Strain and damage monitoring in CFRP fuselage panels using fiber Bragg grating sensors. Part II: Mechanical testing and validation

Abstract

This is the second 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 (FOBG) sensors. The system was developed and tested on the basis of compression after impact response of the panel. The first paper describes the system design, the embedment of fiber sensors in the panel during manufacturing and the impact testing. In the present paper, mechanical testing of the panels and validation of the proposed monitoring system are described. The mechanical tests comprised three load-scenarios: compression to failure of undamaged panel, compression to failure of impacted panel and compression to failure of impacted and fatigued panel. In all cases the panel failed due to buckling. In order to monitor the panels’ behavior during testing and be able to validate the accuracy of the measurements of the FOBGs, a network of strain gauges were used. Impact damage and fatigue damage decreased the failure load and displacement at failure of the panel. The measurements taken from the FOBGs captured all changes in the buckling modes of the panel. Moreover, they showed a good correlation with the measurements taken from strain gauges. These findings validate the proposed monitoring system.

Introduction

Fiber-reinforced polymer composites are revolutionizing the design of large, high-performance structures in the aerospace industry. The use of carbon fiber-reinforced plastics (CFRPs), the most common and lightest of industrial composite materials, is now very widespread; for instance, they comprise almost 50% of the new Boeing 787 Dreamliner and more than 20% of the Airbus A380. The extended use of CFRPs in fuselage structures has increased the need for the use of non-destructive testing techniques to detect damage that could counterbalance the integrity of composite structures. In the last few years, research was devoted to the development of structural health monitoring (SHM) systems, which are one step ahead from conventional non-destructive testing systems as they offer online monitoring of the structures.

SHM systems realize continuous monitoring of structure’s integrity through strategically located sensors coupled with monitoring technology enabling remote interrogation. There are a variety of approaches that can be used. The most common are strain based SHM systems which measure the strain distribution of the structure subject to operational loads via electrical resistive strain gauges or optical fiber sensors. Any damage in the structure, that causes a change in the strain distribution, may be potentially detected by this type of system. Fiber optic sensors are particularly suitable for the health monitoring of large structures, such fuselage panels, because multiple sensors can be attached to a single fiber for distributed sensing over long distances. The small size of fiber optic sensors (<250 μm in diameter) imposes negligible intrusion into the host structure and allows fast interrogation with minimal wiring requirements. The most widely used fiber optic sensor is the Fiber Optic Bragg Grating (FOBG) sensor. This sensor possess superior characteristics such as localized strain measurement ability, relatively small size, high sensitivity, inertness to electric or magnetic inference and multiplexing capability. A fundamental issue in the implementation of a FOBG-based SHM system in composite structures is the embedment of sensors during manufacturing. Although manufacturing methods of composites parts offer the opportunity to build the sensors into the part itself, there are several difficulties to overcome such as the selection of points of ingress and egress of sensor data lines and the possible effect of sensors and ancillaries on the mechanical properties of the composite.

Although major aircraft manufacturers, both civil and military, are keen to introduce SHM systems into future aircrafts, particularly in composite structures, to the date, SHM systems was successfully installed only in experimental applications. In most of these applications FOBG-based SHM systems were applied [1], [2], [3], [4], [5], [6], [7], [8], [9], [10], [11], [12], [13], [14], [15], [16], [17], [18], [19], [20], [21], [22], [23]. In an early work, Takeda et al. [7] have managed to relate governed delamination onset and growth in CFRP laminates with the spectrum of FOBGs. In a latter work, Yashiro and Okabe [20] implemented the methodology proposed in [7] for holed laminates subjected to static and fatigue load. The most important finding of [20] is that fatigue damage in the composite material may not be detected by FOBGs due to debonding of the sensors caused by the cyclic loading. This finding is further analyzed and discussed in [21].

In the majority of reported works, FOBGs were used to detect damage in specimens. There are very few works in which FOBGs were embedded in structural parts [10], [17], [21], [23]. Tsutsui et al. in [10] applied small-diameter optical fiber sensors to stiffened composite panels for the detection of impact damage. Ryu et al. [17] have used multiplexed and multi-channeled built-in FOBGs to monitor the buckling behavior of a composite wing box. Recently, Takeda et al. [23] used FOBG sensors to monitor damage due to compressive load in CFRP stiffened panels.

In the present work, an integrated methodology for monitoring strain and damage in CFRP fuselage panels was designed, implemented and validated on the basis of a series of load-cases representing critical scenarios for the CFRP panel. This is the second paper of a two-paper series describing the monitoring methodology. In [24] the monitoring system design, the embedment of fiber sensors in the panel during manufacturing and the impact testing were described. In the present paper, mechanical testing of the panels and validation of the proposed monitoring system are described.