Aircraft pneumatic pitching moment correction method

Abstract

The invention belongs to the field of aircraft flight load design, particularly relates to a correction method for aircraft aerodynamic pitching moment, relates to a correction method for pitching moment in aircraft aerodynamic load, and is mainly used for correcting unbalanced pitching moment between pressure distribution characteristics and component aerodynamic characteristics in aircraft aerodynamic load calculation and analysis. By adjusting the local attack angle of each grid of the wing body part and reconstructing the pressure distribution characteristics of the wing body grid points, the unbalanced pneumatic pitching moment caused by the pressure distribution characteristics of the aircraft is eliminated, the purpose of load balance is achieved, and the design requirements in engineering can be well met.

Description

Technical Field

[0001] The present application belongs to the field of aircraft flight load design, and specifically relates to a method for correcting an aircraft's aerodynamic pitch moment, and relates to a method for correcting the pitch moment in an aircraft's aerodynamic load, and is mainly used for correcting the unbalanced pitch moment between the pressure distribution characteristics and the aerodynamic characteristics of components in the calculation and analysis of the aircraft's aerodynamic load. 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 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 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 of the aircraft components is zero for every load case.

[0004] 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, especially the unbalanced aerodynamic pitch moment. In view of this, the present application is proposed. 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 pitch moment, comprising:

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

[0009] Step 2: Calculate the velocity components u, v, 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 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, 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 direction, 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 direction, lateral heading and elevator deflection state of the wing body 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 mesh 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 centroid coordinates of the triangular mesh with a total number of N meshes, 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 mesh with a total mesh 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 mesh with a total number of meshes N, and construct the centroid unit vector matrix

[0019] Step 12: Using the aircraft angle of attack α and the coordinates of the centroid of the triangular grid, calculate the local angle of attack of the centroid of the triangular grid with a total number of N grids, and construct the centroid local angle of attack matrix α Po i nt ;

[0020] Step 13: Use the aircraft 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 mesh with a total number of N, and the centroid pressure coefficient matrix D is constructed. MeshCp ;

[0021] Step 14: Calculate the grid aerodynamic force based on the aircraft speed pressure Q, the area of ​​the triangular grid, the unit vector of the grid centroid, and the pressure coefficient of 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) 、The whole 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 step 13 to step 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 pitch moment correction method, step 1 is specifically:

[0026] Calculate the aircraft vacuum speed V and aircraft equivalent speed V based on the aircraft 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 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 body 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 pitch moment correction method, step three is specifically:

[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 is 1%.

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

[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 shape 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 angle.

[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 ith triangle mesh.

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

[0057]

[0058] in:

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

[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 ith 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 pitch moment correction method, step fourteen is specifically:

[0069]

[0070] in,

[0071] is the aerodynamic force of the ith triangular grid.

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

[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 pitch moment correction method, step sixteen is specifically:

[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, the pressure distribution characteristics of the wing-fuselage grid points are reconstructed, and the unbalanced aerodynamic pitching moment of the aircraft caused by the pressure distribution characteristics of the aircraft is eliminated, so as to achieve the purpose of load balance, which can well meet the design requirements in the project. BRIEF DESCRIPTION OF THE DRAWINGS

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

[0082] Figure 2It is a schematic 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 the present embodiment, some contents of the drawings may be omitted, which is only used for illustrative purposes and should not be construed as limiting the present application. DETAILED DESCRIPTION

[0084] In order to make the purpose, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present 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 the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without making creative work belong to 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, so as 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 the need for these specific details. The following description of the embodiments is only 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 all product structures, any improvements, replacements, etc. of the methods covered without departing from the spirit of the present application. In each of the accompanying drawings and the following description, known structures and technologies are not shown to avoid causing unnecessary ambiguity to 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 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 vacuum speed V and its aircraft speed pressure Q.

[0089] Calculate the aircraft vacuum speed V and aircraft equivalent speed V based on the aircraft 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, w in the three directions of the aircraft axis system, and calculate the elevator deflection δ e .

[0092] Based on the aircraft 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 body 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, 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 direction, 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 direction, lateral heading and elevator deflection state of the wing body 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 shape 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 mesh 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 centroid coordinates of the triangular mesh with a total number of N meshes, 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 angle.

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

[0117]

[0118] in,

[0119] S i is the area of ​​the ith 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 mesh with a total number of meshes N, and construct the centroid unit vector matrix

[0121]

[0122] in:

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

[0124] Step 12: Using the aircraft angle of attack α and the coordinates of the centroid of the triangular grid, calculate the local angle of attack of the centroid of the triangular grid with a total number of N grids, and construct the centroid local angle of attack matrix α 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 ith triangle mesh.

[0128] Step 13: Use the aircraft 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 CP (α,β,δe,Ma) The database is interpolated to obtain the pressure coefficients of the centroid points of the triangular mesh with a total number 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 based on the aircraft speed pressure Q, the area of ​​the triangular grid, the unit vector of the grid centroid, and the pressure coefficient of the centroid point.

[0133]

[0134] in,

[0135] is the aerodynamic force of the ith triangular 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 coordinate can be the center of gravity coordinate of the aircraft.

[0141] Step 16: Take the wing aerodynamic pitch moment My (α,β,δe,Ma) 、The whole 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 step 13 to step 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 pitch moment of an aircraft disclosed in the above embodiment initializes the aerodynamic model through parameter initialization and the six-degree-of-freedom dynamic response simulation calculation of the aircraft, establishes the wing aerodynamic force CP database, and initializes the geometric model, establishes the wing geometric grid model, calculates the aerodynamic load and pitch moment of the wing-body assembly, solves the unbalanced moment, and then compares and analyzes the unbalanced moment and the tolerance. The unbalanced aerodynamic pitch moment is eliminated by iterating the local angle of attack of the centroid position of the wing-body assembly grid, and the iterative correction of the aerodynamic pitch moment and the target aerodynamic pitch moment of the whole aircraft are as follows: Figure 2 shown.

[0145] The method for correcting the aerodynamic pitch moment of an aircraft disclosed in the above-mentioned embodiment adjusts the local angle of attack of each grid of the wing-fuselage component, reconstructs the pressure distribution characteristics of the wing-fuselage grid points, eliminates the unbalanced aerodynamic pitch moment of the aircraft caused by the pressure distribution characteristics, achieves the purpose of load balance, and can well meet the design requirements in 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 substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will fall within the scope of protection of the present application.

Claims

1. A method for correcting an aircraft aerodynamic pitch moment, characterized in that: include: Step 1: Calculate the aircraft vacuum speed V and its aircraft speed pressure Q; Step 2: Calculate the velocity components u, v, 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 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, 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 direction, 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 direction, lateral heading and elevator deflection state of the wing body 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 mesh 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 centroid coordinates of the triangular mesh with a total number of N meshes, 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 mesh with a total mesh 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 mesh with a total number of meshes N, and construct the centroid unit vector matrix Step 12: Using the aircraft angle of attack α and the coordinates of the centroid of the triangular grid, calculate the local angle of attack of the centroid of the triangular grid with a total number of N grids, and construct the centroid local angle of attack matrix α Point ; Step 13: Use the aircraft 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 mesh with a total number of N, and the centroid pressure coefficient matrix D is constructed. MeshCp ; Step 14: Calculate the grid aerodynamic force based on the aircraft speed pressure Q, the area of ​​the triangular grid, the unit vector of the grid centroid, and the pressure coefficient of the centroid point; Step 15: Integrate the aerodynamic forces of N grid cells to calculate the wing aerodynamic force Pitch Moment Step 16: Aerodynamic pitch moment of the wing 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 step 13 to step 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 vacuum speed V and aircraft equivalent speed V based on the aircraft 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 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 body 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 is 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 the aerodynamic pitch moment of an aircraft according to claim 6, characterized in that In step seven: in, Cpaero(X mn ,Y mn ,Z mn ) is the pressure distribution shape 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 angle.

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

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

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 ith triangular 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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