Nonlinear Buckling Analysis on Welded Airbus Fuselage Panels P. Reimers IWiS GmbH A. Gorba Airbus Deutschland GmbH
Contents Overview FE - Model Analysis Results Comparison to Test Results Conclusion Discussion
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
A318 rear fuselage welded panel location
Airbus panel test facility at AI-G
panel loading F Co - compression load F Shv1,2 - vertical shear load F Shh - horizontal shear load
typical buckling shapes 100% shear 100% compression
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%
FE panel model, mesh
FE panel model, detail frame clip stringer skin
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)
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
100% compression load, buckling shape
100% compression load, compression stress
100% compression load, v. Mises stresses
100% compression load, weld stresses
100% compression load, weld stresses, detail
100% shear load, buckling shape and deformation
100% shear load, compression stress
100% shear load, shear stress
100% shear load, v. Mises stresses
100% shear load, weld stress
combined load, buckling shape and deformations
combined load, compression stresses
combined load, v. Mises stresses
combined load, weld stresses
combined load, weld stresses, detail
combined load, load vs. displacement
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
FEA characteristics ADINA FE-code used 124.000 DOF 24.000 nodes 8.700 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
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.