Aircraft structure optimization method based on minimum structure weight and suitable for high-Mach-number high-temperature environment

Abstract

The invention discloses an aircraft structure optimization method based on the minimum structure weight and suitable for a high-Mach-number high-temperature environment, and belongs to the technical field of aircraft structure optimization. The problem that in the prior art, according to a traditional aircraft structure optimization method, it is difficult to enable an aircraft to have the light structure heat protection effect under the condition that the aerodynamic characteristics and the long endurance performance are met is solved. According to the method, parametric modeling is carried out to obtain a structure optimization model; calculating the volume of a model design variable as an optimization target of the structure optimization model; obtaining a constraint variable result file of each grid node; extracting the maximum value of a constraint variable; parallel computing is carried out on sample points of the design space, and a minimum optimization result meeting constraint conditions in the design space is obtained; and outputting design variable value data corresponding to the minimum optimization result. According to the method, the lightest structural weight of the high-Mach-number aircraft under the constraint conditions of thermal stress, allowable stress and strain can be obtained, and the method can be applied to aircraft design.

Description

Technical Field

[0001] The present invention relates to an aircraft structure optimization method, in particular to an aircraft structure optimization method based on minimum structural weight and suitable for high Mach number and high temperature environments, belonging to the technical field of aircraft structure optimization. Background Art

[0002] As humans' pursuit of speed continues to increase, various types of high-Mach number aircraft continue to emerge. Their flight characteristics with a wide speed range and large envelope also face many challenges. Among the aerodynamic problems faced by high-Mach number flight, the biggest difference between high-Mach number aircraft and low-Mach number aircraft is the aerodynamic heating phenomenon that cannot be ignored. When a high-Mach number aircraft flies in the atmosphere, the air in front of the aircraft is highly compressed and rubs strongly with the surrounding air. The strong airflow disturbance generates shock waves. After passing through the shock wave, the kinetic energy of the gas is converted into internal energy, generating high temperature in the shock layer to form a "thermal barrier". That is, the aerodynamic heating phenomenon, the "thermal barrier" will heat the aircraft structure. The generation of thermal stress and the attenuation of the thermophysical parameters of the aircraft structure bring greater challenges to the aircraft structure. It is required that high Mach number aircraft must meet the aerodynamic characteristics and long flight performance while also considering the lightweight structure's heat protection capabilities and structural safety. The aircraft structure in the existing technology is difficult to meet the above requirements at the same time. Therefore, it is necessary to conduct in-depth research on the structural characteristics of high Mach number aircraft in high temperature environments, and it is necessary to explore the aerodynamic design, power design and structural design that are most suitable for the aircraft's flight envelope.

[0003] In summary, there is a need for an aircraft structure optimization method based on minimum structural weight and suitable for high Mach number and high temperature environments. Summary of the Invention

[0004] A brief overview of the present invention is provided below to provide a basic understanding of certain aspects of the present invention. It should be understood that this overview is not an exhaustive overview of the present invention. It is not intended to identify key or important aspects of the present invention, nor is it intended to limit the scope of the present invention. Its purpose is simply to present certain concepts in a simplified form as a prelude to the more detailed description discussed later.

[0005] In view of this, in order to solve the problem in the prior art that traditional aircraft structure optimization methods are difficult to make the aircraft have a lightweight structure and heat protection effect while meeting aerodynamic characteristics and long flight time performance, the present invention provides an aircraft structure optimization method based on minimum structural weight and suitable for high Mach number and high temperature environment.

[0006] The technical solution is as follows: A method for optimizing an aircraft structure based on minimum structural weight and suitable for high Mach number and high temperature environments comprises the following steps:

[0007] S1. Parametrically model the model in the modeling software, determine the upper and lower limits and initial values ​​of the model design variables based on the optimization problem, and obtain the structural optimization model;

[0008] S2. Calculate the volume of the model design variable as the optimization target of the structural optimization model, perform minimum optimization on the volume, and obtain the optimization result;

[0009] S3. Use finite element processing software and finite element analysis software to perform finite element pre- and post-processing on the structural optimization model to obtain constraint variable result files for each mesh node;

[0010] S4. Extract the maximum stress and strain values ​​from the result file;

[0011] S5. Based on the structural optimization model and optimization results, perform parallel calculations on the design space sample points, set optimization constraints based on the maximum stress and strain values, and obtain the minimum optimization result under the optimization constraints;

[0012] S6. Based on the minimum optimization result that meets the constraints in the design space, the structural optimization model outputs the design variable value data corresponding to the minimum optimization result.

[0013] Furthermore, in said S1, the research object is parametrically modeled in the modeling software, the number of research objects is selected according to the needs, the research objects are numbered, the position range of each research object is determined as the upper and lower limits of the model design variables in the subsequent structural optimization model, and the established model, i.e., the structural optimization model, is output in the stp format.

[0014] Furthermore, in S2, the calculated volume of the research object under the current research object position combination is used as the optimization target of the structural optimization model. When the material parameters are determined, the smallest volume of the research object means the lightest structure weight, and the smallest volume of the research object is used as the optimization result.

[0015] Furthermore, the step S3 includes the following steps:

[0016] S31. Read the structural optimization model in stp format, mesh the structural optimization model, and form mesh nodes;

[0017] S32. Create corresponding material parameters for different components of the aircraft;

[0018] S33. Create a unit type for the structural optimization model and apply material properties to the structural optimization model;

[0019] S34. Given boundary conditions and after removing the research object, apply unit forces of arbitrary magnitude to the grid nodes and output the pre-processing results.

[0020] S35. Use fluid simulation software to calculate the full-model aerodynamic forces of the aircraft and output the results as a six-column file formatted as "grid node coordinates + aerodynamic forces";

[0021] S36. Interpolate the unit force of the grid node, interpolate the fluid force file calculated by the fluid simulation software into the result file obtained by pre-processing through the rbf interpolation function, and obtain the finite element result file under the actual force conditions of the aircraft;

[0022] S37. Calculate the modified aerodynamic data using finite element analysis software to obtain a finite element file, and obtain a result file containing node stress and strain.

[0023] Furthermore, the step S5 includes the following steps:

[0024] S51. Select the position of the research object as the design variable, and determine the upper and lower limits and initial values ​​of the position change of the research object according to the optimization requirements;

[0025] S52. Using the volume of the research object calculated in step S2 as the optimization target, obtaining an optimization result;

[0026] S53. Setting optimization constraints, including that the maximum stress extracted in step S4 does not exceed the allowable stress of the material, and the maximum strain extracted in step S4 does not exceed the allowable strain of the material;

[0027] S54. Setting an optimization algorithm through the structural optimization model according to the optimization constraints, that is, selecting proxy optimization through the single-objective optimization method, inputting the number of initial sample points, and selecting the EI method as the point addition criterion;

[0028] S55. Based on the optimization results, the structural optimization model is solved to obtain the minimum optimization result that meets the constraints within the design space.

[0029] Furthermore, in S6, steps S1-S4 are used as a calculation process for a sample point, and the sample points in the design space are solved and calculated. After the calculation is completed, the structural optimization model selects a new sample point as the position input of step S1, performs a cyclic calculation, and finally obtains the research object position corresponding to the minimum research object volume that meets the constraint conditions within the design range of the optimization variable.

[0030] The beneficial effects of the present invention are as follows: the present invention provides an aircraft structure optimization method based on minimum structural weight and suitable for high Mach number and high temperature environments. Under the condition of determining the overall aircraft model and material parameters, the lightest structural weight of the high Mach number aircraft under the thermal stress, allowable stress and strain constraints is obtained, making an important contribution to reducing the structural weight of the aircraft, reducing fuel consumption and improving flight efficiency; the present invention can achieve that the optimized aircraft structure in a high Mach number and high temperature environment can make the stress and strain of the surface skin of the aircraft less than the allowable stress and allowable strain while reducing the weight of the aircraft as much as possible, thereby meeting the safety indicators. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] The drawings described herein are used to provide a further understanding of the present invention and constitute a part of the present invention. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation of the present invention. In the drawings:

[0032] Figure 1 The figure is a flow chart of an aircraft structure optimization method based on minimum structural weight and suitable for high Mach number and high temperature environment;

[0033] Figure 2 A logical diagram of the aircraft structure optimization method;

[0034] Figure 3 The present invention is a schematic flow chart of an embodiment of an aircraft structure optimization method based on minimum structural weight and suitable for high Mach number and high temperature environments;

[0035] Figure 4 Schematic diagram of data transmission for aircraft structure optimization method. DETAILED DESCRIPTION

[0036] To make the technical solutions and advantages of the embodiments of the present invention more clearly understood, exemplary embodiments of the present invention are further described in detail below with reference to the accompanying drawings. It should be noted that the embodiments described are only a portion of the embodiments of the present invention, and are not an exhaustive list of all embodiments. It should be noted that the embodiments of the present invention and the features thereof may be combined with each other unless they conflict.

[0037] refer to Figure 1-4 Detailed description of this embodiment, a method for optimizing an aircraft structure based on minimum structural weight and suitable for a high Mach number and high temperature environment, specifically comprising the following steps:

[0038] S1. Perform parameterized modeling on the model, determine the upper and lower limits and initial values ​​of the model design variables according to the optimization problem, and obtain the structural optimization model;

[0039] S2. Calculate the volume of the model design variable as the optimization target of the structural optimization model, perform minimum optimization on the volume, and obtain the optimization result;

[0040] S3. Use finite element processing software and finite element analysis software to perform finite element pre- and post-processing on the structural optimization model to obtain constraint variable result files for each mesh node;

[0041] S4. Extract the maximum stress and strain values ​​from the result file;

[0042] S5. Based on the structural optimization model and optimization results, perform parallel calculations on the design space sample points, set optimization constraints based on the maximum stress and strain values, and obtain the minimum optimization result under the optimization constraints;

[0043] S6. Based on the minimum optimization result that meets the constraints in the design space, the structural optimization model outputs the design variable value data corresponding to the minimum optimization result.

[0044] Furthermore, in said S1, the fuselage annular ribs are parametrically modeled in the modeling software, the number of the fuselage annular ribs is selected according to the requirements, the fuselage annular ribs are numbered, the position sequence of the fuselage annular ribs is in ascending order and cannot be interspersed with each other, the position range of each fuselage annular rib is determined as the upper and lower limits of the model design variables in the subsequent structural optimization model, and the established model, i.e., the structural optimization model, is output in the STP format.

[0045] Furthermore, it is characterized in that, in said S2, the volume of the fuselage annular rib under the current fuselage annular rib position combination obtained by calculation is used as the optimization target of the structural optimization model. When the material parameters are determined, the minimum volume of the fuselage annular rib represents the lightest structural weight, and the minimum volume of the fuselage annular rib is used as the optimization result.

[0046] Furthermore, it is characterized in that said S3 includes the following steps:

[0047] S31. Read the structural optimization model in stp format, mesh the structural optimization model, and form mesh nodes;

[0048] S32. Create corresponding material parameters for different components of the aircraft, ie, create material parameters for the fuselage and the annular ribs respectively;

[0049] S33. Create a unit type for the structural optimization model and apply material properties to the structural optimization model;

[0050] S34. Given boundary conditions, after removing the fuselage annular ribs, apply unit forces of arbitrary magnitude to the mesh nodes and output the pre-processing results.

[0051] S35. Use fluid simulation software to calculate the full-model aerodynamic forces of the aircraft and output the results as a six-column file formatted as "grid node coordinates + aerodynamic forces";

[0052] S36. Interpolate the unit force of the grid node, interpolate the fluid force file calculated by the fluid simulation software into the result file obtained by pre-processing through the rbf interpolation function, and obtain the finite element result file under the actual force conditions of the aircraft;

[0053] S37. Calculate the modified aerodynamic data using finite element analysis software to obtain a finite element file, and obtain a result file containing node stress and strain.

[0054] Furthermore, the step S5 includes the following steps:

[0055] S51. Select the fuselage annular rib position as the design variable and determine the upper and lower limits and initial value of the fuselage annular rib position change according to the optimization requirements;

[0056] S52. The volume of the fuselage annular rib calculated in step S2 is used as the optimization target to obtain the optimization result;

[0057] S53. Setting optimization constraints, including that the maximum stress extracted in step S4 does not exceed the allowable stress of the material, and the maximum strain extracted in step S4 does not exceed the allowable strain of the material;

[0058] S54. Set an optimization algorithm based on the structural optimization model according to the optimization constraints. Specifically, select proxy optimization using a single-objective optimization method. To reduce computation time, use parallel computing as the computation method. Enter the number of initial sample points. Select the EI method as the point addition criterion.

[0059] S55. Based on the optimization results, the structural optimization model is solved to obtain the minimum optimization result that meets the optimization constraints within the design space.

[0060] Specifically, the single-objective optimization method changes the position of the fuselage annular ribs to ensure that the structural mass is the lightest under the constraint condition that the stress and deformation are lower than the allowable stress and strain.

[0061] Furthermore, in S6, steps S1-S4 are used as a calculation process for a sample point, and the sample points in the design space are solved and calculated. After the calculation is completed, the structural optimization model selects a new sample point as the position input of step S1, and performs a cyclic calculation to finally obtain the fuselage annular rib position corresponding to the minimum fuselage annular rib volume that meets the constraint conditions within the design range of the optimization variables.

[0062] Specifically, steps S1-S4 are the calculation process of the maximum stress and maximum strain of a single sample point during the optimization process. The process needs to be automated through a batch file to ensure that the calculation can be completed automatically when the input value of the optimization software changes;

[0063] Constraints are selected based on the calculation requirements. For example, static constraints can be set to ensure that the maximum stress and maximum strain are less than the allowable stress and required strain, and dynamic stability conditions can be set to ensure that flutter requirements are met.

[0064] When pre-processing the structural optimization model in the software, it is necessary to consider the impact of thermal stress on the material properties of the model;

[0065] The aerodynamic forces at the mesh nodes of the structural optimization model only need to be calculated once, and the same aerodynamic force result file is used in subsequent interpolations;

[0066] During the optimization process, select the design variables that need to be optimized, set the upper and lower limits of the optimization variables according to the optimization problem, and in order to reduce the amount of calculation, the closer the initial value is to the optimal value, the better;

[0067] Select the calculated volume of the design variable as the optimization target, select the target minimum optimization, and the weight coefficient is 1. After the material parameters are given, the minimum volume means the minimum structural weight;

[0068] The constraints of allowable stress and allowable strain need to be satisfied at the same time. The output data after the optimization process is the design variable value of the minimum volume that meets the constraints.

[0069] Although the present invention has been described with respect to a limited number of embodiments, it will be apparent to those skilled in the art, having benefit of the foregoing description, that other embodiments are contemplated within the scope of the invention thus described. Furthermore, it should be noted that the language used in this specification has been selected primarily for readability and didactic purposes, rather than for the purpose of explaining or limiting the subject matter of the present invention. Consequently, many modifications and variations will be apparent to those skilled in the art without departing from the scope and spirit of the appended claims. The disclosure of the present invention is intended to be illustrative rather than restrictive of the scope of the invention, which is defined by the appended claims.

Claims

1. A method for optimizing aircraft structure based on minimum structural weight and suitable for high Mach number and high temperature environments, characterized in that: The following steps are involved: S1. Parametrically model the model in the modeling software, determine the upper and lower limits and initial values ​​of the model design variables based on the optimization problem, and obtain the structural optimization model; S2. Calculate the volume of the model design variable as the optimization target of the structural optimization model, perform minimum optimization on the volume, and obtain the optimization result; S3. Use finite element processing software and finite element analysis software to perform finite element pre- and post-processing on the structural optimization model to obtain constraint variable result files for each mesh node; S4. Extract the maximum stress and strain values ​​from the result file; S5. Based on the structural optimization model and optimization results, perform parallel calculations on the design space sample points, set optimization constraints based on the maximum stress and strain values, and obtain the minimum optimization result under the optimization constraints; S6. Based on the minimum optimization result that meets the constraints in the design space, the structural optimization model outputs the design variable value data corresponding to the minimum optimization result.

2. The aircraft structure optimization method based on minimum structural weight and suitable for high Mach number and high temperature environment according to claim 1, characterized in that: In said S1, parametric modeling is performed on the research object in the modeling software, the number of research objects is selected according to the needs, the research objects are numbered, the position range of each research object is determined as the upper and lower limits of the model design variables in the subsequent structural optimization model, and the established model, i.e., the structural optimization model, is output in the stp format.

3. The aircraft structure optimization method based on minimum structural weight and suitable for high Mach number and high temperature environment according to claim 2, characterized in that: In S2, the calculated volume of the research object under the current research object position combination is used as the optimization target of the structural optimization model. When the material parameters are determined, the smallest volume of the research object represents the lightest structure weight, and the smallest volume of the research object is used as the optimization result.

4. The aircraft structure optimization method based on minimum structural weight and suitable for high Mach number and high temperature environment according to claim 3, characterized in that: Said S3 comprises the following steps: S31. Read the structural optimization model in stp format, mesh the structural optimization model, and form mesh nodes; S32. Create corresponding material parameters for different components of the aircraft; S33. Create a unit type for the structural optimization model and apply material properties to the structural optimization model; S34. Given boundary conditions and after removing the research object, apply unit forces of arbitrary magnitude to the grid nodes and output the pre-processing results. S35. Use fluid simulation software to calculate the full-model aerodynamic forces of the aircraft and output the results as a six-column file formatted as "grid node coordinates + aerodynamic forces"; S36. Interpolate the unit force of the grid node, interpolate the fluid force file calculated by the fluid simulation software into the result file obtained by pre-processing through the rbf interpolation function, and obtain the finite element result file under the actual force conditions of the aircraft; S37. Calculate the modified aerodynamic data using finite element analysis software to obtain a finite element file, and obtain a result file containing node stress and strain.

5. The aircraft structure optimization method based on minimum structural weight and suitable for high Mach number and high temperature environment according to claim 4, characterized in that: The S5 comprises the following steps: S51. Select the position of the research object as the design variable, and determine the upper and lower limits and initial values ​​of the position change of the research object according to the optimization requirements; S52. Using the volume of the research object calculated in step S2 as the optimization target, obtaining an optimization result; S53. Setting optimization constraints, including that the maximum stress extracted in step S4 does not exceed the allowable stress of the material, and the maximum strain extracted in step S4 does not exceed the allowable strain of the material; S54. Setting an optimization algorithm through the structural optimization model according to the optimization constraints, that is, selecting proxy optimization through the single-objective optimization method, inputting the number of initial sample points, and selecting the EI method as the point addition criterion; S55. Based on the optimization results, the structural optimization model is solved to obtain the minimum optimization result that meets the constraints within the design space.

6. The aircraft structure optimization method based on minimum structural weight and suitable for high Mach number and high temperature environment according to claim 5, characterized in that: In S6, steps S1-S4 are used as a calculation process for a sample point, and the sample points in the design space are solved and calculated. After the calculation is completed, the structural optimization model selects a new sample point as the position input of step S1, and performs a cyclic calculation to finally obtain the research object position corresponding to the minimum research object volume that meets the constraint conditions within the design range of the optimization variable.

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