In order to estimate the edge stresses at the corners of aircraft doors, a more detailed model of the door area was generated. Several parts, which were incorporated into the skins or frames in the global model, needed to be represented explicitly—due to some parts being wedged between others—where cracks are difficult to notice during visual or dye-penetrant inspections.

All the newly detailed parts were positioned at the layer of adjacent fuselage skins and the connecting fasteners were modelled with zero-length CBUSH elements.

In order to determine the stresses on the door edges, thin bar elements were added between consecutive nodes along each edge. The material properties of these bars were identical with that of the adjoining skins. Depending on location, the length of these elements varied between 0.05 and 0.2 inch. A similar approach was used for other internal parts ending at, or near to, the edges of the door cut-out.

In this new discretisation, the sub-model was solved using MSC Nastran Solution 101 for all fatigue-relevant load cases, and deformations at the door markers were compared against their values in baseline configuration to ensure that any differences were within acceptable limits. Next, the stress data from the bar elements was processed in order to generate the the fatigue stress spectra for the critical mission profile of different locations, and the fatigue strength at the cut-out was conducted using Boeing’s methods and bespoke software. The specific methodology details used during post-processing are proprietary information of the Boeing Commercial Aircraft company.