Damage Assessment of Composite Structures Using Digital Image ...

Appl Compos Mater (2014) 21:91–106 DOI 10.1007/s10443-013-9352-5

Damage Assessment of Composite Structures Using Digital Image Correlation M. A. Caminero · M. Lopez-Pedrosa · C. Pinna · C. Soutis

Received: 2 September 2013 / Accepted: 30 September 2013 / Published online: 22 October 2013 © Springer Science+Business Media Dordrecht 2013

Abstract The steady increase of Carbon-Fiber Reinforced Polymer (CFRP) Structures in modern aircraft will reach a new dimension with the entry into service of the Boeing 787 and Airbus 350. Replacement of damaged parts will not be a preferable solution due to the high level of integration and the large size of the components involved. Consequently the need to develop repair techniques and processes for composite components is readily apparent. Bonded patch repair technologies provide an alternative to mechanically fastened repairs with significantly higher performance, especially for relatively thin skins. Carefully designed adhesively bonded patches can lead to cost effective and highly efficient repairs in comparison with conventional riveted patch repairs that cut fibers and introduce highly strained regions. In this work, the assessment of the damage process taking place in notched (open-hole) specimens under uniaxial tensile loading was studied. Two-dimensional (2D) and three-dimensional (3D) Digital Image Correlation (DIC) techniques were employed to obtain full-field surface strain measurements in carbon-fiber/epoxy T700/M21 composite plates with different stacking sequences in the presence of an open circular hole. Penetrant enhanced X-ray radiographs were taken to identify damage location and extent after loading around the hole. DIC strain fields were compared to finite element predictions. In addition, DIC techniques were used to characterise damage and performance of adhesively bonded patch repairs in composite panels under tensile loading. This part of work relates to strength/stiffness restoration of damaged composite aircraft that becomes more important as composites are used more

M. A. Caminero (B) Escuela Técnica Superior de Ingenieros Industriales, Universidad de Castilla-La Mancha, Campus Universitario s/n, 13071 Ciudad Real, Spain e-mail: [email protected] M. Lopez-Pedrosa · C. Pinna Department of Mechanical Engineering, The University of Sheffield, Sheffield S1 3JD, UK C. Soutis Aerospace Research Institute, The University of Manchester, Manchester M13 9PL, UK

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extensively in the construction of modern jet airliners. The behaviour of bonded patches under loading was monitored using DIC full-field strain measurements. Location and extent of damage identified by X-ray radiography correlates well with DIC strain results giving confidence to the technique for structural health monitoring of bonded patches. Keywords Composites · Stress concentration · Open hole tension · Bonded patch repair · Damage detection · Digital image correlation

1 Background High performance composites are commercially used in the fabrication of aircraft, automotive and marine structures [1–3]. Many composite structures, such as aircraft frames, contain thousands of holes for joining purposes and open cut-outs for access. The development of stress concentrations in composite structures has always been of great concern and many studies have been conducted to investigate the effect of holes on strength [4–14]. The complex damage and failure mechanisms present during the loading stage in a composite laminate are increased due to the presence of a stress concentration, causing a wide range of effects, such as stress or strain gradients, not present in unnotched components [15–17]. It is therefore more desirable when performing experimental investigation on laminated composites to obtain extensive full-field strain data, rather than limited strain or displacement measurements obtained from traditional electrical strain gauges or extensometers [15, 18, 19]. There have been a number of studies in the literature using full-field measurements to examine the mechanical response of notched composite structures under mechanical loading: Strain gauges, photo-elastic coating [20, 21], Moiré interferometry [22–25], electronic speckle pattern interferometry [16], grid method [26–28], laser speckle method [29, 30] or digital image correlation (DIC) method [31–33]. However, only a limited number of studies focuses on the assessment of the damage process in notched composite structures using Digital Image Correlation techniques has been applied before [15, 18, 19, 27, 34–36]. The increasing use of composite materials in aerospace applications has led to increased interest in composite repair technologies. Replacement of the damaged areas is not a preferable solution due to the high level of integration and the large size of the components involved. Consequently the need to develop repair techniques and processes for composite components is readily apparent. Bonded patch repair technologies provide an alternative to mechanically fastened repairs with significantly higher performance, especially for relatively thin skins [34, 35, 37]. Carefully designed adhesively bonded patches can lead to cost effective and highly efficient repairs in comparison with conventional riveted patch repairs that cut fibers and introduce highly strain regions [38]. Experimental studies have shown that an adhesively bonded patch repair can restore up to 80 % of the original strength [34, 35, 39–41]. Experimental methods that provide full-field strain measurements of sufficient sensitivity are required to accurately characterise the effectiveness of the repair. These methods offer several advantages over traditional measurement methods. Full-field strain measurement techniques capture strains over the entire surface of interest. This comprehensive coverage enables complete characterisation of regions with high strain gradients.

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In this work the performance and the assessment of the damage in laminates with an open-hole and adhesively bonded repairs under uniaxial loading are examined. Two-dimensional (2D) and three-dimensional (3D) digital image correlation (DIC) was used to investigate the tensile strain field in laminates made from T700/M21 unidirectional prepreg. Tensile tests with several loading/unloading cycles at different percentages of the failure stress were performed in order to develop damage around the hole. X-radiography was performed to identify the damage localized around the hole. In the second part of the work, DIC technique was used to investigate damage and performance of adhesively bonded patch repairs in composite panels loaded in tension. These DIC results were compared to X-ray images (Off-line technique) for verification.

2 Experimental Procedure 2.1 Open-hole Tensile Specimens The specimens were fabricated from commercially available (Hexcel Composites Ltd.) carbon/epoxy pre-impregnated tapes 0.262 mm thick. The prepreg tapes were made of unidirectional continuous high tensile strength carbon fibers (T700) preimpregnated with Hexcel M21 epoxy resin (35 vol % resin content). The roll was 300 mm wide and 0.262 mm thick. The material was laid up by hand in 200 mm wide by 200 mm long plates with different stacking sequences: Eight-ply unidirectional [0]8 and [90]8 laminates, eight-ply [0/90]2s and [±45]2s laminates with a total thickness of 1.02 mm. These stacking sequences helped to gain an insight into the assessment of different types of damage in notched specimens under uniaxial loading (splitting, matrix cracking, delamination) using DIC. The basic in-plane stiffness and strength of the T700/M21 unidirectional laminate under tensile and compressive loading provided by the materials manufacturer, Hexcel Composite Ltd., are presented in Table 1. The standard cure cycle recommended by Hexcel Composites Ltd. was used for these laminates. The plates were cured at 7 bar autoclave pressure together with slow heating rate (1–3 ◦ C/min), held at 180 ◦ C for 120 min and followed by a cooling rate of 2–5 ◦ C/min. The ASTM D3039 tensile method was followed for testing specimens made from the unidirectional and cross-ply laminates. A circular hole was drilled at the centre of the specimens with carbide tipped drill to characterize notch sensitivity and damage; the hole diameter (D) to specimen width (W) ratio was 0.1. The notched specimens were 190 mm long × 30 mm wide with a hole diameter of 3 mm. Glass fibre tabs (35 mm width and 2.4 mm thick) were bonded to each end

Table 1 Stiffness and strength properties for the T700/M21 carbon fibre/epoxy system

Mechanical properties

T700/M21

E11 (GPa) E22 (GPa) G12 (GPa) ν12 σ11T (MPa) σ11C (MPa) τ12 (MPa)

148 7.8 4.5 0.35 2375 1465 95

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Fig. 1 Scarf repair geometry details and cured scarf repair specimen

of the specimen giving a gauge length of 120 mm. A standard high strength, two component epoxy paste adhesive (Araldite®2015) was used for bonding end-taps to the specimens. All specimens were subjected to tensile loads on a 100 kN electromechanical testing machine at a fixed loading rate of 2 mm/min. 2.2 Adhesively Bonded Repairs: Scarf-type and External Bonded Patches A scarf patch repair was performed on an impact damaged CFRP panel, see Fig. 1. The resulting damage area at the front face was approximately 13 mm in diameter whereas the total damaged area on the back face was close to 25 mm long × 25 mm wide. The panel was made out of woven (2/2 twill) carbon fibre (HTA) with toughened epoxy resin system (M21) supplied by Hexcel Composites and moulded by Hurel-Hispano UK at a nominal 58 % fibre volume fraction. The panel’s lay-up

Table 2 Tensile strength of unnotched and notched specimens and elastic properties of HTA/M21

HTA/M21 (0/90/ ± 45/0/90)3T

Measurements

Unnotched tensile strength (σun ) Notched tensile strength (σn ) Elastic modulus (Ex ) Poisson’s ratio (νxy ) Shear modulus (Gxy )

585 MPa 379 MPa 49.4 GPa 0.26 19.6 GPa

Appl Compos Mater (2014) 21:91–106 Table 3 Average tensile strength and stiffness properties of a T700/M21 laminate

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T700/M21 [0/90/ ± 45/0/90]3T

Measurements

Unnotched tensile strength (σun ) Notched tensile strength (σn ) Elastic modulus (Ex ) Poisson’s ratio νxy  Shear modulus Gxy

996 MPa 576 MPa 67.5 GPa 0.28 26.4 GPa

sequence was (0/90/ ± 45/0/90)3T and its total thickness was 2.8 mm. After impact testing, the damaged area was removed with a rotating pad sander in order to simulate the result of clearance of damaged region. The averaged experimental results (using three unnotched and notched tensile specimens) for the composite system HTA/M21 are shown in Table 2. The gross cross-section of the specimen was used when calculating the resulting tensile strength, see reference [35]. Prior to testing the specimen, C-Scan was performed in order to detect any defects (patch or bond lines defects) or manufacturing faults which could cause premature failure or crack initiation, and it was found to be free of any defects. Glass fibre tabs were bonded to the front and back face of the specimen’s ends with a room temperature cured two component epoxy paste adhesive. With the tabs attached and the panels cut to their final widths, through holes were drilled in the tab region to accommodate the test fixture grips. The assembly was loaded in tension in a hydraulic testing machine. Similar experimental arrangement was developed for the external bonded patch repair analysis. Two composite panels of 200 mm wide by 200 mm long were fabricated by hand lay-up using the same prepreg tapes (T700/M21) following the stacking sequence [0/90/ ± 45/0/90]3T (18 plies and 2.4 mm thick). The panels were cured in an autoclave according to the manufacturer’s curing recommendations. Unnotched and notched strengths and the elastic stiffness properties are shown in Table 3. In order to experimentally simulate a bonded patch repair in a composite structure, a circular hole of 5 mm in diameter was drilled at the centre of the specimen that represents the removing of the damaged region and then external patches were bonded to the surface of the specimen, see Fig. 2. Circular hard patches, 35 mm in diameter, were manufactured from the same T700/M21 material as the substrate but with a stacking sequence of [0/90/ ± 45]s .

Fig. 2 Geometry details of one side and both sides bonded patch repaired specimen

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2.3 On-line Damage Monitoring Using Digital Image Correlation (DIC) DIC is a full-field and non-contact technique for the measurement of 2-dimensional (2D) or 3-dimensional (3D) surface displacement fields of an object. It is based on pattern recognition and exploits the correspondence between the speckle pattern that covers the surface of the tested specimens before and after loading. The speckle pattern can be prepared with white painting and sprayed with a black aerosol. This leads to a random structured aspect, which can be observed with a digital camera. The principle used to measure the whole-field displacements of the specimen surface is as follows: the unique pattern of random speckle is recorded twice, one before loading the specimen, and the other when the specimen is deformed. Then, by correlating sub-images in these two pictures recorded by the cameras, it is possible to determine the surface displacement vector [30]. In comparison with other experimental techniques such as Moiré interferometry and electric speckle interferometry, DIC is simple and robust because complicated surface treatment is not needed, its requirement for testing environment is low [19, 33] and under opportune conditions it shows a precision and versatility that is difficult to obtain using other techniques [42, 43]. When compared with the special requirement of traditional optical techniques, DIC is a versatile method because it takes advantage of the natural speckle pattern on the specimen and only needs a common camera and a computer to store the images to be subsequently analysed. On the other hand, the measurement accuracy of DIC can be affected by many factors, such as sub-pixel optimization algorithm, subset shape function, subset size, sub-pixel intensity interpolation scheme, image noise, camera lens distortion or the quality of a speckle pattern [44, 45]. The DIC images in this study were obtained by two digital cameras with a CCD matrix of more than five million pixels; model DCP 5.0 LIMESS Messtechnik & Software GmbH, see Fig. 3. The typical error of displacement measurement is less than 0.05 pixels, and that of strain measurement is less than 500 μm [30, 33]. The experimental data obtained in this study were processed with VIC-2D and -3D software from Correlated Solutions, Inc. The specimens were prepared with white painting and sprayed with a black aerosol to create a random speckle pattern. The specimen was illuminated by ordinary white light during the experiment. Fig. 3 Experimental set-up for the DIC testing under uniaxial tensile loading

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3 Results and Discussion 3.1 Open Hole Tensile Specimens As discussed earlier, the 3D DIC method was used to obtain the strain fields developed in 0, 90, 0/90 and +45/−45 T700/M21 notched specimens (190 mm × 30 mm with a 3 mm hole diameter) loaded in tension at a rate of 2 mm/min. It is shown that the technique provides quantitative information that can be used to identify possible damage zones, especially in high gradient areas as also reported by other workers [46, 47]. A speckle pattern was applied on the specimen surface using airbrush and white paint in order to perform the DIC measurements; the quality of the speckle pattern (random size and distribution) affects strain results so it needs to be carefully deposited. Prior to performing the tensile experiment, the camera system was calibrated using a uniformly spaced dot pattern with 15 mm spacing and through VIC-Snap software from Correlated Solutions, Inc. The images were recorded using VIC-Snap software from Correlated Solutions, Inc at a rate of 5 Hz. The experimental data obtained in this test were processed with VIC-3D software from Correlated Solutions, Inc. The correlation subset size was large enough to ensure that there was a sufficiently distinctive pattern contained in the area used for correlation. All specimens failed within the gauge length, except the eight-ply [0]8 laminate that failed near the grip and exhibited strength very close to the unnotched value due to the axial splitting that developed at the hole edge. In order to validate the DIC measurements, the experimental results have been compared to the elastic stress (or strain) concentration factor (SCF) in the case of notched specimens. Using the non-isotropic elastic theory [48–50], the SCF at the edge of the hole of an infinite orthotropic plate can be computed by    E yy E yy K∞ = 1 + 2 − ν yx + (1) Exx G yx where E yy , Exx and G yx are respectively the elastic longitudinal, transverse and shear moduli and ν yx the Poisson’s ratio of the considered laminate. Note that the y-axis is oriented along the loading direction, and x-axis is oriented transverse to the loading direction. With the assumption of linear elastic behaviour, the stress (strain) distribution and if normalized by the applied stress (or strain), the SCF along the x-axis can be written as     2  4  8

1 D D D D 6 Kt (x) = 2+ (2) +3 − (K∞ − 3) 5 −7 2 2x 2x 2x 2x where D is the notch diameter and x is the distance from the centre of the hole. In this study, the specimen width (W) to hole diameter (D) ratio was W/D = 10. As a result, it is reasonable to assume an infinite plate since W/D ≥ 8 [51], and hence the FWC (Finite-Width Correction) factor can approximately be equal to 1. The comparison of the Kt values obtained analytically using the elastic solution, Eq. 2 and measured by the DIC method for the notched composite laminates [0]8 and [90]8 subjected to tensile loading are shown in Fig. 4a and b. At and near the hole edge, the FE results differ very much from the DIC measurements due to high strain gradients; these high strains introduce local damage near the hole edge in the form of axial splitting

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Fig. 4 Stress (or strain) concentration factor (SCF) measured by digital image correlation method along the x-axis: a [0]8 notched laminate loaded in tension (applied load 16.8 kN, 32 % OHT) and b [90]8 notched laminate (applied load 0.745 kN, 92 % OHT)

and transverse matrix cracking (as revealed later on by X-ray radiography), which causes stress redistribution and hence reduced SCF. Further away from the hole, the DIC results are in very good agreement with the elastic FE solution since there is no damage. It should be noted that the DIC method may not be accurate enough at low load levels for comparisons with FE calculations that give strains at nodal points. Further studies may be needed on this topic. The notched specimens were all loaded in uniaxial tension at a displacement rate of 2 mm/min and interrupted at pre-determined load levels prior to failure, see Table 4. After stopping the loading at the prescribed stress level, the specimens were examined via X-radiography to assess their damage development. The strain fields obtained by the DIC technique at different applied tensile load levels are shown in Figs. 5, 6 and 7. The results show very high localized strains near the hole, and along directions following the fibre orientation (0, 90, +45/ − 45) that are expected at this

Appl Compos Mater (2014) 21:91–106 Table 4 Stress levels for interrupted testing program

99

T700/M21

% static stress

Failure stress

[0]8 [±45]2s  0/90 2s

36 % (624.2 MPa) 86 % (182.6 MPa) 93.7 % (667 MPa)

1694.4 MPa 212.1 MPa 711.7 MPa

loading level to introduce resin cracks (0◦ splits, 90◦ matrix cracking and 45◦ off-axis cracking) and/or fibre/matrix interface failure. Figures 5–7 show damage development under tensile loading in [0]8 , [±45]2s and [0/90]2s notched laminates using X-radiography and DIC strain fields. The Xradiographs were taken at an applied load of about 36, 86 and 94 % of their respective failure load. It should be noted that X-radiographs have also been taken before any loading was applied and had shown no initial damage due to hole drilling. The comparison of the three radiographs in Figs. 5–7 puts in light the different damage development of the studied composite laminates. Figure 5a depicts X-radiograph of the notched [0]8 , UD-ply specimen at an applied load of 19.1 kN (36 % of failure load). The specimen shows longitudinal splits (fibre/matrix splitting) tangential to the hole that isolate it from the rest of the specimen. Figure 5b, c and d show DIC results in terms of ε yy (loading direction), εxx (transversal direction) and εxy strain plots respectively at the same loading level. These DIC images show high strain  concentrations around the hole ε yy max = 2.4 % >> εnotched and the transverse and shear strain components show high values along narrow bands oriented in the loading direction and tangential to the hole, which correspond to the fibre/matrix splitting observed on the X-radiograph, see Fig. 5a. The notched [±45]2s angle-ply specimen at an applied load of 5.59 kN (86.2 % of failure load) exhibits off-axis cracks at the edge of the hole along the +45 and −45 fibre orientation but also multiple small offaxis cracks in the region around the hole and beyond. The darker region around the

Fig. 5 Damage and strain fields observed in a [0]8 UD-ply notched composite laminate at 36 % of failure load (average failure load = 51.85 kN, hole diameter = 3 mm). a X-ray radiograph and DIC strain distributions: b ε yy (loading direction), c εxx and d εxy distribution

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Fig. 6 Damage and strain fields developed in a [±45]2s angle-ply composite laminate at 86 % of failure load (average failure load = 6.49 kN, hole diameter = 3 mm). a X-ray radiograph and DIC strain distributions: b ε yy (loading direction) and c εxy distribution

hole in Fig. 6a corresponds to local delamination. Figure 6b and c show full-field DIC results in terms of strain distributions at the surface of the notched specimen at the same loading stage as in Fig. 6a. The results show the DIC strain plots correspond well with the damage pattern captured by the X-ray radiography. Figure 6b shows high strain concentrations at the   hole boundaries that result in ±45◦ matrix cracks and delamination ε yy max = 7.4 % . The top-left side presents higher strain concentration than the right side, as in the X-radiograph, and this can be due to specimen misalignment within the test frame or more likely specimen thickness variation along and across the specimen width. Figure 7 shows damage development in a T700/M21 notched [0/90]2s cross-ply specimen at an applied load of 20.41 kN (93.7 % of failure load). In Fig. 7a, 0◦ fibre/matrix splits around the hole and matrix cracks in 90◦ plies (black lines) are developed with small local delaminations that occur where these cracks overlap. Figure 7b and c depict strain plots obtained by the DIC technique. The results

Fig. 7 Damage and strain plots in a [0/90]2s cross-ply composite laminate at 94 % of failure load (average failure load = 21.78 kN, hole diameter = 3 mm). a X-ray radiograph and DIC strain distributions: b ε yy (loading direction) and c εxy distribution

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  show high strain areas ε yy max = 2.15 % in the vicinity of the hole where damage development was more extensive (splitting and 90◦ matrix cracking). The X-ray revealed damage which corresponds very well to the regions where high strain values were developed and successfully captured by the DIC method. 3.2 Adhesively Bonded Patch Repairs 3.2.1 Scarf Repair The scarf repaired panel was loaded in quasi-static tension and DIC images were recorded at 10–20 kN intervals up to 170 kN. At this load, the panel failed prematurely due to cracks that initiated around the holes that were drilled for gripping the specimen. This is a common complication in composite testing since composites exhibit brittle behaviour, hence they cannot undergo plastic flow to relieve the local stresses that initiate around the holes [40]. As a result manufacturing characteristics such as specimen gripping and specimen design should follow appropriate specifications for successful testing under static or dynamic loads; the specimen endtab should be of an appropriate size in relation to the gauge length. Nevertheless, for the characterization of the panel, the amount of data that was collected was sufficient for reaching some useful conclusions in relation to damage monitoring. DIC images obtained in this test were processed with the Vic-3D software from Correlated Solutions, Inc. The selected subset size was large enough to ensure that there is a sufficiently distinctive pattern contained in the area used for correlation. The 3-dimensional DIC analysis provided full-field strain maps of the repaired area for all the considered loading cases. The strain values of interest were obtained in the loading direction (ε yy ). The results of the most representative load levels are illustrated in Fig. 8.

Fig. 8 Full-field   DIC results: strain distribution at the surface of the repaired area in the loading direction ε yy at different loading stages: 40, 80, 120, 140, 150 and 170 (failure) kN

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Table 5 Average tensile strength of T700/M21 multidirectional laminate for single and double patch repair  T700/M21 0/90/ ± 45/0/90 3T Failure stress Single side repair Both sides repair

685 MPa 704 MPa

From the current analysis, an investigation of the local developed stress concentrations was possible in order to identify whether there was any internal damage developed in the repaired area or not. The tensile strength of the scarf repair is expected to be between the tensile strength values of the unnotched and the notched sample. However for precision purposes, it is well accepted that the optimum patch configuration can recover 70–80 % of the laminate’s undamaged strength [37, 39]. Therefore the maximum tensile strength of the scarf repair can be estimated to be approximately between 409–468 MPa. In addition there are several factors that can lead to lower strength such as the discontinuities introduced during the manufacturing process. According to the estimation of the maximum stress that developed at 170 kN locally around the hole, the stress was 415 MPa which is close to the failure stress of the repaired panel. This value was measured through the strain analysis of the DIC figures at an applied load of 170 kN. The current analysis concludes that it is quite possible that even though the load did not reach the ultimate strength of the scarf repair due to the premature grip failure, local internal damage is expected to develop. This assumption is well supported both from the DIC figures and the Lamb wave’s analysis performed in reference [35]. 3.2.2 External Bonded Repair Specimens with a single and double external patch bonded repairs were tested. The average strength of the unnotched specimen was 996 MPa. The average tensile strength of the specimen with a single patch (loaded in the 0◦ direction) was 685 MPa while a value of 704 MPa was recorded for the specimen with patches on both sides, see Table 5, therefore recovering about 70 % of the unnotched strength. At least three specimens were tested for each repair configuration. The specimen failure modes are illustrated in Fig. 9 for both repair configurations. Fig. 9 Failure of the single and double patch repaired specimens loaded in tension

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Fig. 10 DIC strain distributions in terms of Hencky strains at an applied tensile load of 45 kN for a T700/M21 specimen repaired with a single bonded patch. Loading direction is denoted by the y-axis. of the patch repaired region. b Strain distribution ε yy (loading direction) a An X-ray radiography  ε yy max = 4.06 %

The 3D DIC technique was used to examine the performance of the single patch repair when loaded in tension. This was accomplished by applying speckle pattern on both sides using white paint, as explained earlier in the paper. Both sides of the specimen were monitored in order to analyse the parent plate and the performance of the patch. Prior to performing the experiment, the camera system was calibrated accordingly and the specimen was illuminated by ordinary white light. Figure 10 shows 3D-DIC results in terms of Hencky strains for the specimen repaired with a bonded patch on one side at 45 kN. X-ray radiography was used to identify damage after loading. No evidenceof permanent damage is shown in Fig. 10a. The results in the loading direction ε yy show high strain areas at the edge of the patch (upper and lower parts) in the loading direction where damage is expected in the form of patch debonding. These results are in accordance with the failure analysis of references [39, 52]. This failure mode occurs when the parent plate is bonded with relatively strong patches inducing high stress in the bonded region and high shear stress in the adhesive. Hence the adhesive fails first and the patches are detached from the parent plate, leading to increased SCF at the hole edge and final fracture similar to that observed in the notched specimen.

4 Conclusions In the present work, the assessment of damage taking place in composite plates with an open hole when loaded in tension was examined using Digital Image Correlation (DIC), an optical full-field strain measurement technique. In addition, the DIC method was used to investigate damage and performance of adhesively bonded patch repairs in composite panels under tensile loading. It was successfully applied to

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measure surface strains which relate to possible damage that could occur in the composite system. The strain and displacement measurements were compared to analytical predictions with a good agreement. The high localised strains identified by the DIC technique were also in agreement with damaged captured by penetrant enhanced X-ray radiography. It has been shown that resin cracks and delaminations that appear in the radiographs coincide with the high strains developed around the hole region and successfully measured by DIC. Similar success was achieved in the damage assessment of laminates with patch repairs under tensile loading were the DIC full-field strain measurements were in good agreement with results of previous studies [35, 36]. The present work has shown the potential of DIC for on-line monitoring of composite structures. It is revealed as an efficient and simple methodology to study possible damage in laminates with discontinuities and adhesively bonded patch repairs. It is still a challenge though to accurately identify internal damage in the form of resin cracking, fibre/matrix interface failure, delamination [53] and fibre microbuckling, which can be a critical damage mechanism, when the structure is loaded in compression [54]. Further work is required to investigate the ability of the DIC technique to capture the stacking sequence effect on surface displacements and strains, i.e., [0/90]s versus [90/0]s .

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