CN114647895B - Method for setting positioning points for assembly of aircraft structural member - Google Patents
Method for setting positioning points for assembly of aircraft structural member Download PDFInfo
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Abstract
The invention relates to the technical field of aircraft structure assembly, in particular to an aircraft structure assembly positioning point setting method, which comprises the steps of firstly regarding an aircraft structure as a rigid body, setting an initial positioning point by adopting a 3-2-1 positioning principle, and secondly, carrying out finite element analysis on the maximum deformation of the aircraft structure according to the external load and self gravity set boundary conditions and load conditions of the aircraft structure in the assembly process due to the characteristics of the weak rigid body of the aircraft structure; if the maximum deformation of the aircraft structural member exceeds the assembly tolerance requirement, adding a locating point at the maximum deformation position, carrying out finite element analysis again, and repeating the steps 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 assembly process, and simultaneously reduces the number of positioning points, thereby increasing the openness of the assembly space and reducing the manufacturing cost of the assembly tool.
Description
Technical Field
The invention relates to the technical field of aircraft structure assembly, in particular to a method for setting assembly positioning points of aircraft structural parts.
Background
In the aircraft structure assembly, in order to ensure the accuracy of the assembly pose and shape of each part, an assembly fixture positioner is required to accurately position the part. Because of the weak rigidity characteristic of the aircraft parts, the number of positioning points is generally more than that required by the positioning principle of '3-2-1', the more the positioning points are, the better the accuracy of the assembly pose and shape of the aircraft parts is, but the more the required assembly fixture positioners are, the poor assembly space is caused, the assembly operation of workers is not facilitated, and the manufacturing cost of the assembly fixture is increased.
Therefore, it is necessary to optimize the setting of the anchor points so that the fewer the anchor points, the better, under the condition of ensuring the assembly accuracy of the aircraft structural member. Wang Zhongji, huang Jie, kang Yonggang et al published in J4 of mechanical science and technology on the "optimization of positioning strategy for assembly of weak rigid parts of an aircraft based on firefly algorithm" in one text, propose an optimization method for positioning strategy of assembly of weak rigid parts of an aircraft, which uses the maximum deformation of parts of an aircraft as an objective function and adopts firefly algorithm to optimize the distribution of positioning points, but the method uses the minimum maximum deformation of parts of an aircraft as an optimization objective, the number of positioning points is still determined according to engineering experience, and the minimum positioning points and the distribution thereof required for ensuring the assembly accuracy of structural parts of an aircraft cannot be calculated. Dan Zhiyun, liu Yu, and the rest Shi Jian published in "genetic algorithm-based positioning optimization design of flexible sheet stamping parts" in journal of mechanical science and technology, 7 th year 2012, propose a genetic algorithm-based positioning optimization design of flexible sheet stamping parts, but the method is only used for positioning point optimization of flexible sheet stamping parts and is not suitable for positioning point optimization setting required by aircraft structural part assembly.
The Chinese patent with publication number of CN202070885U discloses a multi-point flexible positioning tool for automatic drilling and riveting assembly of a wallboard, wherein the positioning points of the aircraft wallboard are still set by engineering experience, and the distribution of the positioning points is not optimized. The Chinese patent with the publication number of CN204868240U discloses a flexible tool unit and a flexible lattice tool system, wherein the top end of the flexible tool unit is provided with a sucker capable of rotating in a spherical surface, the sucker can be adaptively adjusted according to the state of a contact surface, and the flexible lattice tool system formed by a plurality of flexible tool units arranged in an array manner can be used for clamping and positioning workpieces with various shapes and specifications according to the free lifting of actual needs so as to obtain good adsorption effects; the lattice type flexible positioning system provided by the patent can assist an aircraft structural member to obtain higher assembly pose and shape accuracy in assembly, but has more positioning points, so that the assembly space is poor in openness. The invention discloses an open positioning system for a rear part of an aircraft product, which is disclosed in China patent with publication number CN104029150A, and the movable positioning system for the rear part of the aircraft can be automatically moved away after positioning is completed; although this system improves the openness of subsequent assembly operations, no effective method is proposed to ensure the pose and shape accuracy requirements of the component throughout the assembly process.
At present, the setting of part locating points required by the assembly of an aircraft structure mainly depends on engineering experience of technicians, and the unreasonable setting of the locating points leads to the problems of inaccurate assembly of the aircraft structure, poor openness of assembly space, high manufacturing cost of assembly tools and the like.
Disclosure of Invention
The invention aims at: 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 positioning point setting of parts in the prior art, the method for setting the positioning point of the assembly of the airplane structural part is provided.
In order to achieve the above purpose, the technical scheme adopted by the invention is as follows:
the method for setting the assembly positioning points of the aircraft structural member comprises the following steps of:
Step S1: setting initial positioning points of the aircraft structural member according to a positioning principle of 3-2-1, and adding the initial positioning points into the set P;
Step S2: setting a boundary condition based on the set P; setting load conditions by using external load and gravity of the aircraft structural member;
step S3: based on the load condition and the boundary condition, calculating the maximum deformation of the structural member of the aircraft by adopting finite element analysis;
step S4: judging whether the maximum deformation is smaller than the assembly tolerance requirement, if so, ending the flow; otherwise, enter step S5;
Step S5: adding a locating point at the position of the maximum deformation, and adding the locating 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 preferable scheme of the invention, an aircraft structural member assembly positioning point setting method is characterized in that assembly tolerance is calculated by adopting a tolerance distribution method, and the tolerance distribution method comprises a 3DCS tolerance analysis method.
As a preferable scheme of the invention, in the method for setting the assembly positioning points of the aircraft structural member, in the step S1, the set P comprises a main positioning point which is a position with minimum assembly tolerance requirement based on the assembly tolerance of the aircraft structural member.
As a preferable scheme of the invention, in the method for setting the assembly positioning points of the aircraft structural member, in the step S4, when the maximum deformation is smaller than the assembly tolerance requirement, all positioning points of the set P are used as the positions of the positioners for positioning and compacting the aircraft structural member.
As a preferable scheme of the invention, in the method for setting the assembly positioning points of the aircraft structural member, in the step S5, when the shape, the size, the material, the load condition and the boundary condition of the aircraft structural member are symmetrical, the symmetrical position exists at the position of the maximum deformation, and one positioning point is respectively added at the symmetrical position.
In summary, due to the adoption of the technical scheme, the beneficial effects of the invention are as follows:
According to the invention, the initial positioning points are set by adopting the positioning principle of '3-2-1', and then finite element analysis is carried out according to the external load and self gravity of 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 space openness of assembly operation and reducing the manufacturing cost of the assembly tool.
Drawings
FIG. 1 is a schematic diagram of the method steps of the present invention.
Fig. 2 is a schematic view of an aircraft structural member.
Fig. 3 is a schematic view of a fixture for use in the assembly of aircraft structural members.
FIG. 4 is a graph of the results of a first finite element calculation of the maximum deflection of an aircraft structural member.
Fig. 5 is a graph of the result of the first addition of the anchor points of the structural members of the aircraft.
FIG. 6 is a graph of the results of a second finite element calculation of the maximum deflection of an aircraft structural member.
Fig. 7 is a graph of the result of a second addition of the anchor points of the aircraft structure.
FIG. 8 is a graph of the results of a third finite element calculation of the maximum deflection of an aircraft structural member.
Fig. 9 is a graph of the result of a third addition of the anchor points of the aircraft structure.
FIG. 10 is a graph of the results of a fourth finite element calculation of the maximum deflection of an aircraft structural member.
Fig. 11 is a graph of the result of the fourth addition of the anchor point for the structural member of the aircraft.
FIG. 12 is a graph of the results of a fifth finite element calculation of the maximum deflection of an aircraft structural member.
Fig. 13 is a graph of the result of the fifth addition of the anchor point for the structural member of the aircraft.
Fig. 14 is a graph of the result of the sixth finite element calculation of the maximum deflection of the structural member of the aircraft.
The marks in the figure: 1-ear holes; 2-ear holes; 3-ear holes; 4-ear holes; 5-ear end face; 6-locating pins; 7, positioning surface; 8-a compactor.
Detailed Description
The present invention will be described in detail with reference to the accompanying drawings.
The present invention will be described in further detail with reference to the drawings and examples, in order to make the objects, technical solutions and advantages of the present invention more apparent. It should be understood that the specific embodiments described herein are for purposes of illustration only and are not intended to limit the scope of 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 initial positioning points of the aircraft structural member according to a positioning principle of 3-2-1, wherein the initial positioning points are added into a set P;
Specifically, the known aircraft structural member has weak rigidity characteristics, and the number of positioning points arranged according to the weak rigidity characteristics is generally more than the number arranged on the positioning principle of 3-2-1, so that the aircraft structural member is firstly assumed to be a rigid body, and then the initial positioning points required by the aircraft structural member in the assembly process are arranged on the positioning principle of 3-2-1, thereby reducing the number of positioning points, reducing the number of assembly fixture positioners required in the assembly process, improving the openness of assembly operation space and reducing the manufacturing cost of the assembly fixture.
Furthermore, the initial positioning points are added into the set P, the initial positioning points are distributed and arranged according to the shape of the aircraft structural member to keep the distance as far as possible, and meanwhile, one initial positioning point is selected from the set P to serve as a main positioning point, and the main positioning point is arranged at the position with the minimum assembly tolerance value of the aircraft structural member.
Further, as shown in fig. 2, according to the requirement that the initial positioning points need to be distributed, 4 lugs of the aircraft structural member are distributed at two ends, so that 4 lugs are selected as the initial positioning points.
As shown in fig. 3, the locator is used for the assembly of the aircraft structure, with which the aircraft structure is located. Firstly, calculating to obtain the assembly tolerance of an 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, because a main positioning point is required to be arranged at a position with minimum assembly tolerance requirement based on the assembly tolerance of the aircraft structural member, a lug hole 1 and a lug end face 5 corresponding to the lug hole 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 lug hole 1, the diameter of the positioning pin 6 is the same as that of the lug hole, the lug hole 1 is positioned by adopting transition fit, positioning is realized by attaching a positioning surface 7 of the positioner to the lug end face 5, and further the aircraft structural member is compressed by using a compressor 8, so that the 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 compacting by using a positioner to determine the pitching angle and the rotation angle of the aircraft structural member; finally, the lug holes 3 and the lug holes 4 are selected as secondary positioning points, and the deflection freedom degree of the aircraft structural member is determined by positioning and compacting by a positioner.
Step S2: according to the characteristic of weak rigidity of the aircraft structural member, setting boundary conditions of the aircraft structural member based on the set P; setting load conditions by using external load and gravity parameters of the aircraft structural member; the constraint of the positioner on the support or compression of the positioning points 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 member, and carrying out finite element analysis and calculation to obtain the maximum deformation delta of the aircraft structural member.
Specifically, the load condition refers to the self gravity of the aircraft structural member or the resultant force of the self gravity and the applied load; boundary conditions refer to constraints imposed by the aircraft structure 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, if the maximum deformation delta of the aircraft structural member is smaller than the assembly tolerance requirement, ending the process, and taking the initial positioning point in the set P as the position of the positioner for positioning and compacting the aircraft structural member; if the maximum deformation delta of the aircraft structural member is greater 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 the set P;
Specifically, the position where the maximum deformation delta is located is the maximum deformation position, and since the maximum deformation delta of the aircraft structural member does not meet the requirement of assembly tolerance, a positioning point is additionally arranged at the maximum deformation position of the aircraft structural member, and the positioning point is added into the set P. 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, loading condition and boundary condition of the aircraft structural member are all symmetrical, the maximum deformation position is symmetrical, but the deformation values of the two symmetrical positions are not necessarily absolutely equal due to the finite element analysis and calculation, so that one positioning point is added to each symmetrical position, and the iteration times are reduced.
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 kept unchanged, the boundary condition is corrected according to the new positioning point set P, and the step S3 is returned.
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 to a person skilled in the aircraft manufacturing industry, is a reasonable error of the assembly of the aircraft structural member, and can improve the overall performance of the aircraft structural member; the assembly tolerance requirements of different aircraft structural members are different, as shown in the aircraft structural member in fig. 2, the assembly tolerance is calculated through a tolerance distribution method, and the assembly tolerance requirements of all positions on the aircraft structural member are set to be 0.2mm, so that the aircraft structural member can be assembled reasonably.
Specifically, as shown in fig. 2, the load condition and the boundary condition of the aircraft structural member are input, finite element analysis is performed to obtain the maximum deformation of the aircraft structural member of 28mm, as shown in fig. 4, it can be known that the maximum deformation 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 in the positioning point 1 to position the aircraft structural member through a positioner.
Based on the added positioning point 11, the boundary condition of the aircraft structural member is corrected, finite element analysis is conducted again to obtain the maximum deformation of 18.49mm, the maximum deformation is larger than the assembly tolerance requirement as shown in fig. 6, a positioning point 12 is continuously added at the maximum deformation position, and a positioning hole is formed in the positioning point 2 to position through a positioner as shown in fig. 7.
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 the maximum deformation of 1.49mm, as shown in fig. 8, the maximum deformation is larger than the assembly tolerance requirement, and because the position of the maximum deformation according to the time is symmetrical based on the shape, the size, the load condition and the boundary condition of the aircraft structural member, a positioning point 13 and a positioning point 14 are respectively added near the maximum deformation position and the symmetrical position, as shown in fig. 9, a positioning hole is respectively formed at the two positions of the positioning point 13 and the positioning point 14 to position through a positioner.
Based on the added locating points 13 and 14, the boundary condition of the aircraft structural member is corrected, finite element analysis is carried out again to obtain the maximum deformation of 0.8mm, as shown in fig. 10, the maximum deformation is larger than the assembly tolerance requirement, and because the shape, the size, the load condition and the boundary condition of the aircraft structural member are symmetrical according to the position of the maximum deformation, a locating point 15 and a locating point 16 are respectively added near the maximum deformation position and the symmetrical position, as shown in fig. 11, a locating hole is respectively formed at the two positions of the locating point 15 and the locating point 16 to locate through a locator.
Based on the added positioning points 15 and 16, the boundary condition of the aircraft structural member is corrected, finite element analysis is performed again to obtain the maximum deformation of 0.45mm, as shown in fig. 12, the maximum deformation 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 in the positioning point 17, and positioning is performed through a positioner.
Based on the added positioning points 17, the boundary condition of the aircraft structural member is corrected, finite element analysis is conducted again to obtain the maximum deformation of 0.147mm, as shown in fig. 14, the maximum deformation is smaller than the assembly tolerance requirement, the process is finished, and the initial positioning points 1-4 are added with the added positioning points 11-17 to form the optimal positioning points required by the aircraft structural member to meet the assembly tolerance requirement.
Wherein assembly tolerances are terms of art well known to those skilled in the aircraft manufacturing industry; the locating point is the position of the aircraft structural member which is located and clamped by the assembling fixture locator in the assembling process.
The foregoing description of the preferred embodiments of the invention is not intended to be limiting, but rather is intended to cover all modifications, equivalents, and alternatives falling within the spirit and principles of the invention.
Claims (3)
1. The method for setting the assembly positioning points of the aircraft structural member comprises the aircraft structural member and a positioner and is characterized by comprising the following steps of:
Step S1: setting initial positioning points of the aircraft structural member according to a positioning principle of 3-2-1, wherein the initial positioning points are added into a set P; the set P comprises a main positioning point which is the position corresponding to the minimum assembly tolerance value of the aircraft structural member;
Step S2: setting a boundary condition based on the set P; setting load conditions using external loads and gravity of the aircraft structure;
Step S3: based on the load condition and the boundary condition, calculating the maximum deformation of the aircraft structural member by adopting finite element analysis;
Step S4: judging whether the maximum deformation is smaller than the assembly tolerance requirement, if so, positioning and compacting the aircraft structural member by the positioner; otherwise, enter step S5;
Step S5: adding a positioning point at the position of the maximum deformation, wherein the positioning point is added into the set P;
step S6: based on the set P, the boundary condition is corrected, and the process advances to step S3.
2. An aircraft structural assembly setpoint setting method according to claim 1, wherein the assembly tolerances are calculated using a tolerance distribution method comprising a 3DCS tolerance analysis method.
3. The method for setting positioning points for assembling an aircraft structural member according to claim 1, wherein in step S5, symmetrical positions exist at the positions of the maximum deformation, and a positioning point is added to each symmetrical position.
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Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
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| 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 |
| 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 |
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| 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 |
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| 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 |
| 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 |
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