.2. If the flow requirements are not met, adjust the flow area of the cascade passage of the rear blade of the tandem blades. Modify the throat area of the rear blade of the tandem blades by changing the nominal angle of attack of the rear blade of the tandem blades. The nominal angle of attack of the rear blade of the tandem blades is i AB ; its formula is i AB =β AB -α AB ; in the formula, β AB is the inlet geometric angle of the rear blade of the tandem blades, and α AB is the outlet geometric angle of the front blade of the tandem blades. After changing the flow area of the cascade passage of the tandem blades, it is necessary to re-judge whether the geometric throat of the tandem blades meets the flow requirements to ensure that the geometric throat of the tandem blades meets the flow requirements.

[0065] Step 3: Calculate the flow area of the aerodynamic throat of the tandem blades after aerodynamic correction according to the area correction coefficient of the aerodynamic throat of the tandem blades obtained in 2.1 of Step 2;

[0066] Step 4: Judge whether the aerodynamic throat of the tandem blades meets the flow requirements according to the flow area of the aerodynamic throat of the tandem blades obtained in Step 3;

[0067] 4.1. If the flow requirements are met, determine the range of the nominal angle of attack of the rear blade of the tandem blades under a single condition;

[0068] Establish a single-condition flow state evaluation process with the flow threshold determination as the core. The RANS simulation results of Step 2 are as Figure 8 shown. For the -4.5° scheme, the maximum flow rate is significantly lower than the prototype flow rate, and the flow rate of the front blade hardly changes, while the flow rate of the rear blade decreases significantly; as Figure 9As shown, through the flow field diagram, it can be further determined that a shock wave system structure of front row overflow appears at this angle. Therefore, it can be recognized that the aerodynamic throat of the rear row blades is smaller than the flow demand when the angle is less than -3°. Therefore, the nominal angle of attack of the rear row blades of this tandem blade is the critical value of the flow area of the rear row cascade channel at -3°. Under this working condition, when +2° ≥ i AB ≥ -3°, the plug point flow rate is in a stable state.

[0069] 4.2. If the flow demand is not met, adjust the flow area of the cascade channel of the rear row blades of the tandem blade. By changing the nominal angle of attack of the rear row blades of the tandem blade to modify the throat area of the rear row blades of the tandem blade, the nominal angle of attack of the rear row blades of the tandem blade is i AB ; its formula is i AB = β AB - α AB ; where β AB is the inlet geometric angle of the rear row blades of the tandem blade, and α AB is the outlet geometric angle of the front row blades of the tandem blade. Change the flow area of the cascade channel of the tandem blade, and repeat steps 2 to 4 to re-judge whether the aerodynamic throat of the tandem blade meets the flow demand.

[0070] Step 5. According to the nominal angle of attack range of the rear row blades of the tandem blade determined in 4.1 of step 4, adjust the oncoming flow condition of the tandem blade to determine the relationship between the flow area of the aerodynamic throat of the tandem blade and the shock wave system structure;

[0071] The implementation method is: based on the flow area of the aerodynamic throat of the tandem blade obtained in step 3, by establishing a mapping change relationship between the flow area of the aerodynamic throat of the tandem blade under different oncoming flow conditions and the shock wave system structure of the tandem blade, determine the change range of the nominal angle of attack of the rear row blades of the tandem blade under full working conditions.

[0072] Determine the range of the aerodynamic throat of the rear row blades. By repeating steps 2 to 5 for different oncoming flow conditions, the calculation result is: the range of the nominal angle of attack of the rear row blades under different oncoming flow conditions is +2° ≥ i AB ≥ -3°, and the shock wave system of the blade has a good structure. Therefore, this range is selected as the range of the nominal angle of attack of the rear row blades.

[0073] Step 6. According to the relationship between the flow area of the aerodynamic throat of the tandem blade and the shock wave system structure determined in step 5, output the blade profile optimization result of the supersonic tandem blade;

[0074] Summarize the optimization results of the supersonic tandem blades. The nominal angle of attack of the rear blades is obtained through Step 5, and the one-dimensional blade characteristics such as flow rate, pressure ratio, and efficiency within this range are formed. From this, it can be obtained that within this range, the influence of the tandem blades on the pressure ratio is 4%, and the influence on the flow rate is 5%. Within this interval, a supersonic tandem blade with a stable and controllable shock wave system can be obtained.

[0075] Although embodiments and drawings of the present invention are disclosed for illustrative purposes, those skilled in the art can understand that various substitutions, changes, and modifications are possible without departing from the spirit and scope of the present invention and the appended claims. Therefore, the scope of the present invention is not limited to the content disclosed in the embodiments and drawings.

Claims

1. A flow matching design method for a supersonic tandem blade, characterized in that: The method includes the following steps: Step 1: Select tandem blades, use a geometric throat analysis program, input the shaping parameters of the tandem blades, generate geometric profile coordinate point parameters, and obtain the flow area of the cascade channel of the tandem blades; Step 2: According to the flow area of the cascade channel of the tandem blades obtained in Step 1, determine whether the geometric throat of the tandem blades meets the flow requirements; 2.1 If the flow requirements are met, use RANS simulation to calculate the tandem blades to obtain the area correction coefficient of the aerodynamic throat of the tandem blades; 2.2 If the flow requirements are not met, change the flow area of the cascade channel of the tandem blades; Step 3: According to the area correction coefficient of the aerodynamic throat of the tandem blades obtained in 2.1 of Step 2, calculate the flow area of the aerodynamic throat of the tandem blades after aerodynamic correction; Step 4: According to the flow area of the aerodynamic throat of the tandem blades obtained in Step 3, determine whether the aerodynamic throat of the tandem blades meets the flow requirements; 4.1 If the flow requirements are met, determine the nominal angle of attack range of the rear blades of the tandem blades under single operating conditions; 4.2 If the flow requirements are not met, change the flow area of the cascade channel of the tandem blades; Step 5: According to the nominal angle of attack range of the rear blades of the tandem blades under single operating conditions determined in 4.1 of Step 4, adjust the incoming flow conditions of the tandem blades to determine the relationship between the flow area of the aerodynamic throat of the tandem blades and the shock wave system structure; Step 6: According to the relationship between the flow area of the aerodynamic throat of the tandem blades and the shock wave system structure determined in Step 5, output the blade profile optimization results of the supersonic tandem blades; Step 7: According to the flow area of the cascade channel of the tandem blades changed in 2.2 of Step 2, repeat Steps 2 to 6 to output the blade profile optimization results of the supersonic tandem blades; Step 8: According to the flow area of the cascade channel of the tandem blades changed in 4.2 of Step 4, repeat Steps 2 to 6 to output the blade profile optimization results of the supersonic tandem blades.

2. The flow matching design method of a supersonic tandem blade according to claim 1, characterized in that: The implementation method of the geometric throat analysis program in Step 1 is as follows: By importing the data points on the blade heights of the front and rear blades of the tandem blades respectively, calculate the distance between a point on the suction surface of the front and rear blades of the tandem blades and the intersection point of the normal line of this point and the tangent line of the adjacent pressure surface respectively, so as to form the flow area of the cascade channel of the tandem blades.

3. The flow matching design method of a supersonic tandem blade according to claim 1, characterized in that: The method for obtaining the area correction coefficient of the aerodynamic throat of the tandem blades in 2.1 of Step 2 is as follows: Use RANS simulation to calculate the tandem blades, and quantify the aerodynamic factors such as the development of the boundary layer and the diffusion of the wake through the flow rate and flow field parameters of the tandem blades, and express them with the boundary layer correction coefficient σ1 and the wake correction coefficient σ2.

4. The flow matching design method of a supersonic tandem blade according to claim 1, characterized in that: The implementation method for determining the relationship between the flow area of the aerodynamic throat of the tandem blades and the shock wave system structure in Step 5 is as follows: Based on the flow area of the aerodynamic throat of the tandem blades obtained in Step 3, establish a mapping change relationship between the flow area of the aerodynamic throat of the tandem blades under different incoming flow conditions and the shock wave system structure of the tandem blades, so as to determine the nominal angle of attack change range of the rear blades of the tandem blades under all operating conditions.

5. A flow matching design method for a supersonic tandem blade according to claim 1, characterized in that: In the step 2, the implementation method of changing the flow area of the cascade passage of the tandem blades is as follows: adjust the flow area of the cascade passage of the rear blades of the tandem blades, and define the throat area of the rear blades of the tandem blades by changing the nominal angle of attack of the rear blades of the tandem blades. The nominal angle of attack of the rear blades of the tandem blades is i AB ; its formula is i AB =β AB -α AB ; in the formula, β AB is the inlet geometric angle of the rear blades of the tandem blades, and α AB is the outlet geometric angle of the front blades of the tandem blades.

6. The flow matching design method of a supersonic tandem blade according to claim 1, characterized in that: In step 4.2, the implementation method of changing the flow area of the cascade channel of the tandem blades is as follows: adjust the flow area of the cascade channel of the rear blades of the tandem blades, and define the throat area of the rear blades of the tandem blades by changing the nominal angle of attack of the rear blades of the tandem blades. The nominal angle of attack of the rear blades of the tandem blades is i AB ; its formula is i AB =β AB -α AB ; in the formula, β AB is the inlet geometric angle of the rear blades of the tandem blades, and α AB is the outlet geometric angle of the front blades of the tandem blades.

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