Optimization method for motion deformation of aircraft assembly tool

Abstract

The invention discloses an optimization method for motion deformation of an aircraft assembly tool. The optimization method comprises the following steps: step 1, establishing a finite element geometric model of an initial aircraft assembly tool; 2, determining a tool motion state constraint condition, and calculating first five-order free vibration modes of the tool under the motion state constraint condition; 3, extracting the first two orders of effective free vibration mode frequencies to calculate the tool damping, and assigning the tool damping to the finite element geometric model; 4, motion deformation distribution of the tool under the motion load is calculated, and a motion deformation model of the initial aircraft assembly tool is established; 5, a deformation limiting structure is additionally arranged on the initial aircraft assembly tool, and an aircraft assembly tool optimized structure is formed; and step 6, executing the step 2 to the step 4 on the aircraft assembly tool optimization model, calculating deformation distribution of the aircraft assembly tool optimization structure under the motion load, and evaluating an optimization effect of the added deformation limiting structure on the initial aircraft assembly tool motion deformation.

Description

Technical Field

[0001] This invention relates to, but is not limited to, the field of deformation of aircraft assembly tooling, and specifically to an optimization method for the motion deformation of aircraft assembly tooling. Background Technology

[0002] With the development of intelligent aircraft production lines, there are more and more movement scenarios for aircraft assembly tooling, especially assembly tooling with its own wheels or AGV carrying movement, which has appeared in various assembly production lines and assembly processes.

[0003] Unlike tooling, which has stable load-bearing capacity and boundary conditions when stationary, tooling undergoes significant deformation during motion due to factors such as factory floor construction, tooling start-up and shutdown, and other random stimuli. This deformation can affect aircraft assembly quality. Therefore, limiting the deformation of aircraft assembly tooling during motion and ensuring product assembly stability is an urgent problem to be solved. Summary of the Invention

[0004] The purpose of this invention is to solve the above-mentioned problems by providing an optimization method for the motion deformation of aircraft assembly tooling, so as to solve the problem that the non-negligible deformation of aircraft tooling in motion affects the quality of aircraft structural assembly.

[0005] The technical solution of this invention is as follows: This invention provides an optimization method for the motion deformation of aircraft assembly tooling. The aircraft assembly tooling is configured as a U-shaped frame with equal upper and lower lengths, including an upper beam, a lower beam, and two side columns with a square tube cross-section; it has larger dimensions in the length and height directions and smaller dimensions in the width direction, with the upper beam being narrower than the lower beam. It moves via wheels attached to the lower beam or is carried by an AGV (Automated Guided Vehicle), with the movement direction along the length or width of the tooling. The optimization method includes: Step 1: Establish the finite element geometric model of the initial aircraft assembly tooling. Remove bolts and welds from the finite element geometric model and treat bolt and weld connections as equivalent to bonded connections. Step 2: Assign the material parameters of each component of the aircraft assembly tooling to the finite element geometric model of the aircraft tooling using finite element software, determine the motion state constraints of the assembly tooling, and calculate the first 5 free vibration modes of the aircraft assembly tooling under the motion state constraints; among which, the material parameters include: density, Young's modulus and Poisson's ratio. Step 3: Extract the effective free vibration mode frequencies of the first two free vibration modes from the first five free vibration modes of the aircraft assembly tooling, calculate the damping of the aircraft assembly tooling, and assign the damping of the aircraft assembly tooling to the finite element geometric model of the aircraft assembly tooling through finite element software. Step 4: Determine the motion state load form of the aircraft assembly tooling, and assign the motion state load to the finite element geometric model of the aircraft assembly tooling through finite element software to calculate the motion deformation distribution of the aircraft assembly tooling under the motion load and establish the initial motion deformation model of the aircraft assembly tooling; wherein, the motion deformation model is used to characterize the deformation mode and the degree of deformation of each part of the aircraft assembly tooling. Step 5: Based on the motion deformation model, add a deformation limiting structure to the initial aircraft assembly tooling to form an optimized aircraft assembly tooling structure; wherein, the deformation limiting structure is in the form of "diagonal brace", "vertical brace", or "diagonal brace + vertical brace combination". Step 6: Perform steps 2 to 4 on the aircraft assembly tooling optimization model. Use finite element software to calculate the deformation distribution of the optimized structure of the aircraft assembly tooling under motion load, and establish a motion deformation model of the optimized structure of the aircraft assembly tooling. Compare this model with the motion deformation model of the initial aircraft assembly tooling in step 4 to evaluate the optimization effect of the added deformation limiting structure on the motion deformation of the initial aircraft assembly tooling.

[0006] Optionally, in the optimization method for the motion deformation of aircraft assembly tooling described above, step 3 includes: Step 31: Set the damping type of the aircraft assembly tooling to Rayleigh damping. : (1) in, , Here, is the Rayleigh damping coefficient, and i is the vibration mode order of the aircraft assembly tooling. The frequency of the i-th free vibration mode of the tooling used to assemble the aircraft. Step 32: Based on the closed structure and low-frequency motion of the aircraft assembly tooling, calculate the Rayleigh damping coefficient from the first two effective free vibration mode frequencies of the aircraft assembly tooling, and obtain: (2) in, , These are the first-order and second-order damping ratios, respectively. , These are the first and second effective free vibration mode frequencies, respectively.

[0007] Optionally, in the optimization method for the motion deformation of aircraft assembly tooling described above, if the motion state load of the aircraft assembly tooling is determined to be inertial force or vibration displacement in step 4, then the motion load expression form is: in, Inertial force F: (3) Where m is the mass of the aircraft assembly tooling, and a is the acceleration at the center of mass of the tooling. Vibration displacement U: (4) in, At the moment of vibration displacement excitation, The excitation displacement is a constant value greater than 0. Optionally, in the optimization method for the motion deformation of aircraft assembly tooling described above, step 4 of establishing the initial motion deformation model of the aircraft assembly tooling includes: Step 41: Based on the motion deformation of the aircraft assembly tooling as the upper beam deformation, determine the initial motion deformation model of the aircraft assembly tooling as the upper beam deformation model. Step 42, establish the deformation model of the upper beam of the initial aircraft assembly fixture as follows: When the direction of motion of the aircraft assembly tooling is along its length, the deformation of the aircraft assembly tooling is torsional deformation. Taking the left end point of the upper beam as the origin and the length direction of the upper beam as the x-direction, the torsional deformation can be approximated as: (5) in, The time it takes for the tooling to undergo one torsional deformation. The length deformation coefficient of the aircraft assembly tooling during a single torsion process. This is the deformation amplitude coefficient of the tooling during a single torsion process. The distance from any point on the upper beam to the center point of the upper beam; When the aircraft assembly tooling moves in the width direction, its deformation is a symmetrical bending deformation. Taking half of the upper beam model, the deformation of half of the upper beam model is calculated as follows: (6) in, The time it takes for the tooling to undergo one bending deformation. The bending coefficient of the tooling length during a single bending process. The amplitude deformation coefficient of the tooling during a single torsion process. This is the tooling displacement offset coefficient during a single bending process.

[0008] Optionally, in the optimization method for motion deformation of aircraft assembly tooling as described above, the "diagonal bracing" deformation limiting structure in step 5 includes "K", "V" and "X" shaped structures. The deformation limiting structure is added to the outer end face of one side of the aircraft assembly tooling to connect the upper beam and the lower beam of the tooling; the "vertical bracing" deformation limiting structure is set as an "I" shaped structure, which is vertically connected between the upper beam and the lower beam of the tooling and distributed inside the aircraft assembly tooling. Step 5 involves adding a deformation-limiting structure to the initial aircraft assembly fixture, including: Because the initial aircraft assembly tooling itself has multiple tube beams that conform to the shape of the product during the assembly process, these tube beams are removed in order to reduce the transport weight during transportation. When adding a "vertical brace" deformation limiting structure to the initial aircraft assembly tooling, the pipe beams already installed on the tooling itself are used directly, and they are not removed or only partially removed during transportation. When adding a "slanted brace" deformation limiting structure to the initial aircraft assembly tooling, all pipe beams are removed in the original transportation state, and the "slanted brace" deformation limiting structure is connected to the outer end face of one side of the aircraft assembly tooling. When adding a "diagonal brace + vertical brace combination" deformation restriction structure to the initial aircraft assembly tooling, a portion of the tubular beams are removed from the aircraft assembly tooling, and the remaining tubular beams are used as the primary deformation restriction structure. A secondary "diagonal brace" deformation restriction structure is then arranged between the remaining tubular beams.

[0009] Optionally, in the optimization method for the motion deformation of aircraft assembly tooling described above, The aforementioned "diagonal brace" deformation-limiting structure includes multiple square tubes made of Q235 steel; the total length dimension of all "diagonal brace" deformation-limiting structures is equal to the length of the upper beam of the aircraft assembly tooling; among which... The “K”-shaped structure consists of a vertical square tube and two oblique square tubes. The vertical square tube is replaced by a side column that serves as the initial aircraft assembly tooling, and the two oblique square tubes connect the side column to the upper or lower beam, respectively. The “V” shaped structure is composed of two oblique square tubes. The two ends of the “V” shaped structure are connected to the two outer ends of one end face of the upper beam, and the bottom end is connected to the outer end face of one end face of the lower beam. The “X” shaped structure is formed by the intersection of two diagonally oriented square tubes, with the upper and lower ends of each diagonally oriented square tube connected to the outer side of one end face of the upper beam and the lower beam, respectively.

[0010] Optionally, in the optimization method for the motion deformation of aircraft assembly tooling described above, The cross-sectional dimensions of the square tube in the "diagonal brace" deformation-limiting structure include cross-sectional thickness, cross-sectional length, and cross-sectional width. The cross-sectional thickness of the square tube in the "diagonal brace" deformation-limiting structure is the maximum thickness among the cross-sectional thicknesses of the upper beam, lower beam, and side columns in the aircraft assembly tooling. The cross-sectional length and width of the square tube in the "diagonal brace" deformation-limiting structure are equal to ensure identical moments of inertia in all directions. The ratio of the maximum length to the cross-sectional width of the "diagonal brace" deformation-limiting structure, i.e., the slenderness ratio... Less than 150.

[0011] Optionally, in the optimization method for the motion deformation of aircraft assembly tooling as described above, the method for evaluating the optimization effect in step 6 includes: The optimized motion deformation model of the aircraft assembly tooling structure is the upper beam deformation model. After adding the deformation limiting structure, the deformation of the upper beam at various positions of the aircraft assembly tooling is restricted to varying degrees, and the deformation distribution of the upper beam is more complex. Therefore, for the motion of the optimized aircraft assembly tooling structure along the width direction, the displacement change of the center point of the optimized tooling upper beam is taken as the motion deformation model of the optimized aircraft assembly tooling structure; for the motion of the optimized aircraft assembly tooling structure along the length direction, the displacement change of the forward endpoint of the optimized tooling upper beam is taken as the motion deformation model of the optimized aircraft assembly tooling structure; if the change amplitude is less than the change amplitude of the initial aircraft assembly tooling motion deformation model, the optimization is successful.

[0012] The beneficial effects of this invention: During the movement of aircraft assembly tooling, deformation occurs due to factors such as the flatness of the factory floor, start-up and shutdown, and random vibration excitation. This leads to decreased positioning accuracy and internal stress within the aircraft structure, resulting in reduced assembly quality and lifespan. To address these problems, this invention proposes an optimization method for the motion deformation of aircraft assembly tooling. This method involves establishing an initial finite element geometric model of the aircraft assembly tooling, assigning the material parameters of each component to the finite element geometric model using finite element software, determining the motion state constraints of the assembly tooling, calculating the first five free vibration modes of the aircraft assembly tooling under these constraints, extracting the frequencies of the first two effective free vibration modes from the first five, calculating the damping of the aircraft assembly tooling, and assigning the damping to the finite element geometric model of the aircraft assembly tooling using finite element software. Furthermore, the method utilizes the motion state of the aircraft assembly tooling... A finite element geometric model of the aircraft assembly tooling is subjected to dynamic loads to calculate the kinematic deformation distribution of the aircraft assembly tooling under dynamic loads, establishing an initial kinematic deformation model of the aircraft assembly tooling. Based on the kinematic deformation model, deformation-limiting structures are added to the initial aircraft assembly tooling to form an optimized aircraft assembly tooling structure. The optimized aircraft assembly tooling model is then calculated using finite element software to determine the deformation distribution of the optimized structure under dynamic loads, and a kinematic deformation model of the optimized structure is established and compared with the initial aircraft assembly tooling kinematic deformation model from step 4 to evaluate the optimization effect of the added deformation-limiting structures on the kinematic deformation of the initial aircraft assembly tooling. The optimization method provided by this invention improves the overall torsional and bending stiffness of the aircraft assembly tooling by adding deformation-limiting structures with different distribution forms. The kinematic deformation of the aircraft assembly tooling before and after the improvement is calculated, and the calculation results show that the deformation of the assembly tooling with added deformation-limiting structures is reduced. Clearly, this invention provides an effective and reliable method for optimizing the kinematic deformation of aircraft assembly tooling. Attached Figure Description

[0013] The accompanying drawings are provided to further understand the technical solutions of the present invention and constitute a part of the specification. They are used together with the embodiments of this application to explain the technical solutions of the present invention and do not constitute a limitation on the technical solutions of the present invention.

[0014] Figure 1 This is a schematic diagram of the aircraft assembly tooling structure and connection provided in an embodiment of the present invention; Figure 2 The deformation characteristic curve of the tooling beam provided in the embodiment of the present invention; Figure 3 A schematic diagram of the deformation limiting structure of aircraft assembly tooling is provided for embodiments of the present invention; Figure 4 A schematic diagram of the cross-section of the deformation-limiting structure for aircraft assembly tooling provided in an embodiment of the present invention; Figure 5 This is a flowchart illustrating the optimization method for motion deformation of aircraft assembly tooling provided by the present invention. Detailed Implementation

[0015] To make the objectives, technical solutions, and advantages of the present invention clearer, the embodiments of the present invention will be described in detail below with reference to the accompanying drawings. It should be noted that, unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

[0016] As explained in the background section above, aircraft assembly tooling under static conditions exhibits stable load boundaries and overall deformation. However, under dynamic conditions, the load is highly random, leading to excessive deformation of the aircraft assembly tooling and affecting the assembly quality and lifespan of the aircraft structure.

[0017] Based on the above problems, this invention proposes an optimization method for the motion deformation of aircraft assembly tooling by establishing a finite element model of the initial aircraft assembly tooling structure, determining the load characteristics and boundary constraints of the tooling, calculating the motion deformation of the initial tooling, establishing a motion deformation model of the tooling, adding deformation limiting structures to the initial aircraft assembly tooling, calculating the motion deformation of the optimized tooling, and comparing it with the initial structural deformation. Based on this design idea, this invention proposes an optimization method for the motion deformation of aircraft assembly tooling to reduce the deformation of aircraft assembly tooling during motion.

[0018] The embodiments of the present invention will be described in further detail below with reference to the accompanying drawings and examples. The following examples are for illustrative purposes only and should not be construed as limiting the scope of the invention.

[0019] This invention provides an optimization method for the motion deformation of aircraft assembly tooling. The aircraft assembly tooling is configured as a U-shaped frame of equal length, including an upper beam, a lower beam, and two side columns with a square tube cross-section. It has larger dimensions in the length and height directions and smaller dimensions in the width direction. The upper beam is narrower than the lower beam. Movement is achieved by wheels attached to the lower beam or by being carried by an AGV (Automated Guided Vehicle). The direction of movement is along the length or width of the tooling. The optimization method for the motion deformation of aircraft assembly tooling provided by this invention includes the following steps: Step 1: Establish the finite element geometric model of the initial aircraft assembly tooling. Remove bolts and welds from the finite element geometric model and treat bolt and weld connections as equivalent to bonded connections. Step 2: Assign the material parameters of each component of the aircraft assembly tooling to the finite element geometric model of the aircraft tooling using finite element software, determine the motion state constraints of the assembly tooling, and calculate the first 5 free vibration modes of the aircraft assembly tooling under the motion state constraints; among which, the material parameters include: density, Young's modulus and Poisson's ratio. Step 3: Extract the effective free vibration mode frequencies of the first two free vibration modes from the first five free vibration modes of the aircraft assembly tooling, calculate the damping of the aircraft assembly tooling, and assign the damping of the aircraft assembly tooling to the finite element geometric model of the aircraft assembly tooling through finite element software. Step 4: Determine the motion state load form of the aircraft assembly tooling, and assign the motion state load to the finite element geometric model of the aircraft assembly tooling through finite element software to calculate the motion deformation distribution of the aircraft assembly tooling under the motion load and establish the initial motion deformation model of the aircraft assembly tooling; wherein, the motion deformation model is used to characterize the deformation mode and the degree of deformation of each part of the aircraft assembly tooling. Step 5: Based on the motion deformation model, add a deformation limiting structure to the initial aircraft assembly tooling to form an optimized aircraft assembly tooling structure; wherein, the deformation limiting structure is in the form of "diagonal brace", "vertical brace", or "diagonal brace + vertical brace combination". Step 6: Perform steps 2 to 4 on the aircraft assembly tooling optimization model. Use finite element software to calculate the deformation distribution of the optimized structure of the aircraft assembly tooling under motion load, and establish a motion deformation model of the optimized structure of the aircraft assembly tooling. Compare this model with the motion deformation model of the initial aircraft assembly tooling in step 4 to evaluate the optimization effect of the added deformation limiting structure on the motion deformation of the initial aircraft assembly tooling.

[0020] The following example illustrates the implementation of the optimization method for motion deformation of aircraft assembly tooling provided by the present invention.

[0021] The optimization method for motion deformation of aircraft assembly tooling provided in this embodiment includes the following steps: Step 1: Establish a finite element geometric model of the initial aircraft assembly tooling structure. Bolts and welds are removed from the finite element geometric model, and bolt and weld connections are represented as equivalent to bonded connections, such as... Figure 1 As shown; Step 2: Input the material parameters of each part of the aircraft assembly tooling into the finite element software ABAQUS and assign them to the finite element geometric model of the aircraft tooling. The material parameters include density, Young's modulus, and Poisson's ratio. The material of the aircraft assembly tooling is Q235 steel with a density of 7.85×10^(-9) t / mm. 3 The Young's modulus is 200,000 MPa, and the Poisson's ratio is 0.3. The motion constraints of the assembly tooling are determined. In this embodiment, the tooling moves along the ground via two wheel sets at both ends of the lower beam. The first five free vibration modes under the motion constraints of the aircraft assembly tooling are calculated. In this embodiment, the frequencies of the first 5 free vibration modes are calculated as shown in the table below:

[0022] Step 3: Extract the effective free vibration mode frequencies of the first two free vibration modes from the first five free vibration modes of the aircraft assembly tooling, calculate the damping of the assembly tooling, input the damping of the aircraft assembly tooling into the finite element software and assign it to the finite element geometric model of the tooling. The damping type of the aircraft assembly tooling is Rayleigh damping. The calculation method is as follows: (1) in, , Here, is the Rayleigh damping coefficient, and i is the vibration mode order of the aircraft assembly tooling. The frequencies of each free vibration mode of the aircraft assembly tooling; Based on the closed structure and low-frequency motion of the aircraft assembly tooling, the Rayleigh damping coefficient is calculated from the first two effective free vibration mode frequencies of the aircraft assembly tooling, yielding the following: (2) in, , These are the first-order and second-order damping ratios, respectively. , These are the first and second effective free vibration mode frequencies, respectively.

[0023] In this embodiment, the aircraft is equipped with first-order and second-order damping ratios. , , Calculations yielded .

[0024] Step 4: Determine the load spectrum of the aircraft assembly tooling motion state. The motion deformation of the aircraft assembly tooling is in the category of dynamic deformation. Calculate the motion deformation distribution of the aircraft assembly tooling under motion load and establish the initial aircraft assembly tooling structure motion deformation model. The load spectrum of aircraft assembly tooling in motion is inertial force or vibration displacement, and the form of motion load is as follows: Inertial force F: (3) Where m is the mass of the aircraft assembly tooling, and a is the acceleration at the center of mass of the tooling. Vibration displacement U: (4) in, At the moment of vibration displacement excitation, The excitation displacement is a constant value greater than 0.

[0025] In this embodiment, the deformation of the aircraft assembly tooling during the startup phase is calculated. The motion deformation caused during the startup phase is mainly due to the inertial force during startup. The startup force is set to 1200 N, the friction coefficient between the wheel assembly and the ground is 0.01, and the mass of the assembly tooling is 10 t. The resulting acceleration of the tooling is 0.02 m / s². 2 An acceleration is applied to the center of gravity of the tooling, and the tooling moves along its length and width directions respectively. The displacement of the tooling after 3 seconds of movement is calculated.

[0026] Based on the motion pattern of the aircraft assembly tooling, the structural deformation of the aircraft assembly tooling is the deformation of the upper beam. The initial model of the structural deformation of the aircraft assembly tooling is the deformation model of the upper beam. When the motion direction of the aircraft assembly tooling is the length direction, the structural deformation of the tooling is mainly torsional deformation. Taking the left end point of the upper beam as the origin and the length direction of the upper beam as the x-direction, the torsional deformation can be approximated as: (5) in, The time it takes for the tooling to undergo one torsional deformation. The deformation coefficient of the tooling length during a single torsion process. This is the coefficient of tooling deformation amplitude during a single torsion process. This refers to the distance from a certain point on the upper beam to the center of the upper beam.

[0027] The time for one torsional deformation during the tooling's movement along its length is 0.0597 s. The deformation model is as follows: ; The maximum torsional deformation amplitude is 2.6 mm.

[0028] When the aircraft assembly tooling moves in the width direction, the structural deformation of the tooling mainly involves symmetrical bending deformation. Taking half of the upper beam model of the tooling, the deformation is determined as follows: (6) in, The time it takes for the tooling to undergo one bending deformation. The bending coefficient of the tooling length during a single bending process. The amplitude deformation coefficient of the tooling during a single torsion process. This is the tooling displacement offset coefficient during a single bending process.

[0029] The time for one bending deformation when the tooling moves along the width direction is found to be 0.118s. The deformation model is as follows: ; The maximum bending deformation amplitude is 4.1 mm.

[0030] Deformation characteristics of the upper beam of the tooling along the length and width directions, as follows: Figure 2 As shown.

[0031] Step 5: Add a deformation-limiting structure to the initial model of the aircraft assembly tooling, which is called the optimized model of the aircraft assembly tooling. The deformation-limiting structure can be in the form of "diagonal brace", "vertical brace", or "diagonal brace + vertical brace combination". The "diagonal bracing" deformation-limiting structure is in the form of "K", "V", and "X" shapes, distributed on both outer sides of the aircraft assembly tooling, connecting the upper and lower beams of the tooling. The "vertical bracing" structure is in the form of an "I" shape, perpendicular to the upper and lower beams of the tooling, distributed inside the aircraft assembly tooling, connecting the upper and lower beams of the tooling. Figure 3 As shown.

[0032] For aircraft assembly tooling structures, several tube beams with shapes that follow the product's shape are distributed during the assembly process. However, these tube beams are removed during transportation to reduce transport weight. Therefore, when using a "vertical brace" deformation restriction structure, the existing tube beams can be directly reused, and no removal or partial removal is required during transportation. When using a "diagonal brace" deformation restriction structure, all tube beams are removed according to the original transportation state. The "diagonal brace" structure is connected to the outer side of the aircraft assembly tooling. When using a "diagonal brace + vertical brace combination" deformation restriction structure, a portion of the tube beams of the aircraft assembly tooling are removed, and a secondary "diagonal brace" deformation structure is arranged between the remaining tube beams.

[0033] The cross-section of the "diagonal bracing" deformation limiting structure is a square tube made of Q235 steel. The "K"-shaped deformation limiting structure consists of a vertical square tube and two diagonal square tubes. The vertical tube can be replaced by the columns on both sides of the initial tooling, and the two diagonal square tubes connect to the upper and lower beams of the tooling respectively. The "V"-shaped structure consists of two diagonal square tubes, with the two ends of the "V"-shaped structure connected to the outer side of the upper beam and the bottom end connected to the outer side of the lower beam. The "X"-shaped structure consists of two diagonal square tubes intersecting, with the upper and lower ends connected to the outer sides of the upper and lower beams respectively. The total length dimension of all "diagonal bracing" deformation limiting structures is the length of the upper beam of the aircraft assembly tooling.

[0034] In addition, the cross-sectional dimensions of the "diagonal brace" deformation-limiting structure square tube include the cross-sectional thickness, cross-sectional length, and cross-sectional width. The cross-sectional thickness of the "diagonal brace" deformation-limiting structure is the same as that of the aircraft assembly tooling cross-section. The cross-sectional length and width are equal to ensure that the moments of inertia in all directions are the same. The ratio of the maximum length to the cross-sectional width of the "diagonal brace" deformation-limiting structure is the slenderness ratio. Less than 150 In this embodiment, the tooling section thickness is 10mm, therefore the deformation-limiting structure has a section thickness of 10mm, and a section length and width of 100mm. Figure 4 As shown. The maximum length is 11.18m, resulting in a slenderness ratio of 111.8, which is less than 150, thus meeting the requirements.

[0035] Step 6: Repeat steps 2 to 4 to calculate the deformation distribution of the optimized aircraft assembly tooling model under motion load using finite element software, establish the optimized aircraft assembly tooling motion deformation model, and compare it with the initial aircraft assembly tooling structure motion deformation model.

[0036] This embodiment adopts a "V"-shaped deformation constraint structure. The optimized aircraft assembly tooling motion deformation model is the upper beam deformation model. Due to the constraint and restriction of deformation, the deformation distribution of the optimized aircraft assembly tooling motion deformation model is more complex. For the tooling moving along the width direction, the displacement change of the center point of the upper beam of the optimized tooling is taken as the optimized aircraft assembly tooling motion deformation model. For the tooling moving along the length direction, the displacement change of the endpoints of the upper beam of the optimized tooling along the positive displacement direction is taken as the optimized aircraft assembly tooling motion deformation model. The displacement changes of the left end point of the upper beam of the fixture during the 2s process along the length direction and the displacement changes of the center point of the upper beam of the fixture during the width direction are shown in the table below.

[0037]

[0038] Comparing the maximum amplitude values ​​of the initial model of the tooling in the table, it can be observed that the torsional deformation of the tooling decreases, while the bending deformation remains basically unchanged. That is, the "V"-shaped deformation limiting structure has a greater effect on limiting torsional deformation, while limiting bending deformation is smaller.

[0039] Because aircraft assembly tooling deforms during movement due to factors such as factory floor flatness, start-stop cycles, and random vibration excitation, its positioning accuracy deteriorates and internal stresses are generated within the aircraft structure, leading to a decrease in assembly quality and lifespan. To address these issues, this invention proposes an optimization method for the motion deformation of aircraft assembly tooling. This method involves establishing an initial finite element geometric model of the aircraft assembly tooling, assigning the material parameters of each component to the finite element geometric model using finite element software, determining the motion constraints of the assembly tooling, calculating the first five free vibration modes of the aircraft assembly tooling under these constraints, extracting the frequencies of the first two effective free vibration modes from the first five, calculating the damping of the aircraft assembly tooling, and assigning the damping to the finite element geometric model of the aircraft assembly tooling using finite element software. Furthermore, the method utilizes the motion state of the aircraft assembly tooling... A finite element geometric model of the aircraft assembly tooling is subjected to dynamic loads to calculate the kinematic deformation distribution of the aircraft assembly tooling under dynamic loads, establishing an initial kinematic deformation model of the aircraft assembly tooling. Based on the kinematic deformation model, deformation-limiting structures are added to the initial aircraft assembly tooling to form an optimized aircraft assembly tooling structure. The optimized aircraft assembly tooling model is then calculated using finite element software to determine the deformation distribution of the optimized structure under dynamic loads, and a kinematic deformation model of the optimized structure is established and compared with the initial aircraft assembly tooling kinematic deformation model from step 4 to evaluate the optimization effect of the added deformation-limiting structures on the kinematic deformation of the initial aircraft assembly tooling. The optimization method provided by this invention improves the overall torsional and bending stiffness of the aircraft assembly tooling by adding deformation-limiting structures with different distribution forms. The kinematic deformation of the aircraft assembly tooling before and after the improvement is calculated, and the calculation results show that the deformation of the assembly tooling with added deformation-limiting structures is reduced. Clearly, this invention provides an effective and reliable method for optimizing the kinematic deformation of aircraft assembly tooling.

[0040] While the embodiments disclosed in this invention are as described above, they are merely illustrative of the embodiments to facilitate understanding of the invention and are not intended to limit the invention. Any person skilled in the art to which this invention pertains may make any modifications and variations in the form and details of the implementation without departing from the spirit and scope disclosed herein; however, the scope of patent protection for this invention shall still be determined by the scope defined in the appended claims.

Claims

1. A method for optimization of motion deformation of an aircraft assembly tooling, characterized in that, The aircraft assembly tooling is arranged as a horizontally equal-length mouth-shaped frame, comprising upper beams, lower beams and two side columns in a square tube structure. Step 1, establishing a finite element geometric model of the initial aircraft assembly tooling, removing bolts, welds in the finite element geometric model, and equivalently connecting the bolts, welds as binding connections; Step 2, assigning material parameters of each component of the aircraft assembly tooling to the finite element geometric model of the aircraft tooling through the finite element software, determining the motion state constraint condition of the assembly tooling, and calculating the first five order free vibration modes of the aircraft assembly tooling under the motion state constraint condition; wherein the material parameters include: density, Young's modulus and Poisson's ratio; Step 3, extracting the first two order effective free vibration mode frequencies from the first five order free vibration modes of the aircraft assembly tooling to calculate the damping of the aircraft assembly tooling, and assigning the damping of the aircraft assembly tooling to the finite element geometric model of the aircraft assembly tooling through the finite element software; Step 4, determining the motion state load form of the aircraft assembly tooling, assigning the motion state load to the finite element geometric model of the aircraft tooling through the finite element software to calculate the motion deformation distribution of the aircraft assembly tooling under the motion load, and establishing a motion deformation model of the initial aircraft assembly tooling; wherein the motion deformation model is used to represent the deformation mode and deformation degree of each part of the aircraft assembly tooling; Step 5, based on the motion deformation model, adding a deformation limiting structure to the initial aircraft assembly tooling to form an optimized structure of the aircraft assembly tooling; wherein the deformation limiting structure is in the form of "inclined bracing" or "vertical bracing" or "inclined bracing + vertical bracing combination"; Step 6, performing steps 2 to 4 on the optimized model of the aircraft assembly tooling, calculating the deformation distribution of the optimized structure of the aircraft assembly tooling under the motion load by the finite element software, and establishing a motion deformation model of the optimized structure of the aircraft assembly tooling, and comparing it with the motion deformation model of the initial aircraft assembly tooling in step 4 to evaluate the optimization effect of the added deformation limiting structure on the motion deformation of the initial aircraft assembly tooling.

2. The method for optimization of motion distortion of an aircraft assembly tooling fixture of claim 1, wherein, The step 3 comprises: Step 31, set the aircraft assembly fixture damping form to Rayleigh damping : ；（1） wherein, , is the Rayleigh damping coefficient, i is the vibration mode order of the aircraft assembly tooling, is the frequency of the i-th order free vibration mode of the aircraft assembly tooling; Step 32, based on the closed structure form of the aircraft assembly tooling and the low-frequency motion form, calculating the Rayleigh damping coefficient from the first two order effective free vibration mode frequencies of the aircraft assembly tooling to obtain: ；（2） wherein , are the first and second order damping ratios, respectively, , are the first and second order effective free vibration modal frequencies, respectively.

3. The method for optimization of motion-induced deformation of a tooling assembly for an aircraft as defined in claim 1, wherein, In the step 4, the motion state load form of the aircraft assembly tooling is inertia force or vibration displacement, and the motion load expression form is: Wherein, Inertial force F: ; (3) Wherein, m is the mass of the aircraft assembly tooling, and a is the acceleration at the mass center of the tooling; Vibration displacement U: ；（4） wherein is the moment of the vibration displacement excitation, is the excitation displacement, is a constant greater than 0.

4. The method for optimization of motion distortion of an aircraft assembly tooling fixture of claim 1, wherein, In the step 4, the motion deformation model of the initial aircraft assembly tooling comprises: Step 41, based on the motion deformation of the aircraft assembly tooling being the deformation of the upper beam, determining the motion deformation model of the initial aircraft assembly tooling as an upper beam deformation model; Step 42, establishing the upper beam deformation model of the initial aircraft assembly tooling as: When the movement direction of the aircraft assembly tooling is the length direction, the movement deformation of the aircraft assembly tooling is torsional deformation , taking the left end point of the upper beam as the origin and the length direction of the upper beam as the x direction, the torsional deformation is approximated as ；（5） wherein, is the time for one torsional deformation of the tooling, is the length deformation coefficient of the aircraft assembly tooling during one torsional process, is the deformation amplitude coefficient of the tooling during one torsional process, is the distance of any point of the upper beam from the center point of the upper beam; When the motion direction of the aircraft assembly tooling is the width direction, the motion deformation of the aircraft assembly tooling is symmetrical bending deformation, taking half of the upper beam model, and calculating the deformation of half of the upper beam model as: ；（6） wherein, is the time for one bending deformation of the tool, is the bending coefficient of the tool length in one bending process, is the amplitude deformation coefficient of the tool in one torsion process, is the displacement offset coefficient of the tool in one bending process.

5. The method for optimization of motion-induced deformation of a tooling for aircraft assembly according to claim 1, characterized in that, The "inclined strut type" deformation limiting structure in step 5 includes "K", "V" and "X" structures, which are added to the outer end face of one side of the aircraft assembly tooling for connecting the upper beam and the lower beam of the tooling; the "vertical strut type" deformation limiting structure is provided in the form of an "I" structure and is vertically connected between the upper beam and the lower beam of the tooling and is distributed inside the aircraft assembly tooling; The deformation limiting structure added to the initial aircraft assembly tooling in step 5 includes: Since the initial aircraft assembly tooling is provided with a plurality of pipe beams according to the product shape during assembly, the pipe beams are removed in the transportation state to reduce the transportation weight; When the "vertical strut type" deformation limiting structure is added to the initial aircraft assembly tooling, the pipe beams already provided by the tooling are directly used, and a part of the pipe beams are not removed or removed during transportation; When the "inclined strut type" deformation limiting structure is added to the initial aircraft assembly tooling, all the pipe beams are removed according to the original transportation state, and the "inclined strut type" deformation limiting structure is connected to the outer end face of one side of the aircraft assembly tooling; When the "inclined strut + vertical strut combined type" deformation structure is added to the initial aircraft assembly tooling, a part of the pipe beams are removed from the aircraft assembly tooling, the remaining pipe beams are used as the primary deformation limiting structure, and the secondary "inclined strut type" deformation limiting structure is arranged between the remaining pipe beams.

6. The optimization method for the movement deformation of the aircraft assembly tooling according to claim 5, wherein the "inclined strut type" deformation limiting structure includes a plurality of square tubes made of Q235 steel, and the total length of the "inclined strut type" deformation limiting structure in the length direction is equal to the length of the upper beam of the aircraft assembly tooling; wherein the "K" structure is composed of a vertical square tube and two inclined square tubes, the vertical square tube is replaced by a side column of the initial aircraft assembly tooling, and the two inclined square tubes are connected to the side column and the upper beam or the lower beam, respectively; the "V" structure is composed of two inclined square tubes, the two ends of the "V" structure are connected to the outer sides of one end of the upper beam, and the bottom end is connected to the outer side of one end of the lower beam; and the "X" structure is composed of two intersecting inclined square tubes, and the upper and lower ends of each inclined square tube are connected to the outer sides of one end of the upper beam and the lower beam, respectively.

7. The optimization method for the movement deformation of the aircraft assembly tooling according to claim 6, wherein the method for evaluating the optimization effect in step 6 includes: The movement deformation model of the optimized aircraft assembly tooling optimization structure is the deformation model of the upper beam, the displacement change of the center point of the upper beam of the optimized tooling is taken as the movement deformation model of the optimized aircraft assembly tooling optimization structure for the movement of the aircraft assembly tooling optimization structure in the width direction, the displacement change of the forward end point of the movement of the upper beam of the optimized tooling is taken as the movement deformation model of the optimized aircraft assembly tooling optimization structure for the movement of the aircraft assembly tooling optimization structure in the length direction, and if the change amplitude is smaller than the change amplitude of the movement deformation model of the initial aircraft assembly tooling, the optimization is successful. ​ ​ ​ The section size parameters of the square tube in the "inclined strut type" deformation limiting structure include section thickness, section length and section width; wherein the section thickness of the square tube in the "inclined strut type" deformation limiting structure is the maximum thickness among the section thicknesses of the upper beam, the lower beam and the two side columns in the aircraft assembly tooling, the section length of the square tube in the "inclined strut type" deformation limiting structure is equal to the section width to ensure the same inertia moments in all directions, and the ratio of the maximum length of the "inclined strut type" deformation limiting structure to the section width, i.e. the slenderness ratio is less than 150.

8. The method for optimization of motion distortion of an aircraft assembly tooling fixture of claim 1, wherein, ​ ​

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