I have a global wing model, with plate elements for skin, ribs and spars and beam/bar elements for stringers and other longitudinal members.
which is the usual way for the estimation of the loads of a group of fasteners on a wing spar panel? should I use the nodal interface loads or estimate from the elemental stresses? how?
The panel is delimited between the vertical longeron stiffeners and the longitudinal ones.
which is the usual way for the estimation of the loads of a group of fasteners on a wing spar panel? should I use the nodal interface loads or estimate from the elemental stresses? how?
The panel is delimited between the vertical longeron stiffeners and the longitudinal ones.
