Fibre Optic Sensors for Structural Health Monitoring of Aircraft Composite Structures: Recent Advances and Applications

Abstract

In-service structural health monitoring of composite aircraft structures plays a key role in the assessment of their performance and integrity. In recent years, Fibre Optic Sensors (FOS) have proved to be a potentially excellent technique for real-time in-situ monitoring of these structures due to their numerous advantages, such as immunity to electromagnetic interference, small size, light weight, durability, and high bandwidth, which allows a great number of sensors to operate in the same system, and the possibility to be integrated within the material. However, more effort is still needed to bring the technology to a fully mature readiness level. In this paper, recent research and applications in structural health monitoring of composite aircraft structures using FOS have been critically reviewed, considering both the multi-point and distributed sensing techniques.

Keywords: Brillouin scattering; Rayleigh scattering; aerospace; aircraft; composite materials; fibre Bragg gratings; fibre optic sensors; lamb waves; smart structures; structural health monitoring.

Figure 1

(a) Use of composite materials in the Boeing 787 and (b) evolution of the use of composites in Airbus aircraft

Figure 2

Overview of basic principles and types of fibre optic sensors.

Figure 3

Fibre Bragg Grating’s principle of operation.

Figure 4

Schematic classification of FBG interrogation techniques based on scan rate and applications (adapted from [40]).

Figure 5

Typical spectrum of spontaneous light scattering in a fibre.

Figure 6

Basic OFDR network.

Figure 7

Optical fibre section in (a) unidirectional; (b) cross-ply and (c) woven fabric [83].

Figure 8

Teflon tubes (a) [94] and embedded connector (b).

Figure 9

NASA’s Helios wing (a) and its wreckage in the Pacific (b) (courtesy of NASA).

Figure 10

Photos of the wing-like swept plate experiments with three lines of FBG sensors (property of the NASA Dryden Flight Research Center) [116].

Figure 11

Typical scheme of wing box (a); deployment of the optical fibres in the structure (b).

Figure 12

T38 wing model with the embedded FBG sensors [53].

Figure 13

Safe life or damage tolerant areas of the composite patch (a) and SHM system for monitoring composite repair disbond with FBGs (b) (adapted from [11]).

Figure 14

(a) Specimen with composite lap joint and embedded optical fiber; (b) micrograph of an embedded FBG sensor [132].

Figure 15

(a) STFT (left) and FFT (right) computed from the FBG peak wavelength estimation during fatigue tests; (b) corresponding pulsed-phase thermography phase angle image [132].

Figure 16

Bird strike impact probability (a); impact damage in a flap fairing of a cargo B767 caused by bird strike (b).

Figure 17

Full-spectral FBG response after different strike numbers [133].

Figure 18

(a) Composite specimen representing a typical aeronautical construction; (b) FE model of the specimen [140].

Figure 19

Example of advanced grid structure made of CFRP (a) and its application to an aerospace structure (b).

Figure 20

Advanced structural grid (AGS) with SHM system (adapted from [144]).

Figure 21

(a) Intersection of the calculated ellipses; (b) comparison between the actual and predicted damage location [12].

Figure 22

Scheme of the smart composite wing experiment (adapted from [162]).

Figure 23

Photo of the cantilever plate experiment (a); reconstructed deflection shape of the plate (b) [166].

Figure 24

Position of the optical fiber and foil gage in the clamped fan experiment [65].

Figure 25

Optical fibres deployment and damage around the CFRP bolted joint (a); residual strain distribution along the optical fibres after loading (b) [166].

Figure 26

Impact test layout (a); BGS measured in the point closest to the impact location (b) [166].