An angle-of-attack trim method for aircraft considering nonlinear aeroelastic effects

Abstract

This invention relates to a method for aircraft angle-of-attack trim considering nonlinear aeroelastic effects, belonging to the fields of aerodynamics and aeroelasticity. Addressing the need for angle-of-attack trim in aircraft load calculation, this invention compares the trim results based on high-precision fluid dynamics (CFD) and solid dynamics (CSD) nonlinear aeroelastic calculations, employs a bisection method to adjust the trim angle of attack, iterates repeatedly until the trim converges, and ultimately achieves angle-of-attack trim considering nonlinear aeroelastic effects. This enables accurate calculation of the aircraft trim angle of attack, meeting the high-precision load calculation requirements in modern aircraft calculations.

Description

Technical Field

[0001] This invention relates to the fields of aerodynamics and aeroelasticity, specifically to a method for trimming the angle of attack of an aircraft that considers nonlinear aeroelastic effects. Background Technology

[0002] Aeroelasticity primarily studies the coupling between aerodynamic forces, elastic forces, and inertial forces. Structures deform under aerodynamic loads, and these deformations, in turn, alter the magnitude and distribution of these loads, representing a typical fluid-structure interaction problem. Modern aircraft, such as commercial airliners and flying-wing aircraft, increasingly employ highly flexible wing designs. Under flight loads, these flexible wings undergo significant bending and torsional deformations, exhibiting typical geometrically nonlinear aeroelastic problems that must be considered during aircraft design.

[0003] Accurate calculation of flight loads is a crucial aspect of aircraft design, and achieving precise load calculation is an important guarantee for aircraft safety and reliability. During flight, the aircraft's angle of attack needs to be adjusted to meet the trim requirements under different overload conditions. The load calculation for each flight state must be based on angle-of-attack trim. However, with the increasing flexibility of modern aircraft, they undergo significant structural deformation under aerodynamic forces, leading to significant changes in the aircraft's load distribution. Therefore, aircraft loads need to be calculated under trim conditions that take into account geometric nonlinear effects.

[0004] Given the crucial role of aeroelasticity in aircraft design, a relatively accurate analysis of issues such as deformation and load distribution in civil aircraft is required during the aeroelastic design phase. Solving for aerodynamic forces is one of the core issues in the aeroelastic calculation process. Summary of the Invention

[0005] In view of the above problems, this invention provides a method for aircraft angle-of-attack trim considering nonlinear aeroelastic effects. This invention addresses the need for angle-of-attack trim in aircraft load calculation. Based on nonlinear aeroelastic calculation using high-precision fluid dynamics (CFD) and solid dynamics (CSD), the trim results are compared, and the trim angle of attack is adjusted using a bisection method. The process is iterated repeatedly until the trim converges, ultimately achieving angle-of-attack trim considering nonlinear aeroelastic effects. This enables accurate calculation of the aircraft trim angle of attack, meeting the high-precision load calculation requirements in modern aircraft calculations.

[0006] This invention provides a method for trimming the angle of attack of an aircraft considering nonlinear aeroelastic effects, comprising:

[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 main beam position;

[0008] The entire flow field was meshed using the finite element modeling software ICEM to establish an aerodynamic mesh model of the wing; the point with the maximum structural displacement was selected to observe the displacement convergence.

[0009] Step S2: Preset the convergence error threshold e for nonlinear static aeroelastic calculation. a With angle of attack balancing convergence error threshold e T Among them, the convergence error threshold for the static aerodynamics calculation is the magnitude of the structural deformation, which is set according to the calculation accuracy requirements of the static aerodynamics; the convergence error threshold for the angle of attack trimming is the magnitude of the force difference, which is set according to the calculation accuracy requirements of the trimming angle of attack.

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

[0011] Step S3: Initialize the aerodynamic mesh model 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 aircraft's flight altitude, speed, dynamic pressure, overload coefficient, trim mass and the aircraft's initial calculated angle of attack;

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

[0014] Step S6: Based on the finite element model A of the wing structure t Given the structural interpolation points, solve for the interpolation matrix of the flight angle of attack at time t;

[0015] Based on wing aerodynamic mesh model B t Obtain the aerodynamic grid at the moment of flight angle of attack t;

[0016] Step S7: Perform aerodynamic calculations using CFD based on the flight angle of attack at time t and the aerodynamic grid at the flight angle of attack at time t to obtain the aerodynamic interpolation matrix at the flight angle of attack at time t.

[0017] The aerodynamic force at the flight angle of attack at time t is applied to the flight angle of attack interpolation matrix at time t for force interpolation, thus obtaining the force interpolation matrix at the flight angle of attack at time t.

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

[0019]

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

[0021] Step S8: Based on the flight angle-of-attack force interpolation matrix at time t, perform an analysis on the finite element model A of the wing structure. tTo deform;

[0022] The CSD method of solid dynamics is used to obtain the wing structure deformation at the flight angle of attack at time t.

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

[0024] If convergence is not achieved, the aerodynamic mesh is updated until the static aeroelastic convergence criterion is met, and the flight angle of attack at time t for static aeroelastic convergence is obtained, and the next step is performed.

[0025] Optionally, if convergence is not achieved in step S9, the aerodynamic mesh is updated until the static aeroelastic convergence criterion is met, and the specific steps for proceeding to the next step include:

[0026] If the static aeroelasticity does not converge, displacement interpolation is performed based on the wing structure deformation corresponding to the flight angle of attack at time t. The radial basis function (RBF) is used to deform the wing aerodynamic mesh model to obtain the corresponding displacement. The aerodynamic mesh is then updated based on the corresponding displacement to obtain the flight angle of attack at time t.

[0027] Update the aerodynamic mesh at the flight angle of attack at time t as the aerodynamic mesh at the flight angle of attack at time t+1, let t = t+1, return to step S5, until the static aerodynamic bullet 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 Let d be the magnitude of the structural deformation displacement during the static aeroelastic cycle at time t. t-1 Let e ​​be the magnitude of the structural deformation displacement during the static aeroelastic cycle at time t-1. a The convergence error threshold is calculated for nonlinear aerostatics.

[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 N represents the displacement at the i-th interpolation point; b ω is the total number of interpolation points calculated for mesh deformation. i φ 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 time t with static aeroelastic convergence, and determine whether the flight angle of attack at time t with static aeroelastic convergence is balanced and converged.

[0035] If the balance convergence criterion is not met, the bisection method is used to adjust the upper and lower limits of the flight angle of attack at time t for static aeroelastic convergence, and the updated flight angle of attack at time t is obtained.

[0036] Update the 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 balancing convergence criterion is satisfied, the final flight angle of attack is obtained.

[0037] Optionally, the expression for the balancing convergence criterion is:

[0038] |F t -tMg|≤e T

[0039] Among them, F t Let g be the aerodynamic lift of the t-th trim cycle, g be the gravitational acceleration, and M be gravity.

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

[0041] If the angle of attack at time t with static aeroelastic convergence increases compared to the angle of attack at time t-1 with static aeroelastic convergence, then the angle of attack at time t with static aeroelastic convergence is taken as the lower limit of the angle of attack at time t with static aeroelastic convergence, i.e., α. d,t =α t The updated lower bound of the static aeroelastic convergence flight angle at time t is obtained. Based on the updated lower bound of the static aeroelastic convergence flight angle at time t, the updated flight angle at time t is obtained. If the flight angle at time t is decreasing compared to the static aeroelastic convergence flight angle at time t-1, then the flight angle at time t-1 is taken as the upper bound of the static aeroelastic convergence flight angle at time t, i.e., α. u,t =α t-1 We obtain the updated upper limit of the flight angle of attack for static aeroelastic convergence at time t, and based on the updated upper limit of the flight angle of attack for static aeroelastic convergence at time t, we obtain the updated flight angle of attack for static aeroelastic convergence at time t.

[0042] Optionally, the expression for updating the flight angle of attack at time t with static aeroelastic convergence is:

[0043]

[0044] Where, α t-1 Let α be the angle of attack for static aeroelastic convergence at time t-1. a,tFor the updated flight angle of attack at time t with static aeroelastic convergence, α u,t α is the upper limit of the angle of attack of the flight with static aeroelastic convergence at time t. d,t This is the lower limit of the 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) This invention solves the problem that existing tools cannot meet the high-precision solution requirements for trim angle of attack in the calculation of modern aircraft, especially large flexible aircraft;

[0047] (2) In the process of calculating the angle of attack for balancing, the present invention adopts a nonlinear static aeroelastic solution method coupled with CFD / CSD, which has high calculation accuracy and can effectively consider the aerodynamic nonlinear factors brought about by different working conditions such as transonic speed and the structural nonlinear factors brought about by the increase in structural flexibility. It is also easy to apply in engineering.

[0048] (3) The present invention uses the bisection method criterion to adjust the angle of attack. Its principle is clear and easy to implement in engineering. Moreover, the angle of attack adjustment criterion and the trim convergence judgment criterion can take into account the influence of the overload coefficient on the trim calculation under different flight conditions, which has high engineering practical significance.

[0049] (4) After the trimming is completed, this invention can not only obtain the trimming angle of attack under the current operating conditions, but also directly derive the current structural deformation and aerodynamic load distribution from the nonlinear static aeroelastic calculation structure of the convergence step. Furthermore, the calculation accuracy is high, facilitating subsequent load analysis and calculation for various types of aircraft. These advantages help this method to be applied in aircraft load calculation, achieving high-precision angle-of-attack trimming and load calculation. Attached Figure Description

[0050] The accompanying drawings are for illustrative purposes only and are not intended to limit the scope of the invention.

[0051] Figure 1 This is a schematic diagram of the angle-of-attack balancing calculation process in an embodiment of the present invention;

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

[0053] Figure 3 This is a schematic diagram of the structure of the large flexible wing in an embodiment of the present invention;

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

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

[0056] To better understand the above-described objectives, features, and advantages of the present invention, the invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. It should be noted that, unless otherwise specified, the embodiments of the present invention and the features thereof can be combined with each other. Furthermore, the present invention can be implemented in other ways different from those described herein; therefore, the scope of protection of the present invention is not limited to the specific embodiments disclosed below.

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

[0058] After the balancing calculation begins, a nonlinear static aeroelasticity calculation loop is first entered. The aerodynamic force is calculated using a high-precision CFD method, and the aerodynamic force is interpolated onto the structure and geometric nonlinear structural deformation is calculated. Then, displacement interpolation calculation is performed, and the CFD aerodynamic mesh is updated using the RBF method. The process is iterated repeatedly until the static deformation converges.

[0059] Next, a balance convergence determination is performed. If convergence is not achieved, the calculated angle of attack is adjusted, and the static aeroelasticity calculation is performed again until convergence is achieved, thus obtaining the final balance angle of attack calculation result.

[0060] This method can consider the nonlinear effects of aerodynamic forces under different calculation conditions during the aerodynamic solution process, and consider the dynamic effects brought about by large deformation in the structural calculation. It has high calculation accuracy and can meet the angle-of-attack trim calculation and load calculation requirements of various modern aircraft, especially large flexible aircraft.

[0061] To illustrate the effectiveness of the method proposed in this invention, the following detailed description of the above technical solution is provided through a specific embodiment. The specific implementation steps are as follows:

[0062] This invention provides a method for trimming the angle of attack of an aircraft considering nonlinear aeroelastic effects, comprising:

[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] The entire flow field was meshed using the finite element modeling software ICEM to establish an aerodynamic mesh model of the wing; the point with the maximum structural displacement was selected to observe the displacement convergence.

[0065] Step S2: Preset the convergence error threshold e for nonlinear static aeroelastic calculation. a With angle of attack balancing convergence error threshold eT Among them, the convergence error threshold for the static aerodynamics calculation is the magnitude of the structural deformation, which is set according to the calculation accuracy requirements of the static aerodynamics; the convergence error threshold for the angle of attack trimming is the magnitude of the force difference, which is set according to the calculation accuracy requirements of the trimming angle of attack.

[0066] Preset upper and lower limits for the flight angle of attack;

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

[0068] 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 aircraft's flight altitude, speed, dynamic pressure, overload coefficient, trim mass and the aircraft's initial calculated angle of attack;

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

[0070] Step S6: Based on the finite element model A of the wing structure t Given the structural interpolation points, solve for the interpolation matrix of the flight angle of attack at time t;

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

[0072] Step S7: Perform aerodynamic calculations using CFD based on the flight angle of attack at time t and the aerodynamic grid at the flight angle of attack at time t to obtain the aerodynamic interpolation matrix at the flight angle of attack at time t.

[0073] The aerodynamic force at the flight angle of attack at time t is applied to the flight angle of attack interpolation matrix at time t for force interpolation, thus obtaining the force interpolation matrix at the flight angle of attack at time t.

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

[0075]

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

[0077] Step S8: Based on the flight angle-of-attack force interpolation matrix at time t, perform an analysis on the finite element model A of the wing structure. t To deform;

[0078] The CSD method of solid dynamics is used to obtain the wing structure deformation at the flight angle of attack at time t.

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

[0080] If convergence is not achieved, the aerodynamic mesh is updated until the static aeroelastic convergence criterion is met, and the flight angle of attack at time t for static aeroelastic convergence is obtained, and the next step is performed.

[0081] Optionally, if convergence is not achieved in step S9, the aerodynamic mesh is updated until the static aeroelastic convergence criterion is met, and the specific steps for proceeding to the next step include:

[0082] If the static aeroelasticity does not converge, displacement interpolation is performed based on the wing structure deformation corresponding to the flight angle of attack at time t. The radial basis function (RBF) is used to deform the wing aerodynamic mesh model to obtain the corresponding displacement. The aerodynamic mesh is then updated based on the corresponding displacement to obtain the flight angle of attack at time t.

[0083] Update the aerodynamic mesh at the flight angle of attack at time t as the aerodynamic mesh at the flight angle of attack at time t+1, let t = t+1, return to step S5, until the static aerodynamic bullet convergence criterion is met, and proceed to the next step;

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

[0085] |d t -d t-1 |≤e a

[0086] Where, d t Let d be the magnitude of the structural deformation displacement during the static aeroelastic cycle at time t. t-1 Let e ​​be the magnitude of the structural deformation displacement during the static aeroelastic cycle at time t-1. a The convergence error threshold is calculated for nonlinear aerostatics.

[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 ω is the displacement of the i-th interpolation point; Nb is the total number of interpolation points calculated by mesh deformation; i φ 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 flight angle of attack at time t with static aeroelastic convergence, and determine whether the flight angle of attack at time t with static aeroelastic convergence is balanced and converged.

[0091] If the balance convergence criterion is not met, the bisection method is used to adjust the upper and lower limits of the flight angle of attack at time t for static aeroelastic convergence, and the updated flight angle of attack at time t is obtained.

[0092] Update the 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 balancing convergence criterion is satisfied, the final flight angle of attack is obtained.

[0093] Optionally, the expression for the balancing convergence criterion is:

[0094] |F t -tMg|≤e T

[0095] Among them, F t Let g be the aerodynamic lift of the t-th trim cycle, g be the gravitational acceleration, and M be gravity.

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

[0097] If the angle of attack at time t with static aeroelastic convergence increases compared to the angle of attack at time t-1 with static aeroelastic convergence, then the angle of attack at time t with static aeroelastic convergence is taken as the lower limit of the angle of attack at time t with static aeroelastic convergence, i.e., α. d,t =α t The updated lower bound of the static aeroelastic convergence flight angle at time t is obtained. Based on the updated lower bound of the static aeroelastic convergence flight angle at time t, the updated flight angle at time t is obtained. If the flight angle at time t is decreasing compared to the static aeroelastic convergence flight angle at time t-1, then the flight angle at time t-1 is taken as the upper bound of the static aeroelastic convergence flight angle at time t, i.e., α. u,t =α t-1 We obtain the updated upper limit of the flight angle of attack for static aeroelastic convergence at time t, and based on the updated upper limit of the flight angle of attack for static aeroelastic convergence at time t, we obtain the updated flight angle of attack for static aeroelastic convergence at time t.

[0098] Optionally, the expression for updating the flight angle of attack at time t with static aeroelastic convergence is:

[0099]

[0100] Where, α t-1 Let α be the angle of attack for static aeroelastic convergence at time t-1. a,t For the updated flight angle of attack at time t with static aeroelastic convergence, α u,t α is the upper limit of the angle of attack of the flight with static aeroelastic convergence at time t. d,t This is the lower limit of the angle of attack for static aeroelastic convergence at time t.

[0101] Example 1

[0102] The design employs a large flexible wing with a fully fixed wing root. The wing features are as follows: the large flexible wing adopts a common civil aircraft wing design, with a half-span of 2000mm, a root chord length of 800mm, and a tip chord length of 130mm. It has winglets at the wingtips. A schematic diagram of this large flexible wing model is shown below. Figure 2 As shown.

[0103] An aircraft angle-of-attack trim method considering nonlinear aeroelastic effects according to an embodiment of the present invention includes:

[0104] Step 1: Calculation of operating conditions and reading of calculation files

[0105] The calculation conditions are set as follows: incoming flow velocity Mach 0.6, dynamic pressure 12000Pa, overload factor set to 1 (constant straight-line level flight), trim mass set to 400kg, and initial angle of attack set to 10°.

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

[0107] The entire flow field was structured and meshed using the finite element modeling software ICEM to establish an aerodynamic mesh model of the wing.

[0108] A finite element model of the structure was established using the finite element modeling software MSC.PATRAN, and interpolation points were set along the main beam. The structural model is as follows: Figure 3 As shown.

[0109] Step 2: Specify the calculation error threshold and the upper and lower limits of the bisection method.

[0110] Before starting the calculation, specify the convergence error threshold e for the nonlinear aerostatic elastic calculation. a With angle of attack balancing convergence error threshold e T The convergence error threshold for static aerodynamic calculation is set to 1 mm, and the convergence error threshold for angle of attack balancing is set to 20 N. Furthermore, the upper and lower limits α for the bisection method angle of attack calculation need to be specified based on calculation requirements and engineering experience. u With α d The values ​​are 15° and 0° respectively.

[0111] Step 3: CFD / CSD Nonlinear Static Aeroelasticity Solution

[0112] After the trim calculation begins, it first enters a nonlinear static aeroelastic iterative loop. The input aerodynamic and structural models are read and initialized, and the interpolation matrix is ​​solved based on the set interpolation points for subsequent calculations. After initialization, CFD, CSD, and mesh deformation calculations are called sequentially to perform nonlinear aeroelastic solutions on the wing until the deformation converges and enters the trim determination stage.

[0113] Step 4: Trim convergence determination and update angle of attack

[0114] After the nonlinear static aerodynamic design converges, the solution proceeds to the angle-of-attack trim convergence determination stage. Convergence is determined, and the angle of attack is adjusted using the bisection method. The nonlinear static aerodynamic design is then re-entered into the solution process, iterating repeatedly until final convergence. After five iterations, the trim angle of attack converges, resulting in a final trim angle of attack of 6.875° for the large flexible wing. The wing deformation comparison after final convergence is shown below. Figure 4 The convergence curve of the trim angle of attack of a large flexible wing is as follows: Figure 5 As shown.

[0115] The above description is only a preferred embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any changes or substitutions that can be easily conceived by those skilled in the art within the scope of the technology disclosed in the present invention should be included within the scope of protection of the present invention.

Claims

1. An aircraft angle of attack trim method that takes into account nonlinear aeroelastic effects, characterized in that, The method comprises the following steps: Step S1, establishing a finite element model of the wing structure and setting up structure interpolation points; Step S2, reading the flight attack angle to be measured and its related working conditions; Step S3, setting t=1, which represents the initial time; Step S5, performing aerodynamic force calculation of fluid mechanics CFD based on the flight attack angle working condition at time t and the flight attack angle aerodynamic grid at time t to obtain the flight attack angle aerodynamic force interpolation matrix at time t; Step S4: Based on the finite element model A of the wing structure t The structural interpolation points are used to solve the interpolation matrix of the angle of attack at time t; based on the wing aerodynamic mesh model B t Obtain the aerodynamic grid at the moment of flight angle of attack t; Step S6, obtaining the wing structure deformation amount corresponding to the flight attack angle at time t by using the solid dynamics CSD method; Step S6, interpolating the finite element model A of the wing structure on the basis of the time instant t and the angle of attack force matrix t deformation is performed; Step S7, performing nonlinear static aeroelasticity convergence judgment by using the wing deformation amount corresponding to the flight attack angle at time t, and if the static aeroelasticity converges, the next step is performed; If the convergence is not achieved, the flight attack angle time t aerodynamic grid is updated until the static aeroelasticity convergence standard is met, the flight attack angle of static aeroelasticity convergence at time t is obtained, and the next step is performed; Step S8, obtaining the aerodynamic force of the flight attack angle of static aeroelasticity convergence at time t and judging whether the flight attack angle of static aeroelasticity convergence at time t is trimmed and converged; If the trimming and convergence criterion is met, the flight attack angle of static aeroelasticity convergence at time t is taken as the trimmed flight attack angle; if the trimming and convergence criterion is not met, the upper and lower limits of the flight attack angle of static aeroelasticity convergence at time t are adjusted to obtain the updated flight attack angle at time t; The updated flight attack angle at time t is taken as the flight attack angle at time t+1, t is set to t+1, and the step S3 is returned.

2. The aircraft attack angle trimming method considering nonlinear aeroelasticity effect according to claim 1, wherein the flight attack angle related working conditions comprise aircraft flight height, speed, dynamic pressure, overload coefficient and trimming mass.

3. The aircraft attack angle trimming method considering nonlinear aeroelasticity effect according to claim 1, wherein the specific steps of updating the flight attack angle time t aerodynamic grid until the static aeroelasticity convergence standard is met in step S7 include: The flight attack angle updated aerodynamic grid at time t is taken as the flight attack angle time t+1 aerodynamic grid, t is set to t+1, the step S3 is returned, and the static aeroelasticity convergence standard is met. The specific steps of obtaining the updated flight attack angle at time t in step S8 include: If the flight attack angle of static aeroelasticity convergence at time t is in an increasing state compared with the flight attack angle of static aeroelasticity convergence at time t-1, the flight attack angle of static aeroelasticity convergence at time t is taken as the lower limit of the flight attack angle of static aeroelasticity convergence at time t, and the updated flight attack angle at time t is obtained; if the flight attack angle of static aeroelasticity convergence at time t is in a decreasing state compared with the flight attack angle of static aeroelasticity convergence at time t-1, the flight attack angle of static aeroelasticity convergence at time t-1 is taken as the upper limit of the flight attack angle of static aeroelasticity convergence at time t, and the updated flight attack angle at time t is obtained. If not converged, displacement interpolation is performed according to the wing structure deformation amount corresponding to the flight attack angle at the time t, and a radial basis function RBF is used to update the aerodynamic grid model B t deformation to obtain the corresponding displacement, and the updated aerodynamic grid at the flight attack angle at the time t is obtained based on the corresponding displacement.

5. The aircraft attack angle trimming method considering nonlinear aeroelasticity effect according to claim 1, wherein the expression of the trimming and convergence criterion in step S8 is:

4. The method for determining the angle of attack trim of an aircraft taking into account the nonlinear aeroelastic effects according to claim 1, characterized in that, 6. The aircraft attack angle trimming method considering nonlinear aeroelasticity effect according to claim 3, wherein ​ ​ ​ |d t -d t-1 |≤e a where d t is the structural deformation displacement size of the static aeroelastic cycle at time t, d t-1 is the structural deformation displacement size of the static aeroelastic cycle at time t-1, e a is the nonlinear static aeroelastic calculation convergence error threshold. ​ The expression of the corresponding displacement amount is: 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 the mesh 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. 7.The aircraft angle of attack trim method considering nonlinear aeroelastic effects according to claim 4, characterized in that, The expression of the updated flight attack angle of the static aeroelastic convergence at time t is: where α t-1 is the static aeroelastic converged angle of attack at time t - 1, α a,t is the updated static aeroelastic converged angle of attack at time t, α u,t is the upper limit of the static aeroelastic converged angle of attack at time t, α d,t is the lower limit of the static aeroelastic converged angle of attack at time t, F t is the aerodynamic lift of the tth trim cycle, g is the gravitational acceleration, and M is the weight.

Images