Repairing composites

Collombet, Francis; Grunevald, Yves-Henri; Crouzeix, Laurent; Douchin, Bernard; Zitoune, Rédouane; Davila, Yves; Cerisier, A.; Thevenin, Raphaele

doi:10.1016/b978-1-78242-307-2.00010-5

Cited by 6 publications

(9 citation statements)

References 4 publications

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“…The aim is to illustrate that the maximum loading flow (here the shear flow), when it is not applied in the reinforcement orientation, does not depend on the overlap length. This hypothesis will be used to reduce the repair patch extension [2]. According to [2] along the fibre direction, the overlapping length depends on the chosen step ratio.…”

Section: Test Calculation Comparisonsmentioning

confidence: 99%

“…This hypothesis will be used to reduce the repair patch extension [2]. According to [2] along the fibre direction, the overlapping length depends on the chosen step ratio. In the other directions, this ratio can be reduced down to zero.…”

Section: Test Calculation Comparisonsmentioning

confidence: 99%

“…This patent allows the description and modelling of a minimum volume of material to be removed while ensuring a nominal transfer of efficient stress flow between the parent composite part and the repair patch. To give some orders of magnitude, it was shown that this approach can lead to the reduction of the repair patch grip by more than 70% [2], making the adhesive patch certifiable by aeronautic regulation authorities [3]. Actually today, the bonding repair patch is allowed alone (i.e.…”

Section: Introductionmentioning

confidence: 99%

“…For example, Airbus requires 8 mm overlap for 0.13 mm ply thickness while Boeing 5 mm overlap for the same ply thickness. We have chosen stepped-lap geometry for taking the best performance in terms of patch size reduction [2].…”

Section: Introductionmentioning

confidence: 99%

“…To be close to industrial applications, an original methodology of mechanical characterization (called Multi-Instrumented Technological Evaluator tool box) has been proposed [2,5]. The MITE tool box allows the design of case-by-case specific repair evaluators that can be tested in a multi-axial testing machine under complex loading sequences.…”

Section: Introductionmentioning

confidence: 99%

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

Collombet

Davila

Avila

et al. 2019

Mechanics & Industry

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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.

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Section: Test Calculation Comparisonsmentioning

confidence: 99%

Section: Test Calculation Comparisonsmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

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