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Optical fibre sensors are being investigated since many years as candidates of choice for supporting structural health monitoring (SHM) in aerospace applications. Fibre Bragg grating (FBG) sensors, more specifically, can provide for accurate strain measurements and therefore return useful data about the mechanical strain state of the structure to which they are attached. This functionality can serve the detection of damage in an aircraft structure. However, very few solutions for protecting and bonding optical fibres to a state-of-the-art aircraft composite material have been reported. Most proof-of-principle demonstrations using optical fibre sensors for aerospace SHM-related applications reported in literature indeed rely on unpackaged fibre sensors bonded to isotropic metallic surfaces in a mostly unspecified manner. Neither the operation of the sensor, nor the adhesive material and bonding procedure are tested for their endurance against a full set of standardized in-flight conditions. In this work we propose a specialty coated FBG sensor and its permanent installation on aerospace-grade composite materials, and we demonstrate the compatibility with aerospace in-flight conditions. To do so we thoroughly evaluate the quality of the operation of the FBG sensor by correlating the reflection spectra of the installed sensors before and after exposure to a full set of realistic in-flight conditions. We also evaluate the difference in strain measured by the FBG, since any damage in the adhesive bond line would lead to strain release. The applied test conditions are based on aerospace standards and include temperature cycling, pressure cycling, exposure to humidity and Smart Materials and Structures

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.

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