Next Generation Composite Aircraft Fuselage Materials under Post-crash Fire Conditions
This paper summarizes a series of small-scale tests carried out to evaluate and model the post-crash fire integrity of composite aircraft fuselage structures.
The US Federal Aviation Administration regulations for the penetration of an external fuel fire into an aircraft cabin after crash require a burn-through period of 4min (FAA § 25.856 Appendix F, Part VII). Different candidate structures for the next generation of composite aircraft fuselage, provided by Airbus, were investigated, including CFRP monolithic laminate and a folded-core CFRP sandwich. Those materials were subjected to constant heat flux from a propane gas burner, while being held under compressive load in a small, specially designed compression test rig with anti-buckling guides. The propane burner was calibrated to produce a constant heat flux up to 182kW/m2. The sample time-to-failure was measured, along with the temperatures at various points through the thickness.
Modelling the thermal and structural behaviour under load required the use of a modified version of the Henderson Equation, which describes heat transfer through composites under ablative fire conditions. This has been incorporated into the Com-Fire software model. Kinetic parameters for the resin decomposition reaction were determined from thermo-gravimetric data and other thermal parameters, conductivity and diffusivity were measured experimentally. The paper will compare the behaviour of single and double-skinned structures and will examine measured and modelled behaviour.
KeywordsComposite structures Aerospace Fire resistance Mechanical property
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