Two main aspects are highlighted in this paper: the manufacturing process of composite fuselage stiffened panels and a theoretical approach focused on damage tolerance purpose. New concepts of omega (and blade) stiffened panels for composite fuselage application have been designed, sized, produced and tested. Several manufacturing processes have been tested, such as Automatic fibre placement, pre-forming (hot forming) of the composite omega stiffeners, tools for Outer and Inner mould line concept, different mandrel manufacturing processes, co-curing and co-bonding processes. For damage tolerance demonstration, the criterion of no-growing delamination under imposed loads (fatigue and static ones) is selected as one of the most important criteria to be checked on the produced panels. The proposed approach is based on the Maximum Energy Release Rate in mixed mode (GI/GII) around the delaminated area. Firstly, coupon tests have been performed in order to measure the critical energy release rate in mixed mode. Those tests allow to define a simple relationship to predict the delamination instability for any mixed mode ratio. Secondly, the same concept is extended and applied to the large fuselage panels. These ones are impacted and used as specimens for compression and shear tests. The considered impact damages are Large Visible Impact Damage (LVID) with a high energy impact level, so that they are responsible for composite delamination. The approach is validated through a comparison between the test results and numerical predictions. A good correlation has been found.
展开▼
