An investigation into the fatigue behavior of micro-cracks in resin rich regions of composite laminates produced by liquid infusion

Damage is found in resin rich bead corner radii of RTM 6 epoxy-based composite ribs due to in-service thermal-mechanical loading after an aircraft inspection. The same types of damage are obtained through pure thermal cycling of a single bead. However, thermal cycling is a time-consuming process. Therefore, a faster way of damage investigation is required. A potential way to achieve this objective is by mechanically cycling a coupon specimen with simpler geometry than the bead. To investigate potential solutions to the problem a literature review is performed. The scope of the literature review covers several topics as follows. Laminate fatigue damage modes and their impact on the laminate is researched. It is discovered that it is typical for RTM 6 epoxy-based laminates to build-up high matrix residual tensile stresses after manufacturing. Several differences between thermal and mechanical cycling are discovered. It is found that the most common way to test composites is by the use of ASTM standards. However, there is little available information about the fatigue behavior of laminates with resin rich areas. Investigation of the fatigue behavior of laminates with resin rich areas is performed by the use of FEA and physical tests. Four specimen types labeled from A to D are manufactured. Type A is a dog-bone pure RTM 6 specimen. Types B to D are all composite specimens with the same in-plane geometry and different layups and manufacturing processes. All specimens are tested statically and in fatigue. In the fatigue testing session, fractography of the damage occurring at different test conditions for different specimen types is performed. In addition, two FE models are created. The first model is of the bead. The second model is a harmonized model applicable to all composite specimen types with required layup readjustments for each specimen type. It is discovered by FEA that the maximum principal stress in the resin rich area is perpendicular to the matrix cracks in the bead and the composite specimens. It is also discovered that the maximum principal stress cycle of the matrix at the resin rich layer interface with the fabric is similar in the bead and the specimens. The similar fatigue parameters are the R-ratio and the stress amplitude. The parameter similarity suggests they could potentially drive the resin rich layer fatigue damage initiation. Moreover, a positive through-the-thickness stress gradient is discovered, which suggests the cracks are likely to initiate bellow the resin rich layer surface. This hypothesis is further supported by fractographic observations of cracks not reaching the laminate free surface. Static and fatigue tests are performed. The static test provides the UTS of all specimen types, based on which cyclic load levels are selected. In the specimen fatigue tests several results are observed. In the first place, damage similar to the bead damage is found, namely cracks and delamination. In the second place, the damage is observed to penetrate through the laminate thickness and to be dependent upon the laminate compaction. However, this penetration depth dependency on the compaction might be influenced by the second curing cycle of specimen type C, in which the damage was observed. Finally, reduction in the specimen stiffness is observed due to fatigue damage accumulation in time. Based on the results recommendations are formulated. For design purposes resin rich area formation should be avoided both inside and outside the laminate. If their formation is inevitable, at least the laminate should be kept well compacted and the resin rich area location should be kept only at the surface. For future research, two topics are identified as requiring such. First, is the laminate stiffness reduction. Second, is the influence of the second curing cycle of the well compacted specimen type C.Aerospace EngineeringAerospace Structures & Material