Raghu V Prakash

Flexure behavior of a carbon fiber reinforced plastic (CFRP) prepreg system, AS4/914, with five different layup sequences was investigated under static and fatigue loads to simulate typical loading geometry in an aircraft wing. The laminates were Unidirectional (UD), Cross Ply (CP), Angle Ply (AP), Laminate-1 (L1) and Laminate-2 (L2) with the layup sequences of [0/0]4S, [0/90]4S, [45/−45]5S, [0/45/−45/90]2S and [0//90/45/−45]2S, respectively. The static behavior of the laminates were explained in terms of load–deflection curves, stress–strain curves and these behavioral aspects were compared with the analytical results obtained from the classical lamination theory (CLT). The highest strength with lowest deflection was observed for UD laminates, whereas AP laminates exhibited lowest strength with highest deflection. The CP laminate exhibited good strength, excellent deflection and progressive failure mechanism. The multilayered laminates L1 and L2 exhibited moderate strength. Stiffness degradation model was used to evaluate the progressive fatigue damage in these laminates at a particular load level. The fatigue damage results depicted that performance of L2 was superior in Stage 1, whereas, UD and CP gave steady state growth in Stage 2. The failure mode and the fracture features were evaluated through microstructural studies. A reasonably good correlation could be established between load–deflection curves and macroscopic features of the tested specimens.

A study has been conducted to understand the differences in failure behaviour of a carbon fiber reinforced plastic (CFRP) material fabricated in two different laminate sequences. The material chosen was AS4 with the laminate systems as (0/45/-45/90) and (0/90/45/-45), both of quasi-isotropic (QI) nature. Carefully fabricated test coupons were tested under tensile, compression and flexure load conditions. The tensile and compression test results did not show any influence of the laminate system. However, the flexure loads show a consistent difference in failure behaviour of these two systems, owing to the role of out-of-plane-loading in QI ply. Classical lamination theory (CLT) and the numerical methods were employed to conduct stress analysis. Experimental results are discussed in comparison with the finite element method (FEM) and the analytical results.

The void formation in FRP composites is unavoidable and the void content measurement is very important to study its deleterious effects on the mechanical properties of the material. Generally destructive methods are used to calculate the void volume fraction. But the recent advances in X-ray computed tomography can be used to detect and quantify the void content in the composites in a non-destructive manner. In this study average area method is proposed and validated for the void volume measurement from the X-ray CT image slices through digital image processing. The effects of void size, shape and position in the accuracy of the measurement is studied and presented here. The void volume fraction of a CFRP laminate manufactured with compression molding technique is calculated by this method and found to be around 1%.

The thermo-mechanical response of carbon fiber reinforced polymer (CFRP) laminates subjected to continuous tensile loading and programmed interrupted tensile loading is examined to understand the changes due to damage progression. Quasi-isotropic laminates were prepared using 500 GSM twill weave carbon fabric with LY 556 resin and HY 991 hardener by hand lay-up technique, followed by curing under hot compression. A few specimens were subjected to an impact loading to 23 J and 51 J energy levels using a hemispherical tip to induce low velocity impact damage. Passive thermal imaging of woven CFRP laminates during tensile testing was captured using a TIM 160 Micro-epsilon infrared thermal camera. Temperature response during tensile testing provided a good correlation with deformation mode esp. for specimens impacted with 51 J of energy. Tensile tests were interrupted at periodic loads and unloaded and reloaded to study the thermal response after prior plastic deformation damage in the specimen. Unlike the case of GFRP specimens, distinct changes in thermo-elastic slope due to prior plastic deformation damage could not be clearly identified. As impact damage resulted in de-lamination of some layers, active thermography technique was used to study the rate of cooling of specimen with time when the damage is closer to the camera face as well as when it is away from the camera face. The cooling curves obtained were found to be dependent on the location of the damage, as well as on heating face of the specimen.

Carbon fiber reinforced plastic (CFRP), widely used in aircraft structures is susceptible to accidental low velocity impacts during manufacture or while in service. The damage due to impact, though small initially, can progress due to fatigue loading to cause final fracture. In this study, the CFRP laminates subjected to three different energy levels of low velocity impact were tested under constant amplitude and aircraft fatigue spectrum loading conditions till it reached the failure state. The infrared thermography NDT technique was used to understand the damage distribution across the specimen after fatigue loading of impacted specimens. The cooling response curve of specimens was obtained by the active thermography technique for different experimental conditions, namely, heat transfer by the transmission or reflection mode and impacted or un-impacted surface facing camera. The cooling response obtained when the impact damage as well as heating was on the rear surface provided a good correlation with the damage volume quantified through X-ray CT image processing. The temperature rise observed by passive thermography technique at the fiber/ply breakage during static residual strength tests was found to be proportional to the load drop and extent of pre-existing damage.

Composite materials are used extensively in high-performance applications due to their superior mechanical properties and lower density. The composite materials are susceptible to impact damage, and repair of the damaged components is inevitable. Bonded patch repairs are used extensively in the aerospace industry to repair the low-velocity impact damage. Low-velocity impact on quasi-isotropic, carbon fiber reinforced polymer(CFRP) composites was numerically simulated using finite element analysis ev(FEA). Damage evolution and the failure of the repaired composite laminates under a monotonous uniaxial loading were evaluated using FEA for various patch repair configurations. Load displacement characteristics of pristine, impacted, and repaired specimens were simulated. The analysis indicates that impacted specimen has a lower load-bearing capacity than the pristine one. The specimen with the double-patch displayed superior load-bearing capacity and stiffness compared to the other repair configurations among the repaired specimens. Variation of stress and strain along the longitudinal and transverse directions were evaluated. The stress distribution in the pristine specimen was more uniform than the impacted one. The peak strain values of impacted and repaired specimens were significantly higher than that of pristine ones at the impacted location. A reduction of stress at the impacted location was noticed for the impacted and repaired specimen compared to the pristine one. The current study indicated comparable strength and rigidity for the repaired component and the pristine component. However, the localized stress distributions observed for the pristine, impacted, and repaired specimens were significantly different even though the overall load-bearing characteristics were comparable.

The carbon fiber-reinforced polymeric (CFRP) composite materials are the material of choice for the aircraft structures as the designers require lightweight structures with enhanced mechanical properties. These materials are susceptible to accidental impacts during service and maintenance, and the damage will progress under varying static or dynamic service load conditions leading to the ultimate failure of the component. Recent advancement in non-destructive techniques such as X-ray computed tomography provide excellent details about the presence of damages in 3-Dimension in a component, which is an useful input for failure prediction and remaining life estimation. However, the quality of X-ray CT imaging is dependent on the equipment used, its calibration and image settings which, in turn, may affect the reliability and repeatability of damage quantification, if damage analysis is done in a routine way using binarization algorithms. In this study, the defects as well as the damage present in the low-velocity impacted CFRP laminates subjected to fatigue loading conditions are quantified and analyzed by the analysis of CT scan images obtained from two different CT systems with images of different resolution and contrast. The results of the comparative study show that the damage analysis of polymer composites using X-ray CT depends largely on the image quality and the choice of right threshold level is important for accurate damage estimation.