Alihan Cambaz

On-site repair of carbon fibre-reinforced polymer (CFRP) structures is widely employed when component removal and autoclave-based processing are impractical. However, existing repair guidelines provide limited guidance on the interaction between closely spaced repairs. The objective of this research is to investigate the coupled effects of ligament spacing, patch lay-up, and repair arrangement on the residual tensile strength of dual stepped-lap repaired composite panels. The methodology consists of a shell-based finite element model combining Hashin intralaminar damage and a mixed-mode bilinear cohesive zone model, which was validated against published experiments on damaged, stepped, and wet lay-up repaired laminates, predicting the measured strengths within 6.4%. Subsequently, fifteen repair configurations were analyzed, comprising four ligament spacings, three patch lay-up strategies, and two repair arrangements. Residual strength varied non-monotonically with ligament spacing for repairs arranged transverse to the loading direction. For the baseline patch, the strength retention factor increased from 0.57 for closely spaced repairs to a maximum of 0.85 at 120 mm, and decreased to 0.59 at 180 mm as the repairs approached the panel free edges. The compliant ±45° patch was the most sensitive to spacing (0.42–0.80), whereas the stiffness-matched patch varied only between 0.48 and 0.61. In contrast, in-line repairs showed limited sensitivity to spacing, retaining 0.72–0.77 of the intact strength. In conclusion, the results demonstrate that repair orientation relative to the principal load path governs the interaction between adjacent repairs more strongly than separation distance alone.

The use of carbon fiber-reinforced plastic (CFRP) components has increased significantly in civilian aviation, necessitating effective maintenance and repair strategies to ensure durability and performance. While prior studies have focused on composite repair methods, such as stepped scarf patch and bolted joint repairs, these were limited to specimen and panel levels without addressing full-scale wing models. This study bridges that gap by evaluating stepped-lap repairs on a full-scale composite wing model under realistic loading conditions and exploring various repair scenarios. To reduce computational cost, two-dimensional shell elements were employed to simulate repairs, with results validated using experimental tensile test data from stepped-lap repaired specimens. Numerical models were developed for single regions and two closely located repair regions. For single-region repairs, adding up to two extra layers enhanced mechanical strength, but three extra layers increased strain, diminishing performance. For two closely located repairs, additional layers improved strength, though less effectively than single-region repairs. Square-shaped repairs exhibited higher strain due to stress concentrations at the corners, while circular repairs showed more uniform stress and strain distribution. These findings emphasize the importance of optimizing repair geometry and layer configurations using numerical simulations to ensure optimal structural performance of CFRP components.

The increase in aerospace composites usage for structural components demands advanced repair analysis. Overlay repairs of carbon fiber-reinforced polymer laminates offer an alternative that is easier to perform and less time-consuming to produce than the widely used tapered scarf repair and stepped lap. Composite specimen manufacturing was based on both twill carbon/epoxy prepreg and wet lay-up. The repair was performed with both prepreg and wet extra plies to the parent prepreg structure. However, the design of overlay joints must be carefully investigated to avoid generating stress concentration regions at free edges. This study examined specific extra ply terminations' impact on peak stresses in the adhesive bond line. Linear finite element analysis was performed to conduct a maximum principal stress study with a focus on three joint design parameters: ply material, overply effect, and stacking sequence. FEA accurately predicted experimentally observed responses and provided further insight into the failure behavior of the structure. Results showed that overlay joints have a strong sensitivity to ply material type, the number of overply, and stacking sequence. The introduction of overplies provided protection and stiffness at joint tips, and an overply material behavior was identified. The location of 0̊ plies in the composite laminates was highlighted as an important factor. The analysis was then extended to three-dimensional FE models for verification. In conclusion, results showed that high-stress concentration in overlay joints can be mitigated with the introduction of overplies and appropriate changes in joint design parameters to reduce stress peaks at joint tips and corners.

On-site repairs of carbon fiber reinforced polymer composites, wet layup repairs with heat blanket method play a critical and practical role for the composite defects that occur in production and assembly. The porosity level should be controlled for the repair parts with heat blanket method since the pressure value, which enables ply consolidation, reduce the risk of delamination in the composite layers, is less or zero with the wet layup repaired parts with heat blanket compared to repair parts with autoclave pressure. In this experimental study, an investigation was conducted regarding the tensile strength change of prepreg structures using wet lay-up repair techniques with heat blanket based on the porosity, with a specific focus on stepped-repaired carbon fiber reinforced polymer laminates. This work aims to understand the strength and the associated failure mechanisms of on-site repaired woven carbon fiber reinforced polymer laminates through experiments. The Automatic Ultrasonic Pulse Echo Inspection Method was utilized to see whether porosity level of each repaired samples is within allowable design limits for this purpose. Prepreg structure's repairs using wet lay-up produced according to standardized aerospace procedures were tested under uniaxial tension per ASTM 3039D. The relationship between attenuation difference (ΔdB) and tensile fracture values has been explored, with a focus on investigating the associated failure mechanisms. Initially, a 60% strength recovery was observed for repairs with an 8-decibel difference. However, as the decibel difference increased, the strength recovery gradually decreased, ultimately reaching 45.2%.