A method for correcting aerodynamic roll moment of an aircraft

Abstract

The application belongs to the technical field of aircraft flight load design, and particularly relates to a correction method of an aircraft aerodynamic roll moment, initialization of a calculation parameter, calculation of an aircraft true airspeed, equivalent speed and aircraft speed pressure based on the calculation parameter, calculation of a motion parameter based on a dynamic response of a maneuver simulation, calculation of a full-aircraft target aerodynamic roll moment based on the motion parameter, given of an unbalanced moment tolerance, solving of the unbalanced moment based on the full-aircraft target roll moment, comparison and analysis of a difference value between the unbalanced moment and the unbalanced moment tolerance, and end of the calculation if an absolute value of the unbalanced moment is less than the unbalanced moment tolerance. By adjusting a local angle of attack of each profile of a wing component in a span direction, a wing spanwise pressure distribution characteristic is reconstructed, an unbalanced aerodynamic roll moment of the aircraft caused by the pressure distribution characteristic is eliminated, a load balance purpose is achieved, and a design requirement in engineering is 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] Aircraft flight load calculations involve a wide range of raw data, including the aircraft's overall and component aerodynamic characteristics, weight characteristics, geometry, engine dynamics, control surface control patterns, flight speed, altitude, and flight envelope. Any change in any of these data will significantly impact the aircraft's final design loads.

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

[0004] ∑F(x,y,z)=0

[0005] ∑M(x,y,z)=0

[0006] Simplified load models are a combination of numerous input and output data files and numerous computer programs. Many traditional aerodynamic load correction methods often modify the imbalance relationship through certain coefficients, which can cause distorted component loads. Therefore, it is necessary to develop a method that iterates the component pressure distribution characteristics to correct the imbalance loads generated during flight load calculations. Summary of the Invention

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

[0008] The technical solution of the present application is: a method for correcting an aircraft aerodynamic rolling moment, comprising: initializing calculation parameters, calculating the aircraft's true airspeed, equivalent speed, and aircraft speed pressure based on the calculation parameters, then calculating motion parameters based on a maneuver simulation dynamic response, and calculating the entire aircraft's target aerodynamic rolling moment based on the motion parameters; and setting 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 separately 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 entire 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 entire 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's mass characteristics, aerodynamic characteristics, control 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's 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 the three directions;

[0015] Calculate the target rolling moment Mx of the entire aircraft based on the aircraft speed pressure Q, angle of attack α, sideslip angle β and aerodynamic characteristics data BW ;

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

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

[0018] Q=0.5ρV 2 ;

[0019] The aircraft's 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 force 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 grid node coordinates with a total number of nodes M, and obtain the node coordinate matrix D based on the grid 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 of the triangular mesh with a total number of N meshes, and obtain the coordinate matrix D based on the centroid 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 unit vector matrix of the triangular mesh centroid with a total number of N meshes

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

[0035]

[0036] Coordinate matrix D of the centroid 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] Calculate the local angle of attack α of a certain section based on the aircraft's angle of attack α and the spanwise section position of the wing 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 , interpolation calculates the pressure coefficient Cp of each grid centroid i , according to the pressure coefficient Cp i Get the grid centroid 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 of 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 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 of 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 torque ΔMx is less than the unbalanced torque tolerance R Mx ;

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

[0060] Preferably, if not, then based on the total aerodynamic load of 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.

[0061] Preferably, the calculation is 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 reconstructs the span-wise pressure distribution characteristics of the wing by adjusting the local angle of attack of each span-wise section of the wing component, thereby eliminating the unbalanced aerodynamic rolling moment of the aircraft caused by the pressure distribution characteristics, achieving the purpose of load balance, and meeting the design requirements in the engineering project. BRIEF DESCRIPTION OF THE DRAWINGS

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

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

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

[0067] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. 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 making creative efforts are within the scope of protection of the present invention.

[0068] A correction method for the aerodynamic rolling moment of an aircraft is proposed. The unbalanced aerodynamic rolling moment is eliminated by iteratively changing 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, and aircraft speed pressure based on the calculation parameters, then calculate the aircraft aerodynamic attitude angle and other motion parameters based on the maneuver simulation dynamic response, and calculate the target aerodynamic rolling moment of the entire aircraft based on the motion parameters; and 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 velocity pressure Q as:

[0073] Q=0.5ρV 2(1)

[0074] S12, given the aircraft mass characteristics, aerodynamic characteristics, control characteristics, engine characteristics, and design envelope, establish a six-degree-of-freedom dynamic model of the aircraft, 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, based on the true airspeed V in S11 and the velocity components u, v, and w in the 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 entire aircraft BW Given unbalanced moment tolerance R Mx .

[0080] Step S2: 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;

[0081] Preferably, the method for establishing the wing aerodynamic force 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, according to the aerodynamic model in S21, establish the a The Four Dimensions of Change The database is:

[0085]

[0086] Step S3: Initialize the geometric model parameters of the wing, establish a 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 separately 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 the grid node coordinates with a total number of nodes M, and obtain the node coordinate matrix D based on the grid node coordinates with a total number of nodes M Mpoint ; Among them, the coordinates of each grid node are X k 、Y k 、Z k , node coordinate matrix D Mpoint for:

[0089]

[0090] S32, based on 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 of the triangular mesh with a total number of N meshes, and obtain the coordinate matrix D based on the centroid of the triangular mesh Cmesh ; The coordinates of each grid centroid are x i 、y i 、z i , the coordinate matrix of the centroid of the triangle mesh D 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 unit vector matrix of the triangular mesh centroid with a total number of N meshes for:

[0095]

[0096] Step S4, calculating the wing aerodynamic load and aerodynamic rolling moment based on the motion parameters in step S1, the wing aerodynamic force CP database in step S2, and the triangular grid centroid coordinate matrix in step S3, and solving the unbalanced moment based on the full-aircraft target rolling moment 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 based on 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 profile 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 , interpolation calculates the pressure coefficient Cp of each grid centroid i , according to the pressure coefficient Cp i Get the grid centroid 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 of 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 grid cells 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 full 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: Compare and analyze 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 torque ΔMx is less than the unbalanced torque tolerance R Mx :

[0113]

[0114] S52, if yes, then output the total aerodynamic load of the wing in S44 If not, 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. Other structures can refer to common designs. In the absence of conflicts, 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 scope of protection of the present invention.

Claims

1. A method for correcting an aircraft's aerodynamic rolling moment, characterized in that: include: Initialize the calculation parameters, calculate the true airspeed, equivalent speed, and aircraft speed pressure based on the calculation parameters, then calculate the motion parameters based on the dynamic response of the maneuver simulation, and calculate the target aerodynamic rolling moment of the entire 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 separately 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 entire 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. The specific method for solving the unbalanced torque is: Calculate the local angle of attack α of a certain section based on the aircraft's angle of attack α and the spanwise section position of the wing 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 , triangle mesh centroid coordinate matrix D Cmesh And the current state pressure distribution matrix D of the wing NowCp , interpolation calculates the pressure coefficient Cp of each grid centroid i , according to the pressure coefficient Cp i Get the grid centroid 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 of 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 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.

2. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 1, wherein: 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's mass characteristics, aerodynamic characteristics, control 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's 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 the three directions; Calculate the target rolling moment Mx of the entire aircraft based on the aircraft speed pressure Q, angle of attack α, sideslip angle β and aerodynamic characteristics data BW ; Based on the target rolling moment Mx of the entire aircraft BW Given unbalanced moment tolerance R Mx .

3. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 2, wherein: Aircraft speed pressure Q is: Q=0.5ρV 2 ; The aircraft's 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, wherein: The method for establishing the wing aerodynamic CP database is as follows: 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, wherein: 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 according to claim 1, wherein: 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 grid node coordinates with a total number of nodes M, and obtain the node coordinate matrix D based on the grid 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 of the triangular mesh with a total number of N meshes, and obtain the coordinate matrix D based on the centroid 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 unit vector matrix of the triangular mesh centroid with a total number of N meshes 7. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 6, wherein: Node coordinate matrix D Mpoint for: Coordinate matrix D of the centroid 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 according to claim 1, wherein: 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 of each grid cell for: Total wing aerodynamic load for: The unbalanced moment ΔMx is: ΔMx=Mx Bw -Mx(a,b,d) a ,Ma).

9. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 1, wherein: 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 torque ΔMx is less than the unbalanced torque tolerance R Mx ; If so, output the total aerodynamic load on the wing 10. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 9, wherein: If not, 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.

11. The method for correcting the aerodynamic rolling moment of an aircraft according to claim 9, wherein: 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 。

Images