Airfoil pressure distribution elastic correction method, application, equipment and storage medium

Abstract

The invention discloses an airfoil pressure distribution elastic correction method, application, equipment and a storage medium, and relates to the technical field of aerodynamics, the airfoil pressure distribution elastic correction method comprises the following steps: S1, rigid pressure distribution pretreatment; s2, acquiring an elastic correction amount; s3, correcting the total load of pressure distribution; s4, correcting a pressure distribution spanwise pressure center; s5, after the spanwise pressure center is corrected, solving a new chordwise correction amount; and S6, correcting the chordwise pressure center of pressure distribution, performing static aeroelastic correction on CFD pressure distribution data or pressure measurement test data, obtaining pressure distribution data after considering elastic deformation, providing a basis for structural design, and ensuring flight safety.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aerodynamics, in particular to a method for elastic correction of airfoil pressure distribution, application, device and storage medium. BACKGROUND

[0002] The pressure distribution of an aircraft is the core original input of flight load design, and its accuracy directly affects the reliability of flight load design and structural strength design. Generally, the ways to obtain pressure distribution data include pressure measurement test and computational fluid dynamics (CFD) calculation.

[0003] However, the main airfoil of an aircraft will produce significant elastic deformation under the action of aerodynamic force, and the deformed airfoil shape will cause changes in the pressure distribution. This change has a double impact: on the one hand, it has an unloading effect on the aerodynamic load, which can reduce the load of structural design; on the other hand, the change in pressure distribution may intensify the stress in local areas, thereby posing a potential threat to the safety of the aircraft structure.

[0004] Therefore, in the aircraft design stage, the influence of elastic deformation on pressure distribution must be analyzed to provide accurate data support for the fine design of structural strength, so as to ensure the balance between economy and safety in aircraft design.

[0005] In the prior art, the disclosed related correction methods include: a correction method based on the combination of the lift line theory and the wind tunnel pressure measurement data, such as the patent document with the application number "201911288570.4"; and a correction method based on the engineering surface element method and the wind tunnel model aerodynamic force coefficient, such as the patent document with the application number "201910216036.6". However, the above methods do not involve specific correction methods for airfoil pressure distribution, and there are still obvious defects in the design level. SUMMARY

[0006] The purpose of the present application is to provide a method for elastic correction of airfoil pressure distribution, which aims to correct the CFD pressure distribution data or pressure measurement test data with static aeroelasticity, obtain pressure distribution data considering elastic deformation, and provide a basis for structural design to ensure flight safety.

[0007] The present application is achieved by the following technical solutions: A method for elastic correction of airfoil pressure distribution, comprising the following steps: S1, rigid pressure distribution preprocessing; S2, obtaining an elastic correction amount; S3, total load correction of pressure distribution; S4, spanwise pressure center correction of pressure distribution; S5, after the spanwise pressure center correction, a new chordwise correction amount is obtained; S6, chordwise pressure center correction of pressure distribution.

[0008] Further, in step S1, the rigid pressure distribution preprocessing comprises the following steps: S1.1, three-dimensional space interpolation of initial rigid pressure distribution data according to structured grid; S1.2, ascending arrangement of interpolated data and construction of panelization model; S1.3, calculation of geometric parameters of each panel, including area vector, normal vector and centroid coordinate.

[0009] Further, the rigid pressure distribution preprocessing method is as follows: In step S1.1, wing surface pressure rigid distribution data is obtained, the wing surface pressure rigid distribution data includes upper wing surface pressure distribution data and lower wing surface pressure distribution data, and the wing surface pressure rigid distribution data is derived from pressure test or CFD calculation; The wing surface pressure rigid distribution data is provided with a plurality of sections along the wing surface, each section is provided with a plurality of original data points, and each original data point contains coordinate values (x, y, z) and pressure coefficient Cp ; In step S1.2, if the number of original data points on a plurality of sections is inconsistent, the original data points on each section are reconstructed, specifically including: The original data points on each section are sorted according to x coordinate value, the original data point with the smallest x coordinate value x min is taken as the leading edge point of the section, the original data point with the largest x coordinate value x max is taken as the trailing edge point of the section, and the chord length of the section is calculated as x max -x min ; m+1 reconstruction data points are arranged on the chord length in equal proportion, wherein the x coordinate of each reconstruction data point is determined according to the equal proportion relationship, the y coordinate is the station of the section, and the z coordinate and Cp value are obtained by x coordinate interpolation according to the original data point information of the section; In step S1.3, when all sections are reconstructed, if the total number of sections is n+1, all reconstruction data points form a (n+1)×(m+1) matrix, every four adjacent reconstruction data points form a panel, and the total number of panels is n×m; The center point coordinate, average pressure coefficient, area and normal vector of each panel are calculated n 1ij , n 2ij , n 3ij , wherein: The center point coordinates (X, Y, Z) of the surface element are the average of the coordinates of the four reconstructed data points that constitute the surface element; The average pressure coefficient of the surface element is the sum of the four reconstructed data points constituting that surface element. Cp The average value; The surface element is a quadrilateral, and its area is the sum of the areas of the two triangles that make up the quadrilateral. The area of ​​each triangle is calculated using Heron's formula. The normal vector of the j-th surface element of the i-th profile ( n 1ij , n 2ij , n 3ij The results are obtained by calculation using the preset formulas (1) to (3): (1), (2), (3), Where n3>0 represents the upper wing surface, or n3<0 represents the lower wing surface.

[0010] Furthermore, in step S2, obtaining the elasticity correction amount includes the following steps: S2.1 Static aeroelastic iterative calculations are performed using a CFD / CSD fluid-structure interaction solver. S2.2 Extract the pressure field data after the elastic deformation of the wing surface; S2.3 Calculate the difference between the elastic pressure distribution and the rigid pressure distribution to obtain the total load correction factor and the spanwise core offset.

[0011] Furthermore, in step S2, the influence of elastic deformation on the wing surface load satisfies the following equations (4) to (6). (4), (5), (6), Among them, F 弹性 F 刚性 The total airfoil loads are shown after elastic deformation and without considering elastic deformation, respectively. K To adjust the ratio; Y 弹性 Y 刚性 These are the spanwise pressure centers of the wing after elastic deformation and without considering elastic deformation, respectively. Dy This is the correction amount for the expansion center of gravity; X 弹性 X 刚性 The figures show the chordal pressure center of the wing after elastic deformation and before considering elastic deformation, respectively.Dx is the chordwise pressure center correction amount.

[0012] Further, in step S3, the total load correction of the pressure distribution is performed according to the elastic correction amount on the rigid pressure distribution, so as to ensure that the corrected pressure distribution satisfies the whole machine aerodynamic force balance condition.

[0013] Further, the value of the total load of each surface element is multiplied by a proportional coefficient Cp to obtain a new K value Cp 1.After correction, the total load value of the airfoil is the total load value after considering the elastic deformation, i.e., it satisfies formula (7) as follows, Cp Cp 1 ij = Cp ij ×K (7).

[0014] Further, in step S4, the spanwise pressure center correction of the pressure distribution is to perform spanwise integration on the corrected pressure distribution, and to perform the first correction on the pressure center position based on the spanwise load distribution characteristics of the airfoil.

[0015] Further, in step S4, during the spanwise correction, the total normal force of the airfoil is ensured to be unchanged, and the value of the total load of each surface element of the upper and lower airfoils is multiplied by a parameter Cp 1 K i to obtain a new Cp 2, i.e., the spanwise correction is completed, and the parameter K i takes different values along the spanwise station, and simultaneously satisfies formulas (8) and (9) as follows: (8), (9), wherein the parameter α , b is obtained through formulas (10) and (11); (10); (11); wherein the physical meaning of formula (10) is that the normal force of the airfoil remains unchanged; the physical meaning of formula (11) is that the new bending moment of the airfoil is the original bending moment plus the bending moment increment caused by the change of the spanwise pressure center, wherein C zi and S i are respectively the normal force coefficient of the profile and the profile area, and are obtained through formulas (12) and (13), (12); (13).

[0016] Further, in step S5, the method for calculating the new chordwise correction amount is as follows: according to the pressure distribution corrected in step S4, the chordwise pressure center position is calculated at each station along the spanwise direction, and according to the chordwise pressure center offset rule before and after the elastic deformation, a chordwise correction coefficient function is constructed.

[0017] Further, in step S5, the chordwise pressure center of the airfoil is recalculated according to formula (14) as follows, and the chordwise pressure center correction amount is recalculated according to formula (15) as follows, (14); (15).

[0018] Further, in step S6, the method for correcting the pressure distribution chordwise pressure center is as follows: the chordwise correction coefficient obtained in step S5 is applied to the pressure distribution, so as to ensure that the finally corrected pressure distribution satisfies the spanwise pressure center position and the chordwise pressure center position at each station, and the elastic correction is completed.

[0019] Further, in step S6, when the chordwise correction is performed, the normal force at each section is not changed, and the pressure coefficient increment is superimposed for correction. Pressure distribution increment Δ CP ij obtained through formula (16) as follows, (16), wherein, Xf ij the ratio of the pressure center to the local section chord length is obtained through formula (17); the parameters X , p , q are obtained through formulas (18)-(20) as follows, r (17), (18), (19), (20), wherein, the physical meaning of formula (18) is that the normal force of the section remains unchanged; the physical meaning of formula (19) is that the new torque is equal to the original torque plus the increment caused by the change of the chordwise pressure center; formula (20) is the trailing edge Kutta condition, i.e. Xf ij when =1, Cp the increment is 0; the chordwise pressure center correction amount of each panel is obtained​Cp After increment, the upper and lower wings respectively perform formula (21) and formula (22) to obtain new pressure distribution, that is, the final pressure distribution, The upper wing satisfies the following formula (21) relationship: (21), The lower wing satisfies the following formula (22) relationship: (22).

[0020] The application of the wing pressure distribution elastic correction method according to any one of the preceding embodiments in the design of a UAV and an aircraft wing.

[0021] An apparatus for wing pressure distribution elastic correction, comprising at least one processor and at least one memory in communication connection with the processor; the memory stores program instructions executable by the processor; the processor calling the program instructions can execute the method according to any one of the preceding embodiments.

[0022] A computer readable storage medium, the computer readable storage medium stores computer instructions, the computer instructions are executed to execute the method according to any one of the preceding embodiments.

[0023] Compared with the prior art, the present application has the following advantages and beneficial effects: In the present application, an innovative wing pressure distribution elastic correction method is provided, and the method is reliable and provides a basis for the wings of a UAV and an aircraft. BRIEF DESCRIPTION OF DRAWINGS

[0024] Figure 1 is a flowchart of the present application. DETAILED DESCRIPTION

[0025] The present application will be further described in detail below in combination with embodiments, but the embodiments of the present application are not limited thereto.

[0026] Embodiment 1 In order to facilitate the public to understand the present application, the present embodiment further describes the present scheme in combination with drawings.

[0027] A wing pressure distribution elastic correction method, referring to Figure 1 , comprises the following steps: S1, rigid pressure distribution preprocessing.

[0028] In this step, the wing pressure rigid distribution data is usually obtained from pressure measurement test or CFD calculation, including upper wing pressure distribution and lower wing pressure distribution.

[0029] The pressure distribution form is generally as follows: a plurality of profiles are arranged along the airfoil, and a plurality of data points are provided on each profile, and the data points include coordinate values x, y, z and pressure coefficient Cp (in the coordinate system of the embodiment, x is positive in the backward direction, y is positive in the right direction, and z is positive in the upward direction). Since the number of data points on each profile is not completely consistent, especially if the data is derived from CFD unstructured grid calculation, there will inevitably be a problem of inconsistent number of data points. Therefore, the data points on each profile need to be reconstructed.

[0030] For the original data points of each profile, first, the x values are sorted, and the x value with the smallest value (denoted as x min ) is the leading edge point of the profile; the x value with the largest value (denoted as x max ) is the trailing edge point of the profile. The chord length of the profile is x max -x min . m+1 data points are arranged on the chord length in equal proportions, and the x coordinate of each data point is obtained according to the proportional relationship, the y coordinate is the profile station, and the z coordinate and the Cp value are obtained by interpolation according to the original data point information of the profile and the x coordinate.

[0031] After the above operation is performed on each profile, the number of profiles is defined as n+1, and the data points are a matrix of (n+1)×(m+1). Each adjacent four data points form a face element (a total of n×m face elements), and the center point coordinates (X, Y, Z) of the face element (in the embodiment, the face element center coordinates are represented by capital letters), the average pressure coefficient, the area and the normal vector are calculated. Among them, the face element center point coordinates and the pressure coefficient are the average values of the four data points, the face element is a quadrilateral, which is the sum of two triangles, and the area of each triangle is obtained by Heron's formula. The normal vector of the jth face element of the ith profile (N n 1ij , n 2ij , n 3ij ) is obtained by equations (1)-(3): (1), (2), (3), wherein n 3>0, for the upper surface of the airfoil, or n 3<0, for the lower surface of the airfoil.

[0032] S2, obtaining an elastic correction amount.

[0033] In this step, the CFD / CSD static aeroelastic calculation method is used to calculate the influence amount of the elastic deformation on the airfoil load, which satisfies the following equations (4)-(6): (4), (5), (6), where F 弹性 , F 刚性 are the total load of the airfoil after elastic deformation and without considering elastic deformation, respectively, K is the correction ratio; Y 弹性 , Y 刚性 are the spanwise center of pressure of the airfoil after elastic deformation and without considering elastic deformation, respectively, Dy is the spanwise center of pressure correction; X 弹性 , X 刚性 are the chordwise center of pressure of the airfoil after elastic deformation and without considering elastic deformation, respectively, Dx is the chordwise center of pressure correction.

[0034] S3, total load correction of pressure distribution.

[0035] In this step, for each element of the airfoil, the value of F Cp is multiplied by the correction coefficient K to obtain the new value of F Cp 1 Cp 1, after correction, the total load value of the airfoil is the total load value after considering elastic deformation, that is, it satisfies the following formula (7), Cp 1 ij = Cp ij ×K (7).

[0036] S4, spanwise center of pressure correction of pressure distribution.

[0037] In this step, the center of pressure correction is divided into two steps, and the spanwise correction is performed first. When the spanwise correction is performed, the total normal force of the airfoil is ensured to be unchanged. The value of F Cp 1 K i is multiplied by the parameter Cp 2 K i to obtain the new value of F 2 , that is, the spanwise correction is completed. The parameter takes different values with the spanwise station: where the parameter α , bThe values ​​are obtained using equations (10) and (11). In equation (10), the physical meaning is that the normal force on the airfoil remains constant; in equation (11), the physical meaning is that the new airfoil bending moment is the original bending moment plus the bending moment increment caused by the change in the spanwise center of gravity. C zi and S i The normal force coefficient and cross-sectional area are respectively obtained by equations (12) and (13).

[0038] (10); (11); (12); (13).

[0039] S5. After the spanwise compressive center correction, obtain the new chordwise correction amount.

[0040] In this step, after the spanwise center of gravity is corrected, the chordwise center of gravity of the original airfoil will also change. Therefore, at this time, the chordwise center of gravity of the airfoil needs to be recalculated, as shown in Equation (14), and the new chordwise center of gravity correction amount needs to be obtained, as shown in Equation (15).

[0041] (14); (15).

[0042] S6, Pressure Distribution Chordal Pressure Center Correction.

[0043] Since the spanwise indentation correction has been completed in this step, the magnitude of the normal force on each section should not be changed during the chordwise correction; otherwise, the spanwise indentation will continue to change.

[0044] The chordal correction is achieved by superimposing the pressure coefficient increment. The pressure distribution increment is shown in equation (16). Xf ij For face element X The ratio of the pressure core to the local profile chord length is obtained by equation (17).

[0045] parameter p , q , r The equations (18) to (20) are used to obtain the following equations. Equation (18) means that the normal force of the cross section remains unchanged; Equation (19) means that the new torque is equal to the original torque plus the increment caused by the change in the chordal center of gravity; Equation (20) is the Kutta condition for the trailing edge, i.e. Xf ij When =1, Cp The increment is 0.

[0046] (16), (17), (18), (19), (20), The increments of each surface element are calculated Cp After the increments, the upper and lower wings are respectively executed by formula (21) and formula (22) to calculate the new pressure distribution, that is, the final pressure distribution.

[0047] The upper wing satisfies the following formula (21) relationship: (21), The lower wing satisfies the following formula (22) relationship: (22).

[0048] Taking a certain wing as an example, the normal force is 25166N and the pressure center is (3.403, 2.715) without considering elasticity, and the normal force is 24411N and the pressure center is (3.383, 2.665) after considering elasticity. For comparison, first, multiply the coefficient 0.97 to complete the total load correction, and then correct the pressure center to (3.378, 2.665) through the spanwise pressure center correction, and finally, correct the pressure center to the target value (3.383, 2.665) through the chordwise correction, so as to complete the entire pressure distribution correction.

[0049] Embodiment 2 An apparatus for elastic correction of wing surface pressure distribution, comprising at least one processor and at least one memory connected in communication with the processor; the memory stores program instructions executable by the processor; the processor calling the program instructions can execute the method of embodiment 1.

[0050] Embodiment 3 A computer readable storage medium, the computer readable storage medium stores computer instructions, the computer instructions are executed to execute the method of embodiment 1.

[0051] The elastic correction of wing surface pressure distribution in the application can be applied in the design of unmanned aerial vehicle and aircraft wing.

[0052] The above is only a preferred embodiment of the application, and does not limit the application in any form. Any simple modification or equivalent change based on the technical essence of the application to the above embodiment falls within the protection scope of the application.

Claims

1. A method for elastically correcting pressure distribution on an airfoil, characterized in that, Includes the following steps: S1. Rigid pressure distribution pretreatment; S2, Obtain the elastic correction amount; S3, Total load correction for pressure distribution; S4, Pressure distribution spread towards the center of pressure; S5. After the spanwise centering correction, obtain the new chordwise correction amount; S6, Pressure Distribution Chordal Pressure Center Correction.

2. The method for elastic correction of airfoil pressure distribution according to claim 1, characterized in that, In step S1, the rigid pressure distribution pretreatment includes the following steps: S1.

1. Interpolate the initial rigid pressure distribution data in three-dimensional space according to the structured grid; S1.2 Sort the interpolated data in ascending order and construct a surface-based model; S1.3 Calculate the geometric parameters of each surface element, including the area vector, normal vector, and centroid coordinates.

3. The method for elastic correction of airfoil pressure distribution according to claim 1, characterized in that, The rigid pressure distribution preprocessing method is as follows: In step S1.1, the rigidity distribution data of the airfoil pressure is obtained. The rigidity distribution data of the airfoil pressure includes the pressure distribution data of the upper airfoil and the pressure distribution data of the lower airfoil. The rigidity distribution data of the airfoil pressure is obtained from pressure measurement test or CFD calculation. The airfoil pressure rigidity distribution data is provided by multiple profiles along the airfoil, with multiple raw data points distributed on each profile. Each raw data point includes coordinate values ​​(x, y, z) and a pressure coefficient. Cp ; In step S1.2, if the number of original data points on multiple cross-sections is inconsistent, the original data points on each cross-section are reconstructed, specifically including: The original data points of each profile are sorted according to their x-coordinate values, and the original data point with the smallest x-coordinate value is selected. min As the leading edge point of this profile, the original data point with the largest x-coordinate value is x max As the trailing edge point of this section, calculate the chord length of the section as x. max -x min ; m+1 reconstructed data points are arranged proportionally along the chord length. The x-coordinate of each reconstructed data point is determined according to a proportional relationship, the y-coordinate is the station position of the profile, and the z-coordinate is... Cp The value is obtained by interpolating the x-coordinate based on the original data point information of the profile. In step S1.3, after all profiles have been reconstructed, if the total number of profiles is n+1, then all reconstructed data points form a matrix of (n+1)×(m+1), and every four adjacent reconstructed data points constitute a face element, with a total of n×m face elements. Calculate the center point coordinates, average pressure coefficient, area, and normal vector of each surface element. n 1ij , n 2ij , n 3ij ),in: The center point coordinates (X, Y, Z) of the surface element are the average of the coordinates of the four reconstructed data points that constitute the surface element; The average pressure coefficient of the surface element is the sum of the four reconstructed data points constituting that surface element. Cp The average value; The surface element is a quadrilateral, and its area is the sum of the areas of the two triangles that make up the quadrilateral. The area of ​​each triangle is calculated using Heron's formula. The normal vector of the j-th surface element of the i-th profile ( n 1ij , n 2ij , n 3ij The results are obtained by calculation using the preset formulas (1) to (3): （1）， （2）， （3）， Where n3>0 represents the upper wing surface, or n3<0 represents the lower wing surface.

4. The method for elastic correction of airfoil pressure distribution according to claim 1, characterized in that, In step S2, obtaining the elasticity correction amount includes the following steps: S2.1 Static aeroelastic iterative calculations are performed using a CFD / CSD fluid-structure interaction solver. S2.2 Extract the pressure field data after the elastic deformation of the wing surface; S2.3 Calculate the difference between the elastic pressure distribution and the rigid pressure distribution to obtain the total load correction factor and the spanwise core offset.

5. The method for elastic correction of airfoil pressure distribution according to claim 3, characterized in that, In step S2, The influence of elastic deformation on the wing surface load satisfies the following equations (4) to (6). （4）， （5）， （6）， Among them, F 弹性 F 刚性 The total airfoil loads are shown after elastic deformation and without considering elastic deformation, respectively. K To adjust the ratio; Y 弹性 Y 刚性 These are the spanwise pressure centers of the wing after elastic deformation and without considering elastic deformation, respectively. Dy This is the correction amount for the expansion center of gravity; X 弹性 X 刚性 The figures show the chordal pressure center of the wing after elastic deformation and before considering elastic deformation, respectively. Dx This is the correction amount for the chord-direction pressure center.

6. The method for elastic correction of airfoil pressure distribution according to claim 1, characterized in that, In step S3, the total load correction of the pressure distribution is to correct the total load of the rigid pressure distribution according to the elastic correction amount, so as to ensure that the corrected pressure distribution meets the aerodynamic balance conditions of the whole machine.

7. The method for elastic correction of airfoil pressure distribution according to claim 5, characterized in that, For each face element Cp The value is obtained by multiplying by a scaling factor. K , get new Cp value Cp 1. After correction, the total load value of the wing surface is the total load value after considering elastic deformation, which satisfies the following formula (7). Cp 1 ij = Cp ij ×K （7）。 8. The method for elastic correction of airfoil pressure distribution according to claim 1, characterized in that, In step S4, the spanwise pressure center correction is to perform spanwise integration on the corrected pressure distribution and make the first correction to the pressure center position based on the spanwise load distribution characteristics of the airfoil.

9. The method for elastic correction of airfoil pressure distribution according to claim 7, characterized in that: In step S4, during spanwise correction, the total normal force on the airfoil remains constant. This is achieved by adjusting the force on each surface element of the upper and lower airfoil. Cp 1 multiplied by parameter K i , get new Cp 2, that is, to complete the spanwise correction, parameters K i The station position varies with the exhibition direction, while simultaneously satisfying the following equations (8) and (9): （8）， （9）， Among them, parameters α , b Obtained through equations (10) and (11); （10）； （11）； In this context, equation (10) means that the normal force on the airfoil remains constant; equation (11) means that the new airfoil bending moment is the original bending moment plus the bending moment increment caused by the change in the spanwise pressure center. In the formula C zi and S i These are the normal force coefficient and cross-sectional area of ​​the cross section, respectively, and are obtained through equations (12) and (13). （12）； （13）。 10. The method for elastic correction of airfoil pressure distribution according to claim 1, characterized in that, In step S5, the method for obtaining the new chordal correction amount is as follows: based on the pressure distribution corrected in step S4, calculate the chordal pressure center position at each station along the span, and construct the chordal correction coefficient function based on the chordal pressure center offset law before and after elastic deformation.

11. The method for elastic correction of airfoil pressure distribution according to claim 9, characterized in that, In step S5: recalculate the chordal pressure center of the airfoil according to the following formula (14), and recalculate the chordal pressure center correction amount according to the following formula (15). （14）； （15）。 12. The method for elastic correction of airfoil pressure distribution according to claim 1, characterized in that, In step S6, the method for correcting the chordal pressure center of the pressure distribution is as follows: the chordal correction coefficient obtained in step S5 is applied to the pressure distribution to ensure that the final corrected pressure distribution simultaneously satisfies the spanwise pressure center position and the chordal pressure center position of each station, thus completing the elastic correction.

13. The method for elastic correction of airfoil pressure distribution according to claim 11, characterized in that: In step S6, during chordal correction, the magnitude of the normal force on each cross section is not changed; correction is achieved by superimposing the pressure coefficient increment. Pressure distribution increment Δ CP ij It is obtained through the following formula (16), （16）， in, Xf ij For face element X The ratio of the pressure core to the local profile chord length is obtained through equation (17); parameters p , q , r We obtain the results through equations (18) to (20). （17）， （18）， （19）， （20）， In this context, equation (18) means that the normal force of the profile remains constant; equation (19) means that the new torque is equal to the original torque plus the increment caused by the change in the chordal center of gravity; and equation (20) is the trailing edge Kutta condition, i.e. Xf ij When =1, Cp The increment is 0; Find the extract of each face element. Cp After the increment, equations (21) and (22) are applied to the upper and lower wing surfaces respectively to obtain the new pressure distribution, which is the final pressure distribution. The upper wing surface satisfies the following relationship (21): （21）， The lower wing surface satisfies the following relationship (22): （22）。 14. The application of the elastic correction method for wing pressure distribution as described in any one of claims 1 to 13 in the design of unmanned aerial vehicles and aircraft wings.

15. A device for elastic correction of airfoil pressure distribution, characterized in that: It includes at least one processor and at least one memory communicatively connected to the processor; The memory stores program instructions that can be executed by the processor; the processor can invoke the program instructions to perform the method as described in any one of claims 1 to 13.

16. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores computer instructions that, when executed, can be used to perform the method as described in any one of claims 1 to 13.

Images