Nonlinear Buckling Analysis on Welded Airbus Fuselage Panels

Transcription

1 Nonlinear Buckling Analysis on Welded Airbus Fuselage Panels P. Reimers IWiS GmbH A. Gorba Airbus Deutschland GmbH

2 Contents Overview FE - Model Analysis Results Comparison to Test Results Conclusion Discussion

3 Welded Panel Application Airbus A318 worldwide first application in large scale serial production front and rear fuselage sections Airbus A380 under design other Airbus types

4 A318 rear fuselage welded panel location

5 Airbus panel test facility at AI-G

6 panel loading F Co - compression load F Shv1,2 - vertical shear load F Shh - horizontal shear load

7 typical buckling shapes 100% shear 100% compression

8 load combinations compression load shear load 0,0% 100,0% 41,6% 77,6% 79,7% 47,7% 87,3% 34,8% 94,9% 25,3% 100,0% 0,0%

9 FE panel model, mesh

10 FE panel model, detail frame clip stringer skin

11 criteria of failure load assessment global: collapse of structure load decreases while deformations increase local: material yields local damage growth causes global collapse rivet failure plastic zones (e.g. at top of skin buckles) crack growth (e.g. in welds)

12 weld failure weld opening stresses σ w = (σ zz 2 + τ xz 2 + τ yz2 ) 1/2 procedure locate ε zz, max in stringer weld calculate weld opening stress σ w interpolate global load for σ w = R m, Stringer at ε zz, max location

13 100% compression load, buckling shape

14 100% compression load, compression stress

15 100% compression load, v. Mises stresses

16 100% compression load, weld stresses

17 100% compression load, weld stresses, detail

18 100% shear load, buckling shape and deformation

19 100% shear load, compression stress

20 100% shear load, shear stress

21 100% shear load, v. Mises stresses

22 100% shear load, weld stress

23 combined load, buckling shape and deformations

24 combined load, compression stresses

25 combined load, v. Mises stresses

26 combined load, weld stresses

27 combined load, weld stresses, detail

28 combined load, load vs. displacement

29 comparison of results 100,0% shear / compression interaction curve based on panel failure load 80,0% Test Results Analyses Results shear load 60,0% 40,0% 20,0% 0,0% 0,0% 20,0% 40,0% 60,0% 80,0% 100,0% compression load

30 FEA characteristics ADINA FE-code used DOF nodes elements 8-node shell elements for skin, stringer and frame 3 x 3 Gaussian integration order in-plane 3 Newton-Cotes integration scheme through thickness 1st mode shape of linear buckling analysis used as geometric imperfection nonlinear static buckling analysis with automatic load-displacement ruled step incrementation large displacements and elastic-plastic material behaviour included 3-5 h CPU-time on 1,5 GHz Linux computer

31 Conclusion Fuselage panel tests are valuable in order to optimise the structural design. They are necessary as part of the approving procedure for licensing authorities. FE analysis of test panels provide a more detailed look into the panel behaviour than displacement and strain gauges of limited number can offer. FE analysis serve to minimise test costs by replacing some test load combinations. Good analyst experience is necessary in the subject of nonlinear FE analysis conduction and in modelling the specific boundary conditions of the panel test rig.