Procedia Engineering 41 ( 2012 ) 600 – 606

1877-7058 © 2012 Published by Elsevier Ltd.doi: 10.1016/j.proeng.2012.07.218

International Symposiu

Using Embedded FA

Ramzyzan RamlyaFlight Technology & Test Centre, Faculty o

bPhotonis Laboratory, Institut

Abstract

This paper describes the process of deveof vertical stabilizer was selected and rereproduced were fabricated in accordanmaterials and also the method of construcore, the sandwich panel was cured usorder to make the sandwich panel as embedded between the carbon fiber plieread before and after the FBG arrays wwas a success since the FBG sensors wspecimen will be used for further experi

© 2012 The Authors. Published by of Humanoid Robots and Bio-SenMARA.

Keywords: embedded fiber bragg grating arra

1. Introduction

The aircraft manufacturers starthowever the applications were limplementation of composite materas composite was relatively new afollowing three reasons: fatigue praluminum [2]. The main reason oaluminum in the 1920s was to gecompressed fuselage was developedat the same time adding more speedaerodynamically good performanceTooren [2] had proposed sandwichproposed a smart composite sandw(SHM) embedded in the compositeabout the future of sandwich matercomposite materials in secondary smanufacturers have been using comcomposite, mainly non-metallic, havertical and horizontal stabilizers. Fout of composite materials.

um on Robotics and Intelligent Sensors 2012

iber Bragg Grating (FBG) SensoAircraft Structure Materials

a, Wahyu Kuntjoroa, Mohd. Kamil Abd. R

of Mechancial Engineering, Universiti Teknologi MARA, 40450 Shate of Science, Universiti Teknologi MARA, 40450 Shah Alam, Selan

eloping a smart material with monitoring application to the eproduced using carbon fiber honeycomb core sandwich pannce to the generic sandwich structure and aviation industryuction. Using a carbon fiber from Hexcel as the face-sheet, Nsing Hysol EA9330 resin according to aviation industry sta

smart materials, optical sensor which has fiber bragg gres during the lay-out process. Using an FBG data logger, thewere installed in the sandwich panels. From the initial meawere readable and the difference in the signals was less thaniment for measuring strains and establishing the existence of

Elsevier Ltd. Selection and/or peer-review under respnsor (HuRoBs), Faculty of Mechanical Engineering

ays, optical sensors, honeycomb core sandwich panel

ted to use composite materials in aircraft componelimited to certain components only[1]. The initirials were high development cost including research aat that time. There is a need to replace aluminum wroblem with aluminum, lightness of composites, andof aircraft manufacturers changed the technology froet a lighter, more speed and comfort. The stressedd to be lighter so that the aircraft could fly higher whild to the aircraft [3]. In order to get a lighter, stronger, me aircraft, new composite materials must be adoptedh composite fuselage design for a lighter aircraft andwich structure for aerospace application with structue in order to monitor the integrity of the structure. Herials in commercial aviation. Stickler [5] explained thtructure in aircraft has been in practice since the last

mposite materials in various parts of the aircraft. Theas been used in larger parts in aircraft especially inFigure 1 shows a diagram showing various componen

(IRIS 2012)

ors in Smart

Rahmanb

ah Alam, Selangor, Malaysiagor, Malaysia

aircraft structures. A part nels. The sandwich panels y standards, including the Nomex honeycomb as the andard curing process. In rating arrays, FBG, were e FBG sensor signals were asurement, the experiment n 1 nm. In the future, the f damage in the panel.

onsibility of the Centre , Universiti Teknologi

ents since early 1980s ial problem with the and manufacturing cost

with composites for the d corrosion problem in om steel and wood to -skin construction and le burning less fuel and more fuel efficient, and d and researched. Van d Takeda et.al [3] had ural health monitoring

errmann et.al [4] talked hat the use of advanced t 30 years. The aircraft e honeycomb sandwich n the main box of the nts on the aircraft made

Available online at www.sciencedirect.com

Open access under CC BY-NC-ND license.

Open access under CC BY-NC-ND license.

601 Ramzyzan Ramly et al. / Procedia Engineering 41 ( 2012 ) 600 – 606

Fig 1. Illustration of Airbus A380 showing the components made from composite materials.

1.1. The application of optical sensor in aircraft composite smart materials

The current technology of structural health monitoring (SHM) in composite materials is done using sensors especially optical sensors [6] and [7]. This is due to the fact that the traditional strain gauge is sensitive to lightning, current leakage and corrosion, whereby inaccurate reading will be shown. The main problem with sandwich composite is delamination between the core and the facesheet, and this failure occurs underneath the skin and cannot be seen with naked eye. Therefore, a smart material with embedded fiber bragg grating (FBG) optical sensor has been developed to monitor the delamination between the face-sheet and the honeycomb core. Takeda et.al [3] has been doing an extensive research in smart materials, damage detection using optical sensor [8, 9], health structural monitoring in aircraft honeycomb composite sandwich panel [10] and also the delamination of the face-sheet detection using FBG [11]. Other researchers have also been doing research on the aircraft composite structures monitoring using optical sensor. Riccio et.al [12] was conducting study on the stiffened composite panel, while Giurgiutiu et.al [13] were developing a smart structure with embedded optical sensor for better health monitoring. The recent development in the application of optical sensor especially fiber bragg grating can be seen through the works of Takeda et.al [9], Liu et.al [14] and also Giurgiutiu et.al [14]. K. Diamanti and C. Soutis had compiled the most recent works in structural health monitoring using optical sensor [15]. In Malaysia, FBG optical sensor is used to monitor a civil structure performance [16]. A fully operational airframe fatigue monitoring is conducted in Malaysia to its fighter airplanes although still based on strain gauges [17]. The works reported in this paper is part of an effort to make use FBG as part of the sensor system in aircraft structural integrity monitoring.

The type of FBG sensor used in this research is an array of two sensors in a single optical fiber. It is identified with the wavelength grating of 1550 nm and 1555 nm each. An Optical Spectrum Analyzer, OSA, and FBG scanner were used to measure the signal wavelength of the FBG sensors.

2. The objective of the paper

The objective of this paper is firstly to describe the construction of a smart aircraft structure material, constructed from honeycomb core-carbon fiber facesheet sandwich composite panel embedded with fiber bragg grating (FBG) sensor; and secondly to characterize the wavelength signals of the FBG sensor before and after embedded in the composite sandwich panel.

3. Theory of Fiber Bragg Grating (FBG) Sensors

Fiber Bragg grating, which is an in-fiber component, consists of a periodic modulation of the index of refraction along the fiber core, as shown in Figure 2. FBG functions like a filter when a broad-band light is transmitted into the fiber core, reflecting light at a single wavelength, l, and a single wavelength is filtered in the transmitted light spectrum.

602 Ramzyzan Ramly et al. / Procedia Engineering 41 ( 2012 ) 600 – 606

Fig. 2. Schematic diagram of Fiber Bragg Gratings

When there is a force exerted on the FBG, it will compress or expand, thus the grating spectral response is changed. The wavelength of the light source is related to the Bragg Grating period do through the equation below:

��� 2�

��� (1)

and � � 2��� � 2�

�Δ� (2)

The strain dependence of the Bragg wavelength arises from the physical elongation/retraction of the sensor and the change in refractive index due to photoelastic effects. The wavelength shift due to axial strain can be obtained by equation below:

� � � �� �

�� (3)

where � is the applied strain, and �� is the photoelastic coefficient term given by

��� �

�

�/2 ��

��� ��

��� �

�� � (4)

where ���

coefficients are the Pockel’s coefficient of strain-optic tensor and � is the Poisson’s ratio.

4. The construction of the smart sandwich panel with embedded optical sensor

There were 14 specimens altogether prepared in this research. The specimen was a rectangular shape of size 300 mm long by 200 mm wide and 20 mm thick. It was constructed from a 20 mm thick, 3.2 mm cell size Nomex honeycomb core, sandwiched between two plies of carbon fiber from Hexcel at top and bottom sides. The composite panel was impregnated with Hysol 9330 resin. The fiber bragg grating sensor was laid on between the plies of the carbon fiber on one side of the panel. The specimen was cured at 82.2 degrees Celsius for 90 minutes under 110 kPa of vacuum pressure using Heatcon curing system. Figure 3 below shows the work flow of preparing the smart aircraft structure materials.

Fig 3. The work flow process of preparing the smart aircraft structure

The detail process in the workshop is shown in the Figure 4 below. The process was done in a controlled environment where the room temperature was kept at 25 degrees Celsius. The preparation and curing of the smart structure was done according to the airline standard procedures and in this case all specimens were constructed at Malaysia Airlines Composite Workshop at MAS Engineering Complex A at Subang, Malaysia.

Prepare the wet lay-up of the composite

sandwich panel

Lay the FBG strand on the impregnated

carbon fiber ply before completing

the assembly

Cure the panel assembly

according to the industry standard.

Measure the signals of the FBG sensors

after the specimen cured.

Measure the Signals of FBG optical sensor

before embedding in the specimen

603 Ramzyzan Ramly et al. / Procedia Engineering 41 ( 2012 ) 600 – 606

(a)

(c)Fig 4. Specimen preparation. (a) wet lay-up

process of the panel, (

Before the FBG sensors were emand after cured (Figure 4 (d)), the working in the smart structure. Usiconditions (before and after embeprocedure.

Fig.5. Schematic diagram showing FBG

(b)

(d) p cut-out, (b) laying down the FBG array in the middle of the carbon

(d) completed smart structure sandwich panel with embedded FBG s

mbedded inside the smart structure, the wavelength sisignals were once again taken to ensure that the FBng FBG-SCAN 700/800 scanner from FBGS, the sigdded). Figure 5 below shows the schematic diagram

(a)

(b) G sensor connected to the measuring instruments (a) before embedd

n ply outer layer, (c) curing

sensor

ignals were taken first, BG sensors system was gnals were read in both m of the experimental

ded, (b) after embedded.

604 Ramzyzan Ramly et al. / Procedia Engineering 41 ( 2012 ) 600 – 606

The signals taken were the peak two (2) FBG sensors in a single arraoutput data was taken directly in thafter the sensor was embedded in thcondition of the optical sensors after

Table 1. Sensor condition before and after em

Specimen Number

1 2 3 4 5 6 7 8 9

10 11 12 13 14

5. The results and discussion

The results of the signal readingFigure 6 shows the signal reading othe sandwich composite panel. Figuwere embedded inside the sandwichnegative difference means the senso

Fig 6. G

From Figure 6, it can be seen thThe bar on the left of each couple othe signal taken after embedded. Th14) is the grating wavelength of 155the grating wavelength of 1555 nm.almost identical to the signal beforeading can be seen in the Figure 7.

k wavelength in nanometer, nm, of the FBG in each opay. The optical sensor connector was connected to the

he output computer as shown in Figure 5 (a). The samhe smart structure as shown in Figure 5 (b). The followr they were embedded.

mbedded

Sensor Condition Before Embedded

Sensor Condition Afte

Good One side broken buGood One side broken buGood Good Good Good Good One Side broken buGood Good Good Good Good One side broken buGood Broken in the middle.

from both sidGood One side broken buGood Good Good Good Good Good Good One side broken bu

s of all 14 FBG sensor array is displayed in the Figureof each of the 14 FBG’s before and after the FBG arrure 7 shows the difference of peak wavelength changeh composite panel. Positive difference means the sensor was contracted.

Graph showing the peak wavelength of the 14 FBG arrays

hat the signal reading before and after embedded is noof bars is the signal taken before embedded, while the he first series of arrays for example, An.1 (where n is50 nm, while the second series An.2 (where n is the n. From the figure, it can be seen that the signal readingore it was embedded in the specimen. The exact dif

ptical array. There were e FBG scanner, and the me procedure was done wing Table 1 shows the

er Embedded

ut readableut readable

ut readable

ut readableNeed to read des

ut readable

ut readable

e 6 and Figure 7 below. rays were embedded in es after the FBG arrays sor was stretched while

ot very much different.bar on the right side is

s the number from 1 to number from 1 to 14) is g after the embedded is fferences of the signal

605 Ramzyzan Ramly et al. / Procedia Engineering 41 ( 2012 ) 600 – 606

Fig 7

From Figure 7, it can be seen tsensors show a stretched condition aA7.2, A11.1&A11.2, and sensor Adifference. The compression or extethe main intension of this experimeexample of peak shift using Opticalone before embedded while the righthan the first line shows a power dro

Fig. 8. Graph showing t

6. Conclusion

This paper has described the pstructures. The material being devecomposite panel with embedded Fpreparing the specimen was presemeasured using optical FBG scannethey were embedded in the specimesee the wavelength peak shift beexperiment was a success since the nm. In the future, the specimen wilexistence of damage in the panel.

. The total difference of the FBG signal peak wavelength

that peak wavelength shift generally was less than 1 as indicated by the positive differences while only thre

A 14.1&A14.2) sensor arrays were compressed as indension of the FBG has no significant consequences inent is the sensors are functional as it is supposed to bl Spectrum Analyzer, OSA. In the figure, the left side ht side of each peak is the one after embedded. The seop inside the optical fiber.

the peak shift taken with OSA before and after embedded in the sma

process of developing a smart material in the appleloped in this research was a honeycomb core-carbo

FBG sensor for structural health monitoring. In thisented. Before the FBG sensor arrays were embeder. These signals were once again measured using then. The signals were also recorded using Optical Spectr

efore and after embedded condition. From the iniFBG sensors were readable and the difference in the ll be used for further experiment for measuring strain

nm. Most of the FBG ee FBG (sensor A7.1 & dicated by the negative n this experiment since be. Figure 8 shows the of the each peak is the

econd line is a bit lower

art structure.

lication of the aircraft on fiber skin sandwich paper, the process of

dded, the signals were e same equipment after rum Analyzer, OSA, to itial measurement, the signals was less than 1 ns and establishing the

606 Ramzyzan Ramly et al. / Procedia Engineering 41 ( 2012 ) 600 – 606

Acknowledgements

The authors would like to acknowledge Research Management Institute of Universiti Teknologi MARA for managing the FRGS fund of this research, and the Ministry of Higher Education for providing the FRGS fund. The research use FRGS grant, entitled “Fatigue Load Parameters Measurement By Fiber Bragg Gratings Sensing Under Variable Amplitude Loading”. The authors would like also to thank Malaysia Airlines Bhd who provided the assistance through Composite Workshop, including the materials, expertise, and the facility to set-up the specimens. Special gratitude to the workshop supervisor at Malaysia Airlines who assisted the authors in the making the specimens and his knowledge transfer.

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