This study analyses the mechanical and crack growth behavior of woven carbon fiber reinforced plastics (CPRF) with embedded ceramic sensors. The material studied here is 3K-70-P carbon fiber plain weave with EPOLAM 2015® epoxy resin. The composite is manufactured with vacuum bagging procedure. Later on, the composite Mode I interlaminar fracture toughness (G IC ) is calculated by means of double cantilever beam tests (DCB) for two layout configurations [0/90] and [±45] with and without embedded sensors. Results give an initial approach of the fracture behavior of an instrumented composite facing an interlaminar crack. The interlaminar fracture toughness for the instrumented specimens is lower compared to the noninstrumented coupons. The presence of the sensor and its wire connection has a considerable impact on the damage tolerance of the woven composite, where the sensors surroundings seem to be the more likely region to be affected by an interlaminar fracture. K E Y W O R D S carbon fibers, delamination, fractography, fracture toughness F I G U R E 5 Load-displacement curve for [±45] double cantilever beam coupons [Colour figure can be viewed at wileyonlinelibrary.com] Reference With Peak load = 105 N, Peak load = 45 N, With How to cite this article: Torres M, Tellez RA, Hernández H, Camps T. Mode I interlaminar fracture toughness of carbon-epoxy coupons with embedded ceramic sensors.

The exact solutions to a one-dimensional harmonic oscillator plus a non-polynomial interaction a   x 2  +  b   x 2 /(1 +  c   x 2 ) ( a  > 0, c  > 0) are given by the confluent Heun functions H c ( α , β , γ , δ , η ; z ). The minimum value of the potential well is calculated as V min ( x ) = − ( a + | b | − 2 a   | b | ) / c at x = ± [ ( | b | / a − 1 ) / c ] 1 / 2 (| b | >  a ) for the double-well case ( b  < 0). We illustrate the wave functions through varying the potential parameters a , b , c and show that they are pulled back to the origin when the potential parameter b increases for given values of a and c . However, we find that the wave peaks are concave to the origin as the parameter | b | is increased.

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.

Aircraft composite structures are mostly joined by mechanical fasteners like bolts, pins or screws. However, the effect of the presence of holes in the remaining strength of the composite structures is still being studied extensively. In this work, epoxy/glass laminates with drilled holes of different sizes were tensile tested and from these results, the residual strength was plotted. Strength vs. hole’s diameter at different fiber orientation was obtained. The fracture path and failure mechanism were identified by fractographic examination. The Point Stress Criterion (PSC) was used, in order to establish the stress intensification due to the presence of a drilled hole. A numerical model by Finite Element Method was carried out to verify the experimental results and the analytic failure predictions. A reduction of 50% in laminate strength was observed when diameter-width ratio was 0.12. The principal fracture mechanism observed in composite laminates was interface breakup. FEM results and analytic results by PSC show accuracy of 90% for predicting the damage in drilled composites.

An analysis of the strut‐to‐fuselage and main‐landing‐gear‐to‐fuselage joint is realized in order to evaluate the zones with a high probability of failure. The whole section is analyzed using the finite element method in order to estimate static resistance behavior, therefore it becomes necessary to create a numerical model of the section, to measure the mechanical properties of the C‐Ep composite material, and to calculate the loads applied onto the section. The section is digitized with photogrammetry. Mechanical characterization is achieved with instrumented standardized tensile tests. Results of the analysis show that the zones with higher probability of failure are found around the wing strut and the fuselage joint, with a safety factor higher than necessary even for the most critical condition, and lower in comparison with the average safety factor used on aircrafts built mostly with metals. POLYM. COMPOS., 36:1072–1083, 2015. © 2015 Society of Plastics Engineers