Method and system for setting forced assembly force limit value of aviation composite thin-wall structure

Abstract

The invention provides a method and a system for setting a forced assembly force limit value of an aviation composite thin-wall structure. The method comprises the following steps: S1, obtaining an influence rule of a forced assembly force on internal assembly stress and damage of the composite thin-wall structure; s2, a parametric modeling mode is adopted to express an influence rule, assembly geometry and physical performance indexes are considered cooperatively, and a multi-objective optimization model of forced assembly force parameters is established; s3, solving the assembly force limit value mathematical model, and obtaining a forced assembly force limit value setting interval; and S4, evaluating the effectiveness of assembly force limit value setting by adopting a mode of combining a test with simulation result parameterized modeling, and visually displaying a forced assembly performance result. According to the method, the mapping relation model of the assembly force and the assembly quality is rapidly established in a parametric modeling mode, assembly force limitation setting and rapid visual display of the assembly quality under cooperative guarantee of assembly geometry and physical performance can be achieved, and excessive assembly deformation and structural damage or destruction are avoided.

Description

[Technical field]

[0001] The invention relates to the technical field of quality control of aviation composite structure assembly engineering, and in particular to a method and system for setting a forced assembly force limit value of an aviation composite thin-wall structure. [Background technology]

[0002] In the assembly scenarios of large-scale weak-rigidity structural parts and components of aviation composites, even tiny dimensional deviations, shape and position errors, or surface warping and deformation caused by material properties and manufacturing processes may result in the inability to fit naturally and accurately between components. At this time, it is necessary to intervene with external clamping force to cause the components to undergo necessary displacement or deformation adjustments to ensure that the final assembly not only meets the design specifications in terms of geometric dimensions, but also meets the high standards of high strength, lightness, and durability in terms of physical properties. By forcing the application of external assembly forces, eliminating the fitting clearances between the components of composite thin-walled components, correcting deviations, and forcibly completing precise positioning and clamping tasks, the application of this technology is indispensable for ensuring the aerodynamic shape, structural stability, and safety of the aircraft.

[0003] However, during the forced assembly process, excessive assembly force may cause internal assembly stress between structures, increase the risk of damage or destruction of composite materials, and affect the mechanical service performance of the assembled structure; while if the assembly force is too small, the expected assembly tightness may not be achieved, the fit gap between structural parts cannot be eliminated, and the overall aerodynamic shape performance of the structure may be affected. Therefore, how to accurately set the assembly force limit and develop a scientific and reasonable assembly force limit setting technology is crucial to ensure the assembly quality of thin-walled structures, improve assembly efficiency, and maintain the safety of long-term operation of aircraft. Due to the great difficulties in the study of the formation mechanism of assembly quality of anisotropic composite parts and the lack of scientific modeling design, there are few studies on assembly force limit methods in China. For the practical application of assembly force limit, there is a lack of forced clamping force standard specifications for field application. Most companies rely on hand feeling to apply forced clamping force, and only a few companies refer to foreign specifications to guide actual operations. For example, Boeing's composite parts assembly clamping force limit is 45 pounds / foot, and domestic OEMs set it to 45N / 300mm or 50N / 300mm. However, foreign research on specific processes and methods in this regard has not been published due to confidentiality.

[0004] In this context, this technology came into being. It aims to develop a force limit setting method and system suitable for the forced assembly of aviation composite thin-walled structures through in-depth research on the relationship between the mechanical properties of composite materials, forced assembly process parameters and structural geometry and mechanical response, so as to achieve precise control of the clamping force during the assembly process, avoid geometric shape deviations, excessive internal stress and structural damage during the assembly process, and ensure that the assembly of thin-walled structures is both efficient and safe to meet the high standards of the modern aviation industry.

[0005] Therefore, it is necessary to study a method and system for setting the forced assembly force limit of an aviation composite thin-walled structure to address the deficiencies of the prior art and to solve or alleviate one or more of the above problems. [Summary of the invention]

[0006] In view of this, the present invention provides a method and system for setting the limit of forced assembly force of thin-walled aviation composite structures. The present invention analyzes the influence of forced assembly force on internal damage and assembly stress of thin-walled aviation composites. Due to the structural complexity and material characteristics of thin-walled aviation composite parts, they are prone to deformation during the assembly process, affecting the performance and safety of the final product. If the thin-walled parts are improperly stressed during assembly, they may cause structural damage or destruction due to excessive internal assembly stress. The cost of composite materials is high. Reasonable control of assembly force can avoid material damage and maximize material utilization efficiency. Taking into account factors such as damage, clearance, and assembly force application space, a feasible method for reasonable setting of forced assembly force limits is provided to solve the problems of difficulty and low efficiency in setting assembly force limits.

[0007] In one aspect, the present invention provides a method for setting a limit value of a forced assembly force of an aviation composite thin-walled structure, the method comprising the following steps:

[0008] S1: The method of combining experiment with finite element is used to compare and modify the finite element simulation model and parameters to obtain the influence of assembly force on the internal assembly stress and damage of composite thin-walled structures;

[0009] S2: Use parametric modeling to express the influencing rules, consider the assembly geometry and physical performance indicators in a coordinated manner, perform multi-objective optimization on the forced assembly force parameters of the composite thin-walled structure, and obtain the mathematical model of the composite thin-walled assembly force limit;

[0010] S3: Use intelligent optimization algorithm to solve the mathematical model of assembly force limit of composite thin-walled structure, and obtain the setting range of forced assembly force limit under the coordinated guarantee of assembly geometry and physical performance;

[0011] S4: The effectiveness of setting assembly force limits is evaluated by combining experiments with parametric modeling of simulation results, and the forced assembly geometry and physical performance results are visualized.

[0012] According to the above aspects and any possible implementation, an implementation is further provided, wherein S1 specifically includes:

[0013] S11: Install six-dimensional force sensors at key locations of the flexible clamping unit to record and monitor changes in clamping force during forced assembly in real time;

[0014] S12: By combining terahertz, DIC and ultrasonic measurement technologies, the three work together to dynamically monitor and accurately evaluate the stress and damage state of composite thin-walled structures during forced assembly;

[0015] S13: According to the actual assembly process, obtain the response value of the change of the forced assembly force to the internal assembly stress distribution and damage degree of the composite thin wall;

[0016] S14: Establish a finite element model according to the actual operation process, perform finite element simulation analysis and verification, and compare and correct the finite element simulation model and simulation parameters with the test result data;

[0017] S15: Systematically adjust the size of the assembly force, use finite element simulation to quickly and accurately obtain the mechanical state of the clamping force on the internal structure of the assembly, analyze the assembly surface accuracy, internal stress and damage distribution of the composite wall panel structure, and obtain the influence of the assembly force on the internal damage of the composite thin wall and the assembly stress.

[0018] According to the aspects and any possible implementation methods described above, an implementation method is further provided, wherein S13 is specifically: designing an aviation composite thin-wall assembly test, and gradually adjusting the assembly force acting on the surface of the aviation composite thin-wall according to a regular pattern through a control variable method, and recording in real time the changes in the internal assembly stress and damage state of the composite thin-wall as the assembly force applied to the surface changes.

[0019] According to the aspects and any possible implementations described above, an implementation is further provided, wherein the influence law in S15 is the influence mechanism of the assembly force on the internal microstructure damage state of the aviation composite thin wall and the influence law of the macroscopic internal stress distribution.

[0020] According to the above aspects and any possible implementation manner, an implementation manner is further provided, wherein S2 specifically includes:

[0021] S21: Parameterize the influence law between the forced assembly force obtained by the finite element method verified by the experiment and the thin-wall stress and damage of the aviation composite material, and obtain the input and output corresponding table relationship between the forced assembly force and the assembly quality effect under a given assembly process scheme;

[0022] S22: Consider assembly geometry and physical performance indicators in a coordinated manner, define optimization design variables, including the position coordinates of the forced assembly force and the size of the forced assembly force limit range; set constraints: the forced assembly force is applied within the effective space range, and damage to the thin-walled structure of the aviation composite is not allowed; set target variables: optimize the size of the forced assembly force to minimize the internal stress and damage of the thin-walled aviation composite, and at the same time minimize the height of the geometric fit clearance between parts;

[0023] S23: Based on engineering requirements, multi-objective optimization is carried out on the forced assembly force parameters of composite thin-walled structures, and a mathematical optimization model for the assembly force limit of composite thin-walled structures is constructed.

[0024] According to the above aspects and any possible implementation, an implementation is further provided, wherein S3 specifically includes:

[0025] S31: Using a multi-objective intelligent optimization algorithm, an initial population is created, which consists of multiple different assembly force limit schemes, and the quality of each scheme is evaluated through a fitness function;

[0026] S32: Based on the evaluation results of the fitness function, the high-quality features are mixed to create a better assembly force limit solution, and some features of some individuals in the new population are randomly changed to maintain the diversity of the population and prevent it from falling into the local optimum too early;

[0027] S3: Repeat the genetic operation steps to continuously generate assembly force limit solutions until an assembly force limit design that is superior to other solutions is found, or the preset upper limit of the number of iterations is reached.

[0028] According to the aspects and any possible implementation methods described above, an implementation method is further provided, wherein the fitness function in S31 evaluates the pros and cons of each scheme specifically as follows: a fitness score is assigned to each assembly force limit scheme through a fitness function of the internal stress, damage degree and geometric gap thickness of the aviation composite thin-walled structure to quantify the performance of the scheme.

[0029] According to the above aspects and any possible implementation manner, an implementation manner is further provided, wherein S4 specifically includes:

[0030] S41: Based on the obtained assembly force limit range and spatial position, a series of positioning points are selected at regular intervals to perform an assembly force limit verification test;

[0031] S42: Based on the required assembly force limit range and spatial position, the optimized assembly force limit scheme is applied to the finite element model based on the cohesive force unit for rapid parametric verification;

[0032] S43: Use the finite element model to simulate the stress-strain state between composite thin-walled structural parts, and intuitively visualize the reduction of assembly geometric gaps and possible assembly damage forms;

[0033] S44: Compare and analyze the simulation prediction and measured data to verify the reliability of the model and the rationality of the assembly force limit.

[0034] According to the aspects and any possible implementation methods described above, an implementation method is further provided, wherein S44 is specifically: if the degree of agreement between the simulation data and the measured data is higher than a preset threshold, it indicates that the simulation process of the assembly geometry and physical performance results is accurate; if the results of both are less than the preset threshold, then return to S41 again.

[0035] According to the above aspects and any possible implementation, a system for setting a limit value of forced assembly force of an aviation composite thin-walled structure is further provided, and the system for setting a limit value of forced assembly force of an aviation composite thin-walled structure comprises:

[0036] The influence law acquisition module is used to obtain the influence law of the forced assembly force on the internal damage of the composite thin wall and the assembly stress state by correcting the finite element simulation model and parameters through test results;

[0037] Model building and optimization module, which is used to model and optimize the parametric expression of the forced assembly force of composite thin-walled structures based on the influence law, and to obtain the mathematical model of the composite thin-walled assembly force limit by considering the assembly geometry and physical performance indicators in a coordinated manner;

[0038] A model solving module is used to solve the mathematical model of the assembly force limit of the composite thin-walled structure based on an intelligent optimization algorithm to obtain the setting range of the assembly force limit;

[0039] The visualization and evaluation module is used to visually display the reduction of assembly geometric clearance and possible assembly damage forms, and evaluate the effectiveness of setting assembly force limits by combining experiments with finite element simulation.

[0040] Compared with the prior art, the present invention can achieve the following technical effects:

[0041] 1) The method of verifying and correcting the finite element simulation model using experimental data and theoretical calculation data can accurately obtain the influence of forced assembly force on the internal assembly stress and damage of aviation composite thin-walled structures.

[0042] 2) The influence law between the forced assembly force obtained by finite element method verified by experiment and the stress and damage of thin-walled aerospace composites is parametrically modeled, and a mapping relationship model can be quickly established; based on this, the effective spatial range to which the forced assembly force can be applied and the damage threshold allowed by the aerospace composite thin-wall structure are optimized, and the size of the forced assembly force is adjusted to minimize the assembly damage of the aerospace composite thin-wall structure, while minimizing the gap thickness, and constructing a mathematical optimization model of the relationship between the forced assembly force, gap thickness and damage.

[0043] 3) Through multi-objective optimization and solution, the maximum and minimum values ​​of the assembly force limit are obtained under the constraints of minimum gap thickness, minimum damage degree, and forced assembly force spatial coordinates applied within the effective space range, thereby confirming the assembly force limit range under the coordination of assembly geometric accuracy and assembly physical performance indicators.

[0044] 4) The forced assembly force-stress / damage relationship is tested and visualized by combining finite element simulation with actual measurement to determine the rationality of setting the forced assembly force limit while ensuring assembly quality.

[0045] Of course, any product implementing the present invention does not necessarily need to achieve all of the above-mentioned technical effects at the same time.

Brief Description of the Drawings

[0046] In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0047] Figure 1 This is a technical step diagram for setting the limit of forced assembly force of an aviation composite thin-wall structure according to the present invention;

[0048] Figure 2 is a technical flow chart of setting the forced assembly force limit provided by an embodiment of the present invention;

[0049] Figure 3 This is a description of the structure and adjustable functions of a typical box segment assembly tool of the present invention;

[0050] Figure 4 It is an exploded view of the overall assembly structure of a certain type of composite material wing box provided by an embodiment of the present invention. [Specific implementation method]

[0051] In order to better understand the technical solution of the present invention, the embodiments of the present invention are described in detail below with reference to the accompanying drawings.

[0052] It should be clear that the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0053] The terms used in the embodiments of the present invention are only for the purpose of describing specific embodiments, and are not intended to limit the present invention. The singular forms "a", "said" and "the" used in the embodiments of the present invention and the appended claims are also intended to include plural forms, unless the context clearly indicates other meanings.

[0054] The present invention provides a method for setting a limit value of forced assembly force of an aviation composite thin-wall structure. Figure 2 The specific implementation flow chart of this solution in this embodiment is as follows. The overall idea adopted by the present invention to solve its technical problem is:

[0055] Firstly, the key indicators such as clamping force, stress and damage state of aviation composite thin-walled assembly structures during the assembly process are monitored and experimentally studied: force sensors are installed at the flexible clamping units on the surface of the clamped composite thin-walled materials to collect and monitor in real time the changing data of the assembly force acting on the surface of the aviation composite thin-walled structural parts; the stress and damage state of the composite thin-walled structure during the forced assembly process are dynamically monitored and accurately evaluated by combining terahertz, DIC and ultrasonic measurement technologies; the effects of different assembly force sizes on the internal stress distribution of the aviation composite thin-walled assembly structure and the degree of damage that may be caused are explored through experiments, and on this basis, a finite element model is established based on the actual assembly process of aviation composite thin-walled structures, and the accuracy of the finite element simulation is verified and corrected using experimental data. At the same time, the influence of the assembly force acting on the composite thin-walled material on the internal damage and stress distribution of the composite thin-walled material is obtained through simulation. Secondly, the influence law between the forced assembly force obtained by the finite element method verified by the experiment and the stress and damage of the thin-walled aerospace composites is parameterized to obtain the input and output corresponding table relationship between the forced assembly force and the assembly quality effect under a given assembly process scheme, so as to define the optimization design variables: the position coordinates and size of the assembly force, set reasonable constraints: the assembly force must be applied within the effective space range and not cause damage to the thin-walled composite material, set the optimization target variable: by optimizing the size of the assembly force, it can not only minimize the internal stress and damage of the composite thin-walled structure assembly, but also ensure the minimum assembly gap thickness, and construct a mathematical optimization model for the assembly force limit of the composite thin-walled structure. Thirdly, a multi-objective intelligent optimization algorithm is used for optimization search: a population containing multiple assembly force limit schemes is initialized, and the advantages and disadvantages of each scheme are evaluated through the fitness function. The excellent scheme is selected as the "parent", and crossover and mutation operations are performed to generate a new assembly force limit scheme. At the same time, randomness is introduced to maintain population diversity and avoid local optimization convergence. Iteration continues until the current assembly force limit scheme is found to be significantly better than other schemes or the preset maximum number of iterations is reached. Finally, within the determined range of assembly force limit and spatial position, values ​​are taken at intervals and the series of positioning point values ​​taken are applied to the test verification of the assembly force limit of thin-walled aerospace composites. The assembly geometry and physical performance status of thin-walled aerospace composite structures are calculated through finite element simulation, and the assembly effects such as gap elimination rate, stress-strain distribution, and damage detection are intuitively displayed. The reliability of the model and the rationality of the assembly force limit are verified by comparing the finite element simulation data with the actual measurement data.

[0056] The present invention provides a setting technology for the assembly force limit applied by forced positioning clamping of an aviation composite thin-walled structure. The forced assembly force limit setting technology can be applied in the assembly stage of the aviation composite thin-walled structure, such as optimizing the influence of the tooling clamping force on the internal assembly stress and damage of the composite thin-walled structure, and improving the ability of the forced assembly force to eliminate the assembly gap. Figure 1As shown, the forced assembly force setting technology includes the following steps:

[0057] S1: The experimental method combined with the finite element method is used to compare and modify the finite element simulation model and parameters to accurately obtain the influence of assembly force on the internal assembly stress and damage of composite thin-walled structures.

[0058] Specifically, a six-dimensional force sensor is first installed at the key contact point of the clamping unit on the surface of the aviation composite thin-walled material to grasp the force applied by the flexible clamping unit on the aviation composite thin-walled structural parts during the forced assembly process, and to detect and record the changes in the clamping force acting on the composite thin-walled structural parts in real time.

[0059] Furthermore, by combining terahertz, DIC and ultrasonic measurement technologies, the three work together to dynamically monitor and accurately evaluate the stress and damage state of composite thin-walled structures during forced assembly.

[0060] Furthermore, according to the actual assembly process, an aviation composite thin-wall assembly test was designed. Through the control variable method, the assembly force acting on the surface of the aviation composite thin-wall was gradually adjusted according to the law, and the changes in the internal assembly stress and damage state of the composite thin-wall as the assembly force applied to the surface were recorded in real time.

[0061] Furthermore, a finite element model is established based on the actual operation process of aviation composite thin-wall assembly to simulate the entire composite thin-wall assembly process. The data obtained from the above-mentioned composite thin-wall assembly test is compared and analyzed with the data generated by the finite element simulation. The finite element simulation model and simulation parameters are compared and corrected with the result data obtained from the test to ensure the accuracy and effectiveness of the finite element model.

[0062] Furthermore, through finite element simulation software, the force transmission path and stress distribution pattern in the composite thin-wall assembly process are obtained, and the influence mechanism of assembly force on the internal microstructural damage state of aviation composite thin-wall and the influence law of macroscopic internal stress distribution are analyzed.

[0063] Furthermore, the assembly force is systematically adjusted, and the finite element simulation is used to quickly and accurately obtain the mechanical state of the internal structure of the assembly due to the clamping force. The assembly surface accuracy, internal stress and damage distribution of structural parts such as composite wall panels are analyzed to obtain the influence of assembly force on the internal damage of composite thin walls and assembly stress.

[0064] S2: The influencing rules are expressed by parametric modeling, and the assembly geometry and physical performance indicators are considered synergistically. Multi-objective optimization is carried out on the forced assembly force parameters of composite thin-walled structures to obtain the mathematical model of the composite thin-walled assembly force limit.

[0065] Specifically, the influence law between the forced assembly force obtained by finite element method verified by experiment and the thin-wall stress and damage of aviation composite materials is parameterized to obtain the input and output corresponding tabular relationship between the forced assembly force and the assembly quality effect under a given assembly process scheme, and to clarify the quantitative relationship between the forced assembly force and geometric deviation, stress and damage.

[0066] Furthermore, the assembly geometry and physical performance indicators are considered in a coordinated manner to define the set of design variables X that need to be optimized in the assembly process of aviation composite thin-walled materials, including the key factors that determine the assembly effect: the coordinate position U where the forced assembly force is applied, the limit value F of the assembly force acting on the surface of the aviation composite thin-walled material, and the gap thickness H between the fitting surfaces of the composite thin-walled assembly structure after assembly; considering the actual engineering limitations and safety requirements, set constraints: the forced assembly force must be applied within the effective working space range L, that is, U∈L, and no damage to the aviation composite thin-walled assembly structure is allowed, that is, S(X)<1; establish the optimization goal: minimize the degree of damage to the composite thin-walled structure caused by assembly: min{S(X)}, and minimize the gap thickness H of the composite thin-walled assembly structure after assembly: min{H}.

[0067] Furthermore, based on engineering requirements, the quantitative relationship between the assembly force of thin-walled aviation composites and geometric gap tolerance, stress, and damage is studied, and a mathematical optimization model of the relationship between forced assembly force, gap thickness, and damage is constructed:

[0068] f1(X)={S(X)<1, min{F, H, S(X)}, U∈L}

[0069] f2(X)={S(X)<1, max{F}, min{H, S(X)}, U∈L}

[0070] S3: Use intelligent optimization algorithm to solve the mathematical model of assembly force limit of composite thin-walled structure, and obtain the setting range of forced assembly force limit under the coordinated guarantee of assembly geometry and physical properties.

[0071] Specifically, a multi-objective intelligent optimization algorithm is used to initialize a population containing multiple assembly force limit schemes, each individual representing a possible scheme. A fitness score is assigned to each assembly force limit scheme through a fitness function that comprehensively considers factors such as the internal stress, damage degree, and geometric gap thickness of the aerospace composite thin-walled structure, and the quality of each scheme is evaluated through the fitness function.

[0072] Furthermore, a solution with higher fitness is selected from the current population of assembly force limit configurations as the "parent". By performing a crossover operation on the "parent" solution, key features such as assembly force limit, assembly force application space, and gap thickness of the fitting surface are extracted from the preferred solution and combined into a new offspring solution. Combining the advantages of the "parent" solution, a new generation of solutions with potentially higher fitness is created. In this process, the characteristic parameters of some individuals are randomly changed to increase the probability of searching for the global optimal solution.

[0073] Furthermore, the above genetic operation process is repeated to continuously iterate and generate new assembly force limit solutions. This process will continue until an assembly force limit design that is significantly better than all previous solutions is found, or the algorithm stops when the maximum number of iterations is reached.

[0074] S4: The effectiveness of setting assembly force limits is accurately evaluated by combining experiments with parametric modeling of simulation results, and the forced assembly geometry and physical performance results are visualized.

[0075] Specifically, based on the obtained assembly force limit range and spatial position, according to specific spacing rules, a series of assembly forces are selected in the obtained assembly force limit range, and a series of positioning points are selected at regular intervals to carry out assembly force limit verification tests, which are applied to aviation composite thin-wall assembly structures respectively. The internal assembly stress, geometric gap and possible damage degree of the aviation composite thin-wall are obtained by measuring, and the correctness of the obtained assembly force limit scheme is verified.

[0076] Furthermore, based on the required assembly force limit range and spatial action position, the optimized assembly force limit scheme is applied to the finite element model based on the cohesive force unit for rapid parametric verification.

[0077] Furthermore, the finite element model is called to simulate the stress-strain state between composite thin-walled structural parts, and the finite element analysis software is used to visualize the gap elimination rate, stress-strain curve, and whether there are risk areas that may cause material damage, intuitively displaying the assembly geometry gap reduction and possible assembly damage forms.

[0078] Furthermore, the data obtained from the finite element simulation is compared with the data measured in the actual assembly test. If the simulated data and the measured data are highly consistent, it means that the simulation process is accurate and reliable, thus obtaining the forced assembly force limit that should be followed in actual engineering applications. If the results of both do not meet the engineering requirements, return to repeat the above operation.

[0079] The present invention provides a setting technology for the assembly force limit applied by the forced positioning clamping of aviation composite thin-wall structures, which provides an effective solution to the problems of optimizing the influence of the size, position and other parameters of the forced assembly force limit setting on the internal stress and damage of the composite thin-wall assembly during the aviation composite thin-wall assembly stage, and analyzing the assembly geometry deviation.

[0080] The present invention mainly includes two parts, namely: to optimize and adjust the size of the forced assembly force to minimize the internal stress and damage distribution of the thin-walled aviation composite material and minimize the gap between the internal fitting surfaces of the assembly structure, and to construct a multi-objective mathematical optimization model for assembly force limitation with the effective spatial range to which the forced assembly force can be applied and the damage threshold allowed by the thin-walled aviation composite material structure as constraints; then, the model is optimized and solved and verified based on an intelligent optimization algorithm. The present invention can achieve:

[0081] 1) The method of verifying and correcting the finite element simulation model using experimental data and theoretical calculation data can accurately obtain the influence of forced assembly force on the internal assembly stress and damage of aviation composite thin-walled structures.

[0082] 2) The influence law between the forced assembly force obtained by finite element method verified by experiment and the stress and damage of thin-walled aerospace composites is parametrically modeled, and a mapping relationship model can be quickly established; based on this, the effective spatial range to which the forced assembly force can be applied and the damage threshold allowed by the aerospace composite thin-wall structure are optimized, and the size of the forced assembly force is adjusted to minimize the assembly damage of the aerospace composite thin-wall structure, while minimizing the gap thickness, and constructing a mathematical optimization model of the relationship between the forced assembly force, gap thickness and damage.

[0083] 3) Through multi-objective optimization and solution, the maximum and minimum values ​​of the assembly force limit are obtained under the constraints of minimum gap thickness, minimum damage degree, and forced assembly force spatial coordinates applied within the effective space range, thereby confirming the assembly force limit range under the coordination of assembly geometric accuracy and assembly physical performance indicators.

[0084] 4) The forced assembly force-stress / damage relationship is tested and visualized by combining finite element simulation with actual measurement to determine the rationality of setting the forced assembly force limit while ensuring assembly quality.

[0085] Embodiment 1:

[0086] In this embodiment, a certain type of composite material wing box component assembly application is taken as an example ( Figure 4). First, a controllable assembly force is applied to the composite wing box model through the designed tooling system to reduce the assembly gap between the components, and the changes in stress and damage state in the assembly process are monitored by combining six-dimensional force sensors, terahertz, DIC and ultrasonic measurement technology. Secondly, the finite element software ABAQUS is used to simulate the assembly process in three dimensions and correct the simulation model and parameters to verify the measured data, ensure the accuracy of the rapid prediction of the wing box assembly performance results, and use parametric modeling to clarify the quantitative relationship between the forced assembly force and geometric tolerance, stress and damage. On this basis, combined with engineering constraints, a multi-objective optimization design method is used to establish an assembly force limit optimization model aimed at synergistically reducing damage and minimizing gaps. Subsequently, an intelligent algorithm is used to solve the optimization model, generate and iteratively improve the assembly force limit scheme, in order to find the optimal assembly strategy. Finally, through testing and finite element simulation analysis on a scaled model, the obtained assembly force limit scheme is verified and the assembly performance results are visualized to ensure its effectiveness and safety, thereby providing a scientific basis and technical guidance for the precision assembly of thin-walled structures of aviation composite materials. If the verification result does not meet the requirements, it is necessary to return to the optimization design step for adjustment to form a closed-loop iterative optimization process. The forced assembly force setting technology for the upper and lower wall panels of a certain type of aircraft composite wing box includes the following steps:

[0087] Specifically, by Figure 3 The tooling shown clamps the composite wing box frame into position, and then uses the positioner to fit the upper and lower composite wall panels to the frame beams and wall parts. Due to manufacturing and processing errors, there will be a gap between the composite wall panels and the frame fitting surface. To reduce the gap, a certain clamping force is applied through the positioner. To record and control the clamping force in real time, a six-dimensional force sensor is installed at the key contact point of the clamping unit.

[0088] Furthermore, a scaled-down model of the composite wing box is designed, and assembly force is applied on the composite surface according to a certain rule to reduce the gap. The applied assembly force is obtained through a positioner. Each time a set of assembly force is applied, the changes in internal stress and damage state during the assembly process are monitored by combining terahertz, DIC and ultrasonic measurement technologies.

[0089] Furthermore, through finite element three-dimensional modeling, the wing box was imported into the ABAQUS finite element simulation analysis according to the actual situation. By setting the actual parameters such as the wing box material, the size and position of the assembly force applied to the composite surface, the entire upper and lower wall panel assembly process was simulated. The line graph between the assembly force applied during the assembly process and the internal assembly stress of the wing box was obtained through post-processing, and the relationship between the two was obtained. The data obtained from the assembly test was compared and analyzed to carry out the correction of the finite element simulation model and parameters to ensure the accuracy and effectiveness of the finite element model.

[0090] Furthermore, by setting the output in the finite element simulation software, the influence of the assembly force on the microstructural damage inside the overall frame of the composite wing box and the influence law of the macroscopic stress distribution during the assembly of the upper and lower wall panels are obtained, and the stress and damage on the composite wing box wall panels and the frame structure under the application of assembly force are displayed through cloud maps.

[0091] Furthermore, the influence of the obtained assembly force on the microstructural damage inside the overall frame of the composite wing box and the influence law of the macroscopic stress distribution are parametrically modeled, the stress and damage degree at each location are expressed as a percentage, and the parametric modeling method is used to clarify the quantitative relationship between the forced assembly force and the geometric tolerance, stress and damage.

[0092] Furthermore, based on the quantitative relationship between the assembly force, stress, and damage obtained above, and considering the actual engineering limitations and safety requirements, constraints are set: the forced assembly force must be applied within the effective working space range L, that is, U∈L, and damage to the composite wing box is not allowed, that is, S(X)<1; the optimization goal is established: to minimize the degree of damage to the composite wing box caused by assembly: min{S(X)}, and to minimize the gap thickness H between the composite wall panel and the frame after assembly: min{H}. A mathematical optimization model for the relationship between the maximum and minimum values ​​allowed for the forced assembly force, gap thickness, and damage is constructed:

[0093] f1(X)={S(X)<1, min{F, H, S(X)}, U∈L}

[0094] f2(X)={S(X)<1, max{F}, min{H, S(X)}, U∈L}

[0095] Where U is the coordinate position of the forced assembly force, F is the magnitude of the applied assembly force, and H is the thickness of the gap between the bonding surfaces.

[0096] Furthermore, a multi-objective intelligent optimization algorithm is adopted to first initialize a population containing multiple assembly force limit schemes. The performance of the assembly force limit scheme is quantified through a fitness function that comprehensively considers factors such as the stress distribution and damage degree in the composite wing box and the gap thickness of the bonding surface.

[0097] Furthermore, a solution with higher fitness is selected from the current assembly force limit solution as the "parent". By performing a crossover operation on the "parent" solution, key features such as assembly force limit, assembly force application space, and gap thickness of the fitting surface are extracted from the preferred solution and combined into a new child solution. Combining the advantages of the "parent" solution, a new generation of solutions with potentially higher fitness is created. In this process, the characteristic parameters of some individuals are randomly changed to increase the probability of searching for the global optimal solution.

[0098] Furthermore, the above genetic operation process is repeated to continuously iterate and generate new assembly force limit solutions. This process will continue until an assembly force limit design that is significantly better than all previous solutions is found, or the algorithm stops when the maximum number of iterations is reached.

[0099] Furthermore, the obtained assembly force limit range and spatial position are taken at regular intervals, and the obtained assembly force series values ​​are applied to a scaled composite wing box model for testing. The correctness of the obtained assembly force limit scheme is verified by measuring the internal assembly stress of the composite wing box, the geometric gap of the fitting surface, and the possible degree of damage to the wing box structure corresponding to each applied assembly force.

[0100] Furthermore, based on the desired assembly force limit scheme, ABAQUS is called to apply the desired assembly force limit scheme to the upper and lower panels of the composite wing box. By constructing a finite element wing box model of cohesive force units, the mechanical response of the actual assembly process of aviation composite thin-walled materials is simulated. With the help of ABAQUS, each assembly force limit scheme is observed, and the visualization of the gap elimination rate of the wing box fitting surface and the stress-strain curve, as well as whether there are risk areas that may cause material damage, are used to intuitively display the reduction of the geometric gap of the wing box components and possible assembly damage forms such as delamination and crushing.

[0101] Furthermore, the assembly geometry and physical performance data obtained from the finite element simulation are compared with the data measured in the actual assembly test. If the simulated data and the measured data are highly consistent, it indicates that the simulation process is accurate and reliable, thereby obtaining the forced assembly force limit that should be followed in actual engineering applications. If the results of both do not meet the engineering requirements, return to repeat the above operation.

[0102] The present invention provides a technology for setting the limit of forced assembly force of thin-walled aviation composite structures, which mainly includes two parts: optimizing and adjusting the size of forced assembly force to minimize the internal stress and damage of the assembly of thin-walled aviation composite structures and minimize the gap between the fitting surfaces, and constraining the effective spatial range in which the forced assembly force can be applied and the damage threshold allowed by the thin-walled aviation composites, to construct a multi-objective mathematical optimization model; then, using an intelligent optimization algorithm to solve the model, and verifying the effectiveness of the calculated assembly force limit results. It can achieve: 1. Combining measured data with theoretical modeling calculation data, through finite element simulation models and parameter correction methods, quickly obtain the influence of forced assembly force parameters on the internal structure damage and assembly stress of thin-walled composites. 2. Using parametric modeling processing, quickly establish a mapping relationship model between forced assembly force and stress and damage of thin-walled aviation composites, and through collaborative analysis of assembly geometry and physical properties, based on the engineering requirements of thin-walled composite damage and gap thickness, establish a mathematical model between forced assembly force and internal stress, internal damage, and gap thickness of thin-walled composite structures. ③ Through the multi-objective intelligent optimization solution, the assembly force limit range under the coordinated protection of assembly geometric accuracy and assembly physical performance indicators is determined, while reducing structural damage and failures caused by improper assembly, and realizing the accuracy inspection and visualization of the forced assembly force-stress / damage relationship. ④ A set of standardized force limit setting procedures for the forced clamping link is proposed.

[0103] The above is a detailed introduction to a method and system for setting the limit of forced assembly force of thin-walled composite structures of aviation provided by the embodiment of the present application. The description of the above embodiment is only used to help understand the method and core idea of ​​the present application; at the same time, for those skilled in the art, according to the idea of ​​the present application, there will be changes in the specific implementation method and application scope. In summary, the content of this specification should not be understood as a limitation on the present application.

[0104] For example, certain words are used in the specification and claims to refer to specific components. Those skilled in the art should understand that hardware manufacturers may use different nouns to refer to the same component. This specification and claims do not use differences in names as a way to distinguish components, but use differences in the functions of components as the criteria for distinction. As mentioned throughout the specification and claims, "including" and "comprising" are open-ended terms, so they should be interpreted as "including / including but not limited to". "Approximately" means that within an acceptable error range, those skilled in the art can solve the technical problem within a certain error range and basically achieve the technical effect. The subsequent description of the specification is a preferred embodiment of the present application, but the description is for the purpose of illustrating the general principles of the present application, and is not used to limit the scope of the present application. The scope of protection of the present application shall be determined by the definition of the attached claims.

[0105] It should also be noted that the terms "include", "comprises" or any other variations thereof are intended to cover non-exclusive inclusion, so that a product or system including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such a product or system. In the absence of more restrictions, the elements defined by the sentence "comprises a ..." do not exclude the existence of other identical elements in the product or system including the elements.

[0106] It should be understood that the term "and / or" used in this article is only a description of the association relationship of associated objects, indicating that there can be three relationships. For example, A and / or B can represent: A exists alone, A and B exist at the same time, and B exists alone. In addition, the character " / " in this article generally indicates that the associated objects before and after are in an "or" relationship.

[0107] The above description shows and describes several preferred embodiments of the present application, but as mentioned above, it should be understood that the present application is not limited to the form disclosed herein, and should not be regarded as excluding other embodiments, but can be used in various other combinations, modifications and environments, and can be modified within the scope of the application concept described herein through the above teachings or the technology or knowledge of the relevant field. The changes and modifications made by those skilled in the art do not deviate from the spirit and scope of the present application, and should be within the scope of protection of the claims attached to the present application.

Claims

1. A method for setting the limit of forced assembly force of thin-walled aviation composite structures, characterized in that: The method for setting the forced assembly force limit of the aviation composite thin-walled structure comprises the following steps: S1: The method of combining experiment with finite element is used to compare and modify the finite element simulation model and parameters to obtain the influence of assembly force on the internal assembly stress and damage of composite thin-walled structures; S2: Use parametric modeling to express the influencing rules, consider the assembly geometry and physical performance indicators in a coordinated manner, perform multi-objective optimization on the forced assembly force parameters of the composite thin-walled structure, and obtain the mathematical model of the composite thin-walled assembly force limit; S3: Use intelligent optimization algorithm to solve the mathematical model of assembly force limit of composite thin-walled structure, and obtain the setting range of forced assembly force limit under the coordinated guarantee of assembly geometry and physical performance; S4: The effectiveness of setting assembly force limits is evaluated by combining experiments with parametric modeling of simulation results, and the forced assembly geometry and physical performance results are visualized.

2. The method for setting the forced assembly force limit of an aviation composite thin-walled structure according to claim 1, characterized in that: The S1 specifically includes: S11: Install six-dimensional force sensors at key locations of the flexible clamping unit to record and monitor changes in clamping force during forced assembly in real time; S12: By combining terahertz, DIC and ultrasonic measurement technologies, the three work together to dynamically monitor and evaluate the stress and damage state of composite thin-walled structures during forced assembly; S13: According to the actual assembly process, obtain the response value of the change of the forced assembly force to the internal assembly stress distribution and damage degree of the composite thin wall; S14: Establish a finite element model according to the actual operation process, perform finite element simulation analysis and verification, and compare and correct the finite element simulation model and simulation parameters with the test result data; S15: Systematically adjust the size of the assembly force, use finite element simulation to quickly and accurately obtain the mechanical state of the clamping force on the internal structure of the assembly, analyze the assembly surface accuracy, internal stress and damage distribution of the composite wall panel structure, and obtain the influence of the assembly force on the internal damage of the composite thin wall and the assembly stress.

3. The method for setting the forced assembly force limit of an aviation composite thin-walled structure according to claim 1, characterized in that: The S13 is specifically: designing an aviation composite thin-wall assembly test, and by controlling the variable method, gradually adjusting the size of the assembly force acting on the surface of the aviation composite thin-wall according to a rule, and recording in real time the change rule of the internal assembly stress and damage state of the composite thin-wall with the assembly force applied to the surface.

4. The method for setting the forced assembly force limit of an aviation composite thin-walled structure according to claim 2, characterized in that: The influence law in S15 is the influence mechanism of assembly force on the internal microstructural damage state of aviation composite thin-wall and the influence law of macroscopic internal stress distribution.

5. The method for setting the forced assembly force limit of an aviation composite thin-walled structure according to claim 1, characterized in that: The S2 specifically includes: S21: Parameterize the influence law between the forced assembly force obtained by the finite element method verified by the experiment and the thin-wall stress and damage of the aviation composite material, and obtain the input and output corresponding table relationship between the forced assembly force and the assembly quality effect under a given assembly process scheme; S22: Consider assembly geometry and physical performance indicators in a coordinated manner, define optimization design variables, including the position coordinates of the forced assembly force and the size of the forced assembly force limit range; set constraints: the forced assembly force is applied within the effective space range, and damage to the thin-walled structure of the aviation composite is not allowed; set target variables: optimize the size of the forced assembly force to minimize the internal stress and damage of the thin-walled aviation composite, and at the same time minimize the height of the geometric clearance between parts; S23: Based on engineering requirements, multi-objective optimization is carried out on the forced assembly force parameters of composite thin-walled structures, and a mathematical optimization model for the assembly force limit of composite thin-walled structures is constructed.

6. The method for setting the forced assembly force limit of an aviation composite thin-walled structure according to claim 5, characterized in that: The S3 specifically includes: S31: Using a multi-objective intelligent optimization algorithm, an initial population is created, which consists of multiple different assembly force limit schemes, and the quality of each scheme is evaluated through a fitness function; S32: Based on the evaluation results of the fitness function, the high-quality features are mixed to create a better assembly force limit solution, and some features of some individuals in the new population are randomly changed to maintain the diversity of the population and prevent it from falling into the local optimum too early; S3: Repeat the genetic operation steps to continuously generate assembly force limit solutions until an assembly force limit design that is superior to other solutions is found, or the preset upper limit of the number of iterations is reached.

7. The method for setting the forced assembly force limit of an aviation composite thin-walled structure according to claim 6, characterized in that: The fitness function in S31 evaluates the quality of each solution specifically as follows: a fitness score is assigned to each assembly force limit solution through the fitness function of the internal stress, damage degree and geometric gap thickness of the aviation composite thin-wall structure to quantify the performance of the solution.

8. The method for setting the forced assembly force limit of an aviation composite thin-walled structure according to claim 6, characterized in that: The S4 specifically includes: S41: Based on the obtained assembly force limit range and spatial position, a series of positioning points are selected at regular intervals to perform an assembly force limit verification test; S42: Based on the required assembly force limit range and spatial position, the optimized assembly force limit scheme is applied to the finite element model based on the cohesive force unit for rapid parametric verification; S43: Use the finite element model to simulate the stress-strain state between composite thin-walled structural parts, and intuitively visualize the reduction of assembly geometric gaps and possible assembly damage forms; S44: Compare and analyze the simulation prediction and measured data to verify the reliability of the model and the rationality of the assembly force limit.

9. The method for setting the forced assembly force limit of an aviation composite thin-walled structure according to claim 8, characterized in that: The S44 is specifically as follows: if the degree of agreement between the simulation data and the measured data is higher than a preset threshold, it indicates that the simulation process of the assembly geometry and the physical performance results is accurate; if the results of both are lower than the preset threshold, the process returns to S41 again.

10. A system for setting the limit of forced assembly force of thin-walled composite aviation structures, characterized in that: The forced assembly force limit setting system for aviation composite thin-walled structures includes: The influence law acquisition module is used to obtain the influence law of the forced assembly force on the internal damage of the composite thin wall and the assembly stress state by correcting the finite element simulation model and parameters through test results; Model building and optimization module, which is used to model and optimize the parametric expression of the forced assembly force of composite thin-walled structures based on the influence law, and to obtain the mathematical model of the composite thin-walled assembly force limit by considering the assembly geometry and physical performance indicators in a coordinated manner; A model solving module is used to solve the mathematical model of the assembly force limit of the composite thin-walled structure based on an intelligent optimization algorithm to obtain the setting range of the assembly force limit; The visualization and evaluation module is used to visually display the reduction of assembly geometric clearance and possible assembly damage forms, and evaluate the effectiveness of setting assembly force limits by combining experiments with finite element simulation.

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