CN114818445B - Static force loading point determining method in static force superposition vibration test of airplane structure - Google Patents
Static force loading point determining method in static force superposition vibration test of airplane structure Download PDFInfo
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Abstract
The invention discloses a method for determining a static force loading point in an aircraft structure static force superposition vibration test, which comprises the following steps: firstly, determining the number N of vibration modes considered by the design of an airplane structure test piece; secondly, carrying out finite element modeling on the airplane structure test piece and carrying out modal analysis; thirdly, acquiring a scale factor pre-estimated initial value corresponding to each order of mode; fourthly, calculating initial values of displacement thresholds of all the nodes in the ith order mode; fifthly, acquiring N groups of node sets and solving intersection to obtain node intersection D under the former N-order mode; sixthly, judging whether the node intersection D is an empty set; seventhly, calculating the j-th adjusting value of the corresponding scale factor of each order of mode; eighthly, calculating j adjustment values of displacement thresholds of all nodes in the ith order mode; ninthly, repeating the steps from five to eight until the node intersection D is not an empty set; and ten, selecting the node with the largest distance from the reference point of the test piece in the node intersection D as a static loading point. According to the invention, the position of the static force loading point is obtained through analysis, so that the influence on the dynamic characteristics of the structure is reduced.
Description
Technical Field
The invention belongs to the technical field of static force loading point determination in an airplane structure vibration test, and particularly relates to a method for determining a static force loading point in an airplane structure static force superposition vibration test.
Background
The loads borne by the aircraft in the flying process are very complex, and if the structural strength in the design of the aircraft is insufficient, structural damage can be caused, and a flying accident can be caused in serious cases. Therefore, the aircraft needs to pass ground tests to verify the structural strength before formal sizing. The typical structure of an airplane such as a wing, a vertical fin and the like is subjected to not only steady aerodynamic force, namely static load, but also unsteady aerodynamic force, namely vibration load, so that the steady aerodynamic force, namely the vibration load and the unsteady aerodynamic force need to be considered simultaneously in experimental verification. The loading modes of static force and vibration are different. Static loading is generally carried out by adopting a hydraulic actuator cylinder, vibration loading is usually carried out by adopting a vibration exciter, and two loading modes need to be integrated into one loading system in actual operation to carry out a combined loading test of static superposition vibration. However, the loading mode of the hydraulic actuator cylinder adopted for static loading actually changes the real boundary conditions of the structure, and influences the dynamic characteristics of the structure, namely changes the natural frequency and the vibration mode of the structure, so that the reliability of the test result is reduced, and therefore, the determination of the position of a static loading point in a proper static superposition vibration test of the aircraft structure is very important.
Disclosure of Invention
The technical problem to be solved by the present invention is to provide a method for determining a static force loading point in an aircraft structure static force superposition vibration test, which aims at the defects in the prior art, and comprises the steps of performing finite element modeling on a vibration test structure of an aircraft structure test piece, performing modal analysis on the aircraft structure test piece in a finite element model to obtain a vibration mode of a first N-order mode of the aircraft structure test piece, estimating a corresponding scale factor of each order mode and adjusting a corresponding scale factor of each order mode in an iterative manner, calculating a displacement threshold of each node in an ith-order mode, taking an absolute value of the modal displacement of each order mode and comparing the absolute value with each modal threshold, selecting nodes smaller than each modal threshold to form an ith-order modal node set, performing intersection on the obtained N groups of node sets to obtain a node intersection under the first N-order mode, and enabling the nodes to exist in the intersection, and selecting the node with the largest distance from the reference point of the test piece in the node intersection D as a static loading point, loading the static load in a relatively immobile area of the aircraft structure test piece, and reducing the influence of the static loading on the structure dynamic characteristics, thereby improving the reliability of the test result and being convenient for popularization and use.
In order to solve the technical problems, the invention adopts the technical scheme that: a method for determining a static force loading point in an aircraft structure static force superposition vibration test is characterized by comprising the following steps:
determining the number of vibration modes of an airplane structure test piece design consideration in a static force superposition vibration test as the first N order, wherein N is a positive integer;
carrying out finite element modeling on a vibration test structure of the airplane structure test piece, wherein the vibration test structure comprises the airplane structure test piece, one end of the airplane structure test piece is fixed through a test piece fixer, grid division is carried out on the surface of the airplane structure test piece in a finite element model, the cross point of each grid is a node, and the midpoint of the connection between the airplane structure test piece and the test piece fixer is a test piece reference point;
modal analysis is carried out on the aircraft structure test piece in the finite element model, and the vibration mode of the front N-order mode of the aircraft structure test piece, namely the front N-order modal displacement of all nodes is obtained
Wherein, in the step (A),
is the numbering of the nodes and
，
is the total number of the nodes and,
is numbered for modal order and
；
thirdly, the corresponding scale factor of each order of mode is estimated, and the estimated initial value of the corresponding scale factor of each order of mode is obtained
Wherein, in the step (A),
；
step fourAccording to the formula
Calculating the initial value of the displacement threshold of each node in the ith order mode
Wherein, in the step (A),
is the maximum of the absolute value of the displacement in the i-th mode and
，
is a maximum function;
fifthly, taking absolute values of modal displacement in each order of modes, comparing the absolute values with respective modal threshold values, and selecting nodes smaller than the respective modal threshold values to form ith order modal node set
；
Performing intersection solving on the obtained N groups of node sets to obtain a node intersection D under the former N-order mode;
step six, judging whether the node intersection D is an empty set, and if the node intersection D is the empty set, executing the step seven; if the node intersection D is not an empty set, executing the step ten;
step seven, according to the formula
Calculating the j adjustment value of the corresponding scale factor of each order mode
Wherein, in the step (A),
，
numbers the corresponding scale factor adjusting times of each order mode and
is a positive integer and is a non-zero integer,
adjusting the step size for the scale factor;
step eight, according to the formula
Calculating the j adjustment value of the displacement threshold value of each node in the ith order mode
；
Step nine, repeating the step five to the step eight until the node intersection D is not an empty set;
step ten, if one node exists in the node intersection D, the node is used as a static loading point;
and if the number of the nodes in the node intersection D is larger than 1, selecting the node with the largest distance from the test piece reference point in the node intersection D as the static loading point.
The method for determining the static force loading point in the static force superposition vibration test of the aircraft structure is characterized by comprising the following steps of: the pre-estimated initial value of the corresponding scale factor of each order mode
Taking 0.05-0.1.
The method for determining the static force loading point in the static force superposition vibration test of the aircraft structure is characterized by comprising the following steps of: the scale factor adjustment step size
Taking 0.05-0.1.
The method has the advantages that the method carries out finite element modeling on the vibration test structure of the aircraft structure test piece, carries out modal analysis on the aircraft structure test piece in a finite element model, obtains the vibration mode of the front N-order mode of the aircraft structure test piece, calculates the displacement threshold value of each node in the ith-order mode by predicting the corresponding scale factor of each-order mode and adjusting the corresponding scale factor of each-order mode in an iterative mode, obtains the absolute value of the modal displacement of each-order mode and compares the absolute value with the respective modal threshold value, selects the nodes smaller than the respective modal threshold value to form the ith-order node set, carries out intersection calculation on the obtained N groups of node sets to obtain the node intersection under the front N-order mode, enables the nodes to exist in the intersection, selects the node with the maximum distance from the reference point of the test piece in the node intersection D as the static force loading point, and loads the static force in the relatively motionless area of the aircraft structure test piece, the influence of static loading on the structure dynamics is reduced, so that the reliability of the test result is improved, and the method is convenient to popularize and use.
The technical solution of the present invention is further described in detail by the accompanying drawings and embodiments.
Drawings
FIG. 1 is a schematic view of an installation of a test piece of an aircraft structure and a finite element model diagram.
FIG. 2 is a schematic diagram of a node set of the first N-th order mode according to the present invention.
FIG. 3 is a schematic diagram illustrating intersection results of the previous N-th order modal node set according to the present invention.
FIG. 4 is a block diagram of a method flow of the present invention.
Description of reference numerals:
1-an aircraft structure test piece; 2-a test piece holder; 3-test piece reference point;
4-node.
Detailed Description
As shown in fig. 1 to 4, the method for determining the static loading point in the static superposition vibration test of the aircraft structure of the invention comprises the following steps:
the method comprises the following steps of firstly, determining the number of vibration modes considered by the design of an aircraft structure test piece 1 in a static force superposition vibration test as the first N order, wherein N is a positive integer;
step two, carrying out finite element modeling on a vibration test structure of the aircraft structure test piece 1, wherein the vibration test structure comprises the aircraft structure test piece 1, one end of the aircraft structure test piece 1 is fixed through a test piece fixer 2, the surface of the aircraft structure test piece 1 is divided into grids in a finite element model, a cross point of each grid is a node 4, and a midpoint of the connection between the aircraft structure test piece 1 and the test piece fixer 2 is a test piece reference point 3;
modal analysis is carried out on the aircraft structure test piece 1 in the finite element model, and the vibration mode of the front N-order mode of the aircraft structure test piece 1, namely the front N-order modal displacement of all the nodes 4 is obtained
Wherein, in the step (A),
is the number of node 4 and
，
is the total number of the nodes 4,
is numbered for modal order and
；
thirdly, the corresponding scale factor of each order of mode is estimated, and the estimated initial value of the corresponding scale factor of each order of mode is obtained
Wherein, in the process,
；
step four, according to the formula
Calculating the initial value of the displacement threshold of each node 4 in the ith order mode
Wherein, in the process,
is the maximum of the absolute value of the displacement in the i-th mode and
，
is a maximum function;
fifthly, absolute values of modal displacement in each order of modes are obtained and compared with respective modal threshold values, and nodes smaller than the respective modal threshold values are selected to form an ith order modal node set
；
Performing intersection solving on the obtained N groups of node sets to obtain a node intersection D under the former N-order mode;
step six, judging whether the node intersection D is an empty set, and if the node intersection D is the empty set, executing the step seven; if the node intersection D is not an empty set, executing the step ten;
step seven, according to the formula
Calculating the j-th adjustment value of the corresponding scale factor of each order of mode
Wherein, in the step (A),
，
numbers the corresponding scale factor adjusting times of each order mode and
is a positive integer and is a non-zero integer,
adjusting the step size for the scale factor;
step eight, according to the formula
Calculating the j adjustment value of the displacement threshold value of each node 4 in the ith order mode
；
Step nine, repeating the step five to the step eight until the node intersection D is not an empty set;
step ten, if one node 4 exists in the node intersection D, the node 4 is used as a static force loading point;
and if the number of the nodes 4 in the node intersection D is larger than 1, selecting the node 4 with the largest distance from the reference point 3 of the test piece in the node intersection D as a static force loading point.
In this embodiment, the pre-estimated initial value of the corresponding scale factor of each order mode
Taking 0.05-0.1.
In this embodiment, the scale factor adjusting step size
Taking 0.05-0.1.
When the method is used, the method comprises the steps of carrying out finite element modeling on a vibration test structure of the aircraft structure test piece, carrying out modal analysis on the aircraft structure test piece in a finite element model, obtaining the vibration mode of the front N-order mode of the aircraft structure test piece, calculating the displacement threshold value of each node in the ith-order mode by predicting the corresponding scale factor of each-order mode and adjusting the corresponding scale factor of each-order mode in an iterative mode, taking the absolute value of the modal displacement in each-order mode and comparing the absolute value with the respective mode threshold value, selecting the nodes smaller than the respective mode threshold values to form an ith-order modal node set, solving the intersection of the N groups of node sets to obtain the node intersection under the front N-order mode, enabling the nodes to exist in the intersection, selecting the node with the largest distance from a test piece reference point in the node intersection D as a static force loading point, and enabling a static force load to be loaded in a relatively motionless region of the aircraft structure test piece, the influence of static loading on the structure dynamics is reduced, and the reliability of the test result is improved.
The above description is only a preferred embodiment of the present invention, and is not intended to limit the present invention, and all simple modifications, changes and equivalent structural changes made to the above embodiment according to the technical spirit of the present invention still fall within the protection scope of the technical solution of the present invention.
Claims (3)
1. A method for determining a static force loading point in an aircraft structure static force superposition vibration test is characterized by comprising the following steps:
the method comprises the following steps of firstly, determining the number of vibration modes considered by the design of an aircraft structure test piece (1) in a static force superposition vibration test as the first N order, wherein N is a positive integer;
step two, carrying out finite element modeling on a vibration test structure of the aircraft structure test piece (1), wherein the vibration test structure comprises the aircraft structure test piece (1), one end of the aircraft structure test piece (1) is fixed through a test piece fixer (2), the surface of the aircraft structure test piece (1) is subjected to grid division in a finite element model, a cross point of each grid is a node (4), and a midpoint of the connection between the aircraft structure test piece (1) and the test piece fixer (2) is a test piece reference point (3);
modal analysis is carried out on the aircraft structure test piece (1) in the finite element model, and the mode shape of the front N-order mode of the aircraft structure test piece (1), namely the front N-order modal displacement of all the nodes (4), is obtained
Wherein, in the step (A),
is the number of the node (4) and
，
is the total number of the nodes (4),
is numbered for modal order and
；
thirdly, the corresponding scale factor of each order of mode is estimated, and the estimated initial value of the corresponding scale factor of each order of mode is obtained
Wherein, in the step (A),
；
step four, according to the formula
Calculating the initial value of the displacement threshold of each node (4) in the ith order mode
Wherein, in the step (A),
is the maximum of the absolute value of the displacement in the i-th mode and
，
is a maximum function;
fifthly, taking absolute values of modal displacement in each order of modes, comparing the absolute values with respective modal threshold values, and selecting nodes smaller than the respective modal threshold values to form ith order modal node set
；
Performing intersection solving on the obtained N groups of node sets to obtain a node intersection D under the former N-order mode;
step six, judging whether the node intersection D is an empty set, and if the node intersection D is the empty set, executing the step seven; if the node intersection D is not an empty set, executing the step ten;
step seven, according to the formula
Calculating the j-th adjustment value of the corresponding scale factor of each order of mode
Wherein, in the step (A),
，
numbers the corresponding scale factor adjusting times of each order mode and
is a positive integer and is a non-zero integer,
adjusting the step size for the scale factor;
step eight, according to the formula
Calculating the j adjustment value of the displacement threshold value of each node (4) in the ith order mode
；
Step nine, repeating the step five to the step eight until the node intersection D is not an empty set;
step ten, if one node (4) exists in the node intersection D, taking the node (4) as a static force loading point;
and if the number of the nodes (4) in the node intersection D is larger than 1, selecting the node (4) with the largest distance from the test piece reference point (3) in the node intersection D as a static loading point.
2. The method for determining the static force loading point in the aircraft structure static force superposition vibration test according to claim 1, is characterized in that: the pre-estimated initial value of the corresponding scale factor of each order mode
Taking 0.05-0.1.
3. The method for determining the static force loading point in the aircraft structure static force superposition vibration test according to claim 1, is characterized in that: the scale factor adjustment step size
Taking 0.05-0.1.
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| CN113720707A (en) * | 2021-08-23 | 2021-11-30 | 中国飞机强度研究所 | Design method for large-deformation loading point of structural fatigue test |
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| WO2018117971A1 (en) * | 2016-12-22 | 2018-06-28 | Agency For Science, Technology And Research | Method and system of manufacturing a load-bearing structure and a load-bearing structure manufactured thereof |
| CN112733258A (en) * | 2020-12-29 | 2021-04-30 | 中国航空工业集团公司西安飞机设计研究所 | Analysis method for predicting lug bearing capacity based on nanostran nonlinear technology |
| CN113720707A (en) * | 2021-08-23 | 2021-11-30 | 中国飞机强度研究所 | Design method for large-deformation loading point of structural fatigue test |
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