Aircraft pneumatic rolling torque correction method

Abstract

The invention belongs to the technical field of aircraft flight load design, and particularly relates to an aircraft pneumatic rolling torque correction method, which comprises the steps of initializing calculation parameters, calculating aircraft vacuum speed, equivalent speed and aircraft speed pressure based on the calculation parameters, then calculating motion parameters based on maneuvering simulation dynamic response, and calculating full-aircraft target pneumatic rolling torque based on the motion parameters. Giving unbalanced torque tolerance; solving an unbalanced moment based on the whole-aircraft target rolling moment; and comparing and analyzing the difference value between the unbalanced moment and the unbalanced moment tolerance, and if the absolute value of the unbalanced moment is smaller than the unbalanced moment tolerance, ending the calculation. By adjusting the local attack angle of each section in the spanwise direction of the wing part and reconstructing the pressure distribution characteristic in the spanwise direction of the wing, the unbalanced pneumatic rolling moment of the aircraft caused by the pressure distribution characteristic is eliminated, the purpose of load balance is achieved, and the design requirements in engineering are met.

Description

Technical Field

[0001] The present application belongs to the technical field of aircraft flight load design, and in particular relates to a method for correcting an aircraft aerodynamic rolling moment. Background Art

[0002] The calculation of aircraft flight load involves a wide range of raw data, including the aerodynamic characteristics of the entire aircraft and its components, weight characteristics, geometric shape, engine power characteristics, control surface control laws, flight speed, altitude and flight envelope, etc. Any change in these data will have a significant impact on the final design load of the aircraft.

[0003] Today's computer power allows extensive load studies to be performed, taking into account all weight information (weight, center of gravity, moment of inertia, weight distribution) for the entire aircraft and specific aircraft components, aerodynamic information (aerodynamic pressure, aerodynamic coefficients / derivatives) for different Mach numbers, and static aeroelastic inputs (elastic coefficients and increments) to correct the rigid aerodynamic forces (aerodynamic pressure, aerodynamic coefficients / derivatives) to obtain a balanced load case for the entire aircraft. This means that the sum of the forces and moments of the various aircraft components is zero for each load case and is:

[0004] ∑F(x,y,z)=0

[0005] ∑M(x,y,z)=0

[0006] The simplified load model is a combination of a large number of input and output data files and many computer programs. Many traditional aerodynamic load correction methods often change the unbalanced relationship through a certain coefficient, which will cause the morphological deformation of the component load. Therefore, it is necessary to establish a method to correct the unbalanced load generated in the flight load calculation by iterating the pressure distribution characteristics of the component. Summary of the invention

[0007] The purpose of the present application is to provide a method for correcting the aerodynamic rolling moment of an aircraft, so as to solve the problem that the traditional aerodynamic load correction method changes the unbalanced relationship through a certain coefficient, which may cause the morphological deformity of the component load.

[0008] The technical solution of the present application is: a method for correcting the aerodynamic rolling moment of an aircraft, comprising: initializing calculation parameters, calculating the true airspeed, equivalent speed, and aircraft speed pressure of the aircraft based on the calculation parameters, and then calculating the motion parameters based on the dynamic response of the maneuver simulation, and calculating the target aerodynamic rolling moment of the whole aircraft based on the motion parameters; giving an unbalanced moment tolerance;

[0009] Initialize the aerodynamic model, based on the pressure distribution data Cp of the wing longitudinal, lateral heading and aileron deflection state at different aircraft flight speeds Ma in the aerodynamic model A , establish the wing aerodynamic CP database;

[0010] Initialize the geometric model parameters of the wing, establish the wing triangle mesh model, and calculate the grid node coordinates and node coordinate matrix D with a total number of nodes M based on the wing triangle mesh model. Mpoint ; Based on the grid node coordinates with a total number of nodes M and the node coordinate matrix D Mpoint Solve the triangle mesh centroid coordinate matrix D Cmesh , area matrix D Smesh and the triangle mesh centroid unit vector matrix Based on the motion parameters, the wing aerodynamic CP database and the coordinate matrix of the centroid of the triangular grid, the wing aerodynamic load and aerodynamic rolling moment are calculated, and the unbalanced moment is solved based on the target rolling moment of the whole aircraft;

[0011] Compare and analyze the difference between the unbalanced torque and the unbalanced torque tolerance. If the absolute value of the unbalanced torque is less than the unbalanced torque tolerance, the calculation ends.

[0012] Preferably, the specific calculation method of the motion parameters and the target aerodynamic rolling moment of the whole aircraft is:

[0013] Given the aircraft flight speed Ma, flight altitude H, calculate the aircraft vacuum speed V, aircraft equivalent speed Ve, atmospheric density ρ, and calculate the aircraft speed pressure Q;

[0014] Given the aircraft mass characteristics, aerodynamic characteristics, handling characteristics, engine characteristics, and design envelope, a six-degree-of-freedom dynamic model of the aircraft is established. The fourth-order Runge-Kutta method is used to numerically calculate the velocity components u, v, and w in the three directions of the aircraft body axis system, and the aileron deflection δ is obtained by interpolation calculation. a ; Calculate the aircraft's angle of attack α and sideslip angle β based on the aircraft's true airspeed V and the velocity components u, v, and w in three directions;

[0015] According to the aircraft speed pressure Q, angle of attack α, sideslip angle β and aerodynamic characteristics data, calculate the target rolling moment Mx of the whole aircraft BW ;

[0016] Based on the target rolling moment Mx of the whole aircraft BW Given the unbalanced moment tolerance R Mx .

[0017] Preferably, the aircraft speed pressure Q is:

[0018] Q=0.5ρV 2 ;

[0019] The aircraft angle of attack α and sideslip angle β are:

[0020] α=tan -1 (w / u)

[0021] β=sin-1 (vV).

[0022] Preferably, the method for establishing the wing aerodynamic CP database is:

[0023] Initialize the aerodynamic model, extract the m×n pressure coefficient points on the wing surface in span and chord direction, and calculate the position coordinate matrix W of the wing pressure coefficient points Xaero , W Yaero , W Zaero ;

[0024] According to the aerodynamic model, the a The four dimensions of change database.

[0025] Preferably, the wing pressure coefficient point position coordinate matrix W Xaero , W Yaero , W Zaero for:

[0026]

[0027] Four Dimensions The database is:

[0028]

[0029] Preferably, the specific method of establishing the wing geometric grid model is:

[0030] Initialize the geometric model parameters of the wing, establish the wing triangle mesh model, calculate the mesh node coordinates with a total number of nodes M, and obtain the node coordinate matrix D based on the mesh node coordinates with a total number of nodes M. Mpoint ;

[0031] According to the grid node coordinates with a total number of nodes M and the node coordinate matrix D Mpoint , solve the coordinates of the centroid points of the triangular mesh with a total number of meshes N, and obtain the coordinate matrix D based on the centroid points of the triangular mesh Cmesh ;

[0032] According to the grid node coordinates with a total number of nodes M and the node coordinate matrix D Mpoint , solve the triangular mesh area S with a total mesh size of N i , based on the triangle mesh area S i Get the area matrix D Smesh ;

[0033] According to the three-dimensional grid node information and node coordinate matrix D Mpoint , solve the centroid unit vector matrix of the triangular mesh with a total number of N meshes

[0034] Preferably, the node coordinate matrix D Mpoint for:

[0035]

[0036] Coordinate matrix D of the centroid point of the triangle mesh Cmesh for:

[0037]

[0038] Area Matrix D Smesh for:

[0039]

[0040] Triangle mesh centroid unit vector matrix for:

[0041]

[0042] Preferably, the specific method for solving the unbalanced torque is:

[0043] According to the aircraft angle of attack α and the wing span profile position, calculate the local angle of attack α of a certain profile k , the local angle of attack matrix A with m spanwise sections section ;

[0044] According to the pressure distribution coefficient point coordinate matrix D Xaero , D Zaero , D Zaero , triangle mesh centroid coordinate matrix D Cmesh And the current state pressure distribution matrix D of the wing NowCp , interpolate and calculate the pressure coefficient Cp of each grid centroid i , according to the pressure coefficient Cp i Get the grid centroid point pressure coefficient matrix D MeshCp ;

[0045] According to the grid centroid pressure coefficient matrix D MeshCp , aircraft speed pressure Q, grid area matrix D Smesh , grid centroid unit vector matrix Calculate the aerodynamic force for each grid cell Select the center of gravity of the aircraft as the reference point x ref ,y ref 、z ref , the aerodynamic force of the N grid cells Integrate and calculate the total aerodynamic load of the wing Extract the target rolling moment Mx(α,β,δ a ,Ma), according to the target rolling moment Mx of the whole aircraft BW, solve the unbalanced moment ΔMx.

[0046] Preferably, the pressure distribution matrix D NowCp for:

[0047]

[0048] Grid centroid pressure coefficient matrix D MeshCp for:

[0049]

[0050] According to the unit vector matrix of the triangle mesh centroid Calculate the aerodynamic force for each grid cell for:

[0051]

[0052] Total wing aerodynamic load for:

[0053]

[0054]

[0055] The unbalanced moment ΔMx is:

[0056] ΔMx=Mx Bw -Mx(α,β,δ a ,Ma).

[0057] Preferably, the specific method for comparing the difference between the unbalanced torque and the unbalanced torque tolerance is:

[0058] According to the unbalanced moment ΔMx and the unbalanced moment tolerance R Mx , calculate and determine whether the absolute value of the unbalanced moment ΔMx is less than the unbalanced moment tolerance R Mx ;

[0059] If so, output the total aerodynamic load on the wing

[0060] Preferably, if not, then based on the total wing aerodynamic load Update the local angle of attack matrix A of the wing span profile section Until it is determined that |ΔMx| is less than R Mx , the calculation ends.

[0061] Preferably, the calculation is performed to determine whether the absolute value of the unbalanced moment ΔMx is less than the unbalanced moment tolerance R Mx The formula is:

[0062] |ΔMx|<?R Mx .

[0063] The method for correcting the aerodynamic rolling moment of an aircraft in the present application adjusts the local angle of attack of each span-wise section of the wing component, reconstructs the span-wise pressure distribution characteristics of the wing, eliminates the unbalanced aerodynamic rolling moment of the aircraft caused by the pressure distribution characteristics, achieves the purpose of load balance, and meets the design requirements in the project. BRIEF DESCRIPTION OF THE DRAWINGS

[0064] In order to more clearly illustrate the technical solution provided by the present application, the following is a brief introduction to the accompanying drawings. Obviously, the accompanying drawings described below are only some embodiments of the present application.

[0065] Figure 1 This is a schematic diagram of the overall process of this application;

[0066] Figure 2 This is a schematic diagram of the iteratively corrected aerodynamic roll moment of the aircraft in this application and the target aerodynamic roll moment of the entire aircraft. DETAILED DESCRIPTION

[0067] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0068] A correction method for an aircraft aerodynamic rolling moment eliminates the unbalanced aerodynamic rolling moment by iteratively varying parameters, thereby achieving the purpose of load balance and meeting the design requirements in engineering.

[0069] like Figure 1 , specifically including the following steps:

[0070] Step S1, initialize calculation parameters, including aircraft flight speed, flight altitude, aircraft mass characteristics, etc., calculate the aircraft true airspeed, equivalent speed, aircraft speed pressure based on the calculation parameters, and then calculate the aircraft aerodynamic attitude angle and other motion parameters based on the maneuver simulation dynamic response, calculate the target aerodynamic rolling moment of the whole aircraft based on the motion parameters; give the unbalanced moment tolerance.

[0071] Preferably, if Figure 2 , the specific calculation method of motion parameters and the target aerodynamic rolling moment of the whole aircraft is:

[0072] S11, given the aircraft flight speed Ma, flight altitude H, calculate the aircraft vacuum speed V, aircraft equivalent speed Ve, atmospheric density ρ, and calculate the aircraft speed pressure Q as:

[0073] Q=0.5ρV 2(1)

[0074] S12, given the aircraft mass characteristics, aerodynamic characteristics, control characteristics, engine characteristics, design envelope, establish the aircraft six-degree-of-freedom dynamic model, use the fourth-order Runge-Kutta method to numerically calculate the velocity components u, v, and w in the three directions of the aircraft body axis system, and interpolate to obtain the aileron deflection δ a ;

[0075] S13, according to the true airspeed V of the aircraft in S11 and the velocity components u, v, and w in three directions in S12, calculate the aircraft angle of attack α and sideslip angle β as follows:

[0076] α=tan -1 (w / u) (2)

[0077] β=sin -1 (v / V) (3)

[0078] S14, calculate the target rolling moment Mx of the whole aircraft based on the aircraft speed pressure Q in S11, the angle of attack α and sideslip angle β in S13, and the aerodynamic characteristics data BW ;

[0079] S15, based on the target rolling moment Mx of the whole aircraft BW Given the unbalanced moment tolerance R Mx .

[0080] Step S2, initializing the aerodynamic model, based on the pressure distribution data Cp of the wing longitudinal, lateral heading and aileron deflection state at different aircraft flight speeds Ma in the aerodynamic model A , establish the wing aerodynamic CP database;

[0081] Preferably, the method for establishing the wing aerodynamic CP database is:

[0082] S21, initialize the aerodynamic model, extract the m×n pressure coefficient points on the wing surface in span and chord direction, and calculate the position coordinate matrix W of the wing pressure coefficient points Xaero , W Yaero , W Zaero for:

[0083]

[0084] S22, based on the aerodynamic model in S21, establish the a The four dimensions of change The database is:

[0085]

[0086] Step S3, initializing the geometric model parameters of the wing, establishing a wing triangle mesh model, and calculating the grid node coordinates and node coordinate matrix D with a total number of nodes M based on the wing triangle mesh model. Mpoint ; Based on the grid node coordinates with a total number of nodes M and the node coordinate matrix D Mpoint Solve the triangle mesh centroid coordinate matrix D Cmesh , area matrix D Smesh and the triangle mesh centroid unit vector matrix

[0087] Preferably, the specific method of establishing the wing geometric grid model is:

[0088] S31, initialize the geometric model parameters of the wing, establish the wing triangle mesh model, calculate and obtain the mesh node coordinates with a total number of nodes M, and obtain the node coordinate matrix D based on the mesh node coordinates with a total number of nodes M Mpoint ; where each grid node coordinate is X k , Y k , Z k , node coordinate matrix D Mpoint for:

[0089]

[0090] S32, according to the grid node coordinates with a total number of nodes M in S31 and the node coordinate matrix D Mpoint , solve the coordinates of the centroid points of the triangular mesh with a total number of meshes N, and obtain the coordinate matrix D based on the centroid points of the triangular mesh Cmesh ; The coordinates of each grid centroid are x i ,y i 、z i , the coordinate matrix D of the centroid of the triangle mesh Cmesh for:

[0091]

[0092] S33, according to the grid node coordinates with a total number of nodes M in S31 and the node coordinate matrix D Mpoint , solve the triangular mesh area S with a total mesh size of N i , based on the triangle mesh area S i Get the area matrix D Smesh ; Area matrix D Smesh for:

[0093]

[0094] S34, based on the three-dimensional grid node information and node coordinate matrix D in S31 Mpoint, solve the centroid unit vector matrix of the triangular mesh with a total number of N meshes for:

[0095]

[0096] Step S4, based on the motion parameters in step S1, the wing aerodynamic CP database in step S2, and the coordinate matrix of the centroid of the triangular grid in step S3, the wing aerodynamic load and the aerodynamic rolling moment are calculated, and the unbalanced moment is solved based on the target rolling moment of the whole aircraft in step 1;

[0097] Preferably, the specific method for solving the unbalanced torque is:

[0098] S41, calculate the local angle of attack α of a certain section according to the aircraft angle of attack α and the wing span profile position in S13 k , the local angle of attack matrix A with m spanwise sections section for:

[0099] A section =[α1…α k …α m ] (10)

[0100] S42, according to the aircraft flight speed Ma in S11, the aileron deflection δ in S12 a , the spanwise angle of attack of the wing in S41 is α k , sideslip angle β, for S22 The database interpolation calculates the wing section pressure distribution data Cp in the current state N , and then get the pressure distribution matrix D NowCp for:

[0101]

[0102] S42, according to the pressure distribution coefficient point coordinate matrix D in S21 Xaero , D Zaero , D Zaero 、The coordinate matrix D of the centroid point of the triangle mesh in S32 Cmesh And the current state pressure distribution matrix D of the S41 wing NowCp , interpolate and calculate the pressure coefficient Cp of each grid centroid i , according to the pressure coefficient Cp i Get the grid centroid point pressure coefficient matrix D MeshCp for:

[0103]

[0104] S43, according to the grid centroid pressure coefficient matrix D in S42 MeshCp, aircraft speed pressure Q in S11, grid area matrix D in S33 Smesh 、S34 grid centroid unit vector matrix Calculate the aerodynamic force for each grid cell for:

[0105]

[0106] S44, select the center of gravity of the aircraft as the reference point x ref ,y ref 、z ref , the aerodynamic force of the N grids in S43 Integrate and calculate the total aerodynamic load of the wing for:

[0107]

[0108] S45, extract the whole machine target rolling moment Mx (α, β, δ a ,Ma), according to the whole aircraft target rolling moment Mx of S14 BW , solve the unbalanced moment ΔMx as:

[0109] ΔMx=Mx Bw -Mx(α,β,δ a ,Ma) (16)

[0110] Step S5, comparing and analyzing the difference between the unbalanced torque in step 4 and the unbalanced torque tolerance in step 1, if the absolute value of the unbalanced torque is smaller, the calculation ends.

[0111] Preferably, the specific method for comparing the difference between the unbalanced torque and the unbalanced torque tolerance is:

[0112] S51, according to the unbalanced moment ΔMx in S45, the unbalanced moment tolerance R in step S15 Mx , calculate and determine whether the absolute value of the unbalanced moment ΔMx is less than the unbalanced moment tolerance R Mx :

[0113]

[0114] S52, if yes, then output the total aerodynamic load of the wing in S44 If not, then based on the total aerodynamic load on the wing Update the local angle of attack matrix A of the wing span profile section Until it is determined that |ΔMx| is less than R Mx , the calculation ends.

[0115] In summary, by adjusting the local angle of attack of each span-wise section of the wing component and reconstructing the span-wise pressure distribution characteristics of the wing, the unbalanced aerodynamic rolling moment of the aircraft caused by the pressure distribution characteristics is eliminated, the purpose of load balance is achieved, and the design requirements in the project are met.

[0116] Finally, it should be noted that: the drawings of the embodiments disclosed in the present invention only involve structures related to the embodiments disclosed in the present invention, and other structures can refer to the general design. In the absence of conflict, the same embodiment and different embodiments of the present invention can be combined with each other;

[0117] Finally: The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the protection scope of the present invention.

Claims

1. A method for correcting an aircraft aerodynamic rolling moment, characterized in that: include: Initialize calculation parameters, calculate true airspeed, equivalent speed, and aircraft speed pressure based on the calculation parameters, then calculate motion parameters based on maneuver simulation dynamic response, and calculate the target aerodynamic rolling moment of the whole aircraft based on the motion parameters; Given the unbalanced moment tolerance; Initialize the aerodynamic model, based on the pressure distribution data Cp of the wing longitudinal, lateral heading and aileron deflection state at different aircraft flight speeds Ma in the aerodynamic model A , establish the wing aerodynamic CP database; Initialize the geometric model parameters of the wing, establish the wing triangle mesh model, and calculate the grid node coordinates and node coordinate matrix D with a total number of nodes M based on the wing triangle mesh model. Mpoint ; Based on the grid node coordinates with a total number of nodes M and the node coordinate matrix D Mpoint Solve the triangle mesh centroid coordinate matrix D Cmesh , area matrix D Smesh and the triangle mesh centroid unit vector matrix Based on the motion parameters, the wing aerodynamic CP database and the coordinate matrix of the centroid of the triangular grid, the wing aerodynamic load and aerodynamic rolling moment are calculated, and the unbalanced moment is solved based on the target rolling moment of the whole aircraft; Compare and analyze the difference between the unbalanced torque and the unbalanced torque tolerance. If the absolute value of the unbalanced torque is less than the unbalanced torque tolerance, the calculation ends.

2. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 1, characterized in that: The specific calculation method of motion parameters and the target aerodynamic rolling moment of the whole aircraft is: Given the aircraft flight speed Ma, flight altitude H, calculate the aircraft vacuum speed V, aircraft equivalent speed Ve, atmospheric density ρ, and calculate the aircraft speed pressure Q; Given the aircraft mass characteristics, aerodynamic characteristics, handling characteristics, engine characteristics, and design envelope, a six-degree-of-freedom dynamic model of the aircraft is established. The fourth-order Runge-Kutta method is used to numerically calculate the velocity components u, v, and w in the three directions of the aircraft body axis system, and the aileron deflection δ is obtained by interpolation calculation. a ; Calculate the aircraft's angle of attack α and sideslip angle β based on the aircraft's true airspeed V and the velocity components u, v, and w in three directions; According to the aircraft speed pressure Q, angle of attack α, sideslip angle β and aerodynamic characteristics data, calculate the target rolling moment Mx of the whole aircraft BW ; Based on the target rolling moment Mx of the whole aircraft BW Given the unbalanced moment tolerance R Mx .

3. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 2, characterized in that: Aircraft speed pressure Q is: Q=0.5ρV 2 ; The aircraft angle of attack α and sideslip angle β are: α=tan -1 (w / u) β=sin -1 (vV).

4. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 1, characterized in that: The method to establish the wing aerodynamic CP database is: Initialize the aerodynamic model, extract the m×n pressure coefficient points on the wing surface in span and chord direction, and calculate the position coordinate matrix W of the wing pressure coefficient points Xaero , W Yaero , W Zaero ; According to the aerodynamic model, the a The four dimensions of change database.

5. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 4, characterized in that: Wing pressure coefficient point position coordinate matrix W Xaero , W Yaero , W Zaero for: Four Dimensions The database is:

6. The method for correcting the aerodynamic rolling moment of an aircraft as claimed in claim 1, characterized in that: The specific method of establishing the wing geometric mesh model is: Initialize the geometric model parameters of the wing, establish the wing triangle mesh model, calculate the mesh node coordinates with a total number of nodes M, and obtain the node coordinate matrix D based on the mesh node coordinates with a total number of nodes M. Mpoint ; According to the grid node coordinates with a total number of nodes M and the node coordinate matrix D Mpoint , solve the coordinates of the centroid points of the triangular mesh with a total number of meshes N, and obtain the coordinate matrix D based on the centroid points of the triangular mesh Cmesh ; According to the grid node coordinates with a total number of nodes M and the node coordinate matrix D Mpoint , solve the triangular mesh area S with a total mesh size of N i , based on the triangle mesh area S i Get the area matrix D Smesh ; According to the three-dimensional grid node information and node coordinate matrix D Mpoint , solve the centroid unit vector matrix of the triangular mesh with a total number of N meshes 7. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 6, characterized in that: Node coordinate matrix D Mpoint for: Coordinate matrix D of the centroid point of the triangle mesh Cmesh for: Area Matrix D Smesh for: Triangle mesh centroid unit vector matrix for:

8. The method for correcting the aerodynamic rolling moment of an aircraft as claimed in claim 1, characterized in that: The specific method for solving the unbalanced torque is: According to the aircraft angle of attack α and the wing span profile position, calculate the local angle of attack α of a certain profile k , the local angle of attack matrix A with m spanwise sections section ; According to the pressure distribution coefficient point coordinate matrix D Xaero , D Zaero , D Zaero , the coordinate matrix D of the centroid of the triangle mesh Cmesh And the current state pressure distribution matrix D of the wing NowCp , interpolate and calculate the pressure coefficient Cp of each grid centroid i , according to the pressure coefficient Cp i Get the grid centroid point pressure coefficient matrix D MeshCp ; According to the grid centroid pressure coefficient matrix D MeshCp , aircraft speed pressure Q, grid area matrix D Smesh , grid centroid unit vector matrix Calculate the aerodynamic force for each grid cell Select the center of gravity of the aircraft as the reference point x ref ,y ref 、z ref , the aerodynamic force of the N grid cells Integrate and calculate the total aerodynamic load of the wing Extract the aerodynamic target rolling moment Mx(α,β,δ a ,Ma), according to the target rolling moment Mx of the whole aircraft BW , solve the unbalanced moment ΔMx.

9. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 8, characterized in that: Pressure distribution matrix D NowCp for: Grid centroid pressure coefficient matrix D MeshCp for: According to the unit vector matrix of the triangle mesh centroid Calculate the aerodynamic force for each grid cell for: Total wing aerodynamic load for: The unbalanced moment ΔMx is: ΔMx=Mx Bw -Mx(a,b,d) a ,Ma).

10. The method for correcting the aerodynamic rolling moment of an aircraft as claimed in claim 1, characterized in that: The specific method for comparing the difference between the unbalanced torque and the unbalanced torque tolerance is: According to the unbalanced moment ΔMx and the unbalanced moment tolerance R Mx , calculate and determine whether the absolute value of the unbalanced moment ΔMx is less than the unbalanced moment tolerance R Mx ; If so, output the total aerodynamic load on the wing 11. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 10, characterized in that: If not, then based on the total aerodynamic load on the wing Update the local angle of attack matrix A of the wing span profile section , until it is determined that |ΔMx| is less than R Mx , the calculation ends.

12. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 10, characterized in that: Calculate and determine whether the absolute value of the unbalanced moment ΔMx is less than the unbalanced moment tolerance R Mx The formula is: |ΔMx|<?R Mx 。

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