CN113312815B - Method for calculating overall stability of stringers in fuselage structure - Google Patents
Method for calculating overall stability of stringers in fuselage structure Download PDFInfo
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- 238000004364 calculation method Methods 0.000 claims abstract description 36
- 238000005452 bending Methods 0.000 claims description 22
- 238000004458 analytical method Methods 0.000 claims description 18
- 238000010008 shearing Methods 0.000 claims description 10
- 230000002459 sustained effect Effects 0.000 claims 1
- 230000006835 compression Effects 0.000 description 5
- 238000007906 compression Methods 0.000 description 5
- 238000010586 diagram Methods 0.000 description 2
- 238000011160 research Methods 0.000 description 2
- 230000003068 static effect Effects 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
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- 239000013585 weight reducing agent Substances 0.000 description 1
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- G—PHYSICS
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- G06F30/00—Computer-aided design [CAD]
- G06F30/20—Design optimisation, verification or simulation
- G06F30/23—Design optimisation, verification or simulation using finite element methods [FEM] or finite difference methods [FDM]
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Abstract
According to the invention, the minimum compressive stress of the fuselage stringer is calculated according to the stringer compressive stress calculation formula, and finally the total critical unstability stress of the stringer corresponding to the minimum compressive stress and the ratio of the calculated minimum compressive stress are used to obtain the total stability residual strength of the fuselage stringer.
Description
Technical Field
The invention belongs to the field of aircraft structural strength design, and particularly relates to a method for calculating overall stability of stringers in a fuselage structure.
Background
The invention mainly researches a method for calculating the overall stability of a stringer when a fuselage skin is mainly subjected to pure shearing.
In the current strength check, a checking method is generally adopted for the fuselage stringer, a stringer unit and left and right skin units under severe working conditions are taken as research objects, axial compressive stress of the stringer rod unit is directly extracted from an overall stress analysis model, maximum shearing stress of the skin unit is calculated, the compressive stress and bending stress of the stringer caused by a tension field after shearing instability of the skin unit are summed with the compressive stress of a stringer body, the overall stability of the fuselage stringer is calculated, and the physical geometric plastic coefficient of the stringer during bending and the influence of the compressive stress and the bending stress of the stringer on the physical geometric plastic coefficient are not considered during calculation.
For the fuselage stringers, the overall stability of the fuselage stringers is checked directly by adopting the stress calculation result of the overall stress analysis model, and the accuracy of the obtained result is lower. Therefore, a new calculation method is needed to be explored, and the overall stability of the stringer of the fuselage structure is calculated conveniently, rapidly and accurately.
Disclosure of Invention
The invention aims to provide a method for calculating the overall stability of a stringer in a fuselage structure, which is used for calculating static strength based on a finite element analysis method and aims to improve the calculation accuracy of the overall stability of the stringer when a skin is mainly sheared purely.
A method for calculating the overall stability of a stringer in a fuselage structure, the fuselage structure comprising a stringer, a skin and a fuselage frame, wherein the design digital-analog and load conditions of the fuselage structure are known, is characterized by comprising the following steps: 1) According to the design counting model of the fuselage structure, taking intersection points of the stringers and fuselage frames as references, establishing a finite element model of the fuselage structure, wherein the skin is simplified into plate units, the stringers are simplified into rod units, the fuselage frames are simplified into beam units, a calculation unit is formed between adjacent beam units by the rod units and the plate units on two sides of the rod units, and each calculation unit comprises the stringers in the middle position and the skin on two sides of the stringers; 2) In the finite element model, the frame beam unit of the machine body close to the machine head is simply supported and restrained to be used as a boundary condition of the overall stress analysis model of the machine body structure; 3) According to the boundary conditions and the load working conditions, stress calculation results of all calculation units of the fuselage structure are obtained; 4) According to the stress calculation results of the calculation units, extracting the calculation units with the skins subjected to pure shearing as analysis units; 5) Calculating each analysis unitThe compressive stress of the stringer is compared and analyzed, and the minimum compressive stress sigma of the stringer is extracted c The method comprises the steps of carrying out a first treatment on the surface of the 6) Calculating the critical instability stress sigma of the stringer by utilizing the stringer geometric parameter corresponding to the minimum compressive stress cf The method comprises the steps of carrying out a first treatment on the surface of the 7) Critical buckling stress sigma of stringers corresponding to minimum compressive stress cf And ratio sigma of minimum compressive stress of stringers c As the overall stability residual strength of the fuselage stringers.
In step 5) above, the stringer compressive stress σ of each analysis unit c Comprises three parts: the first part is the stringer compressive stress sigma calculated in the fuselage structure overall finite element model ch The method comprises the steps of carrying out a first treatment on the surface of the The second part is stringer average compressive stress sigma caused by diagonal stretching after skin shear instability st The method comprises the steps of carrying out a first treatment on the surface of the The third part is the bending stress sigma of the stringer at two fulcrums under inward bending load M 。
According to the method for calculating the overall stability of the stringer in the fuselage structure, a finite element model of the fuselage structure is established according to a design digital model of the fuselage structure, calculation units are formed by a rod unit and plate units on two sides of the rod unit, stress calculation results of all calculation units of the fuselage structure are obtained according to boundary conditions and load working conditions, geometric parameters of the calculation units corresponding to the maximum shear stress plate units and the maximum compression stress rod units are utilized, the minimum compression stress of the stringer of the fuselage is calculated according to the stringer compression stress calculation formula, and finally the overall stability residual strength of the stringer of the fuselage is obtained according to the ratio of the overall critical unstability stress of the stringer corresponding to the minimum compression stress and the calculated minimum compression stress.
The beneficial effects of this application lie in: 1) According to the invention, by utilizing a finite element analysis method, through accurate simplification and constraint of the fuselage skin, the fuselage frame and the stringers, the overall stability of the stringers in the fuselage structure is calculated, and a calculation unit of the skin subjected to pure shearing is extracted as an analysis unit, so that the method has the advantages of simple algorithm and accurate result; 2) In the calculation process, the plasticity of the material of the stringers in the analysis unit during bending and the stress redistribution condition after the stringers enter the plasticity are fully considered; 3) According to the actual loading condition of the structure, the influence of the compressive stress and the bending stress of the stringer caused by a tension field on the physical geometric plastic coefficient is fully considered, and a correction method of the physical geometric plastic coefficient of the stringer during bending is provided; 4) The overall stability calculation precision of the stringers when the fuselage structure skin is mainly subjected to pure shearing is improved, and the structural strength can be calculated more accurately and the weight reduction design of the structure can be carried out.
The present application is described in further detail below with reference to the examples and the accompanying drawings.
Drawings
FIG. 1 is a schematic illustration of a typical airframe structure;
FIG. 2 is a schematic diagram of a finite element analysis model of a typical fuselage structure;
fig. 3 is a schematic diagram of a computing unit.
The numbering in the figures illustrates: 1 skin, 2 stringers, 3 fuselage frames, 4 calculation units.
Detailed Description
Referring to the drawings, a typical fuselage structure contains a monolithic skin 1, a plurality of longitudinal stringers 2 and a plurality of transverse fuselage frames 3. The method for calculating the overall stability of the stringers in the fuselage structure comprises the following specific steps:
the static strength calculation is carried out based on a finite element analysis method, and the purpose is to improve the calculation accuracy of the overall stability of the fuselage stringer when the skin is mainly subjected to pure shearing.
According to the design digital model of the fuselage structure, taking the intersection point of the stringer 2 and the fuselage frame 3 as a reference, a finite element model of the fuselage structure is established, wherein the skin 1 is simplified into a plate unit, the stringer 2 is simplified into a rod unit, the fuselage frame 3 is simplified into a beam unit, the rod unit and the adjacent plate units form calculation units 4, and each calculation unit comprises the stringer 2 at the middle position and the skins 1 at two sides of the stringer 2, as shown in figure 3.
During analysis, the frame 3 at the end of the fuselage is simply supported and constrained to serve as a boundary condition of the overall stress analysis model of the fuselage structure.
And obtaining the stress calculation result of each calculation unit of the fuselage structure according to the boundary condition and the load working condition.
Extracting different load conditionsThe lower skin 1 is mainly subjected to pure shearing by a calculating unit 4, the shearing stress of the skin 1 and the calculating stress of the stringers 2 are extracted, and the compressive stress of the stringers calculated by the overall finite element model is recorded as sigma ch The method comprises the steps of carrying out a first treatment on the surface of the The average compressive stress of the stringers caused by diagonal tension after shear instability of the skin is noted as sigma st The method comprises the steps of carrying out a first treatment on the surface of the The third part is that the bending stress of the stringer under inward bending load at two fulcrums is recorded as sigma M Ultimate load bearing stress of the stringer is sigma b 。
According to the geometric plastic coefficient of various section bending of chapter 17-22 of book 9 of the aircraft design manual, the geometric plastic coefficient alpha of the I-beam during bending is obtained, and according to the physical geometric plastic coefficient (K) of chapter 17 of book 9 of the aircraft design manual, chapter 17 of the drawing 17-19 P - α curve) "to derive the physical geometric plasticity coefficient K of the stringer upon bending P 。
Considering stringer compressive stress sigma ch And diagonal tension induced stringer average compressive stress sigma st The maximum bending stress that the stringer can withstand when bending is:
σ max ＝K P σ b -σ st -σ ch ；
considering stringer compressive stress sigma ch And diagonal tension induced stringer average compressive stress sigma st After the influence of (a), the physical geometric plastic correction coefficient of the stringer when bending is:
calculating a minimum calculated stress sigma for the stringers according to the following formula c ：
According to the general stability formula of section bar of chapter 21.4 section bar of book 9 of aircraft design manual, the geometric parameters of stringer 2 are used to calculate the general critical instability stress sigma of the stringer cf 。
Finally, the minimum stress corresponds to the total critical buckling stress sigma of the stringer 2 cf And minimum calculated stress sigma c To yield the overall stability residual strength of the fuselage stringers.
Claims (1)
1. A method for calculating the overall stability of a stringer in a fuselage structure, the fuselage structure comprising a stringer, a skin and a fuselage frame, wherein the design digital-analog and load conditions of the fuselage structure are known, is characterized by comprising the following steps: 1) According to the design counting model of the fuselage structure, taking intersection points of the stringers and fuselage frames as references, establishing a finite element model of the fuselage structure, wherein the skin is simplified into plate units, the stringers are simplified into rod units, the fuselage frames are simplified into beam units, a calculation unit is formed between adjacent beam units by the rod units and the plate units on two sides of the rod units, and each calculation unit comprises the stringers in the middle position and the skin on two sides of the stringers; 2) In the finite element model, the frame beam unit of the machine body close to the machine head is simply supported and restrained to be used as a boundary condition of the overall stress analysis model of the machine body structure; 3) According to the boundary conditions and the load working conditions, stress calculation results of all calculation units of the fuselage structure are obtained; 4) According to the stress calculation results of the calculation units, extracting the calculation units with the skins subjected to pure shearing as analysis units; 5) Calculating the compressive stress of the stringers of each analysis unit, comparing and analyzing the calculation results of the compressive stress of the stringers, and extracting the minimum compressive stress sigma of the stringers c Stringer compressive stress sigma for each analysis unit c Comprises three parts: the first part is the stringer compressive stress sigma calculated in the fuselage structure overall finite element model ch The method comprises the steps of carrying out a first treatment on the surface of the The second part is stringer average compressive stress sigma caused by diagonal stretching after skin shear instability st The method comprises the steps of carrying out a first treatment on the surface of the The third part is the bending stress sigma of the stringer at two fulcrums under inward bending load M The method comprises the steps of carrying out a first treatment on the surface of the 6) Calculating the critical instability stress sigma of the stringer by utilizing the stringer geometric parameter corresponding to the minimum compressive stress cf The method comprises the steps of carrying out a first treatment on the surface of the 7) Critical buckling stress sigma of stringers corresponding to minimum compressive stress cf And minimum compressive stress sigma of stringers c As the overall stability residual strength of the fuselage stringer, the stringer compressive stress sigma is calculated c Consideration should be given to the timePhysical geometric plasticity coefficient K of stringer upon bending P For bending stress sigma M And stringer compressive stress sigma calculated in the fuselage structure overall finite element model ch And stringer mean compressive stress sigma due to diagonal tensile load after skin shear failure st Coefficient of physical geometric plasticity K for stringers when curved P The ultimate load bearing stress of the stringer is sigma b Stringer compressive stress sigma c The specific calculation method of (2) is as follows:
first consider the stringer compressive stress sigma ch And diagonal tension induced stringer average compressive stress sigma st The maximum bending stress sigma that can be sustained when the stringer is bent is calculated according to the following formula max ：
σ max ＝K P σ b -σ st -σ ch ；
Considering the stringer compressive stress sigma ch And diagonal tension induced stringer average compressive stress sigma st The physical geometric plastic correction coefficient K of the stringer during bending is calculated according to the following formula P1 ：
Calculating a minimum calculated stress sigma for the stringers according to the following formula c ：
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- 2021-05-19 CN CN202110547079.XA patent/CN113312815B/en active Active
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| WO2011030079A2 (en) * | 2009-09-14 | 2011-03-17 | Airbus Operation (S.A.S) | Method for the structural analysis of panels consisting of an isotropic material, and stiffened by triangular pockets |
| WO2018028284A1 (en) * | 2016-08-09 | 2018-02-15 | 苏州数设科技有限公司 | Method and device for creating strength model of aircraft structure |
| WO2019050176A1 (en) * | 2017-09-07 | 2019-03-14 | 주식회사 엘지화학 | Structural analysis tool for mono frame and method for designing mono frame |
| CN111914351A (en) * | 2020-07-06 | 2020-11-10 | 西安飞机工业（集团）有限责任公司 | Method for calculating overall stability of reinforced wall plate of fuselage structure |
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