Regular Article

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

¹ Université de Toulouse, UPS, INSA, Mines d'Albi, ISAE, ICA (Institut Clément Ader). Espace Clément Ader, 3 rue Caroline Aigle, 31400 Toulouse, France
² Instituto Politecnico Nacional, ESIME Ticoman. Av. Ticoman N° 600, 07340 Ciudad de Mexico, Mexico
³ Composites, Expertise & Solutions. 131 Traverse de La Penne aux Camoins, 13821 La Penne-Sur-Huveaune, France
⁴ Bayab Industries, 10 allée de Longueterre, 31850 Montrabé, France

^* e-mail: francis.collombet@iut-tlse3.fr

Received: 7 October 2019
Accepted: 16 June 2020

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.