Composite Structures 94 (2012) 813–819

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Composite Structures

journal homepage: www.elsevier .com/locate /compstruct

Damage monitoring of CFRP stiffened panels under compressive load usingFBG sensors

Shin-ichi Takeda ⇑, Yuichiro Aoki, Yosuke NagaoAerospace Research and Development Directorate, Japan Aerospace Exploration Agency, Japan

a r t i c l e i n f o

Article history:Available online 25 February 2011

Keywords:Stiffened panelsCFRPsCompressionFBG sensorsSHM

0263-8223/$ - see front matter Crown Copyright � 2doi:10.1016/j.compstruct.2011.02.020

⇑ Corresponding author. Address: 6-13-1 Osawa MJapan. Tel.: +81 50 3362 6554; fax: +81 422 40 1451.

E-mail address: stakeda@chofu.jaxa.jp (S.-i. Takeda

a b s t r a c t

FBG sensors were embedded in each of two CFRP stiffened panels fabricated by VaRTM. Low-velocityimpacts were applied to one of the panels in order to compare the methods of monitoring impact eventsusing FBG sensors. The main impact damage was an interlaminar delamination inside the skin, whichcould be observed by an ultrasonic C-scan. A monitoring method using the full spectral signals was moreeffective in evaluating the impact damages in detail than that using the center wavelength. Following theimpact tests, buckling behaviors were investigated under compressive loading using FBG sensors and sur-face-attached strain gauges. The FBG sensors could evaluate strain changes resulting from bucklingbehaviors under relatively low compressive loading. They could also evaluate damage growth until thefinal failure and difference of buckling behaviors between panels with and without impact damages.

Crown Copyright � 2011 Published by Elsevier Ltd. All rights reserved.

1. Introduction

Structural Health Monitoring (SHM) technology in compositematerials has increased in importance since the application of Car-bon Fiber-Reinforced Plastics (CFRPs) was expanded to airframestructures. Fiber-optic sensors are one of the most promising sens-ing devices because of their small size, lightweight, and immunityto electromagnetic interference. Numerous applications of fiber-optic sensors to composite materials have been reported [1–21].Some studies have demonstrated methods for evaluating changesin elastic wave propagation and/or eigenfrequency [1–8]. Thesehigh-sensitivity methods are used not only for on-ground inspec-tions but also the identification of impact loadings/damages/loca-tions in real-time given their suitability for transient responses.Other studies have demonstrated methods using Fiber Bragg Grat-ings (FBGs) and/or fiber-optic sensors as strain-sensing devices [9–21]. Given that these sensors directly detect strain changes causedby damage occurrences, they have the advantages of high tolerancefor noises in flight and high reliability of damage detection. Theyalso have the potential to reveal more detailed information aboutthe damages when embedded in composite materials.

In the SHM of composite structures, it is possible to apply fun-damental methods using strain changes to assess structural dam-age concurrently with the high-sensitivity methods. We havepreviously demonstrated the application of FBGs to damage mon-itoring in CFRPs under fatigue loading [15] or low-velocity impact

011 Published by Elsevier Ltd. All r

itaka-shi, Tokyo 181-0015,

).

loading [19]. The processes of manufacturing CFRPs have also beenmonitored by FBGs and by a new distributed strain-sensing systemknown as Pre-Pump Pulse Brillouin Optical Time Domain Analysis(PPP-BOTDA) [20,21]. However, damage monitoring under com-pressive loading has not been studied, and there are few investiga-tions on buckling behaviors in composite stiffened panels [22].

In this study, embedded FBG sensors were used as strain anddamage sensing devices in thin-skin CFRP stiffened panels. The firstobjective is the demonstration of monitoring buckling behavior un-der compressive loading. This is important given the trend of reduc-ing panel thickness to attain lighter structures. Impact damageswere first evaluated using signals from FBGs. Subsequently, an eval-uation method using the spectrum from FBG sensors was comparedwith a method using the center wavelength, an index typically usedto measure strain. Finally, monitoring under compressive loadingwas performed until final failure for panels with and withoutimpact damages.

2. Fabrication of CFRP stiffened panels

Two CFRP stiffened panels were fabricated by Vacuum-assistedResin Transfer Molding (VaRTM), a cost-effective method for mold-ing large-scale composite structures. VaRTM is a flexible moldingmethod for embedding sensors into composite materials becausethe resin is infused into prepared dry fabrics. The stiffened panelsin the present study had thin-skin with three T-shaped stringers asshown in Fig. 1. Each panel was 400 mm in length and 350 mm inwidth. Non-crimp fabrics (SAERTEX Co. KG) made of aerospace-grade carbon fibers were used as materials. The stacking sequencesof the skin and stringer were [(+45/�45)(0/90)]S and [(+45/�45)(0/

ights reserved.

350 mm

400 mm

125 mm

50 mmStringer Side

Fig. 1. Dimensions of CFRP stiffened panel.

Aluminum fixture

Optical fibers Strain gauge line

Fig. 4. Photograph of stiffened panel.

Strain gauge Location

FL1–FL6 Flange

SK1–SK4 Skin

(Stringer side)

SK5–SK8 Skin

(Skin side)

SK1

SK2

SK3

SK4

FL4 FL5

FL6

FBG 1 FBG 2

FBG 3FL1 FL2 FL3

FBG 5

Locations of FBGs correspond to

those of strain gauges in FL series.

Fig. 5. Arrangement of FBGs and strain gauges in panel.

Nonlinear behavior(< 40 kN)

814 S.-i. Takeda et al. / Composite Structures 94 (2012) 813–819

90)]2S, respectively. The panels fell into the category of thin-skinstiffened CFRP panels.

The combined dry fabrics were first placed on the mold withsome vacuum infusion supplies. They were covered by vacuumbagging films, after which epoxy resin (XNR6809/XNH6809, Naga-seEleX Co., LTD) was impregnated into the vacuum bag at 40 �C.The VaRTM resin is an improvement over RTM resin, having thecharacteristics of medium toughness and relatively low viscosity.After confirming impregnation, the resin was cured and post-curedat 80 �C for 6 h and at 120 �C for 3 h, respectively. Finally, themolded panel was taken out of the bag, and aluminum fixtureswere attached to both ends of the panel with epoxy resin.

FBG sensors have a periodic variation in the refractive indexalong a certain length of a single-mode optical fiber. When a broad-band light is injected into the FBG sensor, a narrowband light is re-flected as a result of Bragg diffraction. The relationship betweenthe wavelength of reflected light and the applied strain is foundto be nearly linear [23]. This means that it is easy to assess strainby measuring the shift in the center wavelength of the reflectedlight. Here, the center wavelengths were generally defined as thewavelength at the peak value of the reflected light. FBGs (NeotronCo., Ltd.) with a 10 mm gauge length were used for this study. Be-fore embedding the FBG in the panel, its strain sensitivity was ob-tained by a calibration test using electrical strain gauges. The strainsensitivity used for this study, 1.184 pm/le, was almost the sameas the common value, 1.2 pm/le [24]. Fig. 2 shows the appearanceof FBGs stacked on the dry fabrics before the resin flow process.

FBG

Fig. 2. Photograph of FBGs stacked on dry fabrics.

Optical Fiber

40 mm

25 mm

2.5 mm

1.1 mm(Skin)

1.25 mm(Flange)

Fig. 3. Embedding location of optical fibers.

-3500

-3000

-2500

-2000

-1500

-1000

-500

0

0 20 40 60 80 100

FBG 1FBG 2FBG 3FBG 5

Stra

in (μ

ε)

Load (kN)

Slightly strain changes

Fig. 6. Strain–load curves measured by FBGs.

Table 1Measured residual strains.

FBG1 FBG2 FBG3 FBG5

Strain (le) �143.5 �236.4 �228.0 �118.2

The FBGs were embedded into the interface between the skinand flange portions in order to monitor the damages around thelatter. The embedding location of an FBG is shown in Fig. 3.

(a) 0 kN (b) 40 kN

(c) 50 kN (d) 70 kN (e) 100 kN

Fig. 7. Out-of-plane displacement in skin.

-3500

-3000

-2500

-2000

-1500

-1000

-500

0

0 20 40 60 80 100

FL1FL2FL3FL4FL5FL6

Stra

in (μ

ε)

Load (kN)

Nonlinear behavior(< 40 kN)

Fig. 8. Strain–load curves measured by strain gauges FL1–FL6.

-4000

-3000

-2000

-1000

0

1000

2000

0 20 40 60 80 100

SK1SK2SK3SK4SK5SK6SK7SK8

Stra

in (μ

ε )

Load (kN)

2nd Buckling(53.8 kN)

3rd Buckling(59.8 kN)

1st Buckling

Fig. 9. Strain–load curves measured by strain gauges SK1–SK8.

Strain gauge Location

FL1–FL6 Flange

SK1–SK4 Skin

(Stringer side)

SK5–SK8 Skin

(Skin side)

Locations of FBGs correspond to

those of strain gauges in FL series.

SK1

SK2

SK3

SK4

FL4 FL5

FL6

FBG-A FL1 FL2 FL3

FBG-B

Fig. 10. Arrangement of FBGs and strain gauges in panel for impact test.

S.-i. Takeda et al. / Composite Structures 94 (2012) 813–819 815

3. Monitoring of buckling behavior

A compression test was carried out to investigate the relation-ship between the signals from the FBGs and the buckling behaviorof the panel at relatively low compressive loading. Fig. 4 shows theappearance of the stiffened panel. The loading was applied by ascrew-driven testing machine (INSTRON 5589) under displace-ment control. During the test, out-of-plane displacement of theskin panel was observed by ARAMIS (GOM mgH), an optical 3Ddeformation analysis system. Four FBG sensors were embeddedin the panel, and 14 strain gauges were bonded to the surface ofthe panel. The arrangement of the FBGs and strain gauges are sum-marized in Fig. 5.

Fig. 6 shows the strain–load curves for the FBGs, with the load-ing varied from 0 kN to 100 kN. All FBGs showed compressivestrain at 0 kN because the panel is cooling down from the curingtemperature to room temperature. The residual strains measuredby the FBGs are summarized in Table 1. Although the influenceof location on the residual strain is unclear, the residual strain atthe interface between the skin and the stringer was estimated tobe about �100 to �250 le. The strain–load curves were almost lin-ear up to a compressive loading of 40 kN because global bucklingdid not occur. This behavior was confirmed by measuring the

out-of-plane displacement in the panel. Fig. 7 shows the out-of-plane displacement in the skin panel obtained by ARAMIS, wherethe color contour plots indicate the level of displacement. In thisstudy, the first buckling occurred at 40 kN in a four half-wavemode (Fig. 7b). The buckling mode changed to five half-waves at50 kN (Fig. 7c). Finally, at 70 and 100 kN, the stringer webs startedto buckle and the skin bay buckling mode changed to six half-waves (Fig. 7d and e).

Fixed by Clamp

Skin Side

Fixed by Clamp

Impact BImpact A

40 mm

80 mm

FBG-A

FBG-B

Stringer Side

Fig. 11. Impact locations in CFRP stiffened panel.

1546

1547

1548

1549

1550

-3340

-2505

-1670

-835

0

0 50 100 150

FBG-AFBG-B

Wav

elen

gth

(nm

)

Stra

in (μ

ε)

Time (ms)

Fig. 12. Wavelength and strain changes during impact loading.

(a) Impact A (b) Impact B

Flange

Damages Damages

Flange

Fig. 14. Photographs of impact damages observed by ultrasonic C-scan.

816 S.-i. Takeda et al. / Composite Structures 94 (2012) 813–819

The results for the strain gauges FL1–FL6 are shown in Fig. 8.The first buckling was also confirmed by the nonlinear behaviorof the strain changes. After the nonlinear behavior began, therewere opposite changes between the strains measured by FBGsand those measured by their counterpart strain gauges. For exam-ple, when the compressive strain of FBG3 increased, that of FL3 de-creased due to buckling. This is because the neutral axis of thepanel is located within the flange portion. The results for the straingauges SK1–SK8 are shown in Fig. 9. The drastic strain changesindicated second and third bucklings which began at 53.8 kN and59.8 kN, respectively. These behaviors were confirmed by the

0

0.001

0.002

0.003

0.004

0.005

1544 1546 1548 1550 1552 1554

Opt

ical

Pow

er (W

)

Wavelength (nm)

FBG A

After_Impact

Before_Impact

Fig. 13. Reflection spectra befo

results of out-of-plane displacement (Fig. 7c and d). The embeddedFBG sensors detected strain changes at about 40 kN and 60 kN, cor-responding to the first and third bucklings, respectively.

4. Monitoring of impact damages

Impact damages were monitored in the same CFRP stiffenedpanels as those in the previous section. The arrangement of theFBGs and strain gauges is summarized in Fig. 10. The location ofthe two FBGs and the 14 strain gauges were also the same as thosein the previous section. Two low-velocity impacts were applied tothe panel by a drop-weight impact tester as shown in Fig. 11. Thefixtures of the panel were clamped along the direction of impact.One impact location was the surface of web portion just aboveFBG-A, and another was just above FBG-B. The impact energies

0

0.001

0.002

0.003

0.004

0.005

Opt

ical

Pow

er (W

)

Wavelength (nm)

FBG B

After_Impact

Before_Impact

1544 1546 1548 1550 1552 1554

re and after impact tests.

-8000

-6000

-4000

-2000

0

2000

4000

0 50 100 150 200 0 50 100 150 200

SK1SK2SK3SK4SK5SK6SK7SK8

Stra

in (μ

ε)

Load (kN)

Final failure(185 kN)

(a) Panel without impacts (b) Panel with impacts

-8000

-6000

-4000

-2000

0

2000

4000

SK1SK2SK3SK4SK5SK6SK7SK8

Stra

in (μ

ε)

Load (kN)

Final failure(168 kN)

Divergence

Divergence

Fig. 15. Strain–load curves until final failure measured by strain gauges SK1–SK8.

S.-i. Takeda et al. / Composite Structures 94 (2012) 813–819 817

on FBG-A and FBG-B were 3.35 J/mm (half the value of the stan-dard in ASTM D7136) and 1.675 J/mm (a quarter of the standard),respectively.

The center wavelength of the reflected light has commonly beenused as an indicator in measurements using FBGs. Thus the centerwavelengths were measured by a high-speed monitor measure-ment system (SF3111A, Anritsu Corporation) during impactloading. Fig. 12 shows the wavelength and strain changes forFBG-A and FBG-B. The wavelength was converted to strain usingthe previous strain sensitivity, where 1550 nm corresponds tothe reference value of 0 le. Given that the wavelengths suddenlydropped at 70 ms after measurement began, impact loadings weredetected by changes in the center wavelengths of the FBGs. Thecenter wavelength of FBG-B increased again after impact becausethe impact is relatively small. The difference in wavelength beforeand after impact was about 0.31 nm, corresponding to a strain of262.7 le using the strain sensitivity above.

The axial strain measured by FBG-B was about 2700 le. How-ever, the strain measured by FBG-A was unavailable. This is be-cause the strain-measuring method uses the center wavelengthin FBGs. The reflection spectra before and after impact are shownin Fig. 13. The center wavelength for FBG-A could not be detectedbecause the shape of the spectrum became complicated. It was as-sumed that the impact damage around FBG-A was larger than thataround FBG-B. The impact damages were observed using a 3Dultrasonic inspection system (Matrixeye 64, Toshiba Corporation),as shown in Fig. 14. A characteristic of the impact damages in this

-8000

-6000

-4000

-2000

0

0 50 100 150 200

FL1FL6FL1_impactFL6_impact

Stra

in (μ

ε)

Load (kN)

Final failurefor intact panel

(185 kN)

Final failurefor impacted panel

(168 kN)

Fig. 16. Strain–load curves until final failure measured by strain gauges FL1 andFL6.

study was that interlaminar delaminations mainly occurred insidethe skin. The delaminations around FBG-A grew over the flangeportion and were confirmed to be considerably larger than thosearound FBG-B. As a result, it was found that the strain-measuringmethod using the center wavelength was effective for the detec-tion of impact events. Although this method is a candidate forFBG sensors at this stage, it is possible that loss of the center wave-length in the spectrum was misjudged as the breakage of opticalfibers in the case of relatively large impacts. The monitoring meth-od using the full spectral signal including multiple indicators suchas optical power and the spectral shape was found to have anadvantage over that using the center wavelength.

5. Monitoring of damages in panels during compression loading

Compression tests were performed until final failure for panelswith and without impact damages so that differences in bucklingbehaviors could be investigated. The failure load of the panel withimpact damages was 168 kN, which was less than that of the panelwithout such damages (185 kN). The decrease in the failure loadwas caused by the impact damages stretching across the flangeportion to the skin portion as shown in Fig. 14a. The global behav-iors of the two panels are represented by the strain–load curvesmeasured with the strain gauges SK1–SK8 in Fig. 15. The nonlin-earity of the curves becomes pronounced over 100 kN for the panelwith impacts. Although a pair of gauges in the stringer side (SK1and SK2) and a pair in the skin side (SK5 and SK6) showed similarbehaviors for the panel without impacts, the compressive strain of

-8000

-6000

-4000

-2000

0

2000

0 50 100 150 200

FBG1FBG2FBG3FBG5FBG-AFBG-B

Stra

in (μ

ε)

Load (kN)

Fig. 17. Strain–load curves until final failure measured by FBGs.

0

0.0003

0.0006

0.0009

0.0012

0.0015

1540 1545 1550 1555

0kN (After Impact)100kN110kN0kN (After Compression)

Opt

ical

Pow

er (W

)

Wavelength (nm)

110kN

Fig. 20. Reflection spectra of FBG-A under compressive loadings.

0.001

0.002

0.003

0.004

0.0050kN (After Impact)100kN110kN0kN (After Compression)

Opt

ical

Pow

er (n

W)

818 S.-i. Takeda et al. / Composite Structures 94 (2012) 813–819

SK2 and the tensile strain of SK6 decreased as a result of impactdamages. The divergences between SK3 and SK8 and betweenSK4 and SK7 also increased for the panel with impacts. The impactdamages strongly affected the strains at the skin portion becausethe panel was classified as a thin-skin panel. Fig. 16 showsstrain–load curves by strain gauges FL1 and FL6 for the intactand impacted panels. The locations of FL1 and FL6 correspond tothose of FBG-A and FBG-B, respectively. The differences betweenthe two panels were very small despite the impact damages havingoccurred right beneath the strain gauges. This indicates that theloading paths were maintained at the interface between the skinand flange portions because the delaminations were located insidethe former.

Fig. 17 shows the strain–load curves as measured by the FBGs.The curves indicated buckling initiation at 40 kN, in good agree-ment with the measurements obtained by the strain gauges. Com-pared with the monotonically increasing compressive strains forFBG1, FBG2, and FBG5, the change in FBG3 was exceptional from120 kN to final failure. Figs. 18 and 19 show the reflection spectraof FBG1 and FBG3, respectively. Although both spectra shifted tolower wavelengths with increasing compressive loadings, thespectral shape in the FBG3 clearly changed from 120 kN. Given thatthe strain change resulting from buckling was relatively small, itwas assumed that such a large change was caused by debondingbetween the skin and the flange rather than the growth of inter-laminar delaminations inside the skin. The compressive strain inFBG-A drastically increased to �6000 le at 100 kN and then de-creased again to �1000 le at 110 kN. These strain changes wereconfirmed by the spectral changes in FBG-A as shown in Fig. 20.It had been reported that the shape of the spectrum changed from

0

0.005

0.01

0.015

1540 1542 1544 1546 1548 1550

0kN - 180kN0kN (After Compression)

Opt

ical

Pow

er (W

)

Wavelength (nm)

0kN80kN

100kN

120kN140kN160kN180kN

Fig. 18. Reflection spectra of FBG1 under compressive loadings.

0

0.005

0.01

0.015

1540 1542 1544 1546 1548 1550

0kN - 180kN0kN (After compression)

Opt

ical

Pow

er (W

)

Wavelength (nm)

0kN80kN100kN

120kN140kN160kN180kN

0kN

Fig. 19. Reflection spectra of FBG3 under compressive loadings.

01540 1545 1550 1555

Wavelength (nm)

Fig. 21. Reflection spectra of FBG-B under compressive loadings.

(a) FBG3 (b) FBG1

(c) FBG-A (d) FBG-B

Fig. 22. Photographs of areas around FBGs following compression tests.

S.-i. Takeda et al. / Composite Structures 94 (2012) 813–819 819

a complicated shape to one with a single peak as a result of uni-form strain when the damage grew along the entire gauge lengthof the FBG [14]. Given that the spectrum was restored to its origi-nal shape, the large damage was assumed to be caused bycompression loading. These changes are different from the spectralchanges in FBG-B shown in Fig. 21, and from those in FBG1(Fig. 18).

Following the compression tests, the appearances of the panelswere visually inspected. Photographs of areas around the FBGs areshown in Fig. 22. Debonding between the skin and flange wasfound to occur in FBG3 and FBG-A. In particular, a catastrophic fail-ure through the flange from the skin to the web was observedaround FBG-A. The monitoring method using full spectral signalscould detect damage growth under compressive loading in bothpanels with and without impact damages.

6. Conclusion

FBGs could monitor slight strain changes result from bucklingbehavior in CFRP stiffened panels under compressive loading.Although the measured strain will differ depending on the embed-ding location or pasting position, FBGs can be used in place of elec-trical strain gauges. When FBGs are embedded in the panel, amonitoring method using full spectral signals is effective in evalu-ating the impact damages in detail. FBGs can also evaluate damagegrowth until final failure in panels with and without impact dam-ages. The advantage of the method is that it is applicable not onlyto strain measurement but also to damage detection. We hope thatthe present results will be helpful in practical application giventhat key issues of SHM in CFRP stiffened panels arise in impactdamages and/or the interface between the skin and stringer.

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