A "Computational Design Engine" for multidisciplinary design and optimisation of aeronautical products, specially tailored to the need s of a multi-model, multi-level, multi-site environment, is described. The system is illustrated with an application to the Breguet range optimisation of a Blended Wing Body configuration.

Composite laminates are being increasingly used in a wide variety of industrial applications, but there are difficulties in applying these materials in ways that exploit their full potential, in particular under multi-axial loading. The objective of the present study is to determine by experiments the biaxial failure data for composite laminates produced by Fokker Aerostructures based on the thermoplastic UD carbon reinforced material AS4D/PEKK-FC. A test machine and accompanying cruciform specimens for in-plane biaxial failure tests have been developed. A coupon-level biaxial test program covering various biaxial load combinations in tension-tension, tension-compression and compressionÀcompression has been successfully executed and biaxial failure values for the thermoplastic laminate have been determined. Besides, the experimental biaxial test program, also finite element models and analyses have been used to predict the global outcomes of the biaxial tests and to interpret the test results. Both plain (un-notched) and open-hole (notched) specimens of the thermoplastic laminate have been tested. The biaxial failure data have been collected and further processed in biaxial failure criteria. From the experiments, the failure strains, stresses and loads are determined and a failure envelope is created for both plain and open-hole specimens. Good agreement is found between the theoretically predicted envelopes and the test data. From the findings for biaxial failure criteria from this study, it is expected that structural weight saving can be achieved in the design of multi-axially loaded composite parts as compared to the design with the previous uni-axially based failure criteria.

This paper presents an innovative optimisation method for aircraft fuselage structural design. Detailed local finite element analyses of panel buckling are further processed such that they can be applied as failure constraints in the global level optimisation. The high computational costs involved with the finite element analyses are limited by advanced use of surrogate modelling methods. This yields high flexibility and efficiency in the local level optimisation procedure and allows for efficient gradient based search methods as well as more costly direct search optimisations like genetic algorithms (GAs). The method is demonstrated on a composite fuselage barrel design case considering common structural sizing variables like thicknesses and stringer dimensions. Optimised barrel designs are obtained where the constraints that are derived from the panel buckling analyses are active. The total computational cost for the complete local and global level optimisation procedures is in the order of days on common-performance hardware.

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