Strain and damage monitoring in CFRP fuselage panels using fiber Bragg grating sensors. Part I: Design, manufacturing and impact testing

This is the first paper of a two-paper series describing design, implementation and validation of a strain and damage monitoring system for CFRP fuselage stiffened panels based on fiber optic Bragg grating sensors. The monitoring system was developed and tested on the basis of three load-scenarios: compression to failure of the undamaged panel, compression to failure of the impacted panel and compression to failure of the impacted and fatigued panel. This paper focuses on the design of the fuselage panel, the design of the monitoring system, the embedment of fiber sensors in the panel during manufacturing and the impact testing. The network of the sensors was designed based on a numerical buckling analysis from which the strain field of the panel was computed as a function of the applied compressive load. Embedment of fiber sensors in the panel was done so as to minimize risk of fiber breaking during manufacturing and impact testing and to effectively capture strains that are representative of damage developed in the panel due to compressive load. Barely visible and visible low velocity impact damage sites were created at different locations of the panel using a drop-weight impactor. The panels were inspected using C-scan just after manufacturing, to check quality of the material, and just after impact testing to detect impact damage at each location

Design, numerical and experimental characterization of the composite spar for a regional aircraft

The work is part of the Clean Sky 2 – AIRGREEN2 (AG2) project which aims at designing, analysing, manufacturing and ground-testing a full scale composite Outer Wing Box (OWB) demonstrator for a regional aircraft. In order to achieve this objective, in the current development work a small scale (in length) spar segment has been designed, manufactured and tested in order to validate the related technologies. The OWB has been designed in order to withstand stiffness, strength and stability requirements under critical loading conditions taken from the flight envelope. In particular, a minimum weight design has been obtained with an optimization based on genetic algorithms that explored several thickness distributions and stacking sequences. A further analysis has been performed on the optimized front spar of the OWB. Focusing on the first two (root) bays, a spar segment (approx. 1100 mm long and 400 mm wide, constant section - no tapering) has been considered. The spar segment was manufactured by using the Liquid Resin Infusion (LRI) method, an Out of Autoclave method (OoA), in which Dry Non Crimp Carbon Fabrics are impregnated by epoxy resin (high temperature cure) under the application of vacuum only. The impregnated Carbon fabric material is then cured in a standard oven or self heated tool, heated internally by integrating electrical resistances and externally by heated blankets. In this work, a self heated tool option was selected. The tool was designed as an egg-crate type construction in order to minimize both mass and thermal inertia. CATIA software was used for both the tool and part design. A FEM analysis was performed by implementing progressive failure analysis using the MSC Marc software in order to predict the spar segment structural response during the experimental test. Respective material allowables have been obtained by testing standard coupons manufactured by the same base materials (dry carbon fabrics and resin) and manufacturing method (LRI). Standard hand layup procedures were followed during the layup process of the fabrics. Appropriate auxiliary materials capable to withstand the respective LRI infusion and curing process specifications were used. The spar segment has been tested up to failure in a cantilevered configuration and subjected to a tip in-plane pure shear force. The fixtures and loading system for the test have been designed in order to guarantee the correct clamped boundary conditions and to avoid any torsional effect. The scope of the test was to validate the manufacturing process, as well as the design and FEM analysis in terms of allowables and final failure of the component. A numerical/experimental comparison is presented in terms of load vs displacement response curve and load vs strain (measured at some locations on the spar)

Design, numerical and experimental characterization of the composite rib for a regional aircraft

The work is part of the Clean Sky 2 - AIRGREEN2 (AG2) project which aims at designing, analysing, manufacturing and ground-testing a full scale composite Outer Wing Box (OWB) demonstrator for a regional aircraft. The OWB has been designed in order to withstand stiffness, strength and stability requirements under critical loading conditions taken from the flight envelope. In particular, a minimum weight design has been obtained with an optimization based on genetic algorithms that explored several thickness distributions and stacking sequences. As important step of the project, a composite rib has been designed, manufactured and tested in order to validate the related technologies. The rib, object of this work, corresponds to a middle section rib of the Outer Wing Box (OWB). Static and nonlinear finite element analyses have been performed in order to verify that the proposed rib test geometry would fail at a machine load less than the current maximum available testing load machine. Moreover, buckling analysis has completed the FE analyses as required by the design specifications. The composite rib was manufactured by hand-layup and by using the Liquid Resin Infusion (LRI) method, an Out Of Autoclave method (OoA), in which Dry Non Crimp Carbon Fabrics are impregnated by epoxy resin (high temperature cure) under the application of vacuum only. The impregnated Carbon fabric material was then cured in a standard oven. CATIA software was used for both the tool and part design. Respective material allowables have been obtained by testing standard coupons manufactured by the same base materials (dry carbon fabrics and resin) and manufacturing method (LRI). Standard hand layup procedures were followed during the layup process of the fabrics. Appropriate auxiliary materials capable to withstand the respective LRI infusion and curing process specifications were used. The rib has been tested up to failure under an in-plane shear loading condition. The scope of the test was to validate the manufacturing process, as well as the design and FEM analysis in terms of allowables and final failure of the component