Carbon Fiber Reinforced Plastics (CFRP) have become indispensable structural materials in the aerospace industry.The ability to tailor the stiffness and strength of composite laminates is a unique advantage over metals allowing for optimization at the material level.
The ability to tailor the stiffness and strength towards for applications has been limited by the standardization of laminate design by the aerospace industry. Laminate design of has been standardized into a set of rules, which limit the options available to a designer. Current laminate design allows for alignment of fiber reinforcement along only four standard directions. Other restrictions such as symmetry about the mid plane and balance are also strictly enforced. This thesis studies the effect of using non-standard (NS) angles for the ply orientation when designing a laminate.
To develop NS designs which can be easily compared with standard designs, this study proposes a stiffness matching method. This method allows one to design non-standard laminates that match the in-plane stiffness of standard composite layups. This method has then been validated against a stiffness matching method proposed prior to this work.
NS designs that match the stiffness of a typical wing skin layup were developed based on the stiffness matching method proposed in this thesis. The theoretical strengths of these NS designs were calculated based on the First Ply Failure (FPF) theory. Based on the significantly improved theoretical strengths of the NS designs when compared to the standard wing skin design, physical samples of the standard and NS designs were fabricated and tested.
As notched compressive strength is often the limiting factor in composite structures, Open Hole Compression (OHC)testing has been carried out on non-standard and standard designs. The failure modes and failure strengths for the designs were analyzed. The NS designs were observed to be weaker due to fiber discontinuity.