Aircraft performance optimization method, device and equipment based on efficient dynamic mesh technology

Abstract

This application relates to a method, apparatus, and device for optimizing aircraft performance based on efficient dynamic mesh technology. The method involves interpolating the displacement of the solid domain onto the boundary of a first background mesh using inverse distance weighted interpolation. The inclusion relationship between nodes and elements is determined using the surface vector method, and volume coordinates are calculated. The first background mesh boundary is deformed using a spring method to obtain new coordinates for each node. The original mesh node coordinate information is updated, and the aircraft deformation is calculated based on CFD. A second background mesh and a body mesh for the aircraft are established. The connectivity between the body mesh and the second background mesh is established using inverse distance weighted interpolation, and interpolation is performed. The aircraft performance under different flight angles of attack is calculated using CFD. The aircraft performance is optimized based on the structural deformation and performance under different flight angles of attack. When optimizing an aircraft using this invention, deformation accuracy and computational efficiency can be maintained while improving deformation capability.

Description

Technical Field

[0001] This application relates to the field of computational fluid dynamics technology, and in particular to a method, apparatus and equipment for optimizing aircraft performance based on efficient dynamic mesh technology. Background Technology

[0002] The development of hypersonic vehicles is increasingly recognized by major military powers worldwide for its significant strategic and tactical value in future military confrontations. The design of hypersonic vehicles can be divided into three parts: theoretical calculation, numerical simulation, and experimentation. After the structural parameters of a hypersonic vehicle are designed through theoretical calculations, they are first verified through numerical simulation to determine whether its performance indicators meet the requirements. Then, the structural parameters are improved to obtain the required specifications for experimental testing. It can be seen that high-precision numerical simulation can significantly reduce the number of experiments, thereby reducing design time and costs.

[0003] In existing multiphysics coupled numerical simulations of hypersonic vehicles, dynamic meshing methods organically combine the calculations of the solid and fluid domains. Dynamic meshing methods are mainly divided into mesh deformation methods and overlapping mesh methods. However, among existing mesh deformation methods, commonly used quasi-physical methods, such as the spring analogy, typically input displacement and then iteratively solve a system of equations to obtain the positions of each node in the deformed mesh. This method has poor handling capability when facing large and torsional deformations, and is prone to "bending through" phenomena, making post-deformation calculations impossible. Commonly used interpolation methods can be divided into explicit interpolation methods and implicit interpolation methods. Explicit interpolation methods, such as the inverse distance weighting method, are based on distance calculations and are suitable for large-scale parallel computing. However, the "point-to-point" characteristic of not considering connectivity causes the quality of the deformed mesh to deteriorate rapidly, affecting simulation accuracy. Implicit interpolation methods, such as the radial basis function method, use a function of the radial distance between points, obtain the coefficients of the interpolation function by solving a system of linear equations, and then obtain the positions of the internal nodes. This method achieves good deformation results, but its linear equation system has a scale equal to the square of the number of boundary points, resulting in a large computational load and compromised computational efficiency when applied to 3D problems. Overlapping mesh methods, due to their simplicity and minimal impact on mesh quality, are widely used for calculating the motion of structures in flow fields. Traditional overlapping mesh methods, developed based on CPU architecture, work by dividing the structure into two sets of meshes—a background mesh and a body mesh. The body mesh encloses the structure within the background mesh, and at each node reached, a hole of roughly the same size is created in the background mesh. Then, the boundaries of the background mesh and the body mesh are defined, and information is exchanged between the two boundaries through interpolation to achieve numerical simulation of the structure's motion. However, this method suffers from low computational efficiency and is unsuitable for implementation on GPU architectures due to the serial nature of the labeling and recursive operations on large 3D meshes. Summary of the Invention

[0004] Therefore, it is necessary to provide a method, apparatus, and device for optimizing aircraft performance based on efficient dynamic mesh technology, which is suitable for GPU architecture and can ensure deformation accuracy and computational efficiency while improving deformation capability, in order to address the above-mentioned technical problems.

[0005] A method for optimizing aircraft performance based on efficient dynamic mesh technology, the method comprising:

[0006] Establish the solid domain boundary mesh and the original fluid domain mesh of the aircraft, and establish the first background mesh based on the fluid domain;

[0007] Interpolate the boundary displacement of the solid domain to the boundary of the first background grid; determine the corresponding inclusion relationship between the nodes in the original grid of the fluid domain and the elements in the first background grid of the fluid domain, and calculate the volume coordinates of the original grid nodes in the first background grid elements;

[0008] Deformation calculations are performed on the boundary of the first background mesh to obtain the new coordinate information of each node in the first background mesh after deformation.

[0009] After updating the node coordinate information of the original mesh of the fluid domain according to the volume coordinates and the new coordinate information, the deformation of the aircraft structure and the performance parameters after deformation are obtained based on CFD calculation.

[0010] A second background grid and a body grid for the aircraft are established, and the connection between the body grid and the second background grid is established. After the nodes of the body grid are independently and in parallel interpolated to the second background grid, the aircraft performance under different flight angles of attack is obtained by CFD calculation based on the flight time and corresponding angle of attack of the aircraft.

[0011] The performance of the aircraft is optimized based on the deformation of the aircraft structure, the performance parameters after deformation, and the performance of the aircraft at different flight angles of attack.

[0012] In one embodiment, before establishing the solid domain boundary mesh and the original fluid domain mesh of the aircraft, and before establishing the first background mesh based on the fluid domain, the method further includes:

[0013] Based on the design requirements of the aircraft, the overall design parameters of the aircraft are obtained by analyzing the force balance of the aircraft and calculating based on the aircraft's equation of motion. The basic configuration of the aircraft is then determined based on the overall design parameters.

[0014] Based on the overall design parameters and the basic configuration, a preliminary design of the aircraft is carried out, a physical model of the aircraft is established, and the solid domain boundary mesh and the fluid domain original mesh are divided based on the physical model of the aircraft.

[0015] In one embodiment, the process of obtaining the boundary displacement of the solid domain is as follows: by setting flight conditions, boundary conditions are applied to the structure of the aircraft, and the temperature rise and deformation of the solid domain of the aircraft are calculated by the finite element method to obtain the boundary displacement of the solid domain.

[0016] In one embodiment, before performing deformation calculations on the first background mesh boundary, the method further includes:

[0017] Based on the volume coordinate information, calculate the volumes of the four first tetrahedrons of the cells in the first background mesh and the volumes of the four second tetrahedrons formed by the nodes in the original mesh and the cells in the first background mesh.

[0018] Calculate the absolute value of the sum of the ratios of the volumes of the first tetrahedron and the second tetrahedron, and determine whether the absolute value is greater than a set threshold; if not, perform deformation calculation on the first background mesh.

[0019] In one embodiment, deformation calculation is performed on the boundary of the first background mesh to obtain the new coordinate information of each node in the first background mesh after deformation, including: performing deformation calculation on the boundary of the first background mesh using the spring analogy method;

[0020] First, determine the initial tension of the spring between the two nodes:

[0021] F ij =K ij (x i -x j )

[0022] The resultant force on node i is the resultant force of the tension of all the springs connecting it:

[0023]

[0024] Where i and j represent nodes, x represents the node coordinates, and N i K is the set of all points connected to node i via springs. ij The elastic coefficient is expressed as: Indicates the side length;

[0025] When the boundary of the first background mesh moves, the coordinates of its corresponding boundary mesh nodes change. To maintain the force balance of each node in the background mesh, the new coordinate information of each node after deformation is solved by linear equations:

[0026] AX = F

[0027] Where A represents a matrix, X represents the determinant of the coordinates of each node, and F represents the determinant of the resultant force acting on each node.

[0028] In one embodiment, the new coordinates of each node after deformation are converged, and the iterative formula is as follows:

[0029]

[0030] Where k represents the number of iterations.

[0031] In one embodiment, a second background mesh and a body mesh for the aircraft are established, and the connectivity between the body mesh and the second background mesh is established. The body mesh nodes are then independently and in parallel interpolated to the second background mesh, including:

[0032] The flight path is planned according to the aircraft design requirements and the angle of attack at different times is calculated. The aircraft is divided into a second background grid containing the incoming flow conditions and a body grid that surrounds the aircraft.

[0033] Set boundary conditions for the second background mesh and perform a calculation for one time step, then assign initial values ​​to the second background mesh;

[0034] The correspondence between the body center of the body mesh boundary element and the second background mesh element is solved by the surface vector method. It is determined whether the body center of the body mesh boundary element is in the background mesh element. If so, the relevant information of the background mesh is interpolated onto the body mesh boundary element by the inverse distance weighting method. The body mesh is calculated for one time step with the interpolated information as the boundary condition to obtain the full field value of the body mesh.

[0035] The full-field value is interpolated into the corresponding region of the second background grid using the inverse distance weighting method, and the information of the interpolated unit is fixed so that it does not change with the calculation.

[0036] In one embodiment, the correspondence between the body center of the body-boundary grid element and the second background grid element is solved using the surface vector method, including:

[0037] The judgment expression is expressed as:

[0038]

[0039] Wherein, the volume center coordinates of the boundary element of the body mesh are P(x p y p , z p The midpoint of each face in the second background mesh element is N. i (x i y i , z i The normal unit vector of each face in the second background mesh element is...

[0040] If the A value of each face in the background mesh element is greater than 0, then the body center of the body mesh boundary element is in the background mesh element.

[0041] An aircraft performance optimization device based on efficient dynamic mesh technology, the device comprising:

[0042] The mesh generation module is used to create the boundary mesh of the solid domain and the original mesh of the fluid domain of the aircraft, as well as to create the first background mesh based on the fluid domain;

[0043] The structural deformation analysis module is used to interpolate the boundary displacement of the solid domain onto the boundary of the first background mesh; determine the correspondence between the nodes in the original mesh of the fluid domain and the elements in the first background mesh of the fluid domain, and calculate the volume coordinates of the original mesh nodes in the first background mesh elements; perform deformation calculation on the boundary of the first background mesh to obtain the new coordinate information of each node in the first background mesh after deformation; update the node coordinate information of the original mesh of the fluid domain according to the volume coordinates and the new coordinate information, and then calculate the deformation of the aircraft structure and the performance parameters after deformation based on CFD.

[0044] The angle of attack analysis module is used to establish a second background grid and a body grid for the aircraft, and to establish the connection relationship between the body grid and the second background grid. After interpolating the nodes of the body grid independently and in parallel to the second background grid, the aircraft performance under different flight angles of attack is obtained by CFD calculation based on the flight time and corresponding angle of attack of the aircraft.

[0045] The optimization module is used to optimize the performance of the aircraft based on the structural deformation of the aircraft, the performance parameters after deformation, and the performance of the aircraft at different flight angles of attack.

[0046] A computer device includes a memory and a processor, the memory storing a computer program, and the processor executing the computer program performing the following steps:

[0047] Step 102: Establish the solid domain boundary mesh and the original fluid domain mesh of the aircraft, and establish the first background mesh based on the fluid domain;

[0048] Step 104: Interpolate the boundary displacement of the solid domain onto the boundary of the first background mesh; determine the corresponding inclusion relationship between the nodes in the original mesh of the fluid domain and the elements in the first background mesh of the fluid domain, and calculate the volume coordinates of the original mesh nodes in the first background mesh elements; perform deformation calculation on the boundary of the first background mesh to obtain the new coordinate information of each node in the first background mesh after deformation; update the node coordinate information of the original mesh of the fluid domain according to the volume coordinates and the new coordinate information, and then calculate the deformation of the aircraft structure and the performance parameters after deformation based on CFD.

[0049] Step 106: Establish the second background mesh and the body mesh of the aircraft, and establish the connection relationship between the body mesh and the second background mesh. After interpolating the nodes of the body mesh independently and in parallel to the second background mesh, the aircraft performance under different flight angles of attack is obtained by CFD calculation based on the flight time and corresponding angle of attack of the aircraft.

[0050] Step 108: Optimize the performance of the aircraft based on the deformation of the aircraft structure, the performance parameters after deformation, and the performance of the aircraft at different flight angles of attack.

[0051] The aforementioned method, apparatus, and equipment for optimizing aircraft performance based on efficient dynamic mesh technology establishes a solid domain boundary mesh and a fluid domain original mesh for the aircraft, and establishes a first background mesh based on the fluid domain; interpolates the boundary displacement of the solid domain onto the boundary of the first background mesh; determines the correspondence between nodes in the original fluid domain mesh and elements in the first background mesh, and calculates the volume coordinates of the original mesh nodes in the first background mesh element; performs deformation calculation on the boundary of the first background mesh to obtain the new coordinate information of each node in the first background mesh after deformation; updates the node coordinate information of the original fluid domain mesh according to the volume coordinates and the new coordinate information, and calculates the structural deformation and post-deformation performance parameters of the aircraft based on CFD; establishes a second background mesh and a body mesh for the aircraft, and establishes the connectivity between the body mesh and the second background mesh; interpolates the nodes of the body mesh independently and in parallel to the second background mesh; and calculates the aircraft performance at different flight angles of attack based on the flight time and corresponding angle of attack of the aircraft through CFD; and optimizes the aircraft performance based on the structural deformation, post-deformation performance parameters, and aircraft performance at different flight angles of attack. This invention establishes connectivity between the background mesh and the original mesh through interpolation. Since the background mesh has strong sparse deformation capability, the deformation amount of the background mesh is interpolated onto the original mesh through mesh deformation method, which improves the deformation capability while ensuring the accuracy and efficiency of deformation. At the same time, in order to be suitable for GPU architecture, the "hole-punching" operation in the traditional overlapping mesh method is removed according to the parallel characteristics of GPU architecture. Meanwhile, the connectivity is established through interpolation method to realize independent parallel interpolation of each node, which improves the computational efficiency. Attached Figure Description

[0052] Figure 1 This is a flowchart of an aircraft performance optimization method based on efficient dynamic mesh technology in one embodiment;

[0053] Figure 2 Here is a flowchart of the improved mesh deformation method in one embodiment;

[0054] Figure 3Here is a flowchart of the improved overlapping mesh method in one embodiment;

[0055] Figure 4 This is a structural block diagram of an aircraft performance optimization device based on efficient dynamic mesh technology in one embodiment;

[0056] Figure 5 This is an internal structural diagram of a computer device in one embodiment. Detailed Implementation

[0057] To make the objectives, technical solutions, and advantages of this application clearer, the following detailed description is provided in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative and not intended to limit the scope of this application.

[0058] It should be noted that in this invention, the use of terms such as "first," "second," etc., is for descriptive purposes only and should not be construed as indicating or implying their relative importance or implicitly specifying the number of technical features indicated. Therefore, a feature defined as "first" or "second" may explicitly or implicitly include at least one of that feature. In the description of this invention, "a plurality of" means at least two, such as two, three, etc., unless otherwise explicitly specified.

[0059] In one embodiment, such as Figure 1 As shown, a method for optimizing aircraft performance based on efficient dynamic mesh technology is provided. This method includes the following steps:

[0060] Step 102: Establish the solid domain boundary mesh and the original fluid domain mesh of the aircraft, and establish the first background mesh based on the fluid domain.

[0061] Specifically, based on the aircraft's design requirements, the overall design parameters of the aircraft are obtained by analyzing its force balance and calculating its equations of motion. Based on these parameters, the wings, fuselage, and other components are designed to determine the aircraft's basic configuration. A preliminary design of the aircraft is then performed based on the overall design parameters and basic configuration, establishing a physical model. This physical model is then used to divide the solid domain boundary mesh and the fluid domain original mesh.

[0062] To be more specific, analyzing the force balance of an aircraft, the forces acting on the aircraft can be divided into four parts: the aircraft's own weight G, the lift generated during flight L, the drag generated during flight D, and the thrust of the engine F.

[0063] Taking horizontal uniform linear flight as an example, its force analysis can be obtained as follows: L=G=mg;F=D; In the formula, ρ is the air density, V is the flight speed, S is the wing area, and C is the air density.L C is the lift coefficient of the aircraft. D This represents the aircraft's drag coefficient.

[0064] Based on the aircraft's equations of motion, the main overall design parameters of the aircraft under flight conditions are: takeoff weight m; wing area S; power plant thrust F; and two combined parameters: thrust-to-weight ratio F / (mg); wing loading m / S.

[0065] Taking an aircraft as an example, the overall parameters given by the calculation formula and design requirements are used to design the wings, fuselage, and other parts of the aircraft. First, the layout is selected based on the characteristics of the required parameters. Typical layouts include canard, tailless, and triangular layouts, each with its own characteristics and suitable for different flight applications. After selecting the overall aircraft layout, the wings and tail are designed. Wing airfoils can be classified as laminar flow airfoils, high-lift airfoils, and supercritical airfoils. After selecting the airfoil, the wing geometry is calculated and selected. Assuming b is the wingspan, c... r c is the chord length at the wing root. t Where A is the wingtip chord length, A = b² / S is the wing aspect ratio, Λ₀ is the leading edge sweep angle, and λ = c r / c t The aspect ratio is the tip-to-root ratio. A larger aspect ratio results in a smaller proportion of the wingtip effect's influence area, leading to a higher lift-to-drag ratio and a steeper lift-to-slope. Therefore, for large civilian airliners and military transport aircraft, an aspect ratio of around 10 is chosen to improve the lift-to-drag ratio, while for fighter jets, a ratio between 2.0 and 4.0 is generally selected. Increasing the sweep angle can raise the critical Mach number and delay shock wave generation. The half-apex angle of the wing's leading edge is... The half-apex angle of the disturbance cone is Let n = tanω / tanμ, then n < 1 is the subsonic leading edge, n = 1 is the sonic leading edge, and n > 1 is the supersonic leading edge. When selecting the sweep angle, the case of n = 1 should be avoided. The tip-to-root ratio affects the lift distribution along the span of the wing. Most low-speed straight wings have a tip-to-root ratio of 0.4-0.5, while swept wings have a tip-to-root ratio in the range of 0.167-0.5. Tail configurations include conventional, T-shaped, and cruciform types, and are selected according to the "tail capacity coefficient method."

[0066] The main geometric parameter of the aircraft fuselage is the overall length l F and the maximum cross-sectional area A F At the same time, a derived parameter, the fuselage slenderness ratio λ, is obtained. F =l F / d F In the formula, d F The equivalent diameter of the fuselage is given. The initial design is completed by selecting the slenderness ratio of the fuselage segments.

[0067] After completing the preliminary design of the aircraft based on design requirements and empirical formulas, its physical model is established, and the fluid and solid domains are meshed separately.

[0068] Step 104: Interpolate the boundary displacement of the solid domain onto the boundary of the first background mesh; determine the corresponding inclusion relationship between the nodes in the original mesh of the fluid domain and the elements in the first background mesh of the fluid domain, and calculate the volume coordinates of the original mesh nodes in the first background mesh elements; perform deformation calculation on the boundary of the first background mesh to obtain the new coordinate information of each node in the first background mesh after deformation; update the node coordinate information of the original mesh of the fluid domain according to the volume coordinates and the new coordinate information, and then calculate the deformation of the aircraft structure and the performance parameters after deformation based on CFD.

[0069] Specifically, this invention uses a mesh deformation method combined with multiphysics coupling to analyze structural deformation under different operating conditions. It mainly uses finite element calculation units, mesh deformation units, and CFD calculation units to analyze the impact of structural deformation on aircraft performance after long-endurance flight.

[0070] First, in the finite element calculation unit, by setting reasonable flight conditions, the structure of the aircraft is subjected to boundary conditions such as pressure and heat flow. The temperature rise and deformation of the solid domain of the aircraft are calculated by the finite element method, and the boundary displacement of the solid domain is obtained.

[0071] Then perform mesh deformation, such as Figure 2 As shown, the boundary displacement of the solid domain is interpolated to the first background mesh boundary of the fluid domain and the coupling surface between the fluid and solid domains by the inverse distance weighted interpolation method. It is worth noting that the coupling surface mentioned here is the boundary surface between the fluid and solid domains. The purpose of interpolating the boundary displacement of the solid domain to this coupling surface is mainly to reduce the initial disturbance of deformation and ensure the uniform deformation of the fluid and solid domains in terms of displacement.

[0072] The first background mesh of the fluid domain is deformed. First, the inclusion relationship between nodes in the original fluid domain mesh and elements in the first background mesh is determined using the surface vector method. Then, the volume coordinates of the original mesh nodes in the first background mesh elements are calculated. It is worth noting that during the initial mesh generation process, the relevant information of the nodes in the original mesh and the elements in the first background mesh were already provided. The first background mesh element is a tetrahedral structure with four nodes. During interpolation, the node information of the original mesh is stored in the volume center of the corresponding element in the first background mesh. After determining the inclusion relationship using the surface vector method, the volume coordinates of the corresponding node in the first background mesh are calculated.

[0073] Based on the relevant volume coordinate information, the volumes of the four first tetrahedrons in the first background mesh and the four second tetrahedrons formed by the nodes in the original mesh and the elements in the first background mesh are calculated respectively. Then, the absolute value of the sum of the ratios of the four first tetrahedron volumes to the four second tetrahedron volumes is calculated, and it is determined whether this absolute value is greater than a set threshold. If it is, the second tetrahedron volume is recalculated; otherwise, the first background mesh is deformed using the spring analogy method, so that the background mesh deforms according to the structural changes, and the new coordinates of each node in the first background mesh after deformation are obtained. This threshold is set to 1 based on the relationship between the first and second tetrahedrons.

[0074] When deforming the background mesh using the spring analogy method, it is assumed that each mesh edge is a spring with a spring equilibrium length of 0. First, the resultant force on each node of the original mesh in the fluid domain is calculated. After the mesh is moved, the node positions are adjusted to ensure that the resultant force on the nodes remains constant. The initial spring tension between two nodes i and j is determined using Hooke's law.

[0075] F ij =K ij (x i -x j )

[0076] The resultant force on node i is the resultant force of the tension of all the springs connecting it:

[0077]

[0078] Where x represents the node coordinates, N i K is the set of all points connected to node i via springs. ij The elastic coefficient is expressed as: Indicates the side length.

[0079] The algorithm searches for all points j connected to point i via grid lines, establishing a set. Based on this set, the resultant force equations for all points i are solved, and the corresponding resultant force determinants are established according to the order of points i, used for solving subsequent equations. When the boundary of the first background grid moves, the coordinates of its corresponding boundary grid nodes change accordingly. To maintain the force balance of each node in the background grid, the new coordinate information of each node after deformation is solved using linear equations.

[0080] AX = F

[0081] Where A represents a matrix, X represents the determinant of the coordinates of each node, and F represents the determinant of the resultant force acting on each node. Taking the first row of matrix A as an example, for points in the set N of point 1, there is a value K, and the corresponding elements for other points are 0. The vector resultant force of each point can be obtained through matrix calculation.

[0082] Since the diagonal elements are the set of points j corresponding to point i, all non-zero columns k on the off-diagonal of the row will be considered. Therefore, the diagonal is dominant in this linear equation. The Jacobi iteration method is used to converge the new coordinates of each node after deformation. The iteration formula is as follows:

[0083]

[0084] Where k represents the number of iterations. Since the spring method analysis of a node does not aim to find an exact solution to its force matrix, but only requires a solution that meets certain accuracy requirements, iterative convergence is chosen. After 3-4 iterations, the solution converges to an acceptable accuracy, greatly reducing the computational load.

[0085] The displacement of each point in the background grid is obtained by iterative processing. Then, the displacement of each node in the background grid is reflected onto the original grid of the fluid domain through interpolation. CFD calculations are then performed on the fluid domain using CFD calculation units to obtain the new flow field changes, pressure, and heat flux conditions of the aircraft. Long-duration calculations of the aircraft are then performed to obtain its structural deformation and performance parameters after deformation.

[0086] The improved mesh deformation method provided by this invention uses the spring analogy method to deform the background mesh, establishes a connection between the background mesh and the original mesh through the inverse distance weighted interpolation method, and interpolates the deformation amount of the background mesh onto the original mesh. Since the background mesh is sparse and has strong deformation capability, by combining the spring analogy method with the background mesh and the original mesh of the fluid domain, the accuracy and efficiency of deformation are guaranteed while improving the deformation capability.

[0087] Step 106: Establish the second background grid and the body grid of the aircraft, and establish the connection relationship between the body grid and the second background grid. After interpolating the nodes of the body grid independently and in parallel to the second background grid, the aircraft performance under different flight angles of attack is obtained by CFD calculation based on the flight time and corresponding angle of attack of the aircraft.

[0088] Specifically, such as Figure 3 The diagram shows a flowchart of the improved overlapping mesh method. First, the flight path is planned according to the aircraft design requirements and the angle of attack at different times is calculated. Then, a second background mesh containing the incoming flow conditions and a body mesh surrounding the aircraft are generated for the aircraft.

[0089] Boundary conditions are set for the second background mesh, and a time step calculation is performed to give the second background mesh initial values ​​across the entire field. Relevant information is stored in the cell centers of the second background mesh. The correspondence between the cell centers of the body mesh boundary elements and the second background mesh elements is solved using the surface vector method. It is determined whether the cell centers of the body mesh boundary elements are within the background mesh elements. If so, the relevant information from the background mesh is interpolated onto the body mesh boundary elements using the inverse distance weighted method. The body mesh is then calculated for one time step using the interpolated information as boundary conditions to obtain the full field value of the body mesh. The full field value is interpolated into the corresponding region of the second background mesh using the inverse distance weighted method, and the information of the interpolated cells is fixed so that it does not change during calculation. The second background mesh then performs calculation for the next time step based on the new interpolated information.

[0090] More specifically, assume the volume center coordinates of the boundary element of the body mesh are P(x p y p , z p The midpoint of each face in the second background mesh element is N. i (x i y i , z i The normal unit vector of each face in the second background mesh element is... This yields a judgment expression, expressed as:

[0091]

[0092] If the A value of each face in the background mesh element is greater than 0, then the body center of the body mesh boundary element is determined to be in the background mesh element; otherwise, the body center of the body mesh boundary element is not in the background mesh element.

[0093] Taking a tetrahedron as an example, solve for the four A values ​​respectively. If all of them are greater than 0, it proves that the body center P of the boundary element of the body mesh is inside the second background mesh element; otherwise, it is not.

[0094] According to point P(x) p y p , z p To determine the relationship between the second background mesh element and the second background mesh element, the inverse distance weighted interpolation method is used. The value of the second background mesh element is used to interpolate the body center of the body mesh to find the element containing the point and the element coplanar with this element.

[0095] Taking a hexahedron as an example, after interpolating the body center of the body mesh using the values ​​of the second background mesh elements, the second background mesh elements corresponding to the body center of the body mesh total seven elements. The body center B of these seven elements is then used. i (x i y i , z i Interpolation is performed. Distance Distance N i =1 / L i Interpolation formula In the formula, H O h is the value of the boundary element of the body mesh. n This is the value of the second background grid cell.

[0096] Simultaneously, the body centers of the background mesh elements contained within the boundary elements of the body mesh are determined using the surface vector method, and the correspondence between the second background mesh and the body mesh is established using inverse distance weighted interpolation. Subsequently, the body mesh performs calculations for one time step using the interpolated information as boundary conditions to obtain the full-field value of the body mesh. This full-field value is then interpolated into the corresponding region of the second background mesh using inverse distance weighted interpolation, and the information of the interpolated elements is fixed so that it does not change with the calculation. The second background mesh then performs calculations for the next time step based on the new interpolated information.

[0097] Based on the long-endurance flight time and corresponding angle of attack of the aircraft, a formula for the change of the aircraft's rotation speed over time is established, and the performance of the aircraft under different flight angles of attack is calculated.

[0098] The improved overlapping mesh method provided by this invention, adapted to GPU architecture, improves upon the traditional overlapping mesh method by eliminating the "hole-digging" operation and establishing connectivity between the body mesh and the second background mesh through inverse distance weighted interpolation. With this improved method, all nodes on the boundary of the body mesh can simultaneously perform surface vector calculations to find cells in the second background mesh that contain them, and simultaneously perform inverse distance weighted interpolation calculations. The assignment process of the body mesh to the second background mesh can also be performed using the same method, achieving independent parallel interpolation at each node and improving computational efficiency.

[0099] Step 108: Optimize the performance of the aircraft based on the structural deformation, performance parameters after deformation, and aircraft performance at different flight angles of attack.

[0100] Specifically, the calculation results of steps 104 and 106 are analyzed, the aircraft design parameters in step 102 are optimized, and then repeated calculations and optimizations are performed until the performance of the aircraft at each moment during long-endurance variable angle of attack flight meets the design requirements, thus completing the optimization of the aircraft design.

[0101] It should be understood that, although Figure 1-3 The steps in the flowchart are shown sequentially as indicated by the arrows, but these steps are not necessarily executed in the order indicated by the arrows. Unless otherwise specified herein, there is no strict order in which these steps are executed, and they can be performed in other orders. Figure 1-3At least some of the steps in the process may include multiple sub-steps or multiple stages. These sub-steps or stages are not necessarily completed at the same time, but can be executed at different times. The execution order of these sub-steps or stages is not necessarily sequential, but can be executed in turn or alternately with other steps or at least some of the sub-steps or stages of other steps.

[0102] In one embodiment, such as Figure 4 As shown, an aircraft performance optimization device based on efficient dynamic mesh technology is provided, including: a mesh generation module, a structural deformation analysis module, an angle of attack analysis module, and an optimization module, wherein:

[0103] The mesh generation module is used to create the solid domain boundary mesh and the original fluid domain mesh of the aircraft, as well as to create the first background mesh based on the fluid domain.

[0104] The structural deformation analysis module is used to interpolate the boundary displacement of the solid domain onto the boundary of the first background mesh; determine the corresponding inclusion relationship between the nodes in the original mesh of the fluid domain and the elements in the first background mesh of the fluid domain, and calculate the volume coordinates of the original mesh nodes in the first background mesh elements; perform deformation calculation on the boundary of the first background mesh to obtain the new coordinate information of each node in the first background mesh after deformation; after updating the node coordinate information of the original mesh of the fluid domain according to the volume coordinates and the new coordinate information, the deformation of the aircraft structure and the performance parameters after deformation are obtained based on CFD calculation.

[0105] The angle of attack analysis module is used to establish the second background mesh and the body mesh of the aircraft, and to establish the connection relationship between the body mesh and the second background mesh. After the body mesh nodes are independently and in parallel interpolated to the second background mesh, the aircraft performance under different flight angles of attack is obtained by CFD calculation based on the flight time and corresponding angle of attack of the aircraft.

[0106] The optimization module is used to optimize the performance of the aircraft based on the aircraft's structural deformation, post-deformation performance parameters, and aircraft performance at different flight angles of attack.

[0107] In one embodiment, the structural deformation analysis module includes finite element calculation units, improved mesh deformation units, and CFD calculation units, wherein:

[0108] The finite element calculation unit is used to apply boundary conditions such as pressure and heat flow to the structure of the aircraft by setting reasonable flight conditions, and to calculate the temperature rise and deformation of the solid domain of the aircraft using the finite element method, thereby obtaining the boundary displacement of the solid domain.

[0109] An improved mesh deformation element is used to interpolate the boundary displacement of the solid domain to the boundary of the first background mesh of the fluid domain and the coupling surface between the fluid and solid domains using the inverse distance weighted interpolation method. This deforms the original mesh of the fluid domain. First, the surface vector method is used to determine the corresponding inclusion relationship between nodes in the original fluid domain mesh and elements in the first background mesh, and the corresponding volume coordinates are calculated. Based on the relevant volume coordinate information, the volumes of the four first tetrahedrons of the elements in the first background mesh and the volumes of the four second tetrahedrons formed by the nodes in the original mesh and the elements in the first background mesh are calculated. Then, the absolute value of the sum of the ratios of the four first tetrahedron volumes to the four second tetrahedron volumes is calculated, and it is determined whether this absolute value is greater than a set threshold. If it is, the second tetrahedron volume is recalculated; otherwise, the spring analogy method is used to deform the first background mesh, causing the background mesh to deform according to the structural changes, resulting in new coordinates for each node in the first background mesh after deformation.

[0110] The CFD calculation unit is used to perform CFD calculations on the fluid domain to obtain new flow field changes, pressure, and heat flow conditions of the aircraft.

[0111] Specific limitations regarding the aircraft performance optimization device based on efficient dynamic mesh technology can be found in the limitations of the aircraft performance optimization method based on efficient dynamic mesh technology mentioned above, and will not be repeated here. Each module in the aforementioned aircraft performance optimization device based on efficient dynamic mesh technology can be implemented entirely or partially through software, hardware, or a combination thereof. These modules can be embedded in or independent of the processor in a computer device in hardware form, or stored in the memory of a computer device in software form, so that the processor can call and execute the corresponding operations of each module.

[0112] In one embodiment, a computer device is provided, which may be a server, and its internal structure diagram may be as follows: Figure 5 As shown, the computer device includes a processor, memory, network interface, and database connected via a system bus. The processor provides computational and control capabilities. The memory includes non-volatile storage media and internal memory. The non-volatile storage media stores the operating system, computer programs, and database. The internal memory provides an environment for the operation of the operating system and computer programs in the non-volatile storage media. The database stores data for an aircraft performance optimization method based on efficient dynamic mesh technology. The network interface communicates with external terminals via a network connection. When the processor executes the computer program, it implements an aircraft performance optimization method based on efficient dynamic mesh technology.

[0113] Those skilled in the art will understand that Figure 5The structure shown is merely a block diagram of a portion of the structure related to the present application and does not constitute a limitation on the computer device to which the present application is applied. Specific computer devices may include more or fewer components than those shown in the figure, or combine certain components, or have different component arrangements.

[0114] In one embodiment, a computer device is provided, including a memory and a processor, the memory storing a computer program, the processor executing the computer program to perform the following steps:

[0115] Step 102: Establish the solid domain boundary mesh and the original fluid domain mesh of the aircraft, and establish the first background mesh based on the fluid domain.

[0116] Step 104: Interpolate the boundary displacement of the solid domain onto the boundary of the first background mesh; determine the corresponding inclusion relationship between the nodes in the original mesh of the fluid domain and the elements in the first background mesh of the fluid domain, and calculate the volume coordinates of the original mesh nodes in the first background mesh elements; perform deformation calculation on the boundary of the first background mesh to obtain the new coordinate information of each node in the first background mesh after deformation; update the node coordinate information of the original mesh of the fluid domain according to the volume coordinates and the new coordinate information, and then calculate the deformation of the aircraft structure and the performance parameters after deformation based on CFD.

[0117] Step 106: Establish the second background grid and the body grid of the aircraft, and establish the connection relationship between the body grid and the second background grid. After interpolating the nodes of the body grid independently and in parallel to the second background grid, the aircraft performance under different flight angles of attack is obtained by CFD calculation based on the flight time and corresponding angle of attack of the aircraft.

[0118] Step 108: Optimize the performance of the aircraft based on the structural deformation, performance parameters after deformation, and aircraft performance at different flight angles of attack.

[0119] Those skilled in the art will understand that all or part of the processes in the methods of the above embodiments can be implemented by a computer program instructing related hardware. The computer program can be stored in a non-volatile computer-readable storage medium, and when executed, it can include the processes of the embodiments of the above methods. Any references to memory, storage, databases, or other media used in the embodiments provided in this application can include non-volatile and / or volatile memory. Non-volatile memory can include read-only memory (ROM), programmable ROM (PROM), electrically programmable ROM (EPROM), electrically erasable programmable ROM (EEPROM), or flash memory. Volatile memory can include random access memory (RAM) or external cache memory. By way of illustration and not limitation, RAM is available in various forms, such as static RAM (SRAM), dynamic RAM (DRAM), synchronous DRAM (SDRAM), dual data rate SDRAM (DDRSDRAM), enhanced SDRAM (ESDRAM), synchronous link DRAM (SLDRAM), Rambus direct RAM (RDRAM), direct memory bus dynamic RAM (DRDRAM), and memory bus dynamic RAM (RDRAM), etc.

[0120] The technical features of the above embodiments can be combined in any way. For the sake of brevity, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0121] The embodiments described above are merely illustrative of several implementation methods of this application, and while the descriptions are relatively specific and detailed, they should not be construed as limiting the scope of the invention patent. It should be noted that those skilled in the art can make various modifications and improvements without departing from the concept of this application, and these all fall within the protection scope of this application. Therefore, the protection scope of this patent application should be determined by the appended claims.

Claims

1. A method for optimizing aircraft performance based on efficient dynamic mesh technology, characterized in that, The method includes: Establish the solid domain boundary mesh and the original fluid domain mesh of the aircraft, and establish the first background mesh based on the fluid domain; Interpolate the boundary displacement of the solid domain to the boundary of the first background grid; determine the corresponding inclusion relationship between the nodes in the original grid of the fluid domain and the elements in the first background grid of the fluid domain, and calculate the volume coordinates of the original grid nodes in the first background grid elements; Deformation calculations are performed on the boundary of the first background mesh to obtain the new coordinate information of each node in the first background mesh after deformation. After updating the node coordinate information of the original mesh of the fluid domain according to the volume coordinates and the new coordinate information, the deformation of the aircraft structure and the performance parameters after deformation are obtained based on CFD calculation. A second background grid and a body grid for the aircraft are established, and the connection between the body grid and the second background grid is established. After the nodes of the body grid are independently and in parallel interpolated to the second background grid, the aircraft performance under different flight angles of attack is obtained by CFD calculation based on the flight time and corresponding angle of attack of the aircraft. The performance of the aircraft is optimized based on the deformation of the aircraft structure, the performance parameters after deformation, and the performance of the aircraft at different flight angles of attack. Establish a second background mesh and a body mesh for the aircraft, and establish the connectivity between the body mesh and the second background mesh. Then, independently and in parallel interpolate the nodes of the body mesh to the second background mesh, including: The flight path is planned according to the aircraft design requirements and the angle of attack at different times is calculated. The aircraft is divided into a second background grid containing the incoming flow conditions and a body grid that surrounds the aircraft. Set boundary conditions for the second background mesh and perform a calculation for one time step, then assign initial values ​​to the second background mesh; The correspondence between the body center of the body mesh boundary element and the second background mesh element is solved by the surface vector method. It is determined whether the body center of the body mesh boundary element is in the background mesh element. If so, the relevant information of the background mesh is interpolated onto the body mesh boundary element by the inverse distance weighting method. The body mesh is calculated for one time step with the interpolated information as the boundary condition to obtain the full field value of the body mesh. The full-field value is interpolated into the corresponding region of the second background grid using the inverse distance weighting method, and the information of the interpolated unit is fixed so that it does not change with the calculation; The surface vector method is used to solve the correspondence and inclusion relationship between the body center of the body mesh boundary element and the second background mesh element, including: The judgment expression is expressed as: Wherein, the volume center coordinates of the boundary element of the body mesh are The midpoint of each face in the second background mesh element is The normal unit vector of each face in the second background mesh element is ; If each face in the background mesh unit If all values ​​are greater than 0, then the body center of the body grid boundary element is in the background grid element.

2. The aircraft performance optimization method based on efficient dynamic mesh technology according to claim 1, characterized in that, Before establishing the solid domain boundary mesh and the original fluid domain mesh of the spacecraft, and before establishing the first background mesh based on the fluid domain, the following steps are also included: Based on the design requirements of the aircraft, the overall design parameters of the aircraft are obtained by analyzing the force balance of the aircraft and calculating based on the aircraft's equation of motion. The basic configuration of the aircraft is then determined based on the overall design parameters. Based on the overall design parameters and the basic configuration, a preliminary design of the aircraft is carried out, a physical model of the aircraft is established, and the solid domain boundary mesh and the fluid domain original mesh are divided based on the physical model of the aircraft.

3. The aircraft performance optimization method based on efficient dynamic mesh technology according to claim 1, characterized in that, The process for obtaining the boundary displacement of the solid domain is as follows: By setting flight conditions and applying boundary conditions to the structure of the aircraft, the temperature rise and deformation of the solid domain of the aircraft are calculated using the finite element method, and the boundary displacement of the solid domain is obtained.

4. The aircraft performance optimization method based on efficient dynamic mesh technology according to claim 1, characterized in that, Before performing deformation calculations on the first background mesh boundary, the process also includes: Based on the volume coordinate information, calculate the volumes of the four first tetrahedrons of the cells in the first background mesh and the volumes of the four second tetrahedrons formed by the nodes in the original mesh and the cells in the first background mesh. Calculate the absolute value of the sum of the ratios of the volumes of the first tetrahedron and the second tetrahedron, and determine whether the absolute value is greater than a set threshold; if not, perform deformation calculation on the first background mesh.

5. The aircraft performance optimization method based on efficient dynamic mesh technology according to any one of claims 1 to 4, characterized in that, Deformation calculation is performed on the boundary of the first background mesh to obtain the new coordinate information of each node in the first background mesh after deformation, including: deformation calculation of the boundary of the first background mesh using the spring analogy method; First, determine the initial tension of the spring between the two nodes: node The net force acting on it is the resultant force of the tension of all the springs connecting it: in, and Represents a node. Represents the node coordinates. For all nodes The set of points connected by a spring The elastic coefficient is expressed as: , Indicates the side length; When the boundary of the first background mesh moves, the coordinates of its corresponding boundary mesh nodes change. To maintain the force balance of each node in the background mesh, the new coordinate information of each node after deformation is solved by linear equations: in, Represents a matrix, Represents the determinant of the coordinates of each node. This represents the determinant of the resultant force acting on each node.

6. The aircraft performance optimization method based on efficient dynamic mesh technology according to claim 5, characterized in that, The iterative formula for converging the new coordinates of each node after deformation is as follows: in, Indicates the number of iterations.

7. An aircraft performance optimization device based on high-efficiency dynamic mesh technology, characterized in that, The device for optimizing aircraft performance based on efficient dynamic mesh technology as described in any one of claims 1 to 6 includes: The mesh generation module is used to create the boundary mesh of the solid domain and the original mesh of the fluid domain of the aircraft, as well as to create the first background mesh based on the fluid domain; The structural deformation analysis module is used to interpolate the boundary displacement of the solid domain onto the boundary of the first background mesh; determine the correspondence between the nodes in the original mesh of the fluid domain and the elements in the first background mesh of the fluid domain, and calculate the volume coordinates of the original mesh nodes in the first background mesh elements; perform deformation calculation on the boundary of the first background mesh to obtain the new coordinate information of each node in the first background mesh after deformation; update the node coordinate information of the original mesh of the fluid domain according to the volume coordinates and the new coordinate information, and then calculate the deformation of the aircraft structure and the performance parameters after deformation based on CFD. The angle of attack analysis module is used to establish a second background grid and a body grid for the aircraft, and to establish the connection relationship between the body grid and the second background grid. After interpolating the nodes of the body grid independently and in parallel to the second background grid, the aircraft performance under different flight angles of attack is obtained by CFD calculation based on the flight time and corresponding angle of attack of the aircraft. The optimization module is used to optimize the performance of the aircraft based on the structural deformation of the aircraft, the performance parameters after deformation, and the performance of the aircraft at different flight angles of attack.

8. A computer device comprising a memory and a processor, wherein the memory stores a computer program, characterized in that, When the processor executes the computer program, it implements the steps of the method according to any one of claims 1 to 6.

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