In this paper, we present challenges and achievements in development and use of a compact ultrasonic Phased Array (PA) module with signal processing and imaging technology for autonomous non-destructive evaluation of composite aerospace structures. We analyse two different sets of ultrasonic scan data, acquired from 5 MHz and 10 MHz PA transducers. Although higher frequency transducers promise higher axial (depth) resolution in PA imaging, we face several signal processing challenges to detect defects in composite specimens at 10 MHz. One of the challenges is the presence of multiple echoes at the boundary of the composite layers called structural noise. Here, we propose a wavelet transform-based algorithm that is able to detect and characterize defects (depth, size, and shape in 3D plots). This algorithm uses a smart thresholding technique based on the extracted statistical mean and standard deviation of the structural noise. Finally, we use the proposed algorithm to detect and characterize defects in a standard calibration specimen and validate the results by comparing to the designed depth information.

In this article an attempt has been made to quantitatively assess the extent of delamination in composite laminates, using time-of-flight of fundamental Lamb wave modes, without recourse to baseline data from a healthy structure. An expression has been derived to determine the delamination size, from group velocities of primary Lamb modes in the sub-laminates and time-of-flight of transmitted A_o signal and mode converted A_o signal, which is generated when A_o mode propagates through a delamination. The effectiveness of the expression, when group velocities of primary Lamb modes in the main laminate were used, has been verified through numerical simulations carried out on a quasi-isotropic glass/epoxy laminate with various delamination interfaces. The effectiveness of the expression has also been verified experimentally, on two GFRP cross-ply laminates of [0/90/0] lay-up with 40 mm and 50 mm delamination sizes, using air coupled ultrasonic transducers. The predicted delamination sizes were found to be in good agreement with the actual delamination sizes. Using the proposed technique absolute identification of delamination is possible. A supplementary equation for determining the minimum length of delamination has also been derived and presented in the article.

In this study, a procedure is devised, using the fundamental symmetric Lamb mode (S_o) to estimate the size of an impact-induced delamination at the bottom interface of a composite laminate, which has been subjected to a low-velocity impact. When S _o mode incidents at the edge of delamination, it propagates independently in 0^° layer and the rest of layers. Based on the propagation of S_o mode in the delaminated region, an expression is derived, which can estimate the size of interface delamination using measured arrival times or Time-of-Flight (ToF) of S_o modes. Numerical simulations were carried out on glass/epoxy (GFRP) cross-ply laminates ([0/90₂] _s) to confirm the derived expression and eventually it was validated through experiments carried out on [0/90/0] GFRP laminates. Moreover, the minimum detectable size of delamination through the proposed method was estimated analytically and verified through numerical simulations.

This paper presents an advanced methodology for the detection of damage in aircraft composite materials based on the sensor fusion of two image-based non-destructive evaluation techniques. Both of the techniques, phased-array ultrasonics and infra-red thermography, are benchmarked on an aircraft-grade painted composite material skin panel with stringers. The sensors systems for carrying out the inspections have been developed and miniaturized for being integrated on a vortex-robotic platform inspector, in the framework of a larger research initiative, the Horizon-2020 ‘CompInnova’ project.

Research in quick Non-Destructive Evaluation methods for field maintenance of aircraft has become a focus of attention due to increase in the use of co-cured skin-stiffened composite structures. When Lamb wave propagates over skin-stiffened structure, the occurrence of multiple modes is unavoidable due to structural features such as stringers and stiffeners, which in turn makes the interpretation of the received wave difficult and limits the defect-detection ability of Lamb waves. Using Finite Element simulations, the propagation of incident S₀ Lamb mode in a typical composite T-joint with delamination between the flange and skin (interface) is investigated. Arrival time delays of the out-of-plane mode-converted wave packets are found to be promising indicator for quantitative detection and sizing of the delamination. A novel experimental technique of combining a liquid couplant-free transduction scheme using dry-coupled roller transducer (contact probe) for generating S₀ mode and an air-coupled transducer (non-contact probe) as a receiver is proposed. This transduction scheme helps in selectively separating mode-converted wave packets. This investigation also establishes quantitative B-scan imaging method for characterization of delamination in the T-joint. Experimental results show a good agreement with Finite Element predictions.