CN114512205A - Thermal buckling critical temperature analysis method for aircraft wall panel - Google Patents
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Abstract
The invention provides a thermal buckling critical temperature analysis method for an aircraft panel, and relates to the technical field of aircraft manufacturing. The method comprises the following steps: s1, integral modeling and numerical analysis of the aircraft panel; s2, testing thermal buckling deformation of the airplane wallboard; s3, analyzing the results of the thermal buckling deformation test of the aircraft panel; s4, analyzing the thermal buckling critical temperature of the aircraft panel by a south-well method; and S5, verifying the aircraft panel thermal buckling critical temperature result. Compared with an inflection point method, the south-well method can directly judge according to the full-field test data of the structure, can obtain the result by directly adopting data processing and analysis, and has the advantages of simple method and high reliability.
Description
Technical Field
The invention relates to the technical field of aircraft manufacturing, in particular to a thermal buckling critical temperature analysis method for an aircraft wall plate.
Background
In order to improve market share and model competitiveness, civil aircraft manufacturers at home and abroad are constantly dedicated to research on aspects such as reducing structural weight, improving structural efficiency, prolonging service life of the aircraft and the like. The composite material has excellent specific strength and specific rigidity, and is used in civil aircraft, the most representative new generation of large civil aircraft Boeing 787 and air passenger A350 has composite material consumption of 50% and 52% of the weight of the aircraft body structure, so that the service life of the advanced civil aircraft body, such as A350, B787, etc., reaches over 90000 flight hours, which is far higher than that of domestic branch passenger aircraft ARJ21-700, MA700 and trunk passenger aircraft C919.
The large application of the composite material becomes an important mark for measuring the advancement of a new generation of civil aircraft and is one of the key factors for competing for the market share of a new round of international civil aircraft. One of the key issues in composite and metal aircraft panel design is the thermal stress problem, which is mainly the large difference in thermal expansion coefficient between the composite and the metal.
Generally speaking, the thermal expansion coefficient of the metal is 10 to 20 times of that of the composite material structure. Changes in ambient temperature necessarily result in significant stresses, also referred to as thermal stresses, in the composite-metal aircraft panel. The thermal stress can reach about 40% of the mechanical stress sometimes, neglecting the thermal stress can lead to great potential safety hazard of the composite material-metal aircraft wall plate structure, when the thermal stress reaches a certain value, the composite material-metal aircraft wall plate structure can generate great transverse deformation, so that the composite material-metal aircraft wall plate structure is warped or bulged, and the phenomenon is called thermal buckling.
On one hand, the thermal buckling behavior is sudden, and the stability of the structure is seriously damaged; on the other hand, under high temperature/low temperature environment, the thermal physical property and mechanical property of the material become complex and have obvious nonlinearity, so that the complexity of the thermal buckling behavior is increased. The load-carrying capacity of the aircraft structure is weakened by the thermal buckling behavior, even the structural integrity and even the safety of the aircraft are seriously threatened, and the research on the thermal buckling performance of the aircraft wallboard structure in a thermal environment needs to be carried out by adopting an effective technical means.
When the hybrid structure aircraft panel thermal buckling performance research is carried out, the determination of the thermal buckling critical temperature of the hybrid structure aircraft panel is of great importance for the research of the hybrid structure aircraft panel thermal buckling behavior rule.
Disclosure of Invention
The technical problem solved by the invention is as follows: the thermal buckling behavior weakens the load bearing capacity of the aircraft structure, and poses a serious threat to the structural integrity and even the safety of the aircraft.
In order to solve the problems, the technical scheme of the invention is as follows:
a thermal buckling critical temperature analysis method for an aircraft panel comprises the following steps:
s1, integral modeling and numerical analysis of the aircraft panel, and specifically comprises the following steps:
s1-1, selecting a mixed structure aircraft panel comprising a composite material laminated plate and an aluminum alloy plate reinforcing rib as a test piece, establishing a geometric model of the test piece,
s1-2, assigning the material properties of the composite material laminated plate and the reinforcing rib of the aluminum alloy plate in the calculation process,
s1-3, assuming that the composite material laminated plate in the test piece is completely bonded with the aluminum alloy plate reinforcing rib, carrying out preliminary test piece thermal buckling modal analysis and critical buckling load analysis to obtain the position of the aircraft panel with larger structural deformation;
s2, testing thermal buckling deformation of the aircraft panel: testing the thermal deformation and buckling behavior rules of the structure at the low temperature of-55 ℃, and calculating the three-dimensional deformation measurement of the airplane structure in the low-temperature test by using a non-contact three-dimensional deformation measurement method;
s3, analyzing the results of the aircraft panel thermal buckling deformation test, and specifically comprising the following steps:
s3-1, analyzing the test result of the low-temperature test,
s3-2, analyzing the deformation characteristic point with the maximum deformation in the low-temperature test;
s4, analyzing the thermal buckling critical temperature of the aircraft panel by a south-well method: the thermal buckling critical temperature is judged by utilizing a south-well method, and the ordinate is
The abscissa is
Judging the thermal buckling critical temperature by the reciprocal of the slope of the curve;
s5, verifying the aircraft panel thermal buckling critical temperature result: for the region where buckling deformation occurs most easily, a local flat plate structure is constructed, and the thermal buckling critical load of the local region is calculated.
The South-well method can directly judge according to the structural full-field test data, and has the advantages that the result can be obtained by directly adopting data processing analysis based on the test result, so the processing method is relatively simple, but the requirements on the test process are relatively more, firstly, the area of a test area is large, secondly, the test interval is dense, and the data of the test full process can be ensured to be available.
Further, in step S1-2, the material properties of the composite material laminated plate include: the elastic modulus, Poisson's ratio, thermal conductivity and thermal expansion coefficient of the composite material laminated plate are necessary for analyzing the thermal buckling critical temperature to determine the attribute value of the composite material.
Further, in step S1-2, the material properties of the aluminum alloy plate reinforcing rib include: the elastic modulus, the Poisson ratio, the thermal conductivity and the thermal expansion coefficient of the aluminum alloy material, and the definite property value of the aluminum alloy plate material are necessary for analyzing the thermal buckling critical temperature.
Further, in step S1-3, the thermal buckling modal analysis of the test piece specifically includes the following: introducing a test piece structure into software Abaqus, connecting the composite laminated plate with the aluminum alloy plate reinforcing rib by a binding function in the software Abaqus, simulating that the surfaces of the composite laminated plate and the aluminum alloy plate reinforcing rib are in contact constraint and the periphery of the plate is constrained in a simply supported constraint state, finally calculating a test piece buckling mode and a test piece critical buckling load, wherein in the process of a large-scale aircraft plate heat load test, the aircraft plate is composed of the composite laminated plate and the aluminum alloy plate reinforcing rib, because the thermal expansion coefficients of the composite laminated plate and the aluminum alloy plate are different, the aircraft plate structure deforms under the action of heat stress, the structural buckling instability phenomenon occurs in an area with weak structural rigidity, namely the structure generates out-of-plane displacement, the analysis of the aircraft plate heat buckling mode and the analysis of the critical buckling load are carried out before a low-temperature test, and the position with larger structural deformation of the aircraft plate is obtained, so that the test phase focuses on monitoring the deformation of the area.
Because the high-precision construction of models such as the integral wall plate details of the large-scale aircraft structure is quite complex, the thermal buckling analysis load of the large-scale aircraft wall plate is not very accurate, but the related buckling mode and the initial buckling position of the large-scale aircraft wall plate can provide certain reference for the thermal buckling test of the aircraft wall plate.
Further, step S2 specifically includes the following steps:
s2-1, setting a test area, selecting a plurality of deformation characteristic points in the test area of the test piece according to the results of the thermal buckling modal analysis and the critical buckling load analysis of the test piece in the step S1-3, and continuously observing a deformation curve of the deformation characteristic points in the low-temperature test process;
and S2-2, carrying out low-temperature test, wherein the low-temperature test comprises three stages of cooling, soaking and temperature returning, and each deformation characteristic point is measured by each temperature changing node in the low-temperature test process.
All deformation conditions of each deformation characteristic point in the low-temperature test process can be covered through the steps, and the subsequent analysis of the thermal buckling critical temperature of the test piece is facilitated.
Preferably, in step S2-2, the temperature changing node includes: cooling to 21 ℃, 10 ℃, 0 ℃, -5 ℃, -10 ℃, -15 ℃, -20 ℃, -25 ℃, -30 ℃, -35 ℃, -40 ℃, -45 ℃, -50 ℃, -55 ℃, -50 ℃, -45 ℃, -40 ℃, 0 ℃, 10 ℃ and 21 ℃, covering all temperature variation processes in the low-temperature test process by the nodes, and facilitating the subsequent analysis of the thermal buckling critical temperature of the test piece.
Preferably, step S3-1 specifically includes the following: and acquiring deformation curve graphs of each deformation characteristic point of the test piece measured in the low-temperature test process in the X direction, the Y direction and the Z direction at each temperature point, and analyzing the deformation curve graphs, wherein the Z direction is structure out-of-plane displacement, and multi-angle analysis of the deformation characteristic points is achieved through position deviation in X, Y, Z three directions.
Preferably, step S3-2 specifically includes the following: and analyzing the deformation characteristic point with the maximum deformation in the low-temperature test to obtain a region which is most prone to buckling deformation, and intuitively obtaining the deformation characteristic point through a curve graph.
Further preferably, the step S5 further includes: calculating the thermal buckling critical load of the local area by adopting a thermal buckling theoretical formula (1) of the thin-wall flat plate structure, wherein the thermal buckling theoretical formula (1) of the thin-wall flat plate structure is as follows:
in the formula (1), the reaction mixture is,
in order to obtain the critical load for thermal buckling,
the length of the thin plate is long,
the width of the thin plate is wide,
is the coefficient of thermal expansion of the sheet material,
is the poisson ratio of the sheet material.
The invention has the beneficial effects that:
according to the invention, the thermal buckling critical temperature is analyzed by combining the engineering actual measurement data processing method with the South-well method, the South-well method can be directly judged according to the structural full-field test data, and can be directly used for data processing analysis based on the test result to immediately obtain the result, compared with the prior art, the processing method is simple, the requirements on the test process are relatively more, firstly, the area of the test area must be large, secondly, the test interval requirements are dense, and the real and dense test data enables the analysis result to be closer to the real value;
the method adopts simulation to obtain the approximate integral thermal buckling mode of the aircraft panel at the front part of the method, provides meaningful reference for the layout of monitoring points of subsequent analysis and test, constructs a local flat plate structure, improves the simulation of the structural boundary condition compared with the prior art, and has the advantages of simplifying a model and quickly calculating.
Drawings
FIG. 1 is a flow chart of the present invention;
FIG. 2 is a structural view of a test piece in example 1;
FIG. 3 is a graph of the test area in example 1;
FIG. 4 is a distribution diagram of deformation feature points in example 1;
FIG. 5 is a graph showing the deformation of the deformed characteristic points in the X direction in example 1;
FIG. 6 is a Y-direction deformation curve diagram of the deformation characteristic points in example 1;
FIG. 7 is a Z-direction deformation curve diagram of deformation characteristic points in example 1;
FIG. 8 is a graph of cooling-deformation in the X direction at the deformation characteristic point B in example 1;
FIG. 9 is a Y-direction cooling-deformation curve diagram of the deformation characteristic point B in example 1;
FIG. 10 is a Z-direction cooling-deformation curve diagram of the deformation characteristic point B in example 1;
FIG. 11 is a graph showing the determination of the hot buckling limit temperature South-well in example 1.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention clearer, the present invention will be described in further detail with reference to the accompanying drawings, and it is apparent that the described embodiments are only a part of the embodiments of the present invention, not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
The terminology used in the embodiments of the invention is for the purpose of describing particular embodiments only and is not intended to be limiting of the invention. As used in the examples of the present invention and the appended claims, the singular forms "a", "an", and "the" are intended to include the plural forms as well, unless the context clearly indicates otherwise, and "a plurality" typically includes at least two.
Example 1
The present embodiment is a method for analyzing a critical thermal buckling temperature of an aircraft panel, as shown in fig. 1, including the following steps:
s1, integral modeling and numerical analysis of the aircraft panel, and specifically comprises the following steps:
s1-1, selecting a mixed structure aircraft panel comprising composite material laminated plates and aluminum alloy plate reinforcing ribs as a test piece, establishing a geometric model of the test piece, as shown in figure 2, wherein the composite material laminated plates are 2520mm long, 300mm wide and 3mm thick and are used for simulating the skin of the aircraft panel, each group of aluminum alloy plate reinforcing ribs are 250mm long and 50mm wide, each group of aluminum alloy plate reinforcing ribs are 420mm apart, the composite material laminated plates are connected with the aluminum alloy plate reinforcing ribs through rivets,
s1-2, assigning values to the material properties of the composite material laminated plate and the reinforcing ribs of the aluminum alloy plate in the calculation process, wherein the material properties of the composite material laminated plate comprise: the elastic modulus, Poisson's ratio, thermal conductivity and thermal expansion coefficient of the composite material laminated plate, and the material properties of the aluminum alloy plate reinforcing rib comprise: the aluminum alloy material has elastic modulus, Poisson's ratio, thermal conductivity and thermal expansion coefficient, as shown in tables 1 and 2,
table 1 table of material properties of composite laminates
TABLE 2 Attribute Table of aluminum alloy plate reinforcing rib
S1-3, assuming that the composite material laminated plate in the test piece is completely bonded with the aluminum alloy plate reinforcing rib, carrying out preliminary test piece thermal buckling modal analysis and critical buckling load analysis to obtain the position of the aircraft panel with larger structural deformation, wherein the test piece thermal buckling modal analysis specifically comprises the following contents: introducing a test piece structure into software Abaqus, connecting the composite laminated plate with the aluminum alloy plate reinforcing rib through a binding function in the software Abaqus, simulating that the surface of the composite laminated plate and the surface of the aluminum alloy plate reinforcing rib are in contact constraint, and the periphery of the wall plate is constrained in a simply supported constraint state, and finally calculating a bending mode of the test piece and a critical bending load of the test piece, wherein the bending load is shown in a table 3, the initial bending heat load is-32.3 ℃, and the initial bending temperature during low-temperature test is-11.3 ℃ in a conversion mode;
TABLE 3 aircraft wall thermal buckling load table
S2, testing thermal buckling deformation of the aircraft panel: the method is characterized in that the thermal deformation and buckling behavior rule of the structure is tested under the condition of low temperature of-55 ℃, the three-dimensional deformation measurement of the airplane structure in the low temperature test is calculated by a non-contact three-dimensional deformation measurement method, and the method specifically comprises the following steps:
s2-1, setting a test area, selecting 7 points (A, B, C, D, E, F, G) in the tested area to observe a deformation curve according to the results of the thermal buckling modal analysis and the critical buckling load analysis of the test piece in the step S1-3, continuously observing the deformation curve of the deformation characteristic points in the low-temperature test process, wherein the measurement area is shown in figure 3, the distribution of the deformation characteristic points is shown in figure 4,
s2-2, carrying out low temperature test, wherein the low temperature test comprises three stages of cooling, soaking and temperature returning, each temperature changing node in the low temperature test process measures each deformation characteristic point, and the temperature changing node comprises: cooling to 21 deg.C, 10 deg.C, 0 deg.C, 5 deg.C, 10 deg.C, 15 deg.C, 20 deg.C, 25 deg.C, 30 deg.C, 35 deg.C, 40 deg.C, 45 deg.C, 50 deg.C, 55 deg.C, 50 deg.C, 45 deg.C, 40 deg.C, 30 deg.C, 20 deg.C, 10 deg.C, 21 deg.C;
s3, analyzing the results of the aircraft panel thermal buckling deformation test, and specifically comprising the following steps:
s3-1, analysis of test results of low-temperature test: obtaining deformation curve graphs of each deformation characteristic point of a test piece measured in the low-temperature test process in the X direction, the Y direction and the Z direction at each temperature point, and analyzing the deformation curve graphs, wherein the Z direction is structural out-of-plane displacement, the deformation curve graph of the deformation characteristic point in the X direction is shown in figure 5, the deformation curve graph of the deformation characteristic point in the Y direction is shown in figure 6, the deformation curve graph of the deformation characteristic point in the Z direction is shown in figure 7, according to the analysis of test data, the deformation of the maximum long wall plate structure in the X direction and the Y direction is relatively small and does not exceed 2mm, the out-of-plane displacement (in the Z direction) at the characteristic B point is maximum and reaches 4mm, and the maximum out-of-plane displacement at the test characteristic point B is determined,
s3-2, analyzing the deformation characteristic point with the maximum deformation in the low-temperature test: analyzing a deformation characteristic point B with the maximum deformation in a low-temperature test, generating a structure characteristic point cooling-deformation curve graph of the deformation characteristic point B at different temperature points, wherein the cooling-deformation curve graph of the deformation characteristic point B in the X direction is shown in FIG. 8, the cooling-deformation curve graph of the deformation characteristic point B in the Y direction is shown in FIG. 9, and the cooling-deformation curve graph of the deformation characteristic point B in the Z direction is shown in FIG. 10, so as to obtain an area which is most prone to buckling deformation, and as can be analyzed from FIGS. 8, 9 and 10, the change slope of the deformation of the structure characteristic point B is huge after-15 ℃, and the structure out-of-plane displacement change basically presents a linear relationship;
s4, analyzing the thermal buckling critical temperature of the aircraft panel by using a south-well method: the thermal buckling critical temperature is judged by utilizing a south-well method, and the ordinate is
The abscissa is
Judging the thermal buckling critical temperature by using the reciprocal of the slope of the curve, wherein the slope of the obtained curve is-0.00248 as shown in FIG. 11, the reciprocal of the slope is the thermal buckling critical load, namely-40.3 ℃, and the thermal buckling critical temperature is-19.3 ℃;
s5, verifying the aircraft panel thermal buckling critical temperature result: constructing a local flat plate structure for a region which is most prone to buckling deformation, and calculating the thermal buckling critical load of the local region by adopting a thermal buckling theoretical formula (1) of the thin-wall flat plate structure, wherein the thermal buckling theoretical formula (1) of the thin-wall flat plate structure is as follows:
in the formula (1), the reaction mixture is,
in order to obtain the critical load for thermal buckling,
the length of the thin plate is long,
the width of the thin plate is wide,
is the coefficient of thermal expansion of the sheet material,
is the Poisson's ratio of the sheet material,
input length of 420mm, width of 420mm, thickness of 3mm, α =2.65E-6Mu =0.323, and the critical load of the structure in the region is calculated to be-36.6 ℃, namely the critical temperature of the structure is-15.6 ℃. Therefore, the difference between the thermal buckling critical temperature obtained by adopting an engineering actual measurement data processing method and combining with the south-well method and the result obtained by theoretical calculation is 3.7 ℃, and the error is small.
Claims (8)
1. The method for analyzing the thermal buckling critical temperature of the aircraft wall plate is characterized by comprising the following steps of:
s1, integral modeling and numerical analysis of the aircraft panel, and specifically comprises the following steps:
s1-1, selecting a mixed structure aircraft panel comprising a composite material laminated plate and an aluminum alloy plate reinforcing rib as a test piece, establishing a geometric model of the test piece,
s1-2, assigning the material properties of the composite material laminated plate and the reinforcing rib of the aluminum alloy plate in the calculation process,
s1-3, assuming that the composite material laminated plate in the test piece is completely bonded with the reinforcing rib, carrying out preliminary thermal buckling mode analysis and critical buckling load analysis on the test piece, and obtaining the position of the aircraft panel with larger structural deformation;
s2, testing thermal buckling deformation of the aircraft panel: testing the thermal deformation and buckling behavior rules of the structure at the low temperature of-55 ℃, and calculating the three-dimensional deformation measurement of the airplane structure in the low-temperature test by using a non-contact three-dimensional deformation measurement method;
s3, analyzing the results of the aircraft panel thermal buckling deformation test, and specifically comprising the following steps:
s3-1, analyzing the test result of the low-temperature test,
s3-2, analyzing the deformation characteristic point with the maximum deformation in the low-temperature test;
s4, analyzing the thermal buckling critical temperature of the aircraft panel by using a south-well method: the thermal buckling critical temperature is judged by utilizing a south-well method, and the ordinate is
The abscissa is
Judging the thermal buckling critical temperature by the reciprocal of the slope of the curve;
s5, verifying the aircraft panel thermal buckling critical temperature result: for the region where buckling deformation occurs most easily, a local flat plate structure is constructed, and the thermal buckling critical load of the local region is calculated.
2. The method of analyzing thermal buckling critical temperature of an aircraft panel according to claim 1, wherein in step S1-2, the material properties of the composite laminate include: the composite material laminated plate has elastic modulus, Poisson's ratio, thermal conductivity and thermal expansion coefficient.
3. The method for analyzing the thermal buckling critical temperature of an aircraft panel as claimed in claim 1, wherein in the step S1-2, the material properties of the aluminum alloy reinforcing rib include: the aluminum alloy material has elasticity modulus, Poisson's ratio, heat conductivity and thermal expansion coefficient.
4. The method for analyzing the thermal buckling critical temperature of an aircraft panel according to claim 1, wherein in step S1-3, the thermal buckling modal analysis of the test piece specifically includes the following: and (3) introducing a test piece structure into software Abaqus, connecting the composite laminated plate with the reinforcing rib of the aluminum alloy plate through a binding function in the software Abaqus, simulating the contact constraint of the surface of the composite laminated plate and the surface of the reinforcing rib of the aluminum alloy plate, and calculating the buckling mode of the test piece and the critical buckling load of the test piece, wherein the periphery of the wall plate is constrained in a simply supported constraint state.
5. The method for analyzing the thermal buckling critical temperature of an aircraft panel as claimed in claim 1, wherein the step S2 specifically comprises the steps of:
s2-1, setting a test area, selecting a plurality of deformation characteristic points in the test area of the test piece according to the results of the thermal buckling modal analysis and the critical buckling load analysis of the test piece in the step S1-3, and continuously observing a deformation curve of the deformation characteristic points in the low-temperature test process;
and S2-2, carrying out low-temperature test, wherein the low-temperature test comprises three stages of cooling, soaking and temperature returning, and each deformation characteristic point is measured by each temperature changing node in the low-temperature test process.
6. The method of analyzing thermal buckling critical temperature of an aircraft panel as claimed in claim 5, wherein in step S2-2, said temperature change node comprises: cooling to 21 deg.C, 10 deg.C, 0 deg.C, 5 deg.C, 10 deg.C, 15 deg.C, 20 deg.C, 25 deg.C, 30 deg.C, 35 deg.C, 40 deg.C, 45 deg.C, 50 deg.C, 55 deg.C, 50 deg.C, 45 deg.C, 40 deg.C, 30 deg.C, 20 deg.C, 10 deg.C, 21 deg.C.
7. The method for analyzing the thermal buckling critical temperature of an aircraft panel as claimed in claim 1, wherein the step S3-1 specifically comprises the following steps: and acquiring deformation curve graphs of each deformation characteristic point of the test piece measured in the low-temperature test process in the X direction, the Y direction and the Z direction at each temperature point, and analyzing the deformation curve graphs, wherein the Z direction is structure out-of-plane displacement.
8. The method for analyzing the thermal buckling critical temperature of an aircraft panel as claimed in claim 1, wherein the step S3-2 specifically comprises the following steps: and analyzing the deformation characteristic point with the largest deformation in the low-temperature test to obtain a region which is most likely to generate buckling deformation.
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Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
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