High-load supersonic compressor blade profile excited by coupling multi-point plasmas

Abstract

The invention discloses a high-load supersonic compressor blade profile design method coupled with multi-point plasma excitation, which comprises the following steps: generating high-load large-bend blade profile geometric data on the basis of original geometric data by using a blade profile parameterization modeling method; the method comprises the following steps: acquiring compressor blade channel flow field characteristics under a low Reynolds number working condition through numerical simulation, and determining a blade surface flow separation area under the low Reynolds number; determining a multi-point rapid heating plasma excitation layout mode according to the characteristics of the flow separation area; and analyzing the flow control effect of the multi-point rapid heating plasma excitation on the blade type shock wave / boundary layer interference of the supersonic gas compressor. According to the method, the K-H instability of a shear layer is triggered by applying multi-point rapid heating plasma excitation in a flow field, a spanwise vortex array structure is induced to be formed, small-scale vortexes in a flow separation area are promoted to be mixed and evolved into a spanwise large vortex structure so as to accelerate dissipation, and the effect of inhibiting flow separation is achieved.

Description

Technical Field

[0001] The invention relates to aviation plasma and flow control technology, and in particular to a high-load supersonic compressor blade design method coupled with multi-point plasma excitation. Background Art

[0002] Advanced aircraft engines have high thrust-to-weight ratio, low fuel consumption rate, and high reliability requirements, which constantly push the compression components of aircraft engines, namely the compressor, to increase the stage pressure ratio and pursue higher working efficiency. High-load large-angle blades can increase the single-stage pressure ratio, but too large blade angles can easily cause serious flow separation, resulting in an increase in the total pressure loss coefficient and a decrease in efficiency. Therefore, increasing the stage pressure ratio of the engine while reducing the flow separation loss is of great significance for increasing the thrust-to-weight ratio of aircraft engines, reducing fuel consumption rate, and improving reliability.

[0003] Wide-range compressors will face a prominent low Reynolds number problem when operating at high altitudes. Under this condition, the laminar boundary layer on the surface of the compressor blades will often induce laminar boundary layer separation under the influence of the adverse pressure gradient and form a flow separation zone, resulting in a sharp increase in compressor flow losses and a significant decrease in efficiency and stability margin. Especially for supersonic compressors, the unique shock wave / boundary layer interference problem in the supersonic flow field will make the internal flow of the supersonic compressor more complicated, and the flow separation loss will also be more serious. The existing low-energy consumption rapid heating plasma excitation and regulation of supersonic compressor blade shock waves / boundary layers have problems such as low injection energy and poor flow field disturbance effect. Therefore, it is urgent to develop advanced flow control technology to ensure that supersonic compressors can maintain good aerodynamic performance in a wide range of airspace.

[0004] The present invention combines the above two needs. On the basis of the original supersonic blade geometry data, the present invention uses a blade parameterization modeling method to change the blade bending angle to generate high-load supersonic compressor blade geometry data, and expects to achieve strong interference to the supersonic flow field by optimizing the rapid heating plasma excitation energy injection method, weaken the negative impact of high-load supersonic compressor blade shock wave / boundary layer interference, and achieve significant suppression of high-load supersonic compressor blade surface flow separation under low Reynolds number, providing innovative ideas for the development of high-performance compressors. Summary of the invention

[0005] In view of the problems existing in the prior art, the present invention provides a high-load supersonic compressor blade design method coupled with multi-point plasma excitation, comprising the following steps:

[0006] Step 1: Generate high-load large-angle blade geometry data by using blade parameterization modeling method based on original geometry data;

[0007] Step 2: obtain the flow field characteristics of the compressor blade channel under low Reynolds number conditions through numerical simulation, and determine the flow separation area on the blade surface under low Reynolds number;

[0008] Step 3: Determine the multi-point rapid heating plasma excitation layout mode according to the flow separation area characteristics;

[0009] Step 4: Analysis of the flow control effect of multi-point rapid heating plasma excitation on supersonic compressor blade shock wave / boundary layer interference.

[0010] In one embodiment of the present invention, step 1 is to generate high-load large-angle blade geometry data on a two-dimensional plane by adjusting parameters using a blade parameterization modeling method based on the original blade geometry data; the specific steps are as follows:

[0011] (1) Interpolation of original blade profile data: The blade profile curve includes the leading edge curve, the trailing edge curve, the suction surface curve and the pressure surface curve. These curves are interpolated using the B-spline interpolation algorithm to obtain relatively dense geometric data of the suction surface and pressure surface curves;

[0012] (2) Determine the geometric points and thickness distribution of the mid-arc line; for the suction surface curve, connect the point with all the points on the pressure surface curve, search for the point on the pressure surface curve that is closest to the point, connect the two points, and take the middle point as the point on the mid-arc line; repeat the above steps for all the geometric points on the suction surface curve to obtain the complete distribution of the mid-arc line points, and connect all the mid-arc line points to obtain the mid-arc line curve; while generating the mid-arc line, also obtain the point on the pressure surface curve that is the shortest distance from any point on the suction surface curve, and the midpoint of the line connecting the two points is the mid-arc line. The arc point, and 1 / 2 of the length of the line is defined as the blade thickness at the mid-arc point, thereby obtaining the blade thickness distribution; for the above method of determining the mid-arc and thickness, the angle between the line connecting the suction surface point and the pressure surface point and the tangent at the mid-arc point should be fixed, and the geometric coordinates of the suction surface point and the pressure surface point are used to calculate the slope of the line connecting the two relative to the horizontal direction, and the geometric coordinates of each point on the mid-arc are used to calculate the slope of the tangent direction at the mid-arc point relative to the horizontal direction; since the angle between the two lines is fixed, the difference between their slopes must also be fixed;

[0013] (3) Parametric expression of the center arc geometry: Fitting the center arc points generated in step (2) with a cubic piecewise polynomial to obtain a parametric expression of the center arc;

[0014] (4) parametric expression of thickness distribution; the parameter expression method of the thickness distribution curve adopts a cubic piecewise polynomial method to fit the blade thickness data at each mid-arc point generated in step (2) to obtain a parametric expression of the thickness distribution;

[0015] (5) Generate the geometry of the pressure surface and suction surface; Given the coordinates of any horizontal direction between the leading edge point and the trailing edge point on the suction surface curve of the blade, obtain the coordinates of the mid-arc point by fitting the mid-arc polynomial, and at the same time obtain the blade thickness at the coordinate point; for any mid-arc point on the mid-arc, use its geometric coordinates to calculate the slope of the tangent direction at the mid-arc point relative to the horizontal direction; determine the direction of the line connecting the corresponding suction surface point and pressure surface point according to the slope value; express the polynomial according to the parameterization of the thickness distribution curve, superimpose the thickness on the basis of the mid-arc coordinates, and obtain all the geometric points of the pressure surface and suction surface;

[0016] (6) Generate the leading and trailing edges of the blade; generate a total of four control points of the leading edge and the trailing edge using the original blade geometry data and the distance and angle relationship, and then generate the leading and trailing edge curve expression using a cubic Bezier curve.

[0017] In another embodiment of the present invention, in step 1 (3), a total of 12 segments of cubic polynomials are used to fit the points on the arc, and each segment is evenly distributed in length; for each cubic polynomial, a total of 4 constraints are required to determine its coefficients, and the parameterized expression polynomial is denoted as p i (x), then:

[0018] p i (x) = a 1 (x m ) 3 +a 2 (x m ) 2 +a 3 x m +a 4

[0019] Among them, a 1 、a 2 、a 3 、a 4 represent the coefficients of the cubic polynomial, x m Indicates the axial coordinates of the control points of the mid-arc curve;

[0020] The following are specific instructions:

[0021] The first segment selects three control points. The first control point (x 1 m1 ,y 1 m1 ) is the starting point of the middle arc and also the starting point of the segmented area; the coordinates and slope of the first control point are used as two constraints of the polynomial fitting; the third control point (x 1 m3 ,y 1 m3) is the end point coordinate of the segment. The coordinate of this point is used as a constraint condition of the polynomial fitting. For the second control point (x 1 m2 ,y 1 m2 ), whose value is not given in advance, but is obtained by optimizing the coordinates and slope of the first control point and the coordinates of the third control point with the minimum fitting error as the goal, and using its coordinates as the last constraint of the polynomial fitting; using the above four constraints, the analytical solution of the polynomial coefficients of the parameterized expression of the arc in the first segment is obtained;

[0022]

[0023] From the second to the eleventh segment, three control points are also selected for each segment, denoted as (x i m1 ,y i m1 )(x i m2 ,y i m2 )(x i m3 ,y i m3 ), i=2,…,11; the first control point is the end point of the previous curve, the third control point is the starting point of the next curve, and the second control point is also optimized with the minimum fitting error as the goal; different from the first section, in the arc fitting process from the second to the eleventh section, the slope constraint of the first point in each section is changed to the second-order derivative constraint of the arc curve at the first point to ensure the continuity of the second-order curvature of the arc at the junction of each section, and the other three constraints are the same; the above four constraints are used to obtain the analytical solution of the polynomial coefficients of the parameterized expression of the arc in any intermediate section;

[0024]

[0025] When the last segment is generated, two control points are taken and recorded as (x N m1 ,y N m1 )(x N m2 ,y N m2); the first control point is the end point of the previous curve, the second point is the end point of the mid-arc, and in addition to the coordinates of the above two control points, the second-order derivative at the first point and the first-order derivative at the last point are used as the other two constraints, where the first-order derivative at the last point is specified by the slope value corresponding to the blade profile outlet geometry angle; using the above four constraints, the analytical solution of the polynomial coefficients of the parameterized expression of the last mid-arc is obtained;

[0026]

[0027] In another embodiment of the present invention, in step 1 (4), the parameter expression method of the thickness distribution curve adopts a cubic piecewise polynomial method, and the entire thickness distribution area is divided into four segments. The first and fifth control points are respectively the leading edge point and the trailing edge point, the second and third control points are respectively the starting point and the end point of the pre-compression segment, and the fourth control point is the equal-division point of the line segment between the third control point and the fifth control point in the axial direction; the parameterized expression polynomial of the thickness distribution is denoted as h i (x), then:

[0028] h i (x) = b 1 (x h ) 3 +b 2 (x h ) 2 +b 3 x h +b 4

[0029] Among them, b 1 、b 2 、b 3 、b 4 represent the coefficients of the cubic polynomial, x h Represents the axial coordinate of the control point of the thickness distribution curve;

[0030] The following are specific instructions:

[0031] In the first section, the number of control points is 4. According to the thickness distribution curve, the first point is the starting point of the middle arc, and the fourth point is the starting point of the thickness distribution of the next section. The second and third control points are the same as step (3), which are optimized with the minimum fitting error as the goal, and are evenly spaced between the first and fourth points. The coordinates of the above four points are used as constraints to obtain the analytical solution of the polynomial coefficients of the parameterized expression of the first section thickness distribution curve;

[0032]

[0033] The constraints required for the thickness distribution of the second to fourth sections are different from those of the first section. For each section, the first step is to determine the coordinates of the starting point and the end point of the thickness distribution curve of the corresponding section as two constraints, and then calculate the first-order derivative at the end point of the previous thickness distribution curve as a constraint. After obtaining the above three constraints, the intermediate control point is obtained by optimization with the minimum fitting error as the goal as a constraint. Using the above four constraints, the analytical solution of the polynomial coefficients of the parameterized expression of the thickness distribution curve of the second to fourth sections is obtained;

[0034]

[0035] In another embodiment of the present invention, in step 1 (6), for the front edge curve, the first point of the mid-arc is first set as the center of the front edge (x le0 ,y le0 ), the minimum distance between the leading edge of the original blade and the center of the circle is obtained by traversal method, denoted as R le ; Select the points closest to the suction surface and pressure surface curves from the original blade leading edge geometry points and record them as control points 1(x le1 ,y le1 ) and control point 4(x le4 ,y le4 ), and then solve the slopes between the starting points of the suction surface and pressure surface curves and the corresponding leading edge control points, and find control points 2 and 3 along the slope direction; the distances between control points 2 and 3 and control points 1 and 4 refer to the aforementioned radius R le The specific value is obtained by optimizing with the minimum fitting error as the goal; after obtaining the above four control points, a Bezier curve that can parameterize the front edge curve is generated by these four points.

[0036] In another embodiment of the present invention, step 2 uses a large eddy simulation numerical simulation method to perform high-precision simulation of the compressor blade channel flow under low Reynolds number conditions;

[0037] The computational domain grid is generated using multi-block grid technology. The inlet and outlet sections and the cascade channel are divided using H-type topology grids, and the blade surface is divided using O-type topology grids. The leading and trailing edges of the blade, the blade channel, and the boundary layer of the blade surface are encrypted by adjusting the number of grid nodes.

[0038] During the simulation, in order to simulate the ideal infinite blade cascade model, the translation periodic boundary condition is adopted in the cascade pitch direction; the calculation domain inlet plane is selected at A times the axial chord length from the leading edge of the blade, and the total temperature, total pressure and inlet airflow angle of the blade cascade inlet plane are given according to the actual working conditions; the calculation domain outlet in the numerical simulation is selected at B times the axial chord length from the trailing edge of the blade, and the average static pressure is given as the boundary condition of the blade cascade outlet plane; the calculation domain is extended at the inlet by C times the axial chord length; the ideal gas model is used for the fluid medium, and the viscosity cannot be considered as a constant, and the Sutherland law is used to regulate the viscosity; in the medium model, the nonlinear polynomial fitting method is used to calculate the specific heat capacity C p , the discrete data of thermal conductivity K changing with temperature are fitted into a polynomial function; the blade surface is set as an adiabatic non-slip stationary wall strip, the blade span length is given as D times the axial chord length, and the blade span uses symmetric boundary conditions to ensure that the axial density flow ratio of the blade channel is equal to E; the mid-blade section is selected as the characteristic section to display and analyze the transient structure in the flow channel;

[0039] A variable time step solution is performed, and the Courant number during the solution is not greater than F; when nanosecond pulse plasma excitation is applied, the simulation time step is set to G nanoseconds, and the simulation time step becomes H nanoseconds after the excitation is applied; first, I cycles of excitation are applied to the reference flow field, and after the excitation flow field is stabilized, the duration of I excitation cycles is solved as a statistical sample of the time-averaged data; the time-averaged flow field parameters are obtained by numerically averaging the transient flow field parameters of the last I flow cycles after the excitation flow field is stabilized.

[0040] In a specific embodiment of the present invention,

[0041] The sonic cascade DRL-PAV-1.5 was selected as the research object of numerical simulation. The large eddy simulation method LES was used to perform high-precision simulation of the compressor blade channel flow under low Reynolds number conditions. The analyzed conditions were: the incoming flow Mach number was equal to 1.52, the pressure ratio was 2.2, and the incoming flow Reynolds number was 1.8×10 5 ;

[0042] The dimensionless scale division of the grid along the span direction, normal direction and flow direction is no more than 30, 1 and 15, and the total number of grids is 2.5×10 6 ;

[0043] ANSYS-Fluent 18.0 commercial fluid calculation software was used for simulation, and the simulation model was selected as the LESSmagorinsky-Lilly model;

[0044] A=1.5; B=2; C=0.5; D=0.1; E=1.0; F=0.95; G=5; H=50; I=20.

[0045] In addition, in one embodiment of the present invention, in step 3, nanosecond pulsed plasma excitation Pulsed-DBD plasma excitation is selected as the excitation mode of flow control; a simulation model characterizing the Pulsed-DBD plasma excitation effect is established based on a phenomenological heat source model; the energy Q injected per unit electrode length during a single discharge pulse time by the simulated plasma excitation is calculated by the following formula:

[0046]

[0047] Where ε is the relative dielectric constant of the barrier material, V pp and V bd Represent the excitation voltage amplitude and the breakdown voltage of the medium, l s and d represent the length of plasma column and the thickness of barrier material respectively. Based on the continuous induced vortex phenomenon generated by plasma excitation, the effective excitation frequency f is proposed. eff Determine the criteria, the specific formula is as follows:

[0048]

[0049] Where, T 2 represents the time interval from the excitation application position to the complete dissipation of the vortex structure induced by plasma excitation in the separation zone; T 1 represents the time interval between the DFS moving from the excitation application position to the position where the complete spanwise vortex structure is induced; based on the effective excitation frequency criterion, a preliminary evaluation of the flow field state is performed, and a high-frequency excitation frequency f is selected for calculation, and the corresponding excitation period T 0 =1 / f;

[0050] The plasma actuator is arranged upstream of the pressure surface shock wave / boundary layer interference zone; according to the number of excitation points N and the excitation spacing l, they are arranged in sequence from the excitation reference point to the upstream and named in sequence, namely PS-Actu-1, PS-Actu-2,...

[0051] In another specific embodiment of the present invention, a polyimide material is selected as the blocking medium, and a single discharge pulse time t f =50ns; the reference excitation energy in a single plasma excitation pulse is set to 0.14mJ / cm; the excitation frequency f=f eff =10kHz high-frequency excitation is used for calculation; the multi-point layout excitation reference point PS-Actu-1 is set at a position of 0.27 times the axial chord length downstream of the leading edge of the blade.

[0052] In addition, in one embodiment of the present invention, in step 4, the parameters such as the position of the excitation reference point, the number of excitation points N, the excitation spacing l, the energy of the excitation point, etc. can be adjusted, and the effects of these parameters on the flow control effect are mutually coupled;

[0053] The total pressure loss coefficient averaged over the blade midsection is defined as follows:

[0054]

[0055] In the formula, and are the time-averaged total pressure at the inlet and the local total pressure of the flow field, respectively. is the average static pressure at the inlet of the flow field; if Change to indicate the average total pressure at the outlet The time-averaged total pressure loss coefficient after the grid can be obtained; the relative change of the time-averaged total pressure loss coefficient is defined as Δω=ω D -ω L , where ω D is the time-averaged total pressure loss coefficient under multi-point rapid heating plasma excitation conditions, ω L is the time-averaged total pressure loss coefficient at low Reynolds numbers.

[0056] The present invention triggers the shear layer KH instability and induces the formation of a spanwise vortex structure by applying multi-point rapid heating plasma excitation on the blade pressure surface. The structure promotes the mixing of small-scale vortices in the flow separation zone and evolves into a spanwise large vortex structure to accelerate dissipation, thereby achieving the effect of suppressing flow separation. BRIEF DESCRIPTION OF THE DRAWINGS

[0057] Figure 1 is the average wall axial shear force distribution of the compressor blade surface under low Reynolds number obtained by simulation, where Figure 1 (a) shows the average wall axial shear force distribution on the suction side. Figure 1 (b) shows the average wall axial shear force distribution on the pressure surface;

[0058] Figure 2 It is a schematic diagram of the layout position of the multi-point rapid heating plasma actuator used in the present invention;

[0059] Figure 3 It is the cloud diagram of the change of the total pressure loss coefficient averaged over the middle section of the blade after applying multi-point rapid heating plasma excitation;

[0060] Figure 4 It is the comparison between the distribution curve of the average total pressure loss coefficient along the grating direction during multi-point rapid heating plasma excitation and the reference state;

[0061] Figure 55(a) shows the vorticity cloud diagram of the mid-section of the leaf at different times after applying multi-point rapid heating plasma excitation, wherein 5(b) shows the vorticity cloud diagram of the mid-section of the leaf 100.00μs after applying the excitation, 5(c) shows the vorticity cloud diagram of the mid-section of the leaf 130.05μs after applying the excitation, and 5(d) shows the vorticity cloud diagram of the mid-section of the leaf 290.05μs after applying the excitation. DETAILED DESCRIPTION

[0062] The present invention will be described in detail below with reference to the accompanying drawings.

[0063] Previous researchers have verified that low-energy rapid heating plasma excitation can achieve certain effects in regulating supersonic flow. Here, the present invention proposes a high-load supersonic compressor blade design method coupled with multi-point plasma excitation. Based on the original geometric data, the blade parameterization modeling method is used to generate high-load large-angle blade geometry data, and multi-point plasma excitation is coupled to increase the injected energy power, thereby achieving strong interference to the supersonic flow field, weakening the negative impact of supersonic compressor blade shock wave / boundary layer interference, and ensuring the aerodynamic performance of the blade while improving the single-stage pressure ratio of the compressor. Specifically, the following steps are included.

[0064] 1. Generate high-load large-angle blade geometry data using blade parametric modeling method based on original geometry data;

[0065] The present invention uses a blade parameterization modeling method based on the original blade geometry data to generate high-load large-angle blade geometry data on a two-dimensional plane by adjusting parameters. The main steps are as follows:

[0066] (1) Interpolation of original blade profile data. The blade profile curves are mainly divided into leading edge curves, trailing edge curves, suction surface curves and pressure surface curves. The original blade profile is a relatively discrete geometric point. First, it is necessary to interpolate it using the B-spline interpolation algorithm to obtain relatively dense suction surface and pressure surface curve geometric data. The B-spline interpolation algorithm is well known to those skilled in the art and will not be described again.

[0067] (2) Determine the geometric points and thickness distribution of the mid-arc line. After obtaining enough geometric data of the suction surface and pressure surface curves, take the first point of the suction surface curve as an example, connect this point with all the points on the pressure surface curve, search for the point on the pressure surface curve that is closest to this point, connect these two points, and take the middle point as the point on the mid-arc line. Repeat the above steps for all geometric points on the suction surface curve to obtain the complete distribution of the mid-arc line points, and connect all the mid-arc line points to obtain the mid-arc line curve. While the above method generates the mid-arc line, it also obtains the point on the pressure surface curve that is the shortest distance from any point on the suction surface curve. The midpoint of the line connecting the two points is the mid-arc line point, and 1 / 2 of the length of the line is defined as the blade thickness at the mid-arc line point. In this way, the blade thickness distribution is obtained. For the above method of determining the mid-arc line and thickness, the angle between the line connecting the suction surface point and the pressure surface point and the tangent at the mid-arc line point should be fixed. The geometric coordinates of the suction surface point and the pressure surface point are used to calculate the slope of the line connecting the two relative to the horizontal direction. The geometric coordinates of each point on the mid-arc line are used to calculate the slope of the tangent at the mid-arc line point relative to the horizontal direction. Since the angle between the two lines is fixed, the difference between their slopes must also be fixed.

[0068] (3) Parametric expression of the center arc geometry. Taking into account the accuracy and reliability of the fitting, a cubic piecewise polynomial is used to fit the center arc points generated in step (2) to obtain a parametric expression of the center arc (this method is well known to those skilled in the art). In one embodiment of the present invention, a total of 12 segments of cubic polynomials are used to fit the points on the center arc, and each segment is evenly distributed in length. For each cubic polynomial, a total of 4 constraints are required to determine its coefficients. The parameterized expression polynomial is denoted by p i (x), then:

[0069] p i (x) = a 1 (x m ) 3 +a 2 (x m ) 2 +a 3 x m +a 4

[0070] Among them, a 1 、a 2 、a 3 、a 4 represent the coefficients of the cubic polynomial, x m Indicates the axial coordinate of the midline curve control point.

[0071] The following are specific instructions:

[0072] The first segment selects three control points. The first control point (x1 m1 ,y 1 m1 ) is the starting point of the middle arc and also the starting point of the segmented area. The coordinates and slope of the first control point are used as the two constraints of the polynomial fitting; the third control point (x 1 m3 ,y 1 m3 ) is the end point coordinate of the segment. The coordinate of this point is used as a constraint condition of the polynomial fitting. For the second control point (x 1 m2 ,y 1 m2 ), whose value is not given in advance, but is obtained by optimizing the coordinates and slope of the first control point and the coordinates of the third control point with the minimum fitting error as the goal (this method is well known to those skilled in the art), and its coordinates are used as the last constraint of the polynomial fitting. The above four constraints can be used to obtain the analytical solution of the polynomial coefficients of the parameterized expression of the arc in the first segment.

[0073]

[0074] From the second to the eleventh segment, three control points are also selected for each segment, denoted as (x i m1 ,y i m1 )(x i m2 ,y i m2 )(x i m3 ,y i m3 ), i=2,…,11. The first control point is the end point of the previous curve, the third control point is the starting point of the next curve, and the second control point is also optimized with the minimum fitting error as the goal. Different from the first section, in the process of arc fitting from the second to the eleventh section, the slope constraint of the first point is changed to the second-order derivative constraint of the arc curve at the first point in each section (obtained by the derivation of the expression of the previous section) to ensure the continuity of the second-order curvature of the arc at the junction of each section. The other three constraints are the same (therefore, the second control point is based on the coordinates of the first control point and the second-order derivative of the arc curve, and the coordinates of the third control point, which are also optimized with the minimum fitting error as the goal). The above four constraints can be used to obtain the analytical solution of the polynomial coefficients of the parameterized expression of the arc in any intermediate section.

[0075]

[0076] When the last segment is generated, two control points are taken and recorded as (x N m1 ,y N m1 )(x N m2 ,y N m2 ). The first control point is the end point of the previous curve, and the second point is the end point of the median arc. In addition to the coordinates of the above two control points, the second-order derivative at the first point (derived by the expression in paragraph 11) and the first-order derivative at the last point are used as the other two constraints, where the first-order derivative at the last point is specified by the slope value corresponding to the blade profile outlet geometry angle (this is known to those skilled in the art). Using the above four constraints, the analytical solution of the polynomial coefficients of the parameterized expression of the median arc of the last segment can be obtained.

[0077]

[0078] (4) Parametric expression of thickness distribution. The parameter expression method of the thickness distribution curve also adopts the cubic piecewise polynomial method, which is similar to the mid-arc curve. Since the thickness distribution in the entire interval is relatively smooth, the entire thickness distribution area is divided into four segments. The first and fifth control points are the leading edge point and the trailing edge point, respectively. The second and third control points are the starting point and the end point of the pre-compression segment, respectively. The fourth control point is the equal-division point of the line segment between the third control point and the fifth control point in the axial direction. The parametric expression polynomial of the thickness distribution is denoted as h i (x), then:

[0079] h i (x) = b 1 (x h ) 3 +b 2 (x h ) 2 +b 3 x h +b 4

[0080] Among them, b 1 、b 2 、b 3 、b 4 represent the coefficients of the cubic polynomial, x h Represents the axial coordinate of the control point of the thickness distribution curve.

[0081] The following are specific instructions:

[0082] In the first section, the number of control points is 4. According to the thickness distribution curve, the first point is the starting point of the middle arc, and the fourth point is the starting point of the thickness distribution of the next section. The second and third control points are the same as step (3), which are optimized with the minimum fitting error as the goal. They are evenly spaced between the first and fourth points. Using the coordinates of the above four points as constraints, the analytical solution of the polynomial coefficients of the parameterized expression of the thickness distribution curve of the first section can be obtained.

[0083]

[0084] The constraints required for the thickness distribution of the second to fourth sections are different from those of the first section. For each section, the first step is to determine the coordinates of the starting point and end point of the thickness distribution curve of the corresponding section as two constraints, and then calculate the first-order derivative at the end point of the previous thickness distribution curve as a constraint. After obtaining the above three constraints, the intermediate control point is obtained through optimization with the minimum fitting error as the goal as a constraint. Using the above four constraints, the analytical solution of the polynomial coefficients of the parameterized expression of the thickness distribution curve of the second to fourth sections can be obtained.

[0085]

[0086] (5) Generate the geometry of the pressure surface and suction surface. After obtaining the parametric expression polynomial of the mid-arc line and thickness distribution, given the coordinates of any horizontal direction between the leading edge point and the trailing edge point on the suction surface curve of the blade, the coordinates of the mid-arc line point can be obtained by fitting the mid-arc line polynomial, and the blade thickness at the coordinate point can be obtained at the same time. For any mid-arc line point on the mid-arc line, the slope of the tangent direction at the mid-arc line point relative to the horizontal direction can be calculated using its geometric coordinates. The direction of the line connecting the corresponding suction surface point and the pressure surface point can be determined based on the slope value. Based on the parametric expression polynomial of the thickness distribution curve mentioned above, the thickness is superimposed on the mid-arc line coordinates to obtain all the geometric points of the pressure surface and the suction surface.

[0087] (6) Generate the leading and trailing edges of the blade. Generate four control points for the leading and trailing edges using the original blade geometry data and the distance and angle relationship, and then generate the leading and trailing edge curve expressions using the cubic Bezier curve.

[0088] Taking the front edge curve as an example, first set the first point of the middle arc line to the center of the front edge (x le0 ,y le0 ), the minimum distance between the leading edge of the original blade and the center of the circle is obtained by traversal method, denoted as R le (This method is well known to those skilled in the art.) Select the point closest to the suction surface and pressure surface curves from the original blade leading edge geometric points and record them as control points 1(x le1 ,y le1 ) and control point 4(xle4 ,y le4 ), and then solve the slope between the starting point of the suction surface and pressure surface curve and the corresponding leading edge control point, and find control points 2 and 3 along the slope direction, so as to ensure that the generated leading edge curve is curvature continuous with the suction surface and pressure surface. The distance between control points 2 and 3 and control points 1 and 4 is based on the aforementioned radius R le The specific value is obtained by optimizing with the minimum fitting error as the goal. After obtaining the above four control points, a Bezier curve that can parametrically express the leading edge curve can be generated from these four points (this method is well known to those skilled in the art). The fitting generation method of the trailing edge curve is similar to that of the leading edge curve, and will not be repeated here.

[0089] Second, the flow field characteristics of the compressor blade passage under low Reynolds number conditions are obtained through numerical simulation, and the flow separation area on the blade surface under low Reynolds number is determined;

[0090] The present invention selects the supersonic blade cascade DRL-PAV-1.5 designed by the German Aerospace Center as the research object of numerical simulation, and uses the large eddy simulation (LES) numerical simulation method to perform high-precision simulation of the compressor blade channel flow under low Reynolds number conditions (LES simulation method is well known in the art). The working conditions analyzed by the present invention are: the incoming flow Mach number is equal to 1.52, the pressure ratio is 2.2, and the incoming flow Reynolds number is 1.8×10 5 .

[0091] The computational domain grid of the present invention is generated by multi-block grid technology. The inlet and outlet sections and the cascade channel are divided by H-type topology grids, and the blade surface is divided by O-type topology grids ("cascade HOH multi-block grid generation technology" is well known in the art). In order to accurately capture the channel shock wave / boundary layer interaction and the interference effect of plasma excitation on the flow field, the leading and trailing edges of the blade, the blade channel and the boundary layer of the blade surface are encrypted by adjusting the number of grid nodes (this method is well known to those skilled in the art), ensuring that the dimensionless scale division of the grid along the span direction, normal direction and flow direction is not greater than 30, 1 and 15. It is verified that the LES_IQ value of the grid near the blade surface is greater than 0.8, which meets the grid quality requirements of large eddy simulation. The total number of grids used for numerical simulation is 2.5×10 6 .

[0092] The present invention uses ANSYS-Fluent 18.0 commercial fluid calculation software for simulation, and the simulation model is selected as the LESSmagorinsky-Lilly model. This model determines the eddy viscosity coefficient in a dynamic way and is often used for predictive calculations in solid wall boundary layer flow and shock wave induced separation and other flows. In order to simulate the ideal infinite blade cascade model, the cascade pitch direction adopts a translational periodic boundary condition (the translational periodic boundary condition is well known in the art). The inlet plane of the calculation domain is selected at a distance of 1.5 times the axial chord length from the leading edge of the blade. According to the actual working conditions, the total temperature, total pressure and inlet airflow angle of the blade cascade inlet plane are given. In order to avoid the reflection effect of the outlet boundary from polluting the flow field being calculated, the outlet of the calculation domain in the numerical simulation is selected at a distance of 2 times the axial chord length from the trailing edge of the blade, and the average static pressure is given as the boundary condition of the blade cascade outlet plane. In order to increase the calculation stability and ensure more sufficient flow in the calculation domain, the calculation domain is extended by 0.5 times the axial chord length at the inlet. The ideal gas model is used for the fluid medium. Considering the existence of flow field shock waves and the local high temperature characteristics of plasma excitation discharge, the viscosity cannot be considered as a constant. Therefore, the present invention adopts Sutherland's law to regulate the viscosity. In the medium model, the present invention adopts a nonlinear polynomial fitting method to convert the specific heat capacity C p The discrete data of the variation of thermal conductivity K with temperature are fitted into a polynomial function, and written into a UDF file and imported into the simulation example to study the variation of gas properties with temperature during rapid heating plasma excitation. The blade surface is set as an adiabatic non-slip stationary wall strip, the blade span length is given as 0.1 times the axial chord length, and the blade span uses a symmetric boundary condition to ensure that the axial density flow ratio of the blade channel is equal to 1.0 ("symmetric boundary conditions" are well known to those skilled in the art). In view of the two-dimensionality of the calculation domain of the cascade flow field, the mid-blade section will be selected as the characteristic section to display and analyze the transient structure in the flow channel.

[0093] The present invention writes a UDF file in a Fluent environment to perform variable time step solution, and the Courant number (CFL number) during the solution is not greater than 0.95 (this method is known to those skilled in the art). When nanosecond pulse plasma excitation is applied, the simulation time step is set to 5ns, and the simulation time step is changed to 50ns after the excitation is applied to shorten the numerical simulation cycle. In order to ensure the reliability of the time-averaged results, 20 cycles of excitation are first applied to the reference flow field. After the excitation flow field is stable, the UDF file is called to solve the duration of 20 excitation cycles with a variable time step as a statistical sample of the time-averaged data. By numerically averaging the transient flow field parameters of the last 20 flow cycles after the excitation flow field is stabilized, the time-averaged flow field parameters can be obtained for subsequent analysis.

[0094] For the supersonic compressor blade profile studied, the time-averaged axial shear force distribution of the blade surface under low Reynolds number conditions is as follows: Figure 1 As shown, from Figure 1 In (a), it can be found that the wall friction coefficient of the suction side of the blade surface is x = 55.07%C ax The position changes from positive to negative, and at x = 81.7%C ax The two zero points of the wall friction coefficient curve correspond to the separation point x of the blade suction surface. s-ss and then dot x r-ss In addition, the wall friction coefficient on the suction surface of the blade is x = 94.81%C ax The value becomes negative again from the position to the trailing edge of the blade, which indicates that the boundary layer at the trailing edge of the suction surface has undergone flow separation again under the action of the channel positive shock wave. Figure 1 (b) It is observed that the pressure surface flow separation occurs at x = 33.43%C ax and at x = 72.76%C ax Reattachment occurs at the blade surface, which indicates that serious flow separation occurs on the blade surface under low Reynolds number conditions.

[0095] 3. Determine the layout of multi-point rapid heating plasma excitation according to the characteristics of the flow separation area;

[0096] The present invention selects nanosecond pulsed plasma excitation (Pulsed-DBD) plasma excitation as the excitation method for flow control. Since the plasma phenomenological model does not pay attention to the specific discharge process such as chemical reactions involved in plasma discharge and the movement of charged particles in the electromagnetic field, and only simplifies the plasma into the form of heat and momentum and injects it into the flow field, it has the advantages of small calculation amount and easy coupling with the Navier-Stokes equation. It has been accepted and recognized in the field of plasma flow control. Therefore, the present invention establishes a simulation model to characterize the Pulsed-DBD plasma excitation effect based on the phenomenological heat source model ("phenomenological heat source model" is well known in the art). The experimental data of polyimide material as the blocking medium is selected as a reference. The plasma excitation simulated by the present invention is within a single discharge pulse time (t f =50ns) The energy Q injected per unit electrode length can be calculated by the following formula:

[0097]

[0098] Where ε is the relative dielectric constant of the barrier material, V pp and V bd Represent the excitation voltage amplitude and the breakdown voltage of the medium, l sand d represent the length of the plasma column and the thickness of the barrier material, respectively. The reliability of the model has been verified by experiments and the high-precision plasma simulation program PASSKEy. The baseline excitation energy within a single pulse (50ns) of plasma excitation is set to 0.14mJ / cm, which is at the same order of magnitude as the experimental measurement value. Considering the complexity of the shock wave boundary layer structure of the supersonic flow field and the previous research results, it is necessary to select a suitable excitation frequency to achieve a stable flow control effect. Based on the continuous induced vortex phenomenon generated by plasma excitation, the present invention proposes an effective excitation frequency f eff Determine the criteria, the specific formula is as follows:

[0099]

[0100] Where, T 2 represents the time interval from the excitation application position to the complete dissipation of the vortex structure induced by plasma excitation in the separation zone; T 1 It represents the time interval between the DFS moving from the excitation application position to the position where the complete spanwise vortex structure is induced. Based on the effective excitation frequency criterion, the flow field state is preliminarily evaluated, and the excitation frequency f = f is finally selected. eff =10kHz high frequency excitation for calculation and study, the corresponding excitation period T 0 =1 / f.

[0101] The multi-point layout of plasma excitation needs to consider the position of the excitation reference point, the number of excitation points, and the spacing between the excitation points. According to previous relevant simulation studies, it is better to arrange the plasma exciter upstream of the shock wave / boundary layer interference zone on the pressure surface. At the same time, considering the disturbance intensity and effective action time of the distorted flow structure (DFS) induced by plasma excitation in the supersonic flow field, in order to ensure that the DFS generated at the leading edge of the blade can fully develop in the laminar boundary layer flow area and interact with the shock wave / boundary layer interference position, according to the flow separation characteristics of the blade surface under the aforementioned low Reynolds number, it is finally determined that the multi-point layout excitation reference point PS-Actu-1 is set at 0.27 times the axial chord length downstream of the leading edge of the blade, that is, upstream of the flow separation point on the pressure side of the blade, in order to make the excitation and separation start positions act directly, thereby realizing the regulation of the shock wave boundary layer interference effect. According to the number of excitation points N and the excitation spacing l, they are arranged in sequence from the excitation reference point to the upstream and named in sequence, namely PS-Actu-1, PS-Actu-2, ..., such as Figure 2 shown.

[0102] 4. Analysis of the flow control effect of multi-point rapid heating plasma excitation on supersonic compressor blade shock wave / boundary layer interference;

[0103] In the multi-point excitation layout method proposed in the present invention, parameters such as the position of the excitation reference point, the number of excitation points N, the excitation spacing l, and the energy of the excitation point can be adjusted, and the effects of these parameters on the flow control effect are mutually coupled. Previous studies have found that the spacing between the excitation points and the moving distance of the DFS within a single excitation cycle need to be considered during layout. The larger the spacing between the excitation points, the less likely it is that the DFS will interact with each other. In order to demonstrate the flow control effect of the multi-point rapid heating plasma excitation method, a multi-point layout method with N=3, l=9mm, and the energy of the excitation point as the reference excitation energy is selected as an example.

[0104] The total pressure loss coefficient averaged over the blade midsection is defined as follows:

[0105]

[0106] In the formula, and are the time-averaged total pressure at the inlet and the local total pressure of the flow field, respectively. is the average static pressure at the inlet of the flow field. Change to indicate the average total pressure at the outlet The time-averaged total pressure loss coefficient after the grid can be obtained. The relative change of the time-averaged total pressure loss coefficient is defined as Δω=ω D -ω L , where ω D is the time-averaged total pressure loss coefficient under multi-point rapid heating plasma excitation conditions, ω L is the time-averaged total pressure loss coefficient at low Reynolds numbers. Figure 3 The cloud map of the relative change of the total pressure loss coefficient averaged over the mid-section of the blade after applying multi-point rapid heating plasma excitation is shown. It can be seen from the cloud map that, compared with the operating conditions of low Reynolds numbers, after applying multi-point rapid heating plasma excitation, the excitation suppresses the separation of the boundary layer on the pressure surface of the blade, thereby causing a significant reduction in the total pressure loss in this area. In the position after the positive shock wave in the channel, it can be observed that the total pressure loss in the blade channel has increased.

[0107] In order to further analyze the effect of excitation on the change of total pressure loss in the blade passage, Figure 4 The distribution curve of the average total pressure loss coefficient along the pitch direction of the flow field before and after applying multi-point rapid heating plasma excitation is shown. The flat parts on the left and right sides are shock wave losses, and the peak in the middle is mainly the viscous loss caused by separation. Overall, the KH vortex generated by multi-point rapid heating plasma excitation increases the mixing in the wake zone and reduces the peak loss. It can be observed that while the excitation reduces the boundary layer loss on the pressure surface, it will increase the shock wave loss. At the same time, the boundary layer viscous loss moves toward the suction surface as a whole. This is because the plasma excitation suppresses the separation of the boundary layer on the pressure surface of the blade, resulting in a significant reduction in the total pressure loss in this area.

[0108] Figure 5 In order to extract the vorticity cloud diagram of the middle section of the blade at different times after the multi-point rapid heating plasma excitation, according to Figure 5 (a) After 10.05 μs of excitation, a compression wave induced by the nanosecond pulse excitation can be found, which then quickly spreads and dissipates around, and the plasma excitation induces DFS. Figure 5 (b) After 100.00μs of excitation (at the end of the first excitation cycle), the separation vortex layer can be observed to be winding and breaking, which means that the shear layer is obviously unstable. According to the flow stability theory, this is caused by the strong nonlinear effect after the disturbance in the shear layer enters the nonlinear action region. The instability caused by the disturbance is further amplified, and finally triggers the KH instability of the separation shear after 130.05μs of excitation and induces the formation of a spanwise vortex structure as shown in Figure 2. Figure 5 (c). After applying the excitation for 290.05μs, Figure 5 (d) It can be found that the KH vortex induced by the excitation continuously interacts with the vortex structure in the flow separation zone during the downstream movement, and finally forms a spanwise large vortex structure (LSV) with concentrated vorticity. The LSV continuously sweeps the low-energy fluid downstream. During the movement, it significantly affects the morphology of the positive shock wave in the excitation channel and significantly enhances the mixing between the mainstream and the low-energy fluid in the separation zone. After that, the LSV finally moves to the trailing edge of the blade with the mainstream, the vorticity intensity weakens and gradually merges into the wake zone. At this time, a new LSV begins to form and performs the next round of mixing movement in the separation zone. This cycle forms the alternating movement of the LSV, which promotes the fluid momentum exchange in the shear layer.

[0109] In an example of the present invention, by applying multi-point rapid heating plasma excitation on the pressure surface of the blade, the shear layer KH instability is triggered and the spanwise vortex row structure is induced. The structure promotes the mixing of small-scale vortices in the flow separation zone and evolves into a spanwise large vortex structure to accelerate dissipation, thereby achieving the effect of suppressing flow separation. The present invention provides a high-load supersonic compressor blade design method coupled with multi-point plasma excitation. The simulation case design uses ANSYS-Fluent 18.0 commercial fluid calculation software, and selects nanosecond pulsed plasma excitation (Pulsed-DBD) as the excitation method for flow control. It has the advantages of simple structure, low power consumption, easy processing and implementation, and has broad application prospects in active flow control. Based on the original supersonic blade geometry data, the present invention uses a blade parameterization modeling method to change the blade bending angle and generate high-load supersonic compressor blade geometry data. In view of the problem that the existing rapid heating plasma excitation injection energy is limited and it is difficult to achieve strong interference with the supersonic flow field, the present invention hopes to achieve strong interference with the supersonic flow field by optimizing the rapid heating plasma excitation energy injection method, weaken the negative impact of high-load supersonic compressor blade shock wave / boundary layer interference, and use multi-point rapid heating plasma excitation to achieve significant suppression of flow separation on the surface of high-load supersonic compressor blades under low Reynolds numbers, providing innovative ideas for the development of high-performance compressors.

Claims

1. A high-load supersonic compressor blade design method coupled with multi-point plasma excitation, characterized in that: The following steps are involved: Step 1: Generate high-load large-angle blade geometry data by using blade parameterization modeling method based on original geometry data; Step 2: obtain the flow field characteristics of the compressor blade channel under low Reynolds number conditions through numerical simulation, and determine the flow separation area on the blade surface under low Reynolds number; Step 3: Determine the multi-point rapid heating plasma excitation layout mode according to the flow separation area characteristics; Step 4: Analysis of the flow control effect of multi-point rapid heating plasma excitation on supersonic compressor blade shock wave / boundary layer interference.

2. The high-load supersonic compressor blade design method coupled with multi-point plasma excitation according to claim 1, characterized in that: Step 1 is to use the blade parameterization modeling method based on the original blade geometry data to generate high-load large-angle blade geometry data on a two-dimensional plane by adjusting parameters; the specific steps are as follows: (1) Interpolation of original blade profile data: The blade profile curve includes the leading edge curve, the trailing edge curve, the suction surface curve and the pressure surface curve. These curves are interpolated using the B-spline interpolation algorithm to obtain relatively dense geometric data of the suction surface and pressure surface curves; (2) Determine the geometric points and thickness distribution of the mid-arc line; for the suction surface curve, connect the point with all the points on the pressure surface curve, search for the point on the pressure surface curve that is closest to the point, connect the two points, and take the middle point as the point on the mid-arc line; repeat the above steps for all the geometric points on the suction surface curve to obtain the complete distribution of the mid-arc line points, and connect all the mid-arc line points to obtain the mid-arc line curve; while generating the mid-arc line, also obtain the point on the pressure surface curve that is the shortest distance from any point on the suction surface curve, and the midpoint of the line connecting the two points is the mid-arc line. The arc point, and 1 / 2 of the length of the line is defined as the blade thickness at the mid-arc point, thereby obtaining the blade thickness distribution; for the above method of determining the mid-arc and thickness, the angle between the line connecting the suction surface point and the pressure surface point and the tangent at the mid-arc point should be fixed, and the geometric coordinates of the suction surface point and the pressure surface point are used to calculate the slope of the line connecting the two relative to the horizontal direction, and the geometric coordinates of each point on the mid-arc are used to calculate the slope of the tangent direction at the mid-arc point relative to the horizontal direction; since the angle between the two lines is fixed, the difference between their slopes must also be fixed; (3) Parametric expression of the center arc geometry: Fitting the center arc points generated in step (2) with a cubic piecewise polynomial to obtain a parametric expression of the center arc; (4) parametric expression of thickness distribution; the parameter expression method of the thickness distribution curve adopts a cubic piecewise polynomial method to fit the blade thickness data at each mid-arc point generated in step (2) to obtain a parametric expression of the thickness distribution; (5) Generate the geometry of the pressure surface and suction surface; Given the coordinates of any horizontal direction between the leading edge point and the trailing edge point on the suction surface curve of the blade, obtain the coordinates of the mid-arc point by fitting the mid-arc polynomial, and at the same time obtain the blade thickness at the coordinate point; for any mid-arc point on the mid-arc, use its geometric coordinates to calculate the slope of the tangent direction at the mid-arc point relative to the horizontal direction; determine the direction of the line connecting the corresponding suction surface point and pressure surface point according to the slope value; express the polynomial according to the parameterization of the thickness distribution curve, superimpose the thickness on the basis of the mid-arc coordinates, and obtain all the geometric points of the pressure surface and suction surface; (6) Generate the leading and trailing edges of the blade; Four control points of the leading edge and the trailing edge are generated through the original blade geometry data and the distance and angle relationship, and then the leading and trailing edge curve expressions are generated through the cubic Bezier curve.

3. The high-load supersonic compressor blade design method coupled with multi-point plasma excitation according to claim 2, characterized in that: In step 1 (3), 12 segments of cubic polynomials are used to fit the points on the mid-arc, and each segment is evenly distributed in length. For each cubic polynomial, four constraints are required to determine its coefficients. The parameterized expression polynomial is denoted by p i (x), then: p i (x)=a1(x m ) 3 +a2(x m ) 2 +a3x m +a4 Among them, a1, a2, a3, and a4 represent the coefficients of the cubic polynomial, and x m Indicates the axial coordinates of the control points of the mid-arc curve; The following are specific instructions: The first segment selects three control points. The first control point (x 1 m1 ,y 1 m1 ) is the starting point of the middle arc and also the starting point of the segmented area; the coordinates and slope of the first control point are used as two constraints of the polynomial fitting; the third control point (x 1 m3 ,y 1 m3 ) is the coordinate of the end point of the segment. The coordinate of this point is used as a constraint condition of the polynomial fitting. For the second control point (x 1 m2 ,y 1 m2 ), whose value is not given in advance, but is obtained by optimizing the coordinates and slope of the first control point and the coordinates of the third control point with the minimum fitting error as the goal, and using its coordinates as the last constraint of the polynomial fitting; using the above four constraints, the analytical solution of the polynomial coefficients of the parameterized expression of the arc in the first segment is obtained; From the second to the eleventh segment, three control points are also selected for each segment, denoted as (x i m1 ,y i m1 )(x i m2 ,y i m2 )(x i m3 ,y i m3 ), i=2,…,11; the first control point is the end point of the previous curve, the third control point is the starting point of the next curve, and the second control point is also optimized with the minimum fitting error as the goal; different from the first section, in the arc fitting process from the second to the eleventh section, the slope constraint of the first point in each section is changed to the second-order derivative constraint of the arc curve at the first point to ensure the continuity of the second-order curvature of the arc at the junction of each section, and the other three constraints are the same; the above four constraints are used to obtain the analytical solution of the polynomial coefficients of the parameterized expression of the arc in any intermediate section; When the last segment is generated, two control points are taken and recorded as (x N m1 ,y N m1 )(x N m2 ,y N m2 ); the first control point is the end point of the previous curve, the second point is the end point of the mid-arc, and in addition to the coordinates of the above two control points, the second-order derivative at the first point and the first-order derivative at the last point are used as the other two constraints, where the first-order derivative at the last point is specified by the slope value corresponding to the blade profile outlet geometry angle; using the above four constraints, the analytical solution of the polynomial coefficients of the parameterized expression of the last mid-arc is obtained; 4. The high-load supersonic compressor blade design method coupled with multi-point plasma excitation according to claim 2, characterized in that: In step 1 (4), the parameter expression method of the thickness distribution curve adopts the cubic piecewise polynomial method, and the entire thickness distribution area is divided into four segments. The first and fifth control points are the leading edge point and the trailing edge point, respectively. The second and third control points are the starting point and the end point of the pre-compression segment, respectively. The fourth control point is the equal-division point of the line segment between the third control point and the fifth control point in the axial direction. The parameterized expression polynomial of the thickness distribution is denoted as h i (x), then: h i (x)=b1(x h ) 3 +b2(x h ) 2 +b3x h +b4 Among them, b1, b2, b3, and b4 represent the coefficients of the cubic polynomial, respectively, and x h Represents the axial coordinate of the control point of the thickness distribution curve; The following are specific instructions: In the first section, the number of control points is 4. According to the thickness distribution curve, the first point is the starting point of the middle arc, and the fourth point is the starting point of the thickness distribution of the next section. The second and third control points are the same as step (3), which are optimized with the minimum fitting error as the goal, and are evenly spaced between the first and fourth points. The coordinates of the above four points are used as constraints to obtain the analytical solution of the polynomial coefficients of the parameterized expression of the first section thickness distribution curve; The constraints required for the thickness distribution of the second to fourth sections are different from those of the first section. For each section, the first step is to determine the coordinates of the starting point and the end point of the thickness distribution curve of the corresponding section as two constraints, and then calculate the first-order derivative at the end point of the previous thickness distribution curve as a constraint. After obtaining the above three constraints, the intermediate control point is obtained by optimization with the minimum fitting error as the goal as a constraint. Using the above four constraints, the analytical solution of the polynomial coefficients of the parameterized expression of the thickness distribution curve of the second to fourth sections is obtained; 5. The high-load supersonic compressor blade design method coupled with multi-point plasma excitation according to claim 2, characterized in that: In step 1 (6), for the front edge curve, first set the first point of the mid-arc line to the center of the front edge (x le0 ,y le0 ), the minimum distance between the leading edge of the original blade and the center of the circle is obtained by traversal method, denoted as R le ; Select the points closest to the suction surface and pressure surface curves from the original blade leading edge geometry points and record them as control points 1(x le1 ,y le1 ) and control point 4(x le4 ,y le4 ), and then solve the slopes between the starting points of the suction surface and pressure surface curves and the corresponding leading edge control points, and find control points 2 and 3 along the slope direction; the distances between control points 2 and 3 and control points 1 and 4 refer to the aforementioned radius R le The specific value is obtained by optimizing with the minimum fitting error as the goal; after obtaining the above four control points, a Bezier curve that can parameterize the front edge curve is generated by these four points.

6. The high-load supersonic compressor blade design method coupled with multi-point plasma excitation according to claim 1, characterized in that: Step 2: Using large eddy simulation numerical simulation method to perform high-precision simulation of compressor blade channel flow under low Reynolds number conditions; The computational domain grid is generated using multi-block grid technology. The inlet and outlet sections and the cascade channel are divided using H-type topology grids, and the blade surface is divided using O-type topology grids. The leading and trailing edges of the blade, the blade channel, and the boundary layer of the blade surface are encrypted by adjusting the number of grid nodes. During the simulation, in order to simulate the ideal infinite blade cascade model, the translation periodic boundary condition is adopted in the cascade pitch direction; the calculation domain inlet plane is selected at A times the axial chord length from the leading edge of the blade, and the total temperature, total pressure and inlet airflow angle of the blade cascade inlet plane are given according to the actual working conditions; the calculation domain outlet in the numerical simulation is selected at B times the axial chord length from the trailing edge of the blade, and the average static pressure is given as the boundary condition of the blade cascade outlet plane; the calculation domain is extended at the inlet by C times the axial chord length; the ideal gas model is used for the fluid medium, and the viscosity cannot be considered as a constant, and the Sutherland law is used to regulate the viscosity; in the medium model, the nonlinear polynomial fitting method is used to calculate the specific heat capacity C p , the discrete data of thermal conductivity K changing with temperature are fitted into a polynomial function; the blade surface is set as an adiabatic non-slip stationary wall strip, the blade span length is given as D times the axial chord length, and the blade span uses symmetric boundary conditions to ensure that the axial density flow ratio of the blade channel is equal to E; the mid-blade section is selected as the characteristic section to display and analyze the transient structure in the flow channel; Perform variable time step solution, and the Courant number is not greater than F during the solution; When nanosecond pulse plasma excitation is applied, the simulation time step is set to G nanoseconds, and after the excitation is applied, the simulation time step becomes H nanoseconds; first, I cycles of excitation are applied to the reference flow field, and after the excited flow field is stabilized, the duration of I excitation cycles is solved as a statistical sample of the time-averaged data; the time-averaged flow field parameters are obtained by numerically averaging the transient flow field parameters of the last I flow cycles after the excited flow field is stabilized.

7. The high-load supersonic compressor blade design method coupled with multi-point plasma excitation according to claim 6, characterized in that: The sonic cascade DRL-PAV-1.5 was selected as the research object of numerical simulation. The large eddy simulation method LES was used to perform high-precision simulation of the compressor blade channel flow under low Reynolds number conditions. The analyzed conditions were: the incoming flow Mach number was equal to 1.52, the pressure ratio was 2.2, and the incoming flow Reynolds number was 1.8×10 5 ; The dimensionless scale division of the grid along the span direction, normal direction and flow direction is no more than 30, 1 and 15, and the total number of grids is 2.5×10 6 ; ANSYS-Fluent 18.0 commercial fluid calculation software was used for simulation, and the simulation model was selected as the LESSmagorinsky-Lilly model; A=1.5; B=2; C=0.5; D=0.1; E=1.0; F=0.95; G=5; H=50; I=20.

8. The high-load supersonic compressor blade design method coupled with multi-point plasma excitation according to claim 1, characterized in that: In step 3, nanosecond pulsed plasma excitation Pulsed-DBD plasma excitation is selected as the excitation mode of flow control; a simulation model characterizing the Pulsed-DBD plasma excitation effect is established based on the phenomenological heat source model; the energy Q injected by the simulated plasma excitation per unit electrode length within a single discharge pulse time is calculated by the following formula: Where ε is the relative dielectric constant of the barrier material, V pp and V bd Represent the excitation voltage amplitude and the breakdown voltage of the medium, l s and d represent the length of plasma column and the thickness of barrier material respectively. Based on the continuous induced vortex phenomenon generated by plasma excitation, the effective excitation frequency f is proposed. eff Determine the criteria, the specific formula is as follows: Where, T2 represents the time interval from the excitation application position to the complete dissipation of the vortex structure induced by the plasma excitation in the separation zone; T1 represents the time interval from the movement of the DFS from the excitation application position to the position where the complete spanwise vortex structure is induced; the flow field state is preliminarily evaluated based on the effective excitation frequency criterion, and the high-frequency excitation frequency f is selected for calculation, and the corresponding excitation period T0 = 1 / f; The plasma actuator is arranged upstream of the pressure surface shock wave / boundary layer interference zone; According to the number of excitation points N and the excitation spacing l, they are arranged in sequence from the excitation reference point upstream and named in sequence, namely PS-Actu-1, PS-Actu-2, ...

9. The high-load supersonic compressor blade design method coupled with multi-point plasma excitation according to claim 8, characterized in that: Select polyimide material as the blocking medium, and the single discharge pulse time t f =50ns; the reference excitation energy in a single plasma excitation pulse is set to 0.14mJ / cm; the excitation frequency f=f eff =10kHz high-frequency excitation is used for calculation; the multi-point layout excitation reference point PS-Actu-1 is set at a position of 0.27 times the axial chord length downstream of the leading edge of the blade.

10. The high-load supersonic compressor blade design method coupled with multi-point plasma excitation according to claim 1, characterized in that: In step 4, the parameters such as the position of the excitation reference point, the number of excitation points N, the excitation spacing l, and the energy of the excitation point can be adjusted, and the effects of these parameters on the flow control effect are mutually coupled; The total pressure loss coefficient averaged over the blade midsection is defined as follows: In the formula, and are the time-averaged total pressure at the inlet and the local total pressure of the flow field, respectively. is the average static pressure at the inlet of the flow field; if Change to indicate the average total pressure at the outlet The time-averaged total pressure loss coefficient after the grid can be obtained; the relative change of the time-averaged total pressure loss coefficient is defined as Δω=ω D -ω L , where ω D is the time-averaged total pressure loss coefficient under multi-point rapid heating plasma excitation conditions, ω L is the time-averaged total pressure loss coefficient at low Reynolds numbers.

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