CN112347561B - Method, device, equipment and storage medium for analyzing static aeroelasticity of aircraft - Google Patents

Abstract

The application relates to a static pneumatic elasticity analysis method, a static pneumatic elasticity analysis device, static pneumatic elasticity analysis equipment and a storage medium of an aircraft, which are applied to a high-lift aircraft configuration in a climbing or landing scene. The method comprises the following steps: generating a three-dimensional body-attached grid of each component in the target component; respectively carrying out steady flow field analysis on each component to obtain pneumatic load generated on each component; determining the structural deformation of each component according to the pneumatic load on each component; and deforming the pneumatic surface grid and the pneumatic space grid of each component according to the structural deformation of each component, and outputting the analysis result of each component when the static pneumatic elasticity analysis of each component is determined to reach the preset convergence condition. The method can automatically realize the static pneumatic elasticity analysis of the low-speed high-lift aircraft configuration in a full process, simplify the complexity of the static pneumatic elasticity analysis of the low-speed high-lift aircraft configuration and improve the analysis efficiency. Meanwhile, the accuracy of the analysis result is improved.

Description

Method, device, equipment and storage medium for static aeroelastic analysis of aircraft

Technical Field

The present application relates to the field of aviation design, and in particular to a method, device, equipment and storage medium for static aeroelastic analysis of an aircraft.

Background technique

As we all know, aircraft will deform under the action of aerodynamic force, especially for large aspect ratio aircraft that use a large amount of composite materials. The aerodynamic deformation is quite significant, which will lead to a decrease in the control efficiency of the aircraft and even cause anti-control phenomena. Therefore, it is of great significance to conduct static aeroelastic analysis of aircraft.

Usually, a panel method can be used to perform static aeroelastic analysis on an aircraft. However, for a low-speed, high-lift aircraft configuration, in order to additionally increase the lift of the aircraft, a variety of lift-increasing components are added to the aircraft, such as leading edge inner slats, leading edge outer slats, leading edge droop, trailing edge inner flaps, trailing edge outer flaps, and multi-section flaps, etc., thereby greatly increasing the geometric complexity of the low-speed, high-lift aircraft configuration. However, the panel method mentioned above cannot accurately analyze the static aeroelastic analysis of a low-speed, high-lift aircraft configuration with high complexity. For those skilled in the art, how to perform a more accurate static aeroelastic analysis on a low-speed, high-lift aircraft configuration is a technical problem that needs to be solved urgently.

Summary of the invention

Based on this, a method, device, equipment and storage medium for static aeroelastic analysis of an aircraft are provided to improve the efficiency of static aeroelastic analysis of the aircraft and the accuracy of the analysis results.

In a first aspect, an embodiment of the present application provides a static aeroelastic analysis method for an aircraft, which is applied to a high-lift aircraft configuration in a climbing or landing scenario, and the method comprises:

Generate a three-dimensional body-fitted mesh of each component in the target component, wherein the three-dimensional body-fitted mesh includes an aerodynamic surface mesh and an aerodynamic space mesh;

Performing steady flow field analysis on each of the components respectively to obtain aerodynamic loads generated on each of the components;

Determining the structural deformation of each component according to the aerodynamic load on each component;

According to the structural deformation of each component, the aerodynamic surface grid and the aerodynamic space grid of each component are deformed, and when it is determined that the static aeroelastic analysis of each component reaches a preset convergence condition, the analysis results of each component are output.

In a second aspect, an embodiment of the present application provides a static aeroelastic analysis device for an aircraft, which is applied to a high-lift aircraft configuration in a climbing or landing scenario, and the device includes:

A mesh generation module, used to generate a three-dimensional body-fitting mesh of each component in the target component, wherein the three-dimensional body-fitting mesh includes an aerodynamic surface mesh and an aerodynamic space mesh;

A flow field analysis module, used to perform steady flow field analysis on each of the components to obtain aerodynamic loads generated on each of the components;

A deformation determination module, used to determine the structural deformation of each component according to the aerodynamic load on each component;

A mesh deformation module, used for deforming the aerodynamic surface mesh and the aerodynamic space mesh of each component according to the structural deformation amount of each component;

The result output module is used to output the analysis results of each component when it is determined that the static aeroelastic analysis of each component reaches a preset convergence condition.

In one of the embodiments, optionally, the deformation determination module includes: a load interpolation unit and a structural deformation analysis unit;

A load interpolation unit, used for interpolating the aerodynamic loads on the components to the structural models corresponding to the components;

The structural deformation analysis unit is used to perform structural finite element static deformation analysis on each interpolated structural model to obtain the structural deformation of each component.

In one of the embodiments, optionally, the three-dimensional body-fitting grid further includes: a grid correspondence relationship between the aerodynamic surface grid and the aerodynamic space grid;

The grid deformation module includes: a first deformation unit, a determination unit, and a second deformation unit;

A first deformation unit, used for deforming the aerodynamic surface mesh of each component according to the structural deformation amount of each component;

A determination unit, configured to determine the mesh deformation amount of the aerodynamic space mesh of each component according to the mesh correspondence and the mesh deformation amount of the aerodynamic surface mesh of each component;

The second deformation unit is used to deform the aerodynamic space grid of each component according to the grid deformation amount of the aerodynamic space grid.

In one of the embodiments, optionally, the first deformation unit includes: a grid integration subunit and a deformation interpolation subunit;

A grid integration subunit, used for integrating the aerodynamic surface grids of the components to obtain a target aerodynamic surface grid;

The deformation interpolation subunit is used to interpolate the structural deformation of each component onto the target aerodynamic surface grid.

In one of the embodiments, optionally, the load interpolation unit is specifically used to interpolate the aerodynamic loads on the components to the structural models corresponding to the components through a radial basis function interpolation algorithm.

In one of the embodiments, optionally, the deformation interpolation subunit is specifically configured to interpolate the structural deformation of each component onto the target aerodynamic surface grid by using a radial basis function interpolation algorithm.

Optionally, the target component is a wing and/or a tail of the high-lift aircraft configuration.

In a third aspect, an embodiment of the present application provides a static aeroelastic analysis device for an aircraft, which is applied to a high-lift aircraft configuration in a climbing or landing scenario. The computer device includes a memory and a processor. The memory stores a computer program. When the processor executes the computer program, the steps of the static aeroelastic analysis method for the aircraft provided in the first aspect of the embodiment of the present application are implemented.

In a fourth aspect, an embodiment of the present application provides a computer-readable storage medium, which is applied to a high-lift aircraft configuration in a climbing or landing scenario, and has a computer program stored thereon. When the computer program is executed by a processor, the steps of the static aeroelastic analysis method of the aircraft provided in the first aspect of the embodiment of the present application are implemented.

The static aeroelastic analysis method, device, equipment and storage medium of the aircraft provided by the embodiment of the present application, after generating the three-dimensional body-fitting mesh of each component in the target part, the computer equipment respectively performs steady flow field analysis on each component to obtain the aerodynamic load generated on each component, determines the structural deformation of each component according to the aerodynamic load on each component, and deforms the aerodynamic surface mesh and aerodynamic space mesh of each component according to the structural deformation of each component, and outputs the analysis results of each component when it is determined that the static aeroelastic analysis of each component reaches the preset convergence condition. In the process of static aeroelastic analysis of the low-speed high-lift aircraft configuration, the computer equipment can automatically generate the three-dimensional body-fitting mesh of each component, and can independently complete the flow field calculation of each component, the data transmission between the aerodynamic load and the structural model, and the data transmission between the structural deformation and the aerodynamic model, that is, the computer equipment can realize the static aeroelastic analysis of the low-speed high-lift aircraft configuration in an automated manner, simplifying the complexity of the static aeroelastic analysis of the low-speed high-lift aircraft configuration and improving the analysis efficiency. At the same time, since the three-dimensional body-fitting meshes of each component generated by the computer equipment are more in line with the actual situation of the low-speed high-lift aircraft configuration, the static aeroelastic analysis of the low-speed high-lift aircraft configuration based on the accurate three-dimensional body-fitting mesh also improves the accuracy of the analysis results.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG1 is a schematic diagram of a flow chart of a static aeroelastic analysis method for an aircraft provided in an embodiment of the present application;

FIG2 is another schematic flow chart of a static aeroelastic analysis method for an aircraft provided in an embodiment of the present application;

FIG3 is a schematic diagram of another flow chart of a static aeroelastic analysis method for an aircraft provided in an embodiment of the present application;

FIG4 is a schematic diagram of a principle of a static aeroelastic analysis method for an aircraft provided in an embodiment of the present application;

FIG5 is a schematic structural diagram of a static aeroelasticity analysis device for an aircraft provided in an embodiment of the present application;

FIG6 is a schematic structural diagram of a static aeroelasticity analysis device for an aircraft provided in an embodiment of the present application.

Detailed ways

The static aeroelastic analysis method of an aircraft provided in an embodiment of the present application is applied to a high-lift aircraft configuration in a climbing or landing scenario, that is, applied to a low-speed high-lift aircraft configuration. The wings of the high-lift aircraft configuration include leading edge slats, trailing edge flaps, and main wings. When the aircraft is in the climbing or landing stage, the leading edge slats and trailing edge flaps of the aircraft wings will be opened to enable the aircraft to obtain higher lift. However, during the climbing or landing stage, the wings and tail of the high-lift aircraft configuration will deform under the action of aerodynamic forces. For this application scenario, it is of great significance to perform static aeroelastic analysis on the low-speed high-lift aircraft configuration.

In traditional technology, manual participation is usually required in the division of the grid, and when the configuration and appearance of the low-speed, high-lift aircraft change, it is necessary to combine the grid deformation technology for assistance, and then perform static aeroelastic analysis based on methods such as the panel method. Since the configuration of the low-speed, high-lift aircraft is extremely complex, this configuration has strict requirements on the accuracy and robustness of the grid deformation technology, and it is very easy to have calculation divergence, and the calculation efficiency is low, which cannot meet engineering needs. To this end, the static aeroelastic analysis method, device, equipment and storage medium of the aircraft provided in the embodiments of the present application can realize the static aeroelastic analysis of the low-speed, high-lift aircraft configuration in an automated manner throughout the entire process.

It should be noted that the static aeroelastic analysis method for an aircraft provided in the embodiment of the present application may be executed by a static aeroelastic analysis device for an aircraft, which may be implemented as part or all of a static aeroelastic analysis device for an aircraft (hereinafter referred to as a computer device for the sake of discussion) through software, hardware, or a combination of software and hardware. Optionally, the computer device may be a desktop computer, a mainframe computer, a smart phone, a wearable device, or an independent server and server cluster. The following method embodiments are described by taking a computer device as an example in which the executing subject is used.

In order to make the purpose, technical solution and advantages of this application more clear, the technical solution in the embodiment of this application is further described in detail through the following embodiments and in combination with the accompanying drawings. It should be understood that the specific embodiments described here are only used to explain this application and are not used to limit this application.

FIG1 is a flow chart of a method for analyzing the static aeroelasticity of an aircraft provided in an embodiment of the present application. This embodiment relates to a specific process of how a computer device automatically performs static aeroelasticity analysis of a low-speed, high-lift aircraft configuration. As shown in FIG1 , the method may include:

S101, generating a three-dimensional body-fitting mesh of each component in a target component, wherein the three-dimensional body-fitting mesh includes an aerodynamic surface mesh and an aerodynamic space mesh.

Specifically, the target component is a component that needs to be analyzed for static aeroelasticity in the configuration of a low-speed, high-lift aircraft. In the take-off and landing stages of the low-speed, high-lift aircraft configuration, the wing and empennage of the low-speed, high-lift aircraft configuration are easily deformed under the action of aerodynamic forces, so the target component can be the wing and/or empennage of the low-speed, high-lift aircraft configuration. When the target component is a wing, each assembly in the target component can be a leading edge slat, a main wing, a trailing edge flap, a pylon, and a nacelle, etc. Of course, further refinement can also be performed for leading edge slats and trailing edge flaps, such as further refinement of the leading edge slats into leading edge inner slats, leading edge outer slats, and leading edge droop, etc. When the target component is an empennage, each assembly in the target component can be a horizontal tail and a vertical tail, etc.

Optionally, before S101, the computer device may geometrically divide the target component of the low-speed high-lift aircraft configuration according to the component type. Taking the target component as a wing as an example, the computer device divides the wing of the low-speed high-lift aircraft configuration into five categories: leading edge slats, main wings, trailing edge flaps, pylons, and nacelles.

Next, the computer device automatically generates a three-dimensional body-fitting mesh for each component, and the three-dimensional body-fitting mesh includes an aerodynamic surface mesh and an aerodynamic space mesh. The aerodynamic surface mesh refers to the aerodynamic mesh covering the surface of the component, and the aerodynamic space mesh refers to the aerodynamic mesh surrounding the component. The following specifically introduces the process of the computer device generating a three-dimensional body-fitting mesh for each component. The following takes any target component among the components as an example. The process can be:

S1011. Generate a two-dimensional body-fitting mesh of the target component according to geometric parameters of the target component.

Wherein, the target component is any one of the components in the wing or tail of the high-lift aircraft configuration. For each target component of the high-lift aircraft configuration, its geometric shape can be represented by parameters. For example, for the wing-body assembly configuration, if the target component is an aircraft wing, the wing geometry can be described by the airfoil section data points along the wingspan direction. After the geometry of each target component is described by parameters, for each target component, the computer device generates a two-dimensional body-fitting grid of the target component by a hyperbolic function conformal transformation method based on the geometric parameters of the target component and the C-H grid topology. Of course, the two-dimensional body-fitting grid of the target component can also be generated by an algebraic generation method or a partial differential equation method, which is not limited in this embodiment.

The process of generating a two-dimensional body-fitting mesh of a target component by a computer device using a hyperbolic function conformal transformation method can be as follows: according to the C-H mesh topology and the geometric parameters of the target component, the surface geometry and far-field boundary of the target component are transformed from the real physical domain xy coordinates to the reference domain uv coordinates by a hyperbolic function conformal transformation method, and then the full flow field mesh is interpolated in the reference domain and then transformed back to the physical domain. The two-dimensional body-fitting mesh generated in this way has good orthogonality and body-fitting properties.

S1012: interpolate the two-dimensional body-fitting grid along the span direction or the circumferential direction of the target component to obtain a three-dimensional body-fitting grid of the target component.

Among them, after obtaining the two-dimensional body-fitting mesh of each target component, the computer device interpolates the two-dimensional body-fitting mesh of the target component along the span direction or circumferential direction of the target component, so as to obtain the three-dimensional body-fitting mesh of the target component. Taking the fuselage of an aircraft as an example, the computer device interpolates the two-dimensional body-fitting mesh of the fuselage generated above along the direction from the nose to the tail, so as to obtain the three-dimensional body-fitting mesh of the fuselage of the aircraft; taking the target component as the main wing of the aircraft as an example, the computer device interpolates the two-dimensional body-fitting mesh of the main wing generated above along the span direction, so as to obtain the three-dimensional body-fitting mesh of the main wing of the aircraft; taking the target component as the nacelle of the aircraft as an example, the computer device interpolates the two-dimensional body-fitting mesh of the nacelle generated above along the circumferential direction of the nacelle, so as to obtain the three-dimensional body-fitting mesh of the nacelle of the aircraft. Of course, for the target components such as the leading edge slats, trailing edge flaps and tail wing, the corresponding two-dimensional body-fitting meshes can be interpolated according to the above method to obtain the corresponding three-dimensional body-fitting meshes.

S1013. Determine attribute information of each mesh in the three-dimensional body-fitting mesh of the target component according to the three-dimensional body-fitting mesh of the target component and the three-dimensional body-fitting meshes of adjacent components, and output the three-dimensional body-fitting mesh of the target component including the mesh attribute information.

Among them, the adjacent components are components adjacent to the position where the target component exists. Since the three-dimensional body-fitting grids of the adjacent components and the three-dimensional body-fitting grids of the target component are nested, when performing flow field analysis on a low-speed, high-lift aircraft, it is necessary to consider the flow field influence of the adjacent components on the target component. After obtaining the three-dimensional body-fitting grid of each target component, for each target component, the computer device calculates the attribute information of each grid in the three-dimensional body-fitting grid of the target component according to the positional relationship between the three-dimensional body-fitting grid of the target component and the three-dimensional body-fitting grid of the adjacent components. Among them, the adjacent components of the target component can be one or more, and the attribute information of a certain grid can represent whether the grid is embedded in the three-dimensional body-fitting grid of the adjacent component. In other words, in the process of grid generation, the nested relationship between the components or parts of the configuration of the low-speed, high-lift aircraft is fully considered.

Specifically, for each mesh in the three-dimensional body-fitting mesh of the target component, when the mesh enters the geometric interior of the three-dimensional body-fitting mesh of the adjacent component, the mesh is determined as a geometric interior mesh, and when the mesh does not enter the geometric interior of the three-dimensional body-fitting mesh of the adjacent component, the mesh is determined as a normal mesh; the normal mesh adjacent to each geometric interior mesh is determined as a cross mesh. In this way, each mesh in the three-dimensional body-fitting mesh of the target component is marked with attribute information, and the attribute information of the mesh includes normal mesh, cross mesh and geometric interior mesh. When performing flow field analysis of the high-lift aircraft configuration, the cross mesh is used to transfer the flow field relationship between the three-dimensional body-fitting mesh of the target component and the three-dimensional body-fitting mesh of the adjacent component.

In practical applications, when the target component is the main wing, the fuselage geometry will affect the flow field of the main wing. In order to accurately capture the influence of the fuselage geometry on the flow field of the main wing, the computer equipment also needs to select the main wing root mesh from the 3D body-fitting mesh of the main wing, and project the main wing root mesh onto the fuselage geometry surface to obtain the main wing projected 3D mesh and output the main wing projected 3D mesh.

Among them, the main wing projected 3D mesh is used to calculate the flow field influence of the fuselage on the main wing during flow field analysis and calculation. Since the fuselage shape is composed of a series of quasi-circular geometric sections from the nose to the tail, it is narrow and long, and it is difficult to directly project the main wing root mesh onto the fuselage geometric surface. Therefore, it is necessary to first convert the selected main wing root mesh to the fuselage coordinate system to obtain the main wing root transformation mesh, and then project the main wing root transformation mesh onto the fuselage geometric surface to obtain the main wing projected 3D mesh, and output the obtained main wing projected 3D mesh.

In summary, in the process of generating the three-dimensional body-fitting meshes of each component, the mesh nesting between the components of the low-speed and high-lift aircraft configuration and the influence of the fuselage geometry on the flow field of the main wing are fully considered, so that the generated three-dimensional body-fitting mesh is more accurate, which lays a good foundation for the subsequent static aeroelastic analysis of the low-speed and high-lift aircraft configuration.

S102, performing steady flow field analysis on each of the components respectively to obtain aerodynamic loads generated on each of the components.

Among them, the computer equipment uses the inviscid external flow and boundary layer coupling technology to solve the flow field of each component, so as to obtain the aerodynamic load generated on each component. Specifically, the Euler equation is used for calculation in the external inviscid flow, and the lattice center format finite volume method is used for numerical discretization; at the same time, the implicit upwind format is introduced to improve the convergence criterion, the second-order Roe format is used to calculate the Riemann problem convective flux, and the LU+GMRES method and OpenMP (shared memory parallel programming) parallel technology are used to accelerate the flow field update. Through this technical means, the calculated aerodynamic load is more accurate.

S103: Determine the structural deformation of each component according to the aerodynamic load on each component.

Specifically, after obtaining the aerodynamic loads generated on each component, the computer device needs to transfer the aerodynamic loads on each component to the structural model corresponding to each component, so that each structural model is deformed. Then, the computer device analyzes each deformed structural model through a corresponding analysis method, thereby obtaining the structural deformation amount of each component.

As an optional implementation, the above S103 process may be: interpolating the aerodynamic loads on the components to the structural models corresponding to the components; performing structural finite element static deformation analysis on each interpolated structural model to obtain the structural deformation of each component.

Among them, the computer device can interpolate the aerodynamic loads on each component to the corresponding structural model of each component through the corresponding interpolation algorithm. Optionally, the interpolation algorithm can be a radial basis function interpolation algorithm. The radial basis function interpolation algorithm belongs to a three-dimensional interpolation algorithm, which has the advantages of high precision and strong versatility, and can ensure the accuracy of load interpolation. That is, the computer device can interpolate the aerodynamic loads on each component to the structural model corresponding to each component through the radial basis function interpolation algorithm. Taking each component as the main wing and the leading edge slat as an example, the computer device interpolates the aerodynamic loads on the main wing to the structural model corresponding to the main wing through the radial basis function interpolation algorithm, and interpolates the aerodynamic loads on the leading edge slat to the structural model corresponding to the leading edge slat. For other components, the interpolation of aerodynamic loads can be performed with reference to the main wing method, and this embodiment will not be repeated here.

Specifically, the computer device may interpolate the aerodynamic loads on each component to the structural model corresponding to each component according to the following formula 1.

Formula 1:

Wherein, the above _Fs is the structural load after the aerodynamic load is interpolated and converted, Fa _is the aerodynamic load generated on the component, is the load interpolation parameter.

S104, deforming the aerodynamic surface grid and the aerodynamic space grid of each component according to the structural deformation of each component, and outputting the analysis results of each component when it is determined that the static aeroelastic analysis of each component reaches a preset convergence condition.

Specifically, after obtaining the structural deformation of each component, the computer device needs to transfer the structural deformation of each component to the aerodynamic model corresponding to each component, so that the aerodynamic surface grid and the aerodynamic space grid of the aerodynamic model are deformed.

After deforming the aerodynamic surface mesh and aerodynamic space mesh of each component, the computer device needs to determine whether the static aeroelastic analysis of each component has reached the preset convergence condition. If so, the analysis results of each component are output. If not, the above-mentioned process of S102-S104 is continued until the static aeroelastic analysis of each component reaches the preset convergence condition. Among them, the analysis results may include the corresponding relationship between the aerodynamic load generated on each component and the mesh deformation amount of the three-dimensional body-fitting mesh of each component.

In a specific implementation, the computer device may perform multiple static aeroelastic analyses on each component. When the analysis results of each component tend to be stable, it can be considered that the static aeroelastic analyses of each component have reached a preset convergence condition.

The static aeroelastic analysis method of the aircraft provided by the embodiment of the present application, after generating the three-dimensional body-fitting mesh of each component in the target part, the computer device respectively performs steady flow field analysis on each component to obtain the aerodynamic load generated on each component, determines the structural deformation of each component according to the aerodynamic load on each component, and deforms the aerodynamic surface mesh and aerodynamic space mesh of each component according to the structural deformation of each component, and outputs the analysis results of each component when it is determined that the static aeroelastic analysis of each component reaches the preset convergence condition. In the process of performing static aeroelastic analysis on the low-speed high-lift aircraft configuration, the computer device can automatically generate the three-dimensional body-fitting mesh of each component, and can independently complete the flow field calculation of each component, the data transmission between the aerodynamic load and the structural model, and the data transmission between the structural deformation and the aerodynamic model, that is, the computer device can automatically realize the static aeroelastic analysis of the low-speed high-lift aircraft configuration in the whole process, simplifying the complexity of the static aeroelastic analysis of the low-speed high-lift aircraft configuration and improving the analysis efficiency. At the same time, since the three-dimensional body-fitting meshes of each component generated by the computer equipment are more in line with the actual situation of the low-speed high-lift aircraft configuration, the static aeroelastic analysis of the low-speed high-lift aircraft configuration based on the accurate three-dimensional body-fitting mesh also improves the accuracy of the analysis results.

In one embodiment, a specific process of transferring the structural deformation of each component to the aerodynamic model corresponding to each component is also provided. Based on the above embodiment, optionally, as shown in FIG2 , the above S104 may include:

S201 . Deform the aerodynamic surface mesh of each component according to the structural deformation amount of each component.

After obtaining the structural deformation of each component, the computer device needs to transfer the structural deformation of each component to the aerodynamic surface mesh of each component. As an optional implementation, the computer device can interpolate the structural deformation of each component to the aerodynamic surface mesh of each component through a corresponding interpolation algorithm. The interpolation algorithm can be a radial basis function interpolation algorithm.

In practical applications, since the structural model of the low-speed high-lift aircraft configuration has fewer grid nodes than the aerodynamic model, and components such as the leading edge slats, trailing edge flaps, pylons and nacelles are all in the extrapolated area of the wing box, in order to enable the extrapolated components of the wing to capture the dynamic deformation information, based on the above embodiment, optionally, the process of S201 can be: integrating the aerodynamic surface grids of the components to obtain the target aerodynamic surface grid; and interpolating the structural deformation amounts of the components onto the target aerodynamic surface grid respectively.

Integrating the aerodynamic surface meshes of each component means interpolating the deformation of each component as a whole. Since the radial basis function interpolation algorithm is a three-dimensional interpolation algorithm with the advantages of high precision and strong versatility, the computer device can optionally interpolate the structural deformation of each component onto the target aerodynamic surface mesh through the radial basis function interpolation algorithm.

S202 , determining the mesh deformation amount of the aerodynamic space mesh of each component according to the mesh correspondence and the mesh deformation amount of the aerodynamic surface mesh of each component.

Among them, after the structural deformation of each structural model is interpolated to the aerodynamic surface mesh of the aerodynamic model corresponding to each component, the deformation of the aerodynamic surface mesh will affect the aerodynamic space mesh. Therefore, the computer device predicts the mesh deformation of the aerodynamic space mesh by the mesh deformation of the aerodynamic surface mesh. Optionally, the three-dimensional body-fitting mesh of each component generated by the computer device also includes the mesh correspondence between the aerodynamic surface mesh and the aerodynamic space mesh of each component, that is, the mesh adjacency relationship of the three-dimensional body-fitting mesh of each component is determined. Therefore, the computer device can trace the mesh deformation of each mesh in the aerodynamic space mesh of each component back to the corresponding aerodynamic surface mesh based on the mesh correspondence and the mesh deformation of the aerodynamic surface mesh of each component. Specifically, the computer device can use the traced mesh deformation of the aerodynamic surface mesh as the mesh deformation of the aerodynamic space mesh.

In addition, after the aerodynamic surface mesh of each component is deformed, the computer device can obtain the mesh deformation amount of each mesh in the aerodynamic surface mesh of each component through the following formula 2.

Formula 2: Δr _a = _Cas (C _SS ) ^-1 Δr _s ;

Among them, _Δra is the mesh deformation of the aerodynamic model corresponding to each component, _Δrs is the structural deformation of the structural model corresponding to each component, and _Cas and _Css are deformation interpolation parameters.

S203 , deforming the aerodynamic space grids of the components according to the grid deformation amount of the aerodynamic space grid.

After obtaining the mesh deformation of each mesh in the aerodynamic space mesh of each component, the computer device can deform the aerodynamic space mesh of each component according to the mesh deformation of each mesh, thereby transferring the deformation of the aerodynamic surface mesh of each component to the aerodynamic space mesh, and obtaining the mesh deformation of the aerodynamic surface mesh and the aerodynamic space mesh of each component under the corresponding aerodynamic load. At the same time, when it is determined that the static aeroelastic analysis of each component reaches a preset convergence condition, the computer device can output the mesh deformation of the aerodynamic surface mesh and the aerodynamic space mesh of each component under the corresponding aerodynamic load as the analysis result of the static aeroelastic analysis of each component.

In this embodiment, since the three-dimensional body-fitting grids of each component generated by the computer device include the grid correspondence between the aerodynamic surface grids and the aerodynamic space grids of each component, the computer device can transfer the deformation of the aerodynamic surface grids of each component to the aerodynamic space grid according to the grid correspondence, further simplifying the complexity of the static aeroelastic analysis of the low-speed high-lift aircraft configuration and improving the analysis efficiency. At the same time, during the deformation interpolation process, the computer device integrates the aerodynamic surface grids of each component and interpolates the structural deformation of each component to the integrated target aerodynamic surface grid, so that the extrapolated component of the target component can capture the dynamic deformation information, thereby improving the accuracy of the static aeroelastic analysis.

To facilitate understanding by those skilled in the art, the static aeroelastic analysis method of an aircraft provided by an embodiment of the present application is introduced below by taking a wing as an example of a target component, as shown in FIG. 3 and FIG. 4 . Specifically, the method may include:

S301, respectively generate three-dimensional body-fitting meshes of the slats, main wings, flaps, pylons and nacelles.

Among them, the wings of the low-speed and high-lift aircraft configuration are geometrically divided according to the component type, and the wings are divided into five categories: slats, main wings, flaps, pylons and nacelles.

S302, respectively performing steady flow field analysis on the leading edge slat, the main wing, the trailing edge flap, the pylon and the nacelle to obtain aerodynamic loads generated on the leading edge slat, the main wing, the trailing edge flap, the pylon and the nacelle.

S303, interpolating the aerodynamic loads on the leading edge slat, the main wing, the trailing edge flap, the pylon and the nacelle to the slat structure model, the main wing structure model, the flap structure model, the pylon structure model and the nacelle structure model through a radial basis function interpolation algorithm.

S304, respectively performing structural finite element static deformation analysis on the interpolated slat structure model, the main wing structure model, the flap structure model, the pylon structure model and the nacelle structure model to obtain structural deformation amounts of the slat, the main wing, the flap, the pylon and the nacelle.

S305, interpolating the structural deformations of the slats, main wings, flaps, pylons and nacelles onto the aerodynamic surface grids of the slats, main wings, flaps, pylons and nacelles by using a radial basis function interpolation algorithm.

S306, transferring the mesh deformation on the aerodynamic surface meshes of the slats, main wings, flaps, pylons and nacelles to the aerodynamic space meshes.

Figure 5 is a structural schematic diagram of a static aeroelastic analysis device for an aircraft provided in an embodiment of the present application. The device is applied to a high-lift aircraft configuration in a climbing or landing scenario. As shown in Figure 5, the device may include: a grid generation module 50, a flow field analysis module 51, a deformation determination module 52, a grid deformation module 53 and a result output module 54.

Specifically, the mesh generation module 50 is used to generate a three-dimensional body-fitting mesh of each component in the target component, wherein the three-dimensional body-fitting mesh includes an aerodynamic surface mesh and an aerodynamic space mesh;

The flow field analysis module 51 is used to perform steady flow field analysis on each component to obtain the aerodynamic load generated on each component;

The deformation determination module 52 is used to determine the structural deformation of each component according to the aerodynamic load on each component;

The grid deformation module 53 is used to deform the aerodynamic surface grid and the aerodynamic space grid of each component according to the structural deformation amount of each component;

The result output module 54 is used to output the analysis results of each component when it is determined that the static aeroelastic analysis of each component reaches a preset convergence condition.

The static aeroelastic analysis device for an aircraft provided by the embodiment of the present application, after generating the three-dimensional body-fitting mesh of each component in the target part, the computer device respectively performs steady flow field analysis on each component to obtain the aerodynamic load generated on each component, determines the structural deformation of each component according to the aerodynamic load on each component, and deforms the aerodynamic surface mesh and aerodynamic space mesh of each component according to the structural deformation of each component, and outputs the analysis results of each component when it is determined that the static aeroelastic analysis of each component reaches the preset convergence condition. In the process of performing static aeroelastic analysis on the configuration of a low-speed high-lift aircraft, the computer device can automatically generate the three-dimensional body-fitting mesh of each component, and can independently complete the flow field calculation of each component, the data transmission between the aerodynamic load and the structural model, and the data transmission between the structural deformation and the aerodynamic model, that is, the computer device can automatically realize the static aeroelastic analysis of the configuration of a low-speed high-lift aircraft in the whole process, simplifying the complexity of the static aeroelastic analysis of the configuration of a low-speed high-lift aircraft and improving the analysis efficiency. At the same time, since the three-dimensional body-fitting meshes of each component generated by the computer equipment are more in line with the actual situation of the low-speed high-lift aircraft configuration, the static aeroelastic analysis of the low-speed high-lift aircraft configuration based on the accurate three-dimensional body-fitting mesh also improves the accuracy of the analysis results.

In one of the embodiments, optionally, the deformation determination module 52 includes: a load interpolation unit and a structural deformation analysis unit;

Specifically, the load interpolation unit is used to interpolate the aerodynamic loads on the components to the structural models corresponding to the components;

The structural deformation analysis unit is used to perform structural finite element static deformation analysis on each interpolated structural model to obtain the structural deformation amount of each component.

In one of the embodiments, optionally, the three-dimensional body-fitting grid further includes: a grid correspondence relationship between the aerodynamic surface grid and the aerodynamic space grid;

The grid deformation module 53 includes: a first deformation unit, a determination unit and a second deformation unit;

Specifically, the first deformation unit is used to deform the aerodynamic surface grid of each component according to the structural deformation amount of each component;

The determining unit is used to determine the mesh deformation amount of the aerodynamic space mesh of each component according to the mesh correspondence relationship and the mesh deformation amount of the aerodynamic surface mesh of each component;

The second deformation unit is used to deform the aerodynamic space grid of each component according to the grid deformation amount of the aerodynamic space grid.

In one of the embodiments, optionally, the first deformation unit includes: a grid integration subunit and a deformation interpolation subunit;

Specifically, the grid integration subunit is used to integrate the aerodynamic surface grids of the components to obtain a target aerodynamic surface grid;

The deformation interpolation subunit is used to interpolate the structural deformation of each component onto the target aerodynamic surface grid.

In one of the embodiments, optionally, the load interpolation unit is specifically used to interpolate the aerodynamic loads on the components to the structural models corresponding to the components through a radial basis function interpolation algorithm.

In one of the embodiments, optionally, the deformation interpolation subunit is specifically configured to interpolate the structural deformation of each component onto the target aerodynamic surface grid by using a radial basis function interpolation algorithm.

Optionally, the target component is a wing and/or a tail of the high-lift aircraft configuration.

In one embodiment, a static aeroelasticity analysis device for an aircraft (hereinafter referred to as a computer device) is provided, and the internal structure diagram of the computer device can be shown in Figure 6. The computer device includes a processor, a memory, a network interface, a display screen and an input device connected through a system bus. Among them, the processor of the computer device is used to provide computing and control capabilities. The memory of the computer device includes a non-volatile storage medium and an internal memory. The non-volatile storage medium stores an operating system and a computer program. The internal memory provides an environment for the operation of the operating system and the computer program in the non-volatile storage medium. The network interface of the computer device is used to communicate with an external terminal through a network connection. When the computer program is executed by the processor, a static aeroelasticity analysis method for an aircraft is implemented. The display screen of the computer device can be a liquid crystal display screen or an electronic ink display screen, and the input device of the computer device can be a touch layer covered on the display screen, or a key, trackball or touchpad set on the computer device housing, or an external keyboard, touchpad or mouse, etc.

Those skilled in the art will understand that the structure shown in FIG. 6 is merely a block diagram of a partial structure related to the solution of the present application, and does not constitute a limitation on the computer device to which the solution of the present application is applied. The specific computer device may include more or fewer components than those shown in the figure, or combine certain components, or have a different arrangement of components.

In one embodiment, a static aeroelastic analysis device for an aircraft is provided, which is applied to a high-lift aircraft configuration in a climbing or landing scenario. The device includes a memory and a processor, wherein a computer program is stored in the memory, and when the processor executes the computer program, the following steps are implemented:

Generate a three-dimensional body-fitted mesh of each component in the target component, wherein the three-dimensional body-fitted mesh includes an aerodynamic surface mesh and an aerodynamic space mesh;

Performing steady flow field analysis on each of the components respectively to obtain aerodynamic loads generated on each of the components;

Determining the structural deformation of each component according to the aerodynamic load on each component;

According to the structural deformation of each component, the aerodynamic surface grid and the aerodynamic space grid of each component are deformed, and when it is determined that the static aeroelastic analysis of each component reaches a preset convergence condition, the analysis results of each component are output.

In one embodiment, when the processor executes the computer program, the following steps are also implemented: the aerodynamic loads on the components are interpolated to the structural models corresponding to the components; and structural finite element static deformation analysis is performed on each interpolated structural model to obtain the structural deformation of each component.

In one embodiment, the three-dimensional body-fitting grid further includes: a grid correspondence relationship between the aerodynamic surface grid and the aerodynamic space grid; when the processor executes the computer program, the following steps are also implemented: according to the structural deformation amount of each component, the aerodynamic surface grid of each component is deformed; according to the grid correspondence relationship and the grid deformation amount of the aerodynamic surface grid of each component, the grid deformation amount of the aerodynamic space grid of each component is determined; according to the grid deformation amount of the aerodynamic space grid, the aerodynamic space grid of each component is deformed.

In one embodiment, when the processor executes the computer program, the following steps are further implemented: integrating the aerodynamic surface meshes of the components to obtain a target aerodynamic surface mesh; and interpolating the structural deformation of the components onto the target aerodynamic surface mesh respectively.

In one embodiment, when the processor executes the computer program, the following steps are further implemented: the aerodynamic loads on the components are interpolated to the structural models corresponding to the components by using a radial basis function interpolation algorithm.

In one embodiment, when the processor executes the computer program, the following steps are further implemented: interpolating the structural deformation of each component onto the target aerodynamic surface grid respectively by using a radial basis function interpolation algorithm.

Optionally, the target component is a wing and/or a tail of the high-lift aircraft configuration.

In one embodiment, a computer readable storage medium is provided, which is applied to a high lift aircraft configuration in a climbing or landing scenario, and stores a computer program thereon, and when the computer program is executed by a processor, the following steps are implemented:

Generate a three-dimensional body-fitted mesh of each component in the target component, wherein the three-dimensional body-fitted mesh includes an aerodynamic surface mesh and an aerodynamic space mesh;

Performing steady flow field analysis on each of the components respectively to obtain aerodynamic loads generated on each of the components;

Determining the structural deformation of each component according to the aerodynamic load on each component;

According to the structural deformation of each component, the aerodynamic surface grid and the aerodynamic space grid of each component are deformed, and when it is determined that the static aeroelastic analysis of each component reaches a preset convergence condition, the analysis results of each component are output.

In one embodiment, when the computer program is executed by the processor, the following steps are also implemented: the aerodynamic loads on the components are interpolated to the structural models corresponding to the components; and structural finite element static deformation analysis is performed on each interpolated structural model to obtain the structural deformation of each component.

In one embodiment, the three-dimensional body-fitting grid further includes: a grid correspondence relationship between the aerodynamic surface grid and the aerodynamic space grid; when the computer program is executed by the processor, the following steps are also implemented: according to the structural deformation amount of each component, the aerodynamic surface grid of each component is deformed; according to the grid correspondence relationship and the grid deformation amount of the aerodynamic surface grid of each component, the grid deformation amount of the aerodynamic space grid of each component is determined; according to the grid deformation amount of the aerodynamic space grid, the aerodynamic space grid of each component is deformed.

In one embodiment, when the computer program is executed by the processor, the following steps are further implemented: integrating the aerodynamic surface meshes of the components to obtain a target aerodynamic surface mesh; and interpolating the structural deformation of the components onto the target aerodynamic surface mesh respectively.

In one embodiment, when the computer program is executed by the processor, the following steps are further implemented: the aerodynamic loads on the components are interpolated to the structural models corresponding to the components by using a radial basis function interpolation algorithm.

In one embodiment, when the computer program is executed by the processor, the following steps are further implemented: interpolating the structural deformation of each component onto the target aerodynamic surface grid respectively by using a radial basis function interpolation algorithm.

Optionally, the target component is a wing and/or a tail of the high-lift aircraft configuration.

The static aeroelasticity analysis device, equipment and storage medium of the aircraft provided in the above embodiments can execute the static aeroelasticity analysis method of the aircraft provided in any embodiment of the present application, and have the corresponding functional modules and beneficial effects of executing the method. For technical details not described in detail in the above embodiments, please refer to the static aeroelasticity analysis method of the aircraft provided in any embodiment of the present application.

Those skilled in the art can understand that all or part of the processes in the above-mentioned embodiment methods can be completed by instructing the relevant hardware through a computer program, and the computer program can be stored in a non-volatile computer-readable storage medium. When the computer program is executed, it can include the processes of the embodiments of the above-mentioned methods. Among them, any reference to memory, storage, database or other media used in the embodiments provided in the present 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. As an illustration and not limitation, RAM is available in many forms, such as static RAM (SRAM), dynamic RAM (DRAM), synchronous DRAM (SDRAM), double data rate SDRAM (DDRSDRAM), enhanced SDRAM (ESDRAM), synchronous link (Synchlink) DRAM (SLDRAM), memory bus (Rambus) direct RAM (RDRAM), direct memory bus dynamic RAM (DRDRAM), and memory bus dynamic RAM (RDRAM).

The technical features of the above-described embodiments may be arbitrarily combined. To make the description concise, not all possible combinations of the technical features in the above-described 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.

The above-mentioned embodiments only express several implementation methods of the present application, and the descriptions thereof are relatively specific and detailed, but they cannot be understood as limiting the scope of the present application. It should be pointed out that, for a person of ordinary skill in the art, several variations and improvements can be made without departing from the concept of the present application, and these all belong to the protection scope of the present application. Therefore, the protection scope of the present application shall be subject to the attached claims.

Claims (8)

1. A method of static aeroelastic analysis of an aircraft, for application to a high lift aircraft configuration in a climb or landing scenario, the method comprising:

Generating a three-dimensional body-attached grid of each component in the target component, wherein the three-dimensional body-attached grid comprises a pneumatic surface grid and a pneumatic space grid; the target component is a wing of a high lift aircraft configuration; the components in the target part are a leading edge slat, a main wing, a trailing edge flap, a pylon and a nacelle;

respectively carrying out steady flow field analysis on each component to obtain pneumatic loads generated on each component;

determining the structural deformation of each component according to the pneumatic load on each component;

Deforming the pneumatic surface grid and the pneumatic space grid of each component according to the structural deformation of each component, and outputting the analysis result of each component when the static pneumatic elasticity analysis of each component is determined to reach a preset convergence condition;

The three-dimensional body-contacting grid further comprises: grid correspondence between the aerodynamic surface grid and the aerodynamic spatial grid;

The method for deforming the pneumatic surface grid and the pneumatic space grid of each component according to the structural deformation of each component comprises the following steps:

deforming the pneumatic surface grid of each component according to the structural deformation of each component;

determining the grid deformation of the pneumatic space grid of each component according to the grid corresponding relation and the grid deformation of the pneumatic surface grid of each component;

and deforming the pneumatic space grid of each component according to the grid deformation of the pneumatic space grid.

2. The method of claim 1, wherein said determining the amount of structural deformation of each of said components based on the aerodynamic load on said components comprises:

interpolating pneumatic loads on the components correspondingly to structural models corresponding to the components;

and respectively carrying out structural finite element static deformation analysis on each interpolated structural model to obtain the structural deformation of each component.

3. The method of claim 1, wherein deforming the aerodynamic surface mesh of each component according to the structural deformation of each component comprises:

integrating the pneumatic surface grids of the components to obtain a target pneumatic surface grid;

and respectively interpolating the structural deformation of each component to the target pneumatic surface grid.

4. The method of claim 2, wherein said interpolating pneumatic load correspondences on said components to structural models corresponding to said components comprises:

And interpolating the pneumatic load on each component to the structural model corresponding to each component through a radial basis function interpolation algorithm.

5. A method according to claim 3, wherein said separately interpolating structural deformations of said components onto said target aerodynamic surface mesh comprises:

And respectively interpolating the structural deformation of each component onto the target pneumatic surface grid through a radial basis function interpolation algorithm.

6. A static aeroelastic analysis device for an aircraft, for application in a high lift aircraft configuration in a climb or landing scenario, the device comprising:

The grid generation module is used for generating a three-dimensional body-attached grid of each component in the target component, wherein the three-dimensional body-attached grid comprises a pneumatic surface grid and a pneumatic space grid; the target component is a wing of a high lift aircraft configuration; the components in the target part are a leading edge slat, a main wing, a trailing edge flap, a pylon and a nacelle;

the flow field analysis module is used for respectively carrying out steady flow field analysis on each component to obtain pneumatic loads generated on each component;

The deformation determining module is used for determining the structural deformation of each component according to the pneumatic load on each component;

the grid deformation module is used for deforming the pneumatic surface grid and the pneumatic space grid of each component according to the structural deformation of each component;

The result output module is used for outputting the analysis result of each component when the static pneumatic elasticity analysis of each component is determined to reach the preset convergence condition;

The three-dimensional body-contacting grid further comprises: grid correspondence between the aerodynamic surface grid and the aerodynamic spatial grid;

the grid deformation module includes: a first deforming unit, a determining unit and a second deforming unit;

the first deforming unit is used for deforming the pneumatic surface grid of each component according to the structural deformation of each component;

the determining unit is used for determining the grid deformation of the pneumatic space grid of each component according to the grid corresponding relation and the grid deformation of the pneumatic surface grid of each component;

The second deforming unit is used for deforming the pneumatic space grid of each component according to the grid deformation of the pneumatic space grid.

7. A static pneumatic elasticity analysis device of an aircraft, comprising a memory and a processor, the memory storing a computer program, characterized in that the processor implements the steps of the method of any one of claims 1 to 5 when the computer program is executed.

8. A computer readable storage medium, on which a computer program is stored, characterized in that the computer program, when being executed by a processor, implements the steps of the method according to any one of claims 1 to 5.