Proof of a composite repair concept for aeronautical structures: a simplified method

Abstract

This paper provides an illustration of all stages of primary aeronautical composite structure repair by using industrial tools and scientific methodologies, as well as numerical tools to simplify the cross-over analysis of the mechanical behaviour of the repaired area. Economically and scientifically speaking, one of the main challenges of composite repair (for monolithic long fiber composite parts) consists of promoting a bonded composite patch option without additional riveted doublers. To address this challenge, size reduction of the patch could be mandatory. A patent (jointly owned by ICA, Bayab Industries and CES), entitled “Method for repairing a wall consisting of a plurality of layers”, is devoted to reducing repair patch dimensions of monolithic composite parts provided the bonding zone has a stepped-lap geometry. This patent is based on a simple idea that no overlapping length is required between composite plies for load transfer except in the fiber directions of the plies (unidirectional or biaxial long fiber reinforcements with epoxy matrix). To prove this concept, we consider on one hand, a situation unusual in the literature by studying a composite specimen without fibers aligned along the main loading axis, and on the other hand, a classical situation of where the shape of the specimen is adapted to be studied by uniaxial tension tests. After different manufacturing steps, the studied specimen contains three zones representing both the influence of the total thickness of a repair patch, the stepped-lap area assembled with an adhesive film and the parent composite part. Basically, a simple parent structure consisting of 16 plies of UD Hexply® M21/35%/268/T700GC (close to Airbus composite raw materials on board in A380) is manufactured with a stacking sequence of [+45/−45/−45/+45/+45/−45/−45/+45]s. Then, the parent structure is machined by the Airbus Abrasive Water Jet machine and the final repair area has a stepped-lap geometry by overlapping successive plies of the same nature as the parent plate and after having previously applied an adhesive film (cured at 180 °C). Furthermore, 3 values of overlap length (respectively, 6, 8 and 13 mm) are investigated to include the mean value required by Airbus in the case of the use of the studied prepreg. After abrasive water jet machining of the composite parent part, repair patch manufacturing was performed according to Airbus requirements. The studied specimens were cut from the final plate (involving the parent plate, the stepped lap zone and the zone of the patch itself) and tested in an uniaxial tensile configuration with a loading direction shifted 45° with respect to the fiber direction. Furthermore, studying uniaxial tensile tests on multilayer-pasted interface is innovative in the literature. In this paper, it is shown that the stepped-lap area assembled with an adhesive film is not the weak link of the mechanical response but rather the parent area, i.e. the unrepaired monolithic composite. Numerical calculations confirm this proof of concept by underlying that the level of shear stress in the adhesive film, for these three overlapping values, is below the chosen limit value. These results show that the patch size reduction is possible.

Influence of modelling hypotheses on strength assessment of CFRP stepped repairs

“…A novel machining technique has been developed to perform in-situ stepped repairs. It uses water-jet abrasion instead of traditional hand abrasion, making the machining process faster and more reliable [16,17]. In terms of equivalent scarf angle based on the ratio between total repair thickness and length, tested configurations go from 5.2° to 1.3° angles.…”

Bonded flush repairs for aerospace composite structures: A review on modelling strategies and application to repairs optimization, reliability and durability

“…It was observed that as the number of steps increased, the strength of the repair increased. Collombet et al [46] presented results showing stepped joints with a specific layup only made of +-±45° plies failing by laminate fracture with no debonding at all. They concluded that the stepped joint was not always the weakest part of the assembly.…”

“…Further experimental and numerical work, including damage modelling simulations, is needed to provide an understanding of the differences between failure behaviour of a flush repair and its equivalent joint, and to what extent the equivalent joints hypothesis is valid. This could help saving computation time in strength prediction simulations because even for small coupons, detailed FE models can reach more than 500.000 degrees of freedom [46].…”

“…Research papers tend to show that the strength of a repair coupon can rise until 2° (1:28) [32] or 1° (1:57) [36] and aircraft manufacturers recommend 1:60 (Airbus) and 1:45 (Boeing) [46] Stacking sequences of composite adherents influences joints and repairs strength. Experimental test [84] showed that scarf joints with 0° surface plies have a lower strength than scarf joints with 90° plies.…”