Aircraft attack angle balancing method considering nonlinear aeroelastic effect

Abstract

The invention relates to an aircraft attack angle balancing method considering a non-linear aeroelastic effect, belongs to the technical field of aerodynamics and aeroelastic mechanics, and aims to meet the requirement of attack angle balancing in aircraft load solution. Balancing results are compared on the basis of nonlinear aeroelastic calculation based on high-precision fluid mechanics (CFD) and solid dynamics (CSD), a bisection method is adopted to adjust a balancing attack angle, iteration is repeated until balancing convergence, attack angle balancing considering the nonlinear aeroelastic effect is finally achieved, accurate calculation of the aircraft balancing attack angle is achieved, and the aircraft balancing attack angle is accurately calculated. And the requirement for high-precision solving of the load in modern aircraft calculation is met.

Description

Technical Field

[0001] The present invention relates to the technical fields of aerodynamics and aeroelasticity, and particularly relates to a method for trimming the angle of attack of an aircraft considering non-linear aeroelastic effects. Background Art

[0002] Aeroelasticity mainly studies the coupling problems among aerodynamic forces, elastic forces, and inertial forces. When a structure is subjected to aerodynamic loads, it will deform, and the structural deformation will in turn change the magnitude and distribution of the aerodynamic loads. This is a typical fluid-structure interaction problem. Modern aircraft such as civil airliners and flying wing aircraft increasingly adopt large flexible wing designs. Under the action of flight loads, the large flexible wings will produce large bending and torsional deformations, and there are typical geometric non-linear aeroelastic problems, which must be considered in the aircraft design process.

[0003] The accurate solution of the flight loads of an aircraft is a key link in aircraft design. Achieving an accurate solution of the loads is an important guarantee for the safety and reliability of the aircraft. During the flight of an aircraft, it is necessary to adjust its angle of attack to meet the trimming requirements of different overload states. The load solution for each flight state of the aircraft needs to be based on the trimmed angle of attack. As the flexibility of modern aircraft gradually increases, it will have large structural deformations under aerodynamic forces, which will in turn lead to significant changes in the load distribution of the aircraft. The aircraft loads need to be solved under the trimmed state considering geometric non-linear effects.

[0004] In view of the key role of aeroelasticity in aircraft design, it is necessary to accurately analyze problems such as the deformation and load distribution of civil aircraft during the aeroelastic design stage. The solution of aerodynamic forces is one of the core problems in the aeroelastic calculation process. Summary of the Invention

[0005] In view of the above problems, the present invention provides a method for trimming the angle of attack of an aircraft considering non-linear aeroelastic effects. The present invention is aimed at the need for trimming the angle of attack in the solution of aircraft loads. Based on high-precision computational fluid dynamics (CFD) and computational solid dynamics (CSD), the non-linear aeroelastic calculations are carried out to compare the trimming results. The bisection method is used to adjust the trimmed angle of attack, and the iteration is repeated until the trimming converges. Finally, the trimming of the angle of attack considering non-linear aeroelastic effects is achieved, and the accurate calculation of the trimmed angle of attack of the aircraft is realized, meeting the requirements for high-precision load solution in modern aircraft calculations.

[0006] The present invention provides a method for trimming the angle of attack of an aircraft considering non-linear aeroelastic effects, including:

[0007] Step S1: Use the finite element modeling software MSC.PATRAN to establish a finite element model of the wing structure, and set structural interpolation points along the position of the main beam;

[0008] Use the finite element modeling software ICEM to perform structured mesh division on the entire flow field and establish an aerodynamic grid model of the wing; select the point with the largest structural displacement to observe the displacement convergence situation;

[0009] Step S2: Preset the convergence error threshold e of the non - linear static aeroelastic calculation a and the convergence error threshold e of the angle - of - attack trim T , where the convergence error threshold of the static aeroelastic calculation is the magnitude of the structural deformation, set according to the calculation accuracy requirements of the static aeroelasticity, and the convergence error threshold of the angle - of - attack trim is the magnitude of the force difference, set according to the calculation accuracy requirements of the trimmed angle of attack;

[0010] Preset the upper and lower limits of the flight angle of attack;

[0011] Step S3: Initialize the aerodynamic grid model of the wing and the finite - element model of the wing structure;

[0012] Step S4: Read the flight angle of attack to be measured and its related operating conditions; the related operating conditions of the flight angle of attack include the flight altitude, speed, dynamic pressure, overload coefficient, trimming mass of the aircraft, and the initial calculated angle of attack of the aircraft;

[0013] Step S5: Let t = 1. When t = 1, it represents the initial moment;

[0014] Step S6: According to the structural interpolation points of the finite - element model A t of the wing structure, solve the flight - angle - of - attack interpolation matrix at time t;

[0015] Based on the aerodynamic grid model B t of the wing, obtain the aerodynamic grid at the flight angle of attack at time t;

[0016] Step S7: Based on the flight - angle - of - attack operating conditions at time t and the aerodynamic grid at the flight angle of attack at time t, perform aerodynamic force calculations of computational fluid dynamics (CFD) to obtain the flight - angle - of - attack aerodynamic - force interpolation matrix at time t;

[0017] Load the flight - angle - of - attack aerodynamic force at time t into the flight - angle - of - attack interpolation matrix at time t for force interpolation to obtain the flight - angle - of - attack force interpolation matrix at time t;

[0018] Optionally, the expression of the aerodynamic force at time t is:

[0019]

[0020] where Q is the solution vector, E is the vector of the fluid in the X - direction, J is the source term, F is the vector of the fluid in the Y - direction, and G is the vector of the fluid in the Z - direction.

[0021] Step S8: Based on the flight - angle - of - attack force interpolation matrix at time t, for the finite - element model A tDeform;

[0022] Obtain the wing structure deformation corresponding to the flight angle of attack at time t by using the solid dynamics CSD method;

[0023] Step S9: Use the wing deformation corresponding to the flight angle of attack at time t to determine the convergence of nonlinear static aeroelasticity. If the static aeroelasticity converges, proceed to the next step;

[0024] If it does not converge, update the aerodynamic grid until the static aeroelastic convergence criterion is met, obtain the flight angle of attack with static aeroelastic convergence at time t, and proceed to the next step;

[0025] Optionally, the specific steps of step S9 when it does not converge, update the aerodynamic grid until the static aeroelastic convergence criterion is met, and proceed to the next step include:

[0026] If the static aeroelasticity does not converge, perform displacement interpolation according to the wing structure deformation corresponding to the flight angle of attack at time t, use the radial basis function RBF to deform the wing aerodynamic grid model to obtain the corresponding displacement, and update the aerodynamic grid based on the corresponding displacement at the flight angle of attack at time t;

[0027] Use the updated aerodynamic grid at the flight angle of attack at time t as the aerodynamic grid at the flight angle of attack at time t + 1, let t = t + 1, and return to step S5 until the static aeroelastic convergence criterion is met, and proceed to the next step;

[0028] Optionally, the expression for the convergence determination in step S9 is:

[0029] |d t -d t-1 |≤e a

[0030] where d t is the magnitude of the structural deformation displacement in the static aeroelastic cycle at time t, d t-1 is the magnitude of the structural deformation displacement in the static aeroelastic cycle at time t - 1, and e a is the nonlinear static aeroelastic calculation convergence error threshold.

[0031] Optionally, the expression for the corresponding displacement is:

[0032]

[0033] where s(x) is the displacement of the interpolation target point x, x i is the displacement of the i-th interpolation point; N b is the total number of interpolation points for grid deformation calculation; ω i is the weight coefficient corresponding to the i-th interpolation point; φ(·) is the radial basis function, and ∥·∥ is the Euclidean distance function.

[0034] Step S10: Obtain the aerodynamic force of the flight angle of attack at which the static aeroelasticity converges at time t, and determine whether the flight angle of attack at which the static aeroelasticity converges at time t is trim-converged;

[0035] If the trim-convergence criterion is not satisfied, the upper and lower limits of the flight angle of attack at which the static aeroelasticity converges at time t are adjusted by the bisection method to obtain the updated flight angle of attack at time t;

[0036] Take the updated flight angle of attack at time t as the flight angle of attack at time t + 1, let t = t + 1, and return to step S5; if the trim-convergence criterion is satisfied, obtain the final flight angle of attack.

[0037] Optionally, the expression of the trim-convergence criterion is:

[0038] |F t - tMg| ≤ e T

[0039] where F t is the aerodynamic lift in the t-th trim cycle, g is the acceleration due to gravity, and M is the gravity.

[0040] Optionally, the specific steps to obtain the updated flight angle of attack include:

[0041] If the flight angle of attack at which the static aeroelasticity converges at time t is in an increasing state compared to the flight angle of attack at which the static aeroelasticity converges at time t - 1, then the flight angle of attack at which the static aeroelasticity converges at time t is used as the lower limit of the flight angle of attack at which the static aeroelasticity converges at time t, that is, α d,t = α t , obtain the updated lower limit of the flight angle of attack at which the static aeroelasticity converges at time t, and obtain the updated flight angle of attack at which the static aeroelasticity converges at time t based on the updated lower limit of the flight angle of attack at which the static aeroelasticity converges at time t; if the flight angle of attack at which the static aeroelasticity converges at time t is in a decreasing state compared to the flight angle of attack at which the static aeroelasticity converges at time t - 1, then the flight angle of attack at which the static aeroelasticity converges at time t - 1 is used as the upper limit of the flight angle of attack at which the static aeroelasticity converges at time t, that is, α u,t = α t-1 , obtain the updated upper limit of the flight angle of attack at which the static aeroelasticity converges at time t, and obtain the updated flight angle of attack at which the static aeroelasticity converges at time t based on the updated upper limit of the flight angle of attack at which the static aeroelasticity converges at time t;

[0042] Optionally, the expression of the updated flight angle of attack at which the static aeroelasticity converges at time t is:

[0043]

[0044] where α t-1 is the flight angle of attack at which the static aeroelasticity converges at time t - 1, α a,tThe updated flight angle of attack for static aeroelastic convergence at time t, α u,t The upper limit of the flight angle of attack for static aeroelastic convergence at time t, α d,t The lower limit of the flight angle of attack for static aeroelastic convergence at time t.

[0045] Compared with the prior art, the present invention has at least the following beneficial effects:

[0046] (1) The present invention solves the problem that existing tools cannot meet the requirement of high-precision solution of the trim angle of attack in the calculation of modern aircraft, especially large flexible aircraft;

[0047] (2) In the process of calculating the trim angle of attack, the present invention adopts a non-linear static aeroelastic solution method of CFD / CSD coupling, which has high calculation accuracy, can effectively consider the aerodynamic non-linear factors brought by different working conditions such as transonic speed and the structural non-linear factors brought by the increase of structural flexibility, and is easy to be applied in engineering;

[0048] (3) The present invention adjusts the angle of attack by using the bisection method criterion, which has clear principle and is easy to be realized in engineering. Moreover, the angle of attack adjustment criterion and the trim convergence determination criterion can consider the influence of the overload coefficient under different flight conditions on the trim calculation, and have high engineering practical significance;

[0049] (4) After the trim of the present invention is completed, not only can the trim angle of attack under the current working condition be obtained, but also the current structural deformation situation and the aerodynamic load distribution situation can be directly derived from the non-linear static aeroelastic calculation structure at the convergence step, and the calculation accuracy is high, which is convenient for subsequent carrying out various aircraft load analysis calculations. The above advantages contribute to the application of this method in aircraft load calculation, realizing high-precision angle of attack trim and load calculation. Brief Description of the Drawings

[0050] The drawings are only for the purpose of showing specific embodiments and are not considered as a limitation of the present invention.

[0051] Figure 1 It is a schematic diagram of the angle of attack trim calculation process in the embodiment of the present invention;

[0052] Figure 2 It is a schematic diagram of the appearance of a large flexible wing in the embodiment of the present invention;

[0053] Figure 3 It is a structural schematic diagram of a large flexible wing in the embodiment of the present invention;

[0054] Figure 4 It is a schematic diagram of the deformation of a large flexible wing after convergence in the embodiment of the present invention;

[0055] Figure 5 It is a schematic diagram of the iterative convergence of the trim angle of attack of a large flexible wing in the embodiment of the present invention. Detailed implementation manners

[0056] In order to more clearly understand the above objects, features and advantages of the present invention, the present invention will be further described in detail below with reference to the accompanying drawings and specific implementation manners. It should be noted that, without conflict, the embodiments of the present invention and the features in the embodiments can be combined with each other. In addition, the present invention can also be implemented in other ways different from those described herein. Therefore, the protection scope of the present invention is not limited by the specific embodiments disclosed below.

[0057] A specific embodiment of the present invention is as follows Figures 1-5 , the present invention reads the error threshold and the initial trim angle of attack, specifies the upper and lower limits of the trim angle calculation, and sets the trim calculation conditions and the aircraft overload factor.

[0058] After the trim calculation starts, it first enters the non-linear static aeroelastic calculation loop. The aerodynamic force is calculated by using a high-precision CFD method, interpolated onto the structure and the geometric non-linear structural deformation calculation is performed. Then the displacement interpolation calculation is carried out, and the CFD aerodynamic grid is updated by using the RBF method, and the iteration is repeated until the static deformation converges.

[0059] After that, the trim convergence determination is carried out. If it does not converge, the calculated angle of attack is adjusted and the static aeroelastic calculation is performed again until convergence, and the final trim angle of attack calculation result can be obtained.

[0060] This method can consider the non-linear effects of aerodynamic forces under different calculation conditions during the aerodynamic force solution process, and consider the follower force effects brought by large deformations in the structural calculation. The calculation accuracy is relatively high, and it can meet the angle of attack trim calculation and load calculation requirements of various modern aircraft, especially large flexible aircraft.

[0061] In order to illustrate the effectiveness of the method proposed by the present invention, the above technical solutions of the present invention will be described in detail below through a specific embodiment. The specific implementation steps are as follows:

[0062] The present invention provides a method for trimming the angle of attack of an aircraft considering non-linear aeroelastic effects, including:

[0063] Step S1: Use the finite element modeling software MSC.PATRAN to establish a finite element model of the wing structure, and set structural interpolation points along the main beam position;

[0064] Use the finite element modeling software ICEM to perform structured grid division on the entire flow field to establish a wing aerodynamic grid model; select the point with the largest structural displacement to observe the displacement convergence situation;

[0065] Step S2: Preset the non-linear static aeroelastic calculation convergence error threshold e a and the angle of attack trim convergence error threshold eT Among them, the convergence error threshold for the aeroelastic calculation is the magnitude of the structural deformation, which is set according to the calculation accuracy requirement for aeroelasticity. The convergence error threshold for the trim of the angle of attack is the magnitude of the force difference, which is set according to the calculation accuracy requirement for the trimmed angle of attack;

[0066] Set the upper and lower limits of the preset flight angle of attack;

[0067] Step S3: Initialize the aerodynamic grid model of the wing and the finite element model of the wing structure;

[0068] Step S4: Read the angle of attack to be measured and its related working conditions; the related working conditions of the angle of attack include the flight altitude of the aircraft, speed, dynamic pressure, overload factor, trim mass, and the initial calculated angle of attack of the aircraft;

[0069] Step S5: Let t = 1. When t = 1, it represents the initial moment;

[0070] Step S6: According to the structural interpolation points of the finite element model A of the wing structure t Solve the interpolation matrix of the angle of attack at time t;

[0071] Based on the aerodynamic grid model B of the wing t Obtain the aerodynamic grid at the angle of attack at time t;

[0072] Step S7: Based on the angle of attack working conditions at time t and the aerodynamic grid at the angle of attack at time t, perform the aerodynamic force calculation of computational fluid dynamics (CFD) to obtain the interpolation matrix of the aerodynamic force at the angle of attack at time t;

[0073] Load the aerodynamic force at the angle of attack at time t into the interpolation matrix of the angle of attack at time t for force interpolation to obtain the interpolation matrix of the force at the angle of attack at time t;

[0074] Optionally, the expression of the aerodynamic force at time t is:

[0075]

[0076] Among them, Q is the solution vector, E is the vector of the fluid in the X direction, J is the source term, F is the vector of the fluid in the Y direction, and G is the vector of the fluid in the Z direction.

[0077] Step S8: Based on the interpolation matrix of the force at the angle of attack at time t, deform the finite element model A of the wing structure t ;

[0078] Use the computational solid dynamics (CSD) method to obtain the wing structure deformation corresponding to the angle of attack at time t;

[0079] Step S9: Use the wing deformation corresponding to the angle of attack at time t to determine the convergence of nonlinear static aeroelasticity. If the static aeroelasticity converges, proceed to the next step;

[0080] If not converged, update the aerodynamic grid until the aeroelastic convergence criterion is met, obtain the angle of attack at which the aeroelasticity converges at time t, and proceed to the next step;

[0081] Optionally, the specific steps of "if not converged, update the aerodynamic grid until the aeroelastic convergence criterion is met and proceed to the next step" described in step S9 include:

[0082] If the aeroelasticity is not converged, perform displacement interpolation based on the wing structural deformation corresponding to the angle of attack at time t, deform the wing aerodynamic grid model using the radial basis function RBF to obtain the corresponding displacement, and update the aerodynamic grid at the angle of attack at time t based on the corresponding displacement;

[0083] Take the aerodynamic grid updated at the angle of attack at time t as the aerodynamic grid at the angle of attack at time t+1, let t = t+1, return to step S5, and continue until the aeroelastic convergence criterion is met, then proceed to the next step;

[0084] Optionally, the expression for the convergence determination described in step S9 is:

[0085] |d t -d t-1 |≤e a

[0086] where d t is the magnitude of the structural deformation displacement in the aeroelastic cycle at time t, d t-1 is the magnitude of the structural deformation displacement in the aeroelastic cycle at time t-1, and e a is the nonlinear aeroelastic calculation convergence error threshold.

[0087] Optionally, the expression for the corresponding displacement is:

[0088]

[0089] where s(x) is the displacement of the interpolation target point x, x i is the displacement of the i-th interpolation point; Nb is the total number of interpolation points for grid deformation calculation; ω i is the weight coefficient corresponding to the i-th interpolation point; φ(·) is the radial basis function, and ∥·∥ is the Euclidean distance function.

[0090] Step S10: Obtain the aerodynamic force of the angle of attack at which the aeroelasticity converges at time t, and determine whether the angle of attack at which the aeroelasticity converges at time t is trim-converged;

[0091] If the trim convergence criterion is not met, adjust the upper and lower limits of the angle of attack at which the aeroelasticity converges at time t using the bisection method to obtain the updated angle of attack at time t;

[0092] Update the flight angle of attack at the moment t as the flight angle of attack at the moment t+1, let t = t+1, and return to step S5; if the trim convergence criterion is satisfied, the final flight angle of attack is obtained.

[0093] Optionally, the expression of the trim convergence criterion is:

[0094] |F t -tMg|≤e T

[0095] where F t is the aerodynamic lift in the t-th trim cycle, g is the acceleration due to gravity, and M is the gravity.

[0096] Optionally, the specific steps to obtain the updated flight angle of attack include:

[0097] If the flight angle of attack converged by static aeroelasticity at the moment t is in an increasing state compared with the flight angle of attack converged by static aeroelasticity at the moment t-1, then the flight angle of attack converged by static aeroelasticity at the moment t is used as the lower limit of the flight angle of attack converged by static aeroelasticity at the moment t, that is, α d,t =α t , and the updated lower limit of the flight angle of attack converged by static aeroelasticity at the moment t is obtained. Based on the updated lower limit of the flight angle of attack converged by static aeroelasticity at the moment t, the updated flight angle of attack converged by static aeroelasticity at the moment t is obtained; if the flight angle of attack converged by static aeroelasticity at the moment t is in a decreasing state compared with the flight angle of attack converged by static aeroelasticity at the moment t-1, then the flight angle of attack converged by static aeroelasticity at the moment t-1 is used as the upper limit of the flight angle of attack converged by static aeroelasticity at the moment t, that is, α u,t =α t-1 , and the updated upper limit of the flight angle of attack converged by static aeroelasticity at the moment t is obtained. Based on the updated upper limit of the flight angle of attack converged by static aeroelasticity at the moment t, the updated flight angle of attack converged by static aeroelasticity at the moment t is obtained;

[0098] Optionally, the expression of the updated flight angle of attack converged by static aeroelasticity at the moment t is:

[0099]

[0100] where α t-1 is the flight angle of attack converged by static aeroelasticity at the moment t-1, α a,t is the updated flight angle of attack converged by static aeroelasticity at the moment t, α u,t is the upper limit of the flight angle of attack converged by static aeroelasticity at the moment t, and α d,t is the lower limit of the flight angle of attack converged by static aeroelasticity at the moment t.

[0101] Example 1

[0102] A large flexible wing is adopted, and the wing root is completely fixed. The characteristics of the example wing are as follows: The large flexible wing adopts the design of a common civil aircraft wing, with a semi-span of 2000 mm, a chord length at the wing root of 800 mm, and a chord length at the wing tip of 130 mm. There are winglets at the wing tips. The schematic diagram of the large flexible wing model is as shown in Figure 2 shown.

[0103] According to an embodiment of the present invention, an aircraft angle of attack trimming method considering non-linear aeroelastic effects includes:

[0104] The first step: Calculation conditions and calculation file reading

[0105] Set the calculation conditions as an incoming flow velocity of 0.6 Mach, a dynamic pressure of 12000 Pa, an overload factor set to 1 (steady straight flight), a trimming mass set to 400 kg, and an initial angle of attack set to 10°.

[0106] Establish a finite element model of the large flexible wing.

[0107] Use the finite element modeling software ICEM to perform structured grid division on the entire flow field and establish an aerodynamic grid model of the wing.

[0108] Use the finite element modeling software MSC.PATRAN to establish a structural finite element model and set structural interpolation points along the main beam position. The structural model is as shown in Figure 3 shown.

[0109] The second step: Specify the calculation error threshold and the upper and lower limits of the bisection method calculation

[0110] Before the calculation starts, respectively specify the non-linear static aeroelastic calculation convergence error threshold e a and the angle of attack trimming convergence error threshold e T , where the static aeroelastic calculation convergence error threshold is set to 1 mm, and the angle of attack trimming convergence error threshold is set to 20 N; in addition, according to the calculation requirements and engineering experience, specify the upper and lower limits α u and α d of the bisection method angle of attack calculation to be 15° and 0° respectively.

[0111] The third step: CFD / CSD non-linear static aeroelastic solution

[0112] After the trimming calculation starts, first enter the non-linear static aeroelastic iteration loop. Read the input aerodynamic model and structural model and initialize them. Solve the interpolation matrix according to the set interpolation points for subsequent calculations. After initialization, call CFD, CSD, and grid deformation calculations in sequence to perform non-linear aeroelastic solution on the wing until the deformation converges and enters the trimming determination.

[0113] The fourth step: Trimming convergence determination and update of the angle of attack

[0114] After the convergence of nonlinear aeroelasticity, it enters the angle-of-attack trim convergence determination stage. Convergence determination is carried out and the angle of attack is adjusted by the bisection method, and then it re-enters the nonlinear aeroelasticity solution process, and iterates repeatedly until final convergence. After 5 rounds of iteration, the trim angle of attack converges, and the final trim angle of attack of the large flexible wing is 6.875°. The comparison of wing deformations after final convergence is as Figure 4 , and the convergence curve of the trim angle of attack of the large flexible wing is as Figure 5 shown.

[0115] As mentioned above, the above is only a preferred specific embodiment of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily thought of by those skilled in the art within the technical scope disclosed by the present invention should be covered within the protection scope of the present invention.

Claims

1. A method for balancing the angle of attack of an aircraft considering nonlinear aeroelastic effects, characterized in that: include: Step S1, establishing a finite element model of the wing structure and setting structural interpolation points; Establish the wing aerodynamic mesh model; Step S2, reading the flight attack angle to be measured and its related working conditions; Step S3, let t = 1, when t = 1, it indicates the initial time; Step S4: Based on the finite element model A of the wing structure t The structural interpolation points of the flight angle of attack at time t are solved; based on the wing aerodynamic grid model B t Get the aerodynamic grid at the flight angle of attack time t; Step S5, performing aerodynamic force calculation of fluid mechanics CFD based on the flight angle of attack working condition at time t and the aerodynamic grid at the flight angle of attack at time t, and obtaining the aerodynamic force interpolation matrix at the flight angle of attack at time t; Step S6: Based on the flight angle of attack force interpolation matrix at the time t, the finite element model A of the wing structure is calculated. t To transform; The solid dynamics CSD method is used to obtain the wing structure deformation corresponding to the flight angle of attack at time t; Step S7, using the wing deformation corresponding to the flight angle of attack at the time t to determine the convergence of nonlinear static aeroelasticity, if the static aeroelasticity converges, proceed to the next step; If it has not converged, the aerodynamic grid at the flight angle of attack time t is updated until the static aeroelastic convergence standard is met, and the flight angle of attack at the time t static aeroelastic convergence is obtained, and the next step is performed; Step S8, obtaining the aerodynamic force of the flight angle of attack when the static aeroelasticity converges at time t and determining whether the flight angle of attack when the static aeroelasticity converges at time t is trimmed and converged; If the trim convergence criterion is met, the flight angle of attack at which the static aeroelasticity converges at time t is used as the trimmed flight angle of attack; if the trim convergence criterion is not met, the upper and lower limits of the flight angle of attack at which the static aeroelasticity converges at time t are adjusted to obtain the updated flight angle of attack at time t; The updated flight angle of attack at time t is taken as the flight angle of attack at time t+1, and t=t+1 is set, and the process returns to step S3.

2. The method for balancing the angle of attack of an aircraft considering nonlinear aeroelastic effects according to claim 1, characterized in that: The flight angle of attack related working conditions include the aircraft's flight altitude, speed, dynamic pressure, overload coefficient, and trim mass.

3. The method for balancing the angle of attack of an aircraft considering the nonlinear aeroelastic effect according to claim 1, characterized in that: If the step S7 does not converge, the specific steps of updating the aerodynamic grid at the flight angle of attack time t until the static aerodynamic convergence standard is met include: If it does not converge, the displacement interpolation is performed according to the wing structure deformation corresponding to the flight angle of attack at time t, and the radial basis function RBF is used to calculate the wing aerodynamic mesh model B t Perform deformation to obtain a corresponding displacement, obtain a flight angle of attack at time t based on the corresponding displacement, and update the aerodynamic grid; The updated aerodynamic grid at the flight angle of attack at time t is used as the aerodynamic grid at the flight angle of attack at time t+1, and t=t+1 is set, and the process returns to step S3 until the static aerodynamic convergence standard is met.

4. The method for balancing the angle of attack of an aircraft considering nonlinear aeroelastic effect according to claim 1, characterized in that: The specific steps of obtaining the updated flight angle of attack at time t in step S8 include: If the flight angle of attack of static aeroelastic convergence at time t is increasing compared with the flight angle of attack of static aeroelastic convergence at time t-1, then the flight angle of attack of static aeroelastic convergence at time t is used as the lower limit of the flight angle of attack of static aeroelastic convergence at time t, and the updated flight angle of attack at time t is obtained; if the flight angle of attack of static aeroelastic convergence at time t is decreasing compared with the flight angle of attack of static aeroelastic convergence at time t-1, then the flight angle of attack of static aeroelastic convergence at time t-1 is used as the upper limit of the flight angle of attack of static aeroelastic convergence at time t, and the updated flight angle of attack at time t is obtained.

5. The method for balancing the angle of attack of an aircraft considering nonlinear aeroelastic effect according to claim 1, characterized in that: The expression of the balancing convergence criterion in step S8 is: |d t -d t-1 |≤e a Among them, d t is the structural deformation displacement of the static aeroelastic cycle at time t, d t-1 is the structural deformation displacement of the static aeroelastic cycle at time t-1, e a Compute the convergence error threshold for nonlinear hydrostatic elasticity.

6. The method for balancing the angle of attack of an aircraft considering nonlinear aeroelastic effect according to claim 3, characterized in that: The corresponding displacement expression is: Among them, s(x) is the displacement of the interpolation target point x, x i is the displacement of the i-th interpolation point; N b The total number of interpolation points calculated for the mesh deformation; ω i is the weight coefficient corresponding to the i-th interpolation point; φ(·) is the radial basis function, and ∥·∥ is the Euclidean distance function.

7. The method for balancing the angle of attack of an aircraft considering nonlinear aeroelastic effects according to claim 4, characterized in that: The expression of the updated flight angle of attack when static aeroelasticity converges at time t is: Among them, α t-1 is the flight attack angle at time t-1 when static aeroelasticity converges, α a,t is the updated flight angle of attack at time t when static aeroelasticity converges, α u,t is the upper limit of the flight angle of attack at which static aeroelasticity converges at time t, α d,t is the lower limit of the flight angle of attack at which static aeroelasticity converges at time t, F t is the aerodynamic lift of the tth trim cycle, g is the acceleration due to gravity, and M is the gravity.

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