A method for modifying aerodynamic pitch moment of an aircraft

Abstract

The application belongs to the field of aircraft flight load design, and particularly relates to a correction method for aircraft aerodynamic pitching moment, and relates to a correction method for the pitching moment in aircraft aerodynamic load, which is mainly used for correcting the unbalanced pitching moment between pressure distribution characteristics and component aerodynamic characteristics in aircraft aerodynamic load calculation and analysis, reconstructing wing-body grid point pressure distribution characteristics by adjusting the local angle of attack of each grid of the wing-body component, eliminating the unbalanced aerodynamic pitching moment of the aircraft caused by the pressure distribution characteristics, achieving the purpose of load balance, and well meeting the design requirements in engineering.

Description

Technical Field

[0001] The present application belongs to the field of aircraft flight load design, and specifically relates to a method for correcting the aerodynamic pitch moment of an aircraft, and relates to a method for correcting the pitch moment in the aerodynamic load of an aircraft, which is mainly used to correct the unbalanced pitch moment between the pressure distribution characteristics and the aerodynamic characteristics of the components in the calculation and analysis of the aerodynamic load of the aircraft. 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 capabilities allow for extensive load studies that consider all weight information for the entire aircraft and specific aircraft components (weight, center of gravity, moments of inertia, weight distribution), aerodynamic information for different Mach numbers (aerodynamic pressures, aerodynamic coefficients / derivatives), and static aeroelastic inputs (elastic coefficients, increments) to correct for the rigid aerodynamic forces (aerodynamic pressures, aerodynamic coefficients / derivatives) to obtain a balanced load case for the entire aircraft, meaning that the sum of the forces and moments on the aircraft components is zero for each load case.

[0004] The simplified load model is a combination of numerous input and output data files and numerous computer programs. Many traditional aerodynamic load correction methods often modify the imbalance relationship using certain coefficients, which can cause distorted component loads. Therefore, it is necessary to establish a method that iterates component pressure distribution characteristics to correct the unbalanced loads generated in flight load calculations, especially the unbalanced aerodynamic pitching moment. This application is based on this. Summary of the Invention

[0005] The purpose of this application is to provide a method for correcting the aerodynamic pitch moment of an aircraft, which eliminates the unbalanced aerodynamic pitch moment by iteratively changing parameters, thereby achieving the purpose of load balance and meeting the design requirements in engineering.

[0006] The technical solution of this application is:

[0007] A method for correcting an aircraft aerodynamic pitching moment, comprising:

[0008] Step 1: Calculate the aircraft's vacuum speed V and its aircraft speed pressure Q;

[0009] Step 2: Calculate the velocity components u, v, and w in the three directions of the aircraft axis system, and calculate the elevator deflection δ e ;

[0010] Step 3: 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 of the aircraft's axis system.

[0011] Step 4: Calculate the aircraft target pitch moment My based on the aircraft speed pressure Q, aircraft angle of attack α and sideslip angle β and aerodynamic characteristics data BW ;

[0012] Step 5: Determine the unbalanced moment tolerance R My ;

[0013] Step 6: Initialize the aircraft aerodynamic model, extract the m×n pressure coefficient points on the wing-body surface in span and chord directions, and obtain the wing-body pressure coefficient point position coordinate matrix BW Xaero , BW Yaero , BW Zaero ;

[0014] Step 7: Using the aircraft aerodynamic model, obtain the pressure distribution data Cp corresponding to the longitudinal, lateral and elevator deflection states of the wing at different aircraft flight Mach numbers Ma. aero , establish the e Changing four-dimensional CP (α,β,δe,Ma) database;

[0015] Step 8: Initialize the geometric model parameters of the wing-body, establish the wing-body triangle mesh model, calculate the grid node coordinates with a total number of nodes M, and construct the node coordinate matrix D Mpoint ;

[0016] Step 9: Based on the three-dimensional grid node information and node coordinate matrix D Mpoint , calculate the coordinates of the centroid points of the triangular grid with a total number of grids N, and construct the centroid coordinate matrix D Cmesh ;

[0017] Step 10: Based on the three-dimensional grid node information and node coordinate matrix D Mpoint , calculate the area of ​​the triangular grid with a total grid size of N, and construct the area matrix D Smesh ;

[0018] Step 11: Based on the three-dimensional grid node information and node coordinate matrix D Mpoint , solve the centroid unit vector of the triangular grid with a total number of grids N, and construct the centroid unit vector matrix

[0019] Step 12: Calculate the local angle of attack of the triangle mesh centroid with a total number of N grids using the aircraft angle of attack α and the coordinates of the triangle mesh centroid, and construct the local angle of attack matrix α of the centroid Po i nt ;

[0020] Step 13: Use the aircraft's flight Mach number Ma and elevator deflection δ e , the coordinate matrix BW of the local angle of attack, sideslip angle β and wing-body pressure coefficient point of the triangular grid centroid Xaero , BW Yaero , BW Zaero , for the four-dimensional The database is interpolated to obtain the pressure coefficients of the centroid points of the triangular grid with a total of N, and the centroid pressure coefficient matrix D is constructed. MeshCp ;

[0021] Step 14: Calculate the grid aerodynamic force using the aircraft speed pressure Q, the area of ​​the triangular grid, the unit vector of the grid centroid, and the pressure coefficient at the centroid point;

[0022] Step 15: Integrate the aerodynamic forces of N grid cells to calculate the wing aerodynamic force pitch moment

[0023] Step 16: Take the wing aerodynamic pitch moment My (α,β,δe,Ma) 、Full-aircraft target pitch moment My BW , calculate the unbalanced moment △My;

[0024] Step 17: Determine whether the unbalanced torque △My is less than the unbalanced torque tolerance R My If not, update the local angle of attack matrix α of the centroid point Po i nt , repeat steps 13 to 17 until the unbalanced moment △My is less than the unbalanced moment tolerance R My , and the wing aerodynamic force is obtained pitch moment

[0025] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitching moment correction method, step 1 is specifically as follows:

[0026] Calculate the aircraft's vacuum speed V and equivalent speed V based on the aircraft's flight Mach number Ma and flight altitude H. e , atmospheric density ρ, and then calculate the aircraft speed pressure Q:

[0027] Q=0.5ρV 2 .

[0028] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitch moment correction method, step 2 is specifically as follows:

[0029] Based on 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 elevator deflection δ is obtained by interpolation calculation. e .

[0030] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitching moment correction method, step three is specifically as follows:

[0031] α=tan -1 (wu);

[0032] β=sin -1 (vV).

[0033] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitch moment correction method, in step 5, the unbalanced moment tolerance R is determined. My 1%.

[0034] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitch moment correction method, in step 6:

[0035]

[0036]

[0037]

[0038] in,

[0039] X mn 、Y mn , Z mn is the position coordinate of the m×nth wing-body pressure coefficient point.

[0040] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitch moment correction method, in step seven:

[0041]

[0042] in,

[0043] Cpaero(X mn ,Y mn ,Z mn ) is the pressure distribution at the m×nth wing-body pressure coefficient point.

[0044] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitch moment correction method, in step eight:

[0045]

[0046] in,

[0047] X k 、Y k , Z k is the coordinate of the kth grid node.

[0048] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitch moment correction method, in step nine:

[0049]

[0050] in,

[0051] x i 、y i 、z i is the coordinate of the centroid point of the i-th grid corner.

[0052] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitch moment correction method, in step 10:

[0053]

[0054] in,

[0055] S i is the area of ​​the i-th triangle mesh.

[0056] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitching moment correction method, in step 11:

[0057]

[0058] in:

[0059] is the unit vector of the centroid of the i-th grid.

[0060] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitch moment correction method, in step 12:

[0061] α Point =[α1…α i …α N ];

[0062] in,

[0063] α i is the local angle of attack of the centroid of the i-th triangle mesh.

[0064] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitch moment correction method, in step thirteen:

[0065] ;

[0066] in,

[0067] Cp i is the pressure coefficient of the centroid of the i-th triangle mesh.

[0068] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitching moment correction method, step fourteen is specifically as follows:

[0069]

[0070] in,

[0071] is the aerodynamic force of the i-th triangle grid.

[0072] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitching moment correction method, step fifteen is specifically as follows:

[0073]

[0074]

[0075] in,

[0076] x ref 、y ref 、z ref is the test point coordinate, specifically the center of gravity coordinate of the aircraft can be selected.

[0077] According to at least one embodiment of the present application, in the above-mentioned aircraft aerodynamic pitching moment correction method, step 16 is specifically as follows:

[0078] △My=My Bw -My (α,β,δe,Ma) .

[0079] This application has at least the following beneficial technical effects:

[0080] A method for correcting the aerodynamic pitching moment of an aircraft is provided. By adjusting the local angle of attack of each grid of the wing-fuselage component and reconstructing the pressure distribution characteristics of the wing-fuselage grid points, the unbalanced aerodynamic pitching moment of the aircraft caused by the pressure distribution characteristics is eliminated, and the purpose of load balance is achieved, which can well meet the design requirements in engineering. BRIEF DESCRIPTION OF THE DRAWINGS

[0081] Figure 1 Schematic diagram of a method for correcting an aircraft aerodynamic pitching moment provided in an embodiment of the present application;

[0082] Figure 2Schematic diagram of the iteratively corrected aerodynamic pitch moment and the target aerodynamic pitch moment of the entire aircraft provided in an embodiment of the present application.

[0083] In order to better illustrate this embodiment, some contents of the drawings may be omitted. They are only used for illustrative purposes and should not be construed as limiting the present application. DETAILED DESCRIPTION

[0084] To make the purpose, technical solutions, and advantages of this application more clear, the technical solutions in the embodiments of this application will be clearly and completely described below in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of this application, not all of them. Based on the embodiments in this application, all other embodiments obtained by ordinary technicians in this field without making any creative work are within the scope of protection of this application.

[0085] The features of various aspects of the embodiments of the present application will be described in detail below. In the detailed description below, many specific details are proposed to fully understand the present application. However, it is obvious to those skilled in the art that the present application can also be implemented without these specific details. The following description of the embodiments is merely for a better understanding of the present application by illustrating the examples of the present application. The present application is not limited to any specific settings and methods provided below, but covers any improvements, replacements, etc. of all product structures and methods covered without departing from the spirit of the present application. In the various drawings and the following description, well-known structures and technologies are not shown to avoid unnecessary ambiguity in the present application.

[0086] It should be noted that, in the absence of conflict, the embodiments of the present application and the features in the embodiments can be combined with each other, and the various embodiments can refer to and quote each other. The present application will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0087] A correction method for the aerodynamic pitch moment of an aircraft, such as Figure 1 shown.

[0088] Step 1: Calculate the aircraft's true airspeed V and its aircraft speed pressure Q.

[0089] Calculate the aircraft's vacuum speed V and equivalent speed V based on the aircraft's flight Mach number Ma and flight altitude H. e , atmospheric density ρ, and then calculate the aircraft speed pressure Q:

[0090] Q=0.5ρV 2 …………1

[0091] Step 2: Calculate the velocity components u, v, and w in the three directions of the aircraft axis system, and calculate the elevator deflection δ e .

[0092] Based on 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 elevator deflection δ is obtained by interpolation calculation. e .

[0093] Step 3: 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 of the aircraft's axis system.

[0094] α=tan -1 (wu)……………2

[0095] β=sin -1 (vV)…………3

[0096] Step 4: Calculate the aircraft target pitch moment My based on the aircraft speed pressure Q, aircraft angle of attack α and sideslip angle β and aerodynamic characteristics data BW .

[0097] Step 5: Determine the unbalanced moment tolerance R My , which can be selected according to actual conditions, for example, it can be 1%.

[0098] Step 6: Initialize the aircraft aerodynamic model, extract the m×n pressure coefficient points on the wing-body surface in span and chord directions, and obtain the wing-body pressure coefficient point position coordinate matrix BW Xaero , BW Yaero , BW Zaero .

[0099]

[0100]

[0101]

[0102] in,

[0103] X mn 、Y mn , Z mn is the position coordinate of the m×nth wing-body pressure coefficient point.

[0104] Step 7: Using the aircraft aerodynamic model, obtain the pressure distribution data Cp corresponding to the longitudinal, lateral and elevator deflection states of the wing at different aircraft flight Mach numbers Ma. aero , establish the e Changing four-dimensional CP (α,β,δe,Ma) database.

[0105]

[0106] in,

[0107] Cpaero(X mn ,Y mn ,Z mn ) is the pressure distribution at the m×nth wing-body pressure coefficient point.

[0108] Step 8: Initialize the geometric model parameters of the wing-body, establish the wing-body triangle mesh model, calculate the grid node coordinates with a total number of nodes M, and construct the node coordinate matrix D Mpoint .

[0109]

[0110] in,

[0111] X k 、Y k , Z k is the coordinate of the kth grid node.

[0112] Step 9: Based on the three-dimensional grid node information and node coordinate matrix D Mpoint , calculate the coordinates of the centroid points of the triangular grid with a total number of grids N, and construct the centroid coordinate matrix D Cmesh .

[0113]

[0114] in,

[0115] x i 、y i 、z i is the coordinate of the centroid point of the i-th grid corner.

[0116] Step 10: Based on the three-dimensional grid node information and node coordinate matrix D Mpoint , calculate the area of ​​the triangular grid with a total grid size of N, and construct the area matrix D Smesh .

[0117]

[0118] in,

[0119] S i is the area of ​​the i-th triangle mesh.

[0120] Step 11: Based on the three-dimensional grid node information and node coordinate matrix D Mpoint , solve the centroid unit vector of the triangular grid with a total number of grids N, and construct the centroid unit vector matrix

[0121]

[0122] in:

[0123] is the unit vector of the centroid of the i-th grid.

[0124] Step 12: Calculate the local angle of attack of the triangle mesh centroid with a total number of N grids using the aircraft angle of attack α and the coordinates of the triangle mesh centroid, and construct the local angle of attack matrix α of the centroid Po i nt .

[0125] α Point =[α1…α i …α N ]……………10

[0126] in,

[0127] α i is the local angle of attack of the centroid of the i-th triangle mesh.

[0128] Step 13: Use the aircraft's flight Mach number Ma and elevator deflection δ e , the coordinate matrix BW of the local angle of attack, sideslip angle β and wing-body pressure coefficient point of the triangular grid centroid Xaero , BW Yaero , BW Zaero , for four-dimensional CP (α,β,δe,Ma) The database is interpolated to obtain the pressure coefficients of the centroid points of the triangular grid with a total of N, and the centroid pressure coefficient matrix D is constructed. MeshCp .

[0129]

[0130] in,

[0131] Cp i is the pressure coefficient of the centroid of the i-th triangle mesh.

[0132] Step 14: Calculate the grid aerodynamic force using the aircraft speed pressure Q, the area of ​​the triangular grid, the unit vector of the grid centroid, and the pressure coefficient at the centroid point.

[0133]

[0134] in,

[0135] is the aerodynamic force of the i-th triangle grid.

[0136] Step 15: Integrate the aerodynamic forces of N grid cells to calculate the wing aerodynamic force pitch moment

[0137]

[0138]

[0139] in,

[0140] x re f、y re f、z re f is the coordinate of the test point, and the specific coordinates can be the center of gravity coordinates of the aircraft.

[0141] Step 16: Take the wing aerodynamic pitch moment My (α,β,δe,Ma) 、Full-aircraft target pitch moment My BW , calculate the unbalanced moment △My.

[0142] △My=My Bw -My (α,β,δe,Ma) …………15

[0143] Step 17: Determine whether the unbalanced torque △My is less than the unbalanced torque tolerance R My If not, update the local angle of attack matrix αP of the centroid point o i nt , repeat steps 13 to 17 until the unbalanced moment △My is less than the unbalanced moment tolerance R My , and the wing aerodynamic force is obtained pitch moment

[0144] The method for correcting the aerodynamic pitching moment of an aircraft disclosed in the above embodiment initializes the aerodynamic model through parameter initialization and six-degree-of-freedom dynamic response simulation calculation of the aircraft, establishes a wing aerodynamic force CP database, and initializes the geometric model, establishes a wing geometric grid model, calculates the aerodynamic load and pitching moment of the wing-body assembly, solves the unbalanced moment, and then compares and analyzes the unbalanced moment with the tolerance. The unbalanced aerodynamic pitching moment is eliminated by iterating the local angle of attack of the centroid position of the wing-body assembly grid. The iterative correction of the aerodynamic pitching moment is compared with the target aerodynamic pitching moment of the entire aircraft. Figure 2 shown.

[0145] The method for correcting the aerodynamic pitching moment of an aircraft disclosed in the above embodiment adjusts the local angle of attack of each grid of the wing-fuselage component and reconstructs the pressure distribution characteristics of the wing-fuselage grid points, thereby eliminating the unbalanced aerodynamic pitching moment of the aircraft caused by the pressure distribution characteristics, achieving the purpose of load balance, and can well meet the design requirements in the engineering.

[0146] So far, the technical solution of the present application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of the present application is obviously not limited to these specific embodiments. Without departing from the principles of the present application, those skilled in the art can make equivalent changes or replacements to the relevant technical features, and the technical solutions after these changes or replacements will fall within the scope of protection of the present application.

Claims

1. A method for correcting an aircraft's aerodynamic pitching moment, characterized in that: include: Step 1: Calculate the aircraft's vacuum speed V and its aircraft speed pressure Q; Step 2: Calculate the velocity components u, v, and w in the three directions of the aircraft axis system, and calculate the elevator deflection δ e ; Step 3: 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 of the aircraft's axis system. Step 4: Calculate the aircraft target pitch moment My based on the aircraft speed pressure Q, aircraft angle of attack α and sideslip angle β and aerodynamic characteristics data BW ; Step 5: Determine the unbalanced moment tolerance R My ; Step 6: Initialize the aircraft aerodynamic model, extract the m×n pressure coefficient points on the wing-body surface in span and chord directions, and obtain the wing-body pressure coefficient point position coordinate matrix BW Xaero , BW Yaero , BW Zaero ; Step 7: Using the aircraft aerodynamic model, obtain the pressure distribution data Cp corresponding to the longitudinal, lateral and elevator deflection states of the wing at different aircraft flight Mach numbers Ma. aero , establish the e The Four Dimensions of Change database; Step 8: Initialize the geometric model parameters of the wing-body, establish the wing-body triangle mesh model, calculate the grid node coordinates with a total number of nodes M, and construct the node coordinate matrix D Mpoint ; Step 9: Based on the three-dimensional grid node information and node coordinate matrix D Mpoint , calculate the coordinates of the centroid points of the triangular grid with a total number of grids N, and construct the centroid coordinate matrix D Cmesh ; Step 10: Based on the three-dimensional grid node information and node coordinate matrix D Mpoint , calculate the area of ​​the triangular grid with a total grid size of N, and construct the area matrix D Smesh ; Step 11: Based on the three-dimensional grid node information and node coordinate matrix D Mpoint , solve the centroid unit vector of the triangular grid with a total number of grids N, and construct the centroid unit vector matrix Step 12: Calculate the local angle of attack of the triangle mesh centroid with a total number of N grids using the aircraft angle of attack α and the coordinates of the triangle mesh centroid, and construct the local angle of attack matrix α of the centroid Point ; Step 13: Use the aircraft's flight Mach number Ma and elevator deflection δ e , the coordinate matrix BW of the local angle of attack, sideslip angle β and wing-body pressure coefficient point of the triangular grid centroid Xaero , BW Yaero , BW Zaero , for the four-dimensional The database is interpolated to obtain the pressure coefficients of the centroid points of the triangular grid with a total of N, and the centroid pressure coefficient matrix D is constructed. MeshCp ; Step 14: Calculate the grid aerodynamic force using the aircraft speed pressure Q, the area of ​​the triangular grid, the unit vector of the grid centroid, and the pressure coefficient at the centroid point; Step 15: Integrate the aerodynamic forces of N grid cells to calculate the wing aerodynamic force pitch moment Step 16: Using the wing aerodynamic pitch moment The target pitch moment of the whole aircraft My BW , calculate the unbalanced moment △My; Step 17: Determine whether the unbalanced torque △My is less than the unbalanced torque tolerance R My If not, update the local angle of attack matrix α of the centroid point Point , repeat steps 13 to 17 until the unbalanced moment △My is less than the unbalanced moment tolerance R My , and the wing aerodynamic force is obtained pitch moment 2. The method for correcting the aircraft aerodynamic pitching moment according to claim 1, characterized in that: Step 1 is as follows: Calculate the aircraft's vacuum speed V and equivalent speed V based on the aircraft's flight Mach number Ma and flight altitude H. e , atmospheric density ρ, and then calculate the aircraft speed pressure Q: Q=0.5ρV 2 。 3. The method for correcting the aircraft aerodynamic pitching moment according to claim 2, characterized in that: Step 2 is as follows: Based on 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 elevator deflection δ is obtained by interpolation calculation. e .

4. The method for correcting the aircraft aerodynamic pitching moment according to claim 3, characterized in that: Step three is as follows: α=tan -1 (w / u); β=sin -1 (vV).

5. The method for correcting the aircraft aerodynamic pitching moment according to claim 4, characterized in that: In step 5, determine the unbalanced moment tolerance R My 1%.

6. The method for correcting the aircraft aerodynamic pitching moment according to claim 5, characterized in that: In step six: in, X mn 、Y mn , Z mn is the position coordinate of the m×nth wing-body pressure coefficient point.

7. The method for correcting an aircraft aerodynamic pitching moment according to claim 6, wherein in step seven: in, Cpaero(X mn ,Y mn ,Z mn ) is the pressure distribution at the m×nth wing-body pressure coefficient point.

8. The method for correcting the aircraft aerodynamic pitching moment according to claim 7, characterized in that: In step eight: in, X k 、Y k , Z k is the coordinate of the kth grid node.

9. The method for correcting the aircraft aerodynamic pitching moment according to claim 8, characterized in that: In step nine: in, x i 、y i 、z i is the coordinate of the centroid point of the i-th grid corner.

10. The method for correcting the aircraft aerodynamic pitching moment according to claim 9, characterized in that: In step 10: in, S i is the area of ​​the i-th triangle mesh.

11. The method for correcting the aircraft aerodynamic pitching moment according to claim 10, characterized in that: In step 11: in: R Vmeshi is the unit vector of the centroid of the i-th grid.

12. The method for correcting the aircraft aerodynamic pitching moment according to claim 11, characterized in that: In step 12: α Point ＝[α1…α i …α N ]； in, α i is the local angle of attack of the centroid of the i-th triangle mesh.

13. The method for correcting the aircraft aerodynamic pitching moment according to claim 12, characterized in that: In step 13: in, Cp i is the pressure coefficient of the centroid of the i-th triangle mesh.

14. The method for correcting the aircraft aerodynamic pitching moment according to claim 13, characterized in that: Step 14 is as follows: in, is the aerodynamic force of the i-th triangle grid.

15. The method for correcting the aircraft aerodynamic pitching moment according to claim 14, characterized in that: Step 15 is as follows: in, x ref 、y ref 、z ref is the test point coordinate, specifically the center of gravity coordinate of the aircraft can be selected.

16. The method for correcting the aircraft aerodynamic pitching moment according to claim 15, characterized in that: Step 16 is as follows:

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