This paper demonstrates the results of an experimental study on cross ply carbon/epoxy composite laminates fabricated from high temperature hardener HT972 subjected to impact loading at different velocities and temperatures. The carbon fiber reinforced plastic (CFRP) samples were impacted at velocities 1.5 m/s and 2.5 m/ s, each at a temperature level of 308C, 608C, 908C, and 1208C. The impact response of the material towards various velocities and temperatures was determined using impact parameters like peak force, absorbed energy, maximum deflection, and rebound velocity. Result reveals that the velocity and temperature play a significant role in the impact response of the material. The variation in the trend of Flexural After Impact (FAI) strength of composite laminates at different velocities and temperatures was determined using FAI test and these results were further correlated with impact results. The dominating failure modes affecting the residual strength of the samples were found using acoustic emission (AE) monitoring. POLYM. COMPOS.,

The Flexural After Impact (FAI) behaviour of epoxy and vinyl-ester based carbon fiber reinforced composite laminates was investigated at elevated temperatures. Carbon/Epoxy (CE) and Carbon/Vinyl-ester (CV) laminates with cross-ply configuration (0/90/90/0)3S were manufactured via a compression moulding technique and subjected to Low-Velocity Impact (LVI) s at ∼1.5 and ∼2.5 m/s under temperatures 30, 60 and 90°C. The flexural behaviour of composite laminates was investigated via three-point bending tests. The non-impacted and impacted CE and CV samples' failure profile during the flexural tests was examined using the real-time Acoustic Emission (AE) monitoring technique via peak frequency analysis of AE events. Flexural after impact strength of CE samples at both the velocities were higher than that of the CV ones. For CE and CV samples, the flexural after impact strength increases at 60°C, and decreases when approaching 90°C. At 90°C, flexural strength degradation was considerably higher in the CE ones because the Co-efficient of Thermal Expansion (CTE) of the epoxy matrix occurs at a much higher rate than the vinyl-ester, which generates higher residual stress at the carbon fiber-epoxy interfaces. Acoustic emission (AE) monitoring allowed to capture the interface among the fiber and matrix due to the exposure temperature and impact velocity.

Generally, the aircraft structural parts are economically high in cost so the materials need to be inspected for defects or damages using various non-destructive testing (NDT) methods like ultrasonic, thermography and acoustic emission. The aim of this project is to characterize the defects in composite laminates before and after the flexural loading using infra-red thermography NDT method. GFRP and hybrid (GFRP+CFRP) composite laminates are fabricated with different orientation such as uni-directional, cross ply, anti-symmetric and angle ply and then tested under flexural loading according to ASTM D790 standard. The volume fraction of the fibre and matrix needs to be found out to know the void content and the mixing ratio of reinforcement and binder.