CN114647895A - Method for setting assembly positioning points of aircraft structural part - Google Patents
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Abstract
The invention relates to the technical field of aircraft structure assembly, in particular to a method for setting an assembly positioning point of an aircraft structural part, which comprises the steps of firstly regarding the aircraft structural part as a rigid body, setting an initial positioning point by adopting a '3-2-1' positioning principle, secondly setting boundary conditions and load conditions according to external load and self gravity borne by the aircraft structural part in the assembly process due to the characteristics of a weak rigid body of the aircraft structural part, and carrying out finite element analysis on the maximum deformation of the aircraft structural part; and if the maximum deformation of the aircraft structural member exceeds the assembly tolerance requirement, adding a positioning point at the maximum deformation position, performing finite element analysis again, and repeating the step until the maximum deformation of the part is smaller than the assembly tolerance requirement. The method ensures the pose accuracy and the shape accuracy of the aircraft structural member in the assembling process, and reduces the number of positioning points, thereby increasing the openness of the assembling space and reducing the manufacturing cost of the assembling tool.
Description
Technical Field
The invention relates to the technical field of aircraft structure assembly, in particular to a method for setting an assembly positioning point of an aircraft structural part.
Background
In the assembly of the airplane structure, in order to ensure the accuracy of the assembly pose and the shape of each part, an assembly tool positioner is required to be adopted to accurately position the part. Due to the weak rigidity of the airplane parts, the number of the positioning points is generally more than that required by the 3-2-1 positioning principle, the more the positioning points are, the better the accuracy of the assembly pose and the shape of the airplane parts is, but the more the required assembly tool positioners are, the poor assembly space openness is caused, the assembly operation of workers is not facilitated, and the manufacturing cost of the assembly tool is increased.
Therefore, it is necessary to optimize the setting of the positioning points so that the fewer the positioning points, the better, under the conditions that ensure the assembly accuracy of the aircraft structural member. The Wangzhongqi, Huangjie, Kangminggang and the like are published in the journal of mechanical science and technology in 4 th stage of 2016, namely the 'firefly algorithm-based aircraft weak rigid part assembling and positioning strategy optimization' article, and an aircraft weak rigid part assembling and positioning strategy optimization method is provided. The 'optimization design of the positioning position of the flexible thin plate stamping part based on the genetic algorithm' is published in the journal of mechanical science and technology in 7 th stage of 2012 in the article "Shishixiyu, Liuyu, and the rest of the time, and the optimization design of the positioning position of the flexible thin plate stamping part based on the genetic algorithm is proposed, but the method is only used for the optimization of the positioning point of the flexible thin plate stamping part and is not suitable for the optimization setting of the positioning point required by the assembly of the aircraft structural part.
Chinese patent publication No. CN202070885U discloses a multi-point flexible positioning tool for automatic drilling and riveting assembly of wall panels, wherein the positioning points of the aircraft wall panels are still set by engineering experience, and the distribution of the positioning points is not optimized. Chinese patent publication No. CN204868240U discloses a flexible tool unit and a flexible dot matrix tool system, wherein a sucker capable of rotating in a spherical surface is arranged at the top end of the flexible tool unit, and can be adjusted in a self-adaptive manner according to the state of a contact surface, and can be freely lifted and lowered according to actual needs to obtain a good adsorption effect, and the flexible dot matrix tool system composed of a plurality of flexible tool units arranged in an array manner can be used for clamping and positioning workpieces of various shapes and specifications; although the lattice type flexible positioning system provided by the patent can assist the aircraft structural member to obtain higher assembly pose and shape accuracy in assembly, the number of positioning points is large, so that the assembly space is poor in openness. The Chinese patent with publication number CN104029150A discloses an open type positioning system for rear-section components of aircraft products, and the movable positioning system for the rear-section components of the aircraft provided by the invention can be automatically moved away after the positioning is finished; although the system improves the openness of subsequent assembly operation, an effective method is not provided to ensure the requirements of the pose and shape accuracy of the components in the whole assembly process.
At present, the arrangement of part positioning points required by the assembly of an airplane structure mainly depends on engineering experience of technicians, and the problems of inaccurate assembly of the airplane structure, poor openness of an assembly space, high manufacturing cost of an assembly tool and the like caused by unreasonable arrangement of the positioning points are solved.
Disclosure of Invention
The invention aims to: aiming at the problems of inaccurate assembly of an airplane structure, poor openness of an assembly space, high manufacturing cost of an assembly tool and the like caused by unreasonable arrangement of part positioning points in the prior art, the method for arranging the assembly positioning points of the machine structural part is provided.
In order to achieve the purpose, the invention adopts the technical scheme that:
the method for setting the assembly positioning point of the aircraft structural part comprises the aircraft structural part and a positioner, and comprises the following steps of:
step S1: setting an initial positioning point of the aircraft structural part according to a 3-2-1 positioning principle, wherein the initial positioning point is added into the set P;
step S2: setting boundary conditions based on the set P; setting a load condition by using an external load and gravity of the aircraft structural member;
step S3: calculating the maximum deformation of the aircraft structural part by adopting finite element analysis based on the load condition and the boundary condition;
step S4: judging whether the maximum deformation is smaller than the assembly tolerance requirement or not, and if so, ending the process; otherwise, go to step S5;
step S5: adding a positioning point at the position of the maximum deformation, and adding the positioning point into the set P;
step S6: the boundary condition is corrected based on the set P, and the process advances to step S3.
As a preferred scheme of the invention, the method for setting the assembly positioning points of the aircraft structural part is characterized in that the assembly tolerance is calculated by adopting a tolerance distribution method, and the tolerance distribution method comprises a 3DCS tolerance analysis method.
As a preferred aspect of the present invention, in step S1, the set P includes a main positioning point, where the main positioning point is a position where the assembly tolerance requirement is minimum based on the assembly tolerance of the aircraft structural member.
As a preferred scheme of the present invention, in step S4, when the maximum deformation amount is smaller than the assembly tolerance requirement, all the positioning points of the set P are positions where the positioner positions and compresses the aircraft structural member.
As a preferable aspect of the present invention, in step S5, when the shape, size, material, load condition, and boundary condition of the aircraft structural member are symmetrical, and there is a symmetrical position at the position where the maximum deformation amount exists, a positioning point is added at each of the symmetrical positions.
In summary, due to the adoption of the technical scheme, the invention has the beneficial effects that:
according to the invention, the initial positioning points are set by adopting a '3-2-1' positioning principle, and then finite element analysis is carried out according to the external load and self gravity borne by the aircraft structural member in the assembly process, so that the number of the positioning points can be reduced under the condition of ensuring the assembly accuracy of the aircraft structural member, thereby improving the openness of the assembly operation space and reducing the manufacturing cost of the assembly tool.
Drawings
FIG. 1 is a schematic diagram of the process steps of the present invention.
Fig. 2 is a schematic view of an aircraft structure.
Figure 3 is a schematic view of a locator for use in the assembly of aircraft structural members.
FIG. 4 is a graph of the results of a first finite element calculation of maximum deflection of an aircraft structural member.
FIG. 5 is a graph of the results of a first addition of an aircraft structural member anchor point.
FIG. 6 is a graph of the results of a second finite element calculation of maximum deflection of an aircraft structural member.
FIG. 7 is a graph of the results of a second addition of an aircraft structural member anchor point.
FIG. 8 is a graph of the maximum deflection of an aircraft structural member from a third finite element calculation.
FIG. 9 is a graph illustrating the results of a third addition of an anchor point for a structural member of an aircraft.
FIG. 10 is a graph illustrating the maximum deflection of the aircraft structural member from a fourth finite element calculation.
FIG. 11 is a graph illustrating the results of a fourth addition of an aircraft structural member anchor point.
FIG. 12 is a graph illustrating the results of a fifth finite element calculation of maximum deflection for an aircraft structural member.
FIG. 13 is a graph illustrating the results of a fifth addition of an aircraft structural member setpoint.
FIG. 14 is a graph of the results of a sixth finite element calculation of maximum deflection for an aircraft structural member.
The marks in the figure are: 1-ear piece hole; 2-ear piece hole; 3-ear piece hole; 4-ear piece hole; 5-lug end face; 6-positioning pins; 7-a positioning surface; 8-a compactor.
Detailed Description
The present invention will be described in detail below with reference to the accompanying drawings.
In order to make the objects, technical solutions and advantages of the present invention more apparent, the present invention is described in further detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative of the invention and are not intended to limit the invention.
Example 1
As shown in FIG. 1, the invention provides a method for setting an assembly positioning point of an aircraft structural member, which comprises the following steps:
step S1: setting an initial positioning point of the aircraft structural part according to a 3-2-1 positioning principle, wherein the initial positioning point is added into a set P;
specifically, the known aircraft structural member has a weak rigidity characteristic, and the number of positioning points arranged according to the weak rigidity characteristic is generally more than that arranged according to the 3-2-1 positioning principle, so that the aircraft structural member is assumed to be a rigid body, and then the 3-2-1 positioning principle is adopted to arrange initial positioning points required by the aircraft structural member in the assembly process, so that the number of the positioning points is reduced, and meanwhile, the assembly tool positioners required in the assembly process are reduced, so that the openness of an assembly operation space is improved, and the manufacturing cost of the assembly tool is reduced.
Furthermore, the initial positioning points are added into the set P, in order to improve the assembly precision of the aircraft structural member, the initial positioning points are dispersedly arranged according to the shape of the aircraft structural member, the distance between the initial positioning points and the aircraft structural member is kept as far as possible, meanwhile, one initial positioning point is selected as a main positioning point based on the set P, and the main positioning point is arranged at the position where the assembly tolerance value of the aircraft structural member is minimum.
Further, as for the aircraft structural member shown in fig. 2, according to the requirement that the initial positioning points need to be distributed dispersedly, since 4 tab holes of the aircraft structural member are distributed at two ends, 4 tabs are selected as the initial positioning points.
As shown in fig. 3, a locator for use in assembling an aircraft structure is used to locate the aircraft structure. Firstly, calculating the assembly tolerance of the aircraft structural member by using a tolerance distribution method, wherein the tolerance distribution method can adopt a 3DCS tolerance analysis method and the like, as shown in FIG. 2, as a main positioning point needs to be arranged at a position based on the assembly tolerance of the aircraft structural member and the assembly tolerance requirement is minimum, an ear piece hole 1 and a corresponding ear piece end face 5 thereof are selected as the main positioning point according to the assembly tolerance, a positioning pin 6 of a positioner penetrates through the middle of the ear piece hole 1, the diameter of the positioning pin 6 is the same as that of the ear piece hole, the positioning of the ear piece hole 1 is realized by adopting transition fit, the positioning is realized by attaching a positioning surface 7 of the positioner to the ear piece end face 5, and a compactor 8 is further used for compacting, so that X, Y, Z space coordinates of the aircraft structural member are determined; secondly, selecting the lug holes 2 as secondary positioning points, and positioning and pressing by using a positioner to determine the pitch angle and the rotation angle of the aircraft structural part; and finally, selecting the lug holes 3 and the lug holes 4 as secondary positioning points, and positioning and pressing by using a positioner to determine the deflection freedom degree of the airplane structural member.
Step S2: setting boundary conditions of the aircraft structural member based on the set P according to the characteristic of weak rigidity of the aircraft structural member; setting a load condition by using external load and gravity parameters of the aircraft structural part; the constraint of the positioner on the supporting or pressing of the positioning point of the aircraft structural member is the boundary condition of the aircraft structural member.
Step S3: and inputting the load condition and the boundary condition of the aircraft structural part to perform finite element analysis and calculation to obtain the maximum deformation delta of the aircraft structural part.
Specifically, the loading condition refers to the self-gravity of the aircraft structural member or the resultant force of the self-gravity and an external load; the boundary conditions refer to the constraints to which the aircraft structural member is subjected after being positioned and clamped by the positioner.
Step S4: judging whether the maximum deformation delta of the aircraft structural member is smaller than the assembly tolerance requirement or not, if so, ending the process, and at the moment, the initial positioning point in the set P is the position where the positioner positions and compresses the aircraft structural member; if the maximum deformation amount delta of the aircraft structural member is larger than or equal to the assembly tolerance requirement, the step S5 is carried out;
step S5: adding a positioning point at the position of the maximum deformation delta, wherein the positioning point is added into a set P;
specifically, the position of the maximum deformation delta is the maximum deformation position, and because the maximum deformation delta of the aircraft structural member does not meet the assembly tolerance requirement, a positioning point is additionally arranged at the maximum deformation position of the aircraft structural member, and the set P is added. If the objective condition does not allow the positioning point to be set at the maximum deformation position of the part, a positioning point is set at the position closest to the maximum deformation position.
It should be noted that, when the shape, size, material, load condition and boundary condition of the aircraft structural member are all symmetrical, the maximum deformation position is also symmetrical, and only because the deformation values of the two symmetrical positions are not necessarily absolutely equal due to finite element analysis calculation, a positioning point is respectively added on the symmetrical positions, thereby reducing the number of iterations.
Step S6: the boundary condition is corrected based on the set P, and the process advances to step S3.
Specifically, the load condition of the aircraft structure is maintained, the boundary condition is corrected based on the new anchor point set P, and the process returns to step S3.
Repeating the steps S3-S6 until the maximum deformation delta of the aircraft structural member is smaller than the assembly tolerance requirement, ending the process and achieving the purpose of the assembly accuracy of the aircraft structural member; the assembly tolerance is a technical term well known by technicians in the aircraft manufacturing industry, and is a reasonable error of the aircraft structural member assembly, and the reasonable assembly tolerance can improve the overall performance of the aircraft structural member; the assembly tolerance requirements of different aircraft structural parts are different, such as the aircraft structural part shown in fig. 2, the assembly tolerance is obtained through calculation by a tolerance allocation method, and the assembly tolerance requirements of all positions on the aircraft structural part are set to be 0.2mm, so that the reasonable assembly of the aircraft structural part is realized.
Specifically, as shown in fig. 2, the load condition and the boundary condition of the aircraft structural member are input, and finite element analysis is performed to obtain that the maximum deformation amount of the aircraft structural member is 28mm, as shown in fig. 4, it can be known that the maximum deformation amount is greater than the assembly tolerance requirement, so that a positioning point 11 is added near the maximum deformation position, as shown in fig. 5, a positioning hole is formed at the positioning point 1 and is positioned by a positioner.
Based on the additional positioning point 11, the boundary condition of the aircraft structural member is corrected, finite element analysis is carried out again to obtain that the maximum deformation amount is 18.49mm, as shown in FIG. 6, the maximum deformation amount is larger than the assembly tolerance requirement, a positioning point 12 is continuously added at the maximum deformation position, as shown in FIG. 7, a positioning hole is formed at the positioning point 2, and positioning is carried out through a positioner.
Based on the added positioning point 12, the boundary condition of the aircraft structural member is corrected, finite element analysis is carried out again to obtain that the maximum deformation amount is 1.49mm, as shown in fig. 8, the maximum deformation amount is larger than the assembly tolerance requirement, because the position of the maximum deformation amount is according to the time, the shape, the size, the load condition and the boundary condition of the aircraft structural member are symmetrical, a positioning point 13 and a positioning point 14 are respectively added near the maximum deformation position and the symmetrical position thereof, as shown in fig. 9, two positioning holes are respectively formed at the two positions of the positioning point 13 and the positioning point 14, and positioning is carried out through a positioner.
Based on the added positioning points 13 and 14, the boundary condition of the aircraft structural member is modified, finite element analysis is carried out again to obtain that the maximum deformation amount is 0.8mm, as shown in fig. 10, the maximum deformation amount is larger than the assembly tolerance requirement, and because the position of the maximum deformation amount is according to the current time, the shape, the size, the load condition and the boundary condition of the aircraft structural member are symmetrical, a positioning point 15 and a positioning point 16 are respectively added at the maximum deformation position and the vicinity of the symmetrical position thereof, as shown in fig. 11, a positioning hole is respectively arranged at the two positions of the positioning point 15 and the positioning point 16, and the positioning is carried out through a positioner.
Based on the additional positioning points 15 and 16, the boundary condition of the aircraft structural member is modified, finite element analysis is conducted again to obtain that the maximum deformation amount is 0.45mm, as shown in fig. 12, the maximum deformation amount is larger than the assembly tolerance requirement, then a positioning point 17 is added at the maximum deformation position, as shown in fig. 13, a positioning hole is formed at the positioning point 17, and positioning is conducted through a positioner.
Based on the added positioning points 17, the boundary conditions of the aircraft structural member are modified, finite element analysis is carried out again to obtain the maximum deformation amount of 0.147mm, as shown in fig. 14, the maximum deformation amount is smaller than the assembly tolerance requirement, the process is finished, and the initial positioning points 1-4 and the added positioning points 11-17 form the optimal positioning points required by the aircraft structural member to meet the assembly tolerance requirement.
Where assembly tolerances are a term of art well known to those skilled in the aircraft manufacturing industry; the positioning point is a position where the aircraft structural part is positioned and clamped by the assembly tool positioner in the assembly process.
The above description is only for the purpose of illustrating the preferred embodiments of the present invention and is not to be construed as limiting the invention, and any modifications, equivalents and improvements made within the spirit and principle of the present invention are intended to be included within the scope of the present invention.
Claims (5)
1. The method for setting the assembly positioning point of the aircraft structural part comprises the aircraft structural part and a positioner, and is characterized by comprising the following steps of:
step S1: setting an initial positioning point of the aircraft structural part according to a 3-2-1 positioning principle, wherein the initial positioning point is added into a set P;
step S2: setting a boundary condition based on the set P; setting a loading condition by using an external load and gravity of the aircraft structural member;
step S3: calculating the maximum deformation of the aircraft structural part by adopting finite element analysis based on the load condition and the boundary condition;
step S4: judging whether the maximum deformation is smaller than the assembly tolerance requirement or not, and if so, ending the process; otherwise, go to step S5;
step S5: adding a positioning point at the position of the maximum deformation, wherein the positioning point is added into a set P;
step S6: the boundary condition is corrected based on the set P, and the process advances to step S3.
2. The method for setting the assembly positioning point of the structural part of the airplane as claimed in claim 1, wherein the assembly tolerance is calculated by a tolerance distribution method, and the tolerance distribution method comprises a 3DCS tolerance analysis method.
3. The method for setting the assembly positioning point of the aircraft structure according to claim 1, wherein in step S1, the set P includes a main positioning point, and the main positioning point is a position corresponding to a minimum assembly tolerance value of the aircraft structure.
4. The method as claimed in claim 1, wherein in step S4, when the maximum deformation amount is smaller than the assembly tolerance requirement, all the positioning points in the set P are the positions where the positioner performs positioning and pressing on the aircraft structural member.
5. The method for setting the positioning point of the structural member of the aircraft as claimed in claim 1, wherein in step S5, if there is a symmetrical position at the position where the maximum deformation exists, then a positioning point is added at each of the symmetrical positions.
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Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN107628267A (en) * | 2017-07-31 | 2018-01-26 | 成都飞机工业（集团）有限责任公司 | A kind of deep camber airframe assembles double track locator unit |
| CN111504283A (en) * | 2020-04-29 | 2020-08-07 | 南京航空航天大学 | Method for calibrating point position of airplane assembly measurement field |
| CN113742979A (en) * | 2021-09-16 | 2021-12-03 | 山东大学深圳研究院 | Positioning point optimal arrangement method for clamping thin-wall arc-shaped piece |
-
2022
- 2022-03-25 CN CN202210299885.4A patent/CN114647895A/en active Pending
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN107628267A (en) * | 2017-07-31 | 2018-01-26 | 成都飞机工业（集团）有限责任公司 | A kind of deep camber airframe assembles double track locator unit |
| CN111504283A (en) * | 2020-04-29 | 2020-08-07 | 南京航空航天大学 | Method for calibrating point position of airplane assembly measurement field |
| CN113742979A (en) * | 2021-09-16 | 2021-12-03 | 山东大学深圳研究院 | Positioning point optimal arrangement method for clamping thin-wall arc-shaped piece |
Non-Patent Citations (6)
| Title |
|---|
| 刘东红;赵建才;张翊翔;: "基于有限元分析的高精密柔性夹具优化方案设计", 工程设计学报, no. 03, 15 June 2010 (2010-06-15), pages 36 - 41 * |
| 孙明凯;钱炜;: "薄壁件辅助支承布局优化设计", 农业装备与车辆工程, no. 05, 10 May 2020 (2020-05-10), pages 5 - 9 * |
| 李西宁;赵志浩;仝梦佳;王守川;: "基于0-1整数规划的航空薄壁件定位布局优化", 航空制造技术, no. 07, 1 April 2018 (2018-04-01), pages 30 - 35 * |
| 王仲奇;黄杰;康永刚;杨勃;杨元;: "基于萤火虫算法的飞机弱刚性件装配定位策略优化", 机械科学与技术, no. 04, 30 April 2016 (2016-04-30), pages 140 - 143 * |
| 王少锋;洪军;王建国;李宝童;杨朝晖;: "大型薄壁件的多点支承/定位方法研究", 西安交通大学学报, no. 06, 30 June 2016 (2016-06-30), pages 128 - 135 * |
| 王鹏;姜坤;周雄辉;: "基于粒子群算法的测量夹具定位点的优化设计", 模具工业, no. 04, 15 April 2013 (2013-04-15), pages 23 - 28 * |
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