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Study on impact damage mechanisms and TAI capacity for the composite scarf repair of the primary load-bearing level
Composite Structures 181 (2017) 183–193
Contents lists available at ScienceDirect
Composite Structures journal homepage: www.elsevier.com/locate/compstruct
Study on impact damage mechanisms and TAI capacity for the composite scarf repair of the primary load-bearing level Bin Liu a,⇑, Fei Xu a, Jian Qin b, Zhixian Lu a a b
School of Aeronautics, Northwestern Polytechnical University, Youyi West Road, No. 127, Xi’an, China Eastern Airlines Technic Co., Ltd Northwest Branch, Xi’an, China
a r t i c l e
i n f o
Article history: Received 18 April 2017 Revised 29 July 2017 Accepted 29 August 2017 Available online 31 August 2017 Keywords: Impact damage Microcrack Composite repair Bonding Damage tolerance
a b s t r a c t As composite material plays a leading role in aircraft, composite bonding repair has been extensively applied. Among composite bonding repairs, the scarf bonding repair is widely adopted and has high repair efficiency especially in primary load-bearing structures. However, the impact damage tolerance and impact damage mechanisms were not considered for repaired structure integrity design yet. This paper experimentally and numerically studied the scarfed bonding repair of the advanced CFRP, which may suffer a low velocity impact load in service. At the central location of adhesive zone, impact energy and response regularity were studied to reveal the competition failure mechanism for inner kinds of materials. In the impact procedure, double force peaks phenomenon and four typical phases were found. Tension after impact (TAI) capacities were also tested to explain the impact damage effects on residual strength. The adhesive damage has strong influence over tension after impact capability. The most easily broken location in the bonded zone is the feathered tip on the back of impact point. The critical impact energy 23 J exists for this size of specimen. When the impact energy is higher than the critical 23 J, except for the composites damage, the adhesive damage can be observed at the second force dropping. The scarfed adhesive damage occurred at the scarf feathered tip of back side. Ó 2017 Elsevier Ltd. All rights reserved.
1. Introduction Advanced CFRP (Carbon Fiber Reinforced Plastics) application in aircraft has turned from secondary structures to primary load bearing structures. High strength and middle modulus CFRP promote composites application to load bearing structures, such as T700, T800, T1100 and MJ series. Damage of aircraft load bearing level structures always occurred in impacting and maintaining. As composites have low interlaminar toughness, low velocity impact leads to an invisible inner damage. For the upgraded composites, the mechanical characteristics would change. Hence, the problems of composite repair exist in load bearing structures and are complicated and important [1,2]. At present, composite repair technique has become one of key factors for further composite applications. The existing aircraft structures design considered only the static strength, stiffness recovering level, but didn’t include the second impact in the service of repaired structures. When stepped and scarf repaired structures are impacted, the innerlaminar, interlaminar and adhesive will suffer damage in some extent. The competed failure mecha⇑ Corresponding author. E-mail address: [email protected] (B. Liu). http://dx.doi.org/10.1016/j.compstruct.2017.08.087 0263-8223/Ó 2017 Elsevier Ltd. All rights reserved.
nisms of different materials must be revealed clearly to obtain the impact tolerance and resistance for composite repairs. Low velocity impact has strong influence to composite stepped and scarf repairs [3–5]. In 2006, U.K. Vaidya [6] found that outplane load leads to higher peel stress and concentration than inplane load case, and adhesive crack initiates as mixture mode and transforms to mode II. In 2007, I. Takahashi [7] applied health monitoring technique to detect scarf adhesive damage. A.B. Harman [8] discovered that the impact damage tolerance for composite scarf repair structure reduced comparing with laminate plates. H.C.H. Li adopted quasi-static out-plane load to substitute impact dynamic load and found that in-plane prestrain affects bending stiffness. In 2012, M.K. Kim [9] studied the impact damage of laminate plates and composite scarf repair under combined in-plane load and out-plane impact, and gave the conclusion of that impact resistance increase obviously as the prestrain increases. Berrin Gunaydın [10] investigated the effects of composite repair patches and number of patch layers on the fatigue behavior of surfacenotched composite pipes. In 2015, C.H. Wang [11] experimentally studied CAI (Compression after Impact) mechanical performance for 2 mm thickness stepped repair. C.H. Wang [12] proposed that the existed design methodologies consider only loading capacity of integrated repair structures without including adhesive damage
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and delamination. J.J. Andrew [13] conducted experiments of multiple low velocity impact and CAI for GFRP bonding lap joints. Effects of number of patch layers on the burst pressure of low velocity impact damaged tubes that have been repaired with composite patches were investigated [14]. In 2016, G. Balaganesan [15] found that the patch of GFRP stepped repair structures could absorb 50% to 80% impact energy. In 2017, O. Balcı [16] studied the honey comb GFRP structures repaired by lap joint and concluded 3 J impact energy produce surface damage and 8 J is the critical penetrated energy. S.R.M. Coelho [17] concerned the mechanical performance and the damage for single patch and double patches with multiple impacts. In summary, impact problems of composite scarfed and stepped repair were studied by few researchers, but impact response, damage mechanisms and damage modes were not researched systematically. The conjunction studies of macroscopic damage phenomenon and microscopic failure cracks [18–20] were conducted much less and needed to be carried out. 2. Specimens and tests setup 2.1. Specimens As shown in Fig. 1, the specimen length, width and thickness are 300 mm, 50 mm and 4 mm [5]. Scarfed surfaces with the angle of 5° were cut and polished by machines. The scarf bond length is 45.7 mm and the bond surface thickness is 0.2 mm. The supplements of both ends are clamped area for tensile test machine
and the size is 50 50 mm. To avoid the stress concentration, the clamped supplements were machined with the angle of 15°. The composite scarf adherends are hard patch repair which were pre-cured. The composite adherends were fabricated using the layup consisting of 32 plies of high performance T700/LT03A carbon/epoxy prepregs. The stack sequence of composites is [45/ 0/-45/90/0/45/0/-45/0/0/90/0/-45/0/45/0]S. The percentage of 0°, 45° and 90° is [50/37.5/12.5]%. This kind of layup being composed of 0° of 50% is a representative for primary-load bearing composite structures [1]. The prepreg material has a nominal ply with thickness of 0.125 mm and curing temperature of 120 °C. The adhesive used to joint two composite adherends which were made from epoxy of Cytec FM73M with thickness of approximately 0.2 mm and curing temperature of 120 °C. FM73M is an aerospace high-performance film adhesive. 2.2. Tests setup On account of that the existing impact standards for composites didn’t include relative narrow specimen, we designed the narrow impact setup for composite scarf repair by referring to A.B. Harman and A.N. Rider [5]. Fig. 2(a) shows the schematic of impact fixture and specimen. Fig. 2(b) shows the experimental setup coinciding to Fig. 2(a). The drop-testing machine Instron-9250HV was used and the tip diameter was 16 mm. The quality of the punch is 5.067 kg. During impact process, the responses of impact load, deflection, velocity, absorbed energy and time can be obtained by the testing machine. To investigate the effect of impact energy on composite
Fig. 1. Specimen schematic of the composite scarf repair.
Fig. 2. The impact fixture and specimen: (a) schematic; (b) setup.
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scarf repair, 4 J, 6 J, 8 J, 12 J, 16 J, 20 J, 23 J, 25 J and 30 J were chosen. The impact initiation velocities were 1.194 m/s, 1.462 m/s, 1.689 m/s, 2.069 m/s, 2.389 m/s, 2.671 m/s, 2.986 m/s, and 3.271 m/s for the nine impact energy levels respectively. Tension and tension after impact tests were operated on the static testing machine with the displacement control of 0.5 mm/min by ASTMD3039/D3039M-008. 3. Experimental results and discussion 3.1. Impact response On the basis of the experimental results with different impact energy, the graphs of impact load vs. time show double peaks. We gave the impact load and deflection with time for different impact energy levels of 8 J and 25 J respectively as
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shown in Fig. 3(a) and Fig. 3(b). From load-time graphs, we provided the schematic of load-time in Fig. 4(a). The first peak of force denoted by Fmax1 keeps constant in 5000 N–6000 N, and the second peak of force denoted by Fmax2 rises from 4000 N to 12000 N as the impact energy increases as shown in Fig. 4(b). The regularity research of impact load and deflection were conducted as follows. The typical 4 phases can be summarized from the studies. Firstly, the example of 8 J impact energy was provided as shown in Fig. 5(a). Phase I, linear elastic phase: the impact load varied from 0 to Fmax1. In this phase, the repair structure suffered an elastic deformation. Phase II, impact load dropping: the impact load dropped suddenly after Fmax1. The damage of repair structure could occur. In this short time, the partial strain energy of structure was released
Fig. 3. Impact load and deflection with time of (a) 8 J and (b) 25 J.
Fig. 4. (a) Schematic of impact load with time and (b) variation of load peak Fmax1 and Fmax2 with impact energy.
Fig. 5. Impact load-deflection curves (a) 8 J (b) 25 J.
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by the damage. Thus the impact load dropped suddenly and the stiffness of structure also decreased. Phase III, impact load rising again: the graph of load-deflection vibrated and tended to being stable in this phase till the impact load reached to Fmax2. Phase IV, the impactor rebounding: the strain energy was released completely and translated to impactor kinetic energy. After Fmax2, the impact load decreased non-linearly to zero. However, as the impact energy rose to being greater than 23 J, the second peak of impact load Fmax2 dropped suddenly. Here we
Fig. 6. The front impact dent and the back adhesive crack with 25 J impact energy.
gave the relative distinct graph of load-deflection with 25 J impact energy as shown in Fig. 5(b). 3.2. Impact damage In order to reveal the damage mechanism of impact and tension after impact, we studied three kinds of damage which respectively were visual visible damage, microscope visible damage and ultrasonic inspecting damage. Firstly, the visual visible damage showed that when impact energy rose greater than 23 J, the dent depth of front surface became clearly and the scarfed tip of back surface appeared obvious adhesive fracture. The front surface and back surface of 25 J impacted specimen are illustrated in Fig. 6. In order to get the inner damage details, the specimens were cut along the longitudinal central line. To avoid producing new damage for impacted specimens, we chose diamond cutting blade with rotate speed of 1500 r/min and feed speed of 300 mm/min. From the central cut section, by using microscope, the damage details of 25 J were indicated in Fig. 7. The damage modes of 25 J included adhesive damage, matrix crack of 90° plies, and the interlaminar delamination 0°–90° and 0°–±45°. Among all the damage modes, adhesive damage was most obvious and located from the back scarfed tip to the tenth ply counting from the back. The adhesive damage amount was totally 31% of the scarf bonding length. The microscope visible damage of adhesive under 20 J and 23 J was not found in the cut section. As the impact energy increases to 30 J, not only the adhesive had more obvious damage (37.5%) than 25 J (31%), but also the
Fig. 7. The microscopic damage image captured from central cut section for the specimen with 25 J impact energy.
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Fig. 8. The microscopic damage image captured from central cut section for the specimen with 30 J impact energy.
fiber on back side existed fracture. Meanwhile, a plastic deformation appeared in composite materials and adhesive under the impact point, as shown in Fig. 8. Summarizing Fig. 7 and Fig. 8, Fig. 9 and Fig. 10 showed the visualized schematics of the central cut section damage found by microscope for 25 J and 30 J impact energy respectively. In the schematics, the white color denotes 0° plies, gray color ±45°, black color 90°, and red lines illustrate the crack path. Fig. 11 shows the ultrasonic C-scan images of impacted damage for different impact energy. The different colors denote the damage
Fig. 11. The ultrasonic C-scan images of impacted damage for different impact energy.
depth and the blue color denotes no damage exists. The length and width can be also red from Fig. 11. Due to the shortage of C-scan, we couldn’t get the inner adhesive damage information. However, the damage size gotten by C-scan revealed the composite materials damage variation with impact energy. Damage length and width
Fig. 9. The schematic of microscopic visible damage from central section with 25 J impact energy.
Fig. 10. The schematic of microscopic visible damage from central section with 30 J impact energy.
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The two parts of adherends of specimens impacted were separated from each other under tension load. The two parts firstly were combined together, then, by means of microscope, were observed to reveal the mechanisms for tension after impact. Here, the microscope images of 25 J impact energy was provided as illustrated in Fig. 14. The damage details in Fig. 14 can be shown as the schematic type in Fig. 15. The damage modes included compositeadhesive interface fracture, 90°–0° and ±45°–0° delamination, adhesive cohesive damage, 90° plies and 45° plies matrix crack. Compared with specimens impacted without tension, the delamination and matrix crack length and width of tension after impact became larger. It explains that tension load made the impacted damage severe and adhesive damage was the major reason for load capacity reducing.
Fig. 12. The variation of length and width of damage size with impact energy.
varied nonlinearly as the impact energy increased. We adopted the least square method and obtained the fitted quadratic polynomial to descript damage size variation with impact energy. The damage of the composite scarf repair propagates easily along length direction compared to width (Fig. 12). 3.3. Tension after impact After studying the impact damage mechanisms for composite scarf repair, we tried to combine impact damage mechanisms with load capacity of tension after impact. Below the threshold impact energy, there was no obvious decreasing of load capacity of tension after impact as shown in Fig. 13. When impact energy increased to 25 J, load capacity of tension after impact decreased obviously. As impact energy equaled to 30 J, load capacity of tension after impact continued to decline. Thus 25 J can be considered as the critical energy or threshold level of energy. The original average strength T0 of no impact (0 J) repairs is 69.9 kN. TE delegates TAI strength under impact energy E. So the TAI capability reduction can be calculated by equation (T0 TE)/ T0 100%. Compared with no impact specimens, the load capacity T25 of TAI with 25 J dropped 23% and 41% relatively for two specimens while T30 with 30 J dropped 29% and 40%. As what was mentioned before, adhesive damage percentage were 31% and 37.5% for 25 J and 30 J. We can clearly see that TAI load capacity decreasing has nearly the same amount as the adhesive damage percentage. Therefore, we could draw a conclusion that TAI load capacity is closely related to adhesive damage for composite scarf repair.
Fig. 13. Load capacity of TAI with discrete impact energies.
4. Simulation On account of that impact procedure time is very short, and C-scan and microscope tools cannot capture the sequence of different materials mechanisms, simulation is needed to recruit with experimental studies. The mechanical properties of the composite material and the adhesive, as provided by the manufactures, are presented in Tables 1–3. The whole FEM model mainly includes impactor, support, press head, and specimen as shown in Fig. 16. From Fig. 16, the FEM model of composite scarf repair includes composite innerlayer 3D elements, composite interlamina cohesive zone elements, and scarfed adhesive cohesive zone elements. The composite innerlayer elements are 8 nodes hexahedral reduced integration elements and 6 nodes pentahedral reduced integration elements. The composite interlamina elements and scarfed adhesive elements are 8 nodes hexahedral cohesive zone elements. Support, press, and impactor include 4 nodes tetrahedral, 6 nodes pentahedral and 8 nodes hexahedral reduced integration elements. The typical 4 phases of impact load vs. displacement was found from the experimental studies. In the 4 phases, the reason of load dropping of phase II and phase IV and the competed damage sequence of materials cannot be obtained in experiments. We tried to explain and reveal them by FEM simulation of two kinds of impact energy 8 J and 25 J. As shown in Fig. 17 and Fig. 18, impact load and time relationship of simulation coincides well with the experimental results. The simulation indicated that the first impact load dropping of 8 J and 25 J was caused by composite delamination and the second impact load dropping of 25 J was caused by scarfed adhesive damage. In the whole impact procedure, composite material damage certainly contained matrix crack, but the strain energy was mainly released by composite delamination. For low impact energy such as 8 J below 23 J, materials damage sequence can be divided into three phases of delamination initiation, delamination propagation and no delamination propagation. For relative high impact energy such as 25 J, materials damage sequence can be divided, by the time t1, t2, t3, and t4, into five phases of delamination initiation, delamination propagation, no delamination propagation, scarfed adhesive damage, and no adhesive propagation. The phases were divided by the time t1, t2, t3, and t4. In phase I, linear process had no materials damage till the first force peak at the time of t1. In phase II, impact load dropped suddenly for the reason of that composite delamination initiating and propagating fleetly leaded the structure strain energy releasing and shear stiffness declining in short time t1-t2. In phase III, the impact load climbed again with the composite delamination propagation. When the impact load reached to the second force peak, for the impact energy above 23 J, the scarfed adhesive appearing damage leaded to the second impact load dropping.
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Fig. 14. The section microscope damage for specimen of tension after 25 J impact.
The critical points of t1, t2, t3 and t4 separated the impact procedure into different materials damage included that the composites delamination dominated the first force dropping, and that the
adhesive damage dominated the second force dropping. The delamination of initiation and propagation got by simulations was presented from two views: top view, xoy plane; and side view,
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Fig. 15. The schematic of crack path for specimen of tension after 25 J impact.
Table 1 Mechanical properties of T700/LT03A and FM73M. Material
E11/GPa
E22 = E33/GPa
m
G12 = G23 = G13/GPa
q
T700/LT03A FM73M
128 2.27
8.46
0.322 0.35
3.89 0.84
1560 kg/m3 1200 kg/m3
Table 2 Damage related properties of T700/LT03A and FM73 (Unit: MPa). Material
XT
XC
YT
YC
ZT
S
sf
T700/LT03A FM73M
2372
1234
50
178
50 7.3
107
80.7 46.1
Table 3 Fracture related properties of T700/LT03A and FM73 (Unit: mJ/mm2). Material T700/LT03A FM73M
G1c 0.9
G2c 1.8
G3c
Fig. 17. Impact load and time relationship of simulation and tests with 8 J.
Gft
Gfc
Gmt
Gmt
91.6
79.9
0.22
1.1
1.8
xoz plane (enlargement). Figs. 19–21 are the simulation results of 8 J. Figs. 22–26 illustrate the simulation results of 25 J. At time of t1 and t2, the model possesses the same size of composites delamination. From view of xoy plane, it could be found that delamination initiated at the impact location and then propagated around. From view of xoz plane, detail delamination location along the thickness
direction z and length direction x were implied. The largest of delamination occurred between 10th layer 0° and 11th layer 90°, and was labeled in red slash. The final impact delamination size for 8 J and 25 J were 34 mm(x direction) 30 mm(y direction) and 64 mm 50 mm by c-scan respectively. The delamination of simulation coinciding with the c-scan results explains that the damages detected by c-scan are the main delamination. The difference of 25 J from relative low energy 8 J was that delamination continued propagating in the process of t2-t3 till t4. For 25 J case, scarfed adhesive damage appeared at the time t4. Fig. 26 shows
Fig. 16. The impact model and FEM mesh illustration.
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that the scarfed adhesive damage occurred at the scarf feathered tip of back side. On the contrary, the adhesive under the impact point wasn’t damaged. It reveals that the most easily broken location is the feathered tip at the back of impact point. 5. Conclusion The impact damage mechanisms, responses and load-carrying capacity of TAI for the composite scarf repair were studied in this paper.
Fig. 18. Impact load and time relationship of simulation and tests with 25 J.
(1). In the impact procedure, it has double force peaks phenomenon and four typical phases. The first force peak keeps constant in 5000 N–6000 N, and the second force peak rises from 4000 N to 12000 N as the impact energy increases.
Fig. 19. Delamination initiation with 8 J at the time of t1: (a) from xoy plane, and (b) from xoz plane (enlargement).
Fig. 20. Delamination propagation with 8 J at the time of t2: (a) from xoy plane, and (b) from xoz plane (enlargement).
Fig. 21. The final delamination with 8 J at the time of t3: (a) from xoy plane, (b) c-scan, and (c) from xoz plane (enlargement).
Fig. 22. Delamination initiation with 25 J at the time of t1: (a) from xoy plane, and (b) from xoz plane (enlargement).
Fig. 23. Delamination propagation with 25 J at the time of t2: (a) from xoy plane, and (b) from xoz plane (enlargement).
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Fig. 24. Delamination propagation with 25 J at the time of t2-t3: (a) from xoy plane, and (b) from xoz plane (enlargement).
Fig. 25. The final delamination with 25 J at the time of t3: (a) from xoy plane, (b) c-scan, and (c) from xoz plane (enlargement).
(3). When the impact energy is below the critical 23 J, composite delamination and matrix cracks can be found in the process of first force dropping. When the impact energy is higher than the critical 23 J, except for the composites damage, the adhesive damage can be observed at the second force dropping. (4). The scarfed adhesive damage occurs at the scarf feathered tip of back side. It reveals that the most easily broken location is the feathered tip on the back of impact point. The adhesive damage has strong influence over tension after impact capability.
Acknowledgements
Fig. 26. The final scarfed adhesive damage with 25 J at the time of t4 from xoy plane.
Phase Ⅰ, linear elastic phase: the impact load varied from 0 to Fmax1. In this phase, the repair structure suffered the elastic deformation. When the impact load reached to Fmax1, large amount of delamination occurred in this short time. Phase II, impact load dropping: the impact load dropped suddenly after Fmax1. The damage of repair structure could occur. In this short time, the partial strain energy of structure was released by the damage. In the force dropping procedure, the materials, which were being damaged, were not elastic status. But other part of materials still suffered the elastic deformation. Phase III, impact load rising again: the graph of load-deflection vibrated and tended to be stable in this phase till the impact load reached to Fmax2.Phase Ⅳ, the impactor rebounding: the strain energy was released completely and translated to impactor kinetic energy after Fmax2. (2). The critical impact energy exists for the composite scarf repair, which would be 23 J for this size of specimen. For relative impact energy below 8 J, phase I was no damage linear phase; phase II of force dropping was the composites damage phase. When impact rises greater than 8 J, phase III started to make the composites damage arising.
The authors give sincere thanks to master Yanan Li, Zhuo Sun and Wei Feng in the School of Aeronautics, Northwestern Polytechnical University for the helpful experimental works. The research work was supported by ‘‘the Fundamental Research Funds for the Central Universities”, No. 3102017zy046. References [1] Liu B, Xu F, Feng W, Yan R, Xie W. Experiment and design methods of composite scarf repair for primary-load bearing structures. Compos A 2016;88:27–38. [2] Liang K, Sun Q. Buckling and post-buckling analysis of the delaminated composite plates using the Koiter-Newton method. Compos Struct 2017;168:266–76. [3] Kim MK, Elder DJ, Wang CH, Feih S. Interaction of laminate damage and adhesive disbonding in composite scarf joints subjected to combined in-plane loading and impact. Compos Struct 2012;94:945–53. [4] Goh JY, Georgiadis S, Orifici AC, Wang CH. Effects of bondline flaws on the damage tolerance of composite scarf joints. Compos A 2013;55:110–9. [5] Harman AB, Rider AN. Impact damage tolerance of composite repairs to highlyloaded, high temperature composite structures. Compos A 2011;42:1321–34. [6] Vaidya UK, Gautam ARS, Hosur M, Dutta P. Experimental-numerical studies of transverse impact response of adhesively bonded lap joints in composite structures. Int J Adhes Adhes 2006;26:184–98. [7] Takahashi I, Ito Y, Takeda S, Iwahori Y, Takatubo J, Takeda N. Impact damage detection on scarf-repaired composites using lamb wave sensing. In 16th international conference on composites materials, Japan; 2007. [8] Harman AB, Wang CH. Damage tolerance and impact resistance of composite scarf joints. In: 16th international conference on composites materials, Japan; 2007. [9] Kim MK, Elder DJ, Wang CH, Feih S. Interaction of laminate damage and adhesive disbonding in composite scarf joints subjected to combined in-plane loading and impact. Compos Struct 2012;94:945–53.
B. Liu et al. / Composite Structures 181 (2017) 183–193 [10] Gunaydın Berrin, Daghan Behcet, Avci Ahmet. Fatigue behavior of surfacenotched composite pipes repaired by composite patches. Int J Damage Mech 2012;22(4):490–8. [11] Wang CH, Venugopal V, Peng L. Stepped flush repairs for primary composite structures. J Adhes 2015;91:95–112. [12] Wang CH, Venugopal V, Peng L. Stepped flush repairs for primary composite structures. J Adhes 2015;91:95–112. [13] Andrew JJ, Arumugam V, Saravanakumar K, Dhakal HN, Santulli C. Compression after impact strength of repaired GFRP composite laminates under repeated impact loading. Compos Struct 2015;133:911–20. [14] Kara Memduh, Uyaner Mesut, Avci Ahmet. Repairing impact damaged fiber reinforced composite pipes by external wrapping with composite patches. Compos Struct 2015;123:1–8. [15] Balaganesan G, Khan VC. Energy absorption of repaired composite laminates subjected to impact loading. Compos B 2016;98:39–48.
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[16] Balcı O, Çoban O, Borab MÖ, Akagündüz E, Yalçin EB. Experimental investigation of single and repeated impacts for repaired honeycomb sandwich structures. Mater Sci Eng, A 2017;682:23–30. [17] Coelho SRM, Reis PNB, Ferreira JAM, Pereira AM. Effects of external patch configuration on repaired composite laminates subjected to multi-impacts. Compos Struct 2017;168:259–65. [18] Fan XL, Sun Q, Wang TJ. Damage evolution of sandwich composite structure using a progressive failure analysis methodology. Procedia Eng 2011;10:530–5. [19] Liu B, Xu Fei, Yan Rui, Ji Zhe, Li Wenying. Parameters sensitivity and optimization for composite scarf repair. J Reinf Plast Compos 2014;33 (23):2164–73. [20] Fan XL, Zhang WX, Wang TJ, Liu GW, Zhang JH. Investigation on periodic cracking of elastic film/substrate system by the extended finite element method. Appl Surf Sci 2011;257(15):6718–24.
Contents lists available at ScienceDirect
Composite Structures journal homepage: www.elsevier.com/locate/compstruct
Study on impact damage mechanisms and TAI capacity for the composite scarf repair of the primary load-bearing level Bin Liu a,⇑, Fei Xu a, Jian Qin b, Zhixian Lu a a b
School of Aeronautics, Northwestern Polytechnical University, Youyi West Road, No. 127, Xi’an, China Eastern Airlines Technic Co., Ltd Northwest Branch, Xi’an, China
a r t i c l e
i n f o
Article history: Received 18 April 2017 Revised 29 July 2017 Accepted 29 August 2017 Available online 31 August 2017 Keywords: Impact damage Microcrack Composite repair Bonding Damage tolerance
a b s t r a c t As composite material plays a leading role in aircraft, composite bonding repair has been extensively applied. Among composite bonding repairs, the scarf bonding repair is widely adopted and has high repair efficiency especially in primary load-bearing structures. However, the impact damage tolerance and impact damage mechanisms were not considered for repaired structure integrity design yet. This paper experimentally and numerically studied the scarfed bonding repair of the advanced CFRP, which may suffer a low velocity impact load in service. At the central location of adhesive zone, impact energy and response regularity were studied to reveal the competition failure mechanism for inner kinds of materials. In the impact procedure, double force peaks phenomenon and four typical phases were found. Tension after impact (TAI) capacities were also tested to explain the impact damage effects on residual strength. The adhesive damage has strong influence over tension after impact capability. The most easily broken location in the bonded zone is the feathered tip on the back of impact point. The critical impact energy 23 J exists for this size of specimen. When the impact energy is higher than the critical 23 J, except for the composites damage, the adhesive damage can be observed at the second force dropping. The scarfed adhesive damage occurred at the scarf feathered tip of back side. Ó 2017 Elsevier Ltd. All rights reserved.
1. Introduction Advanced CFRP (Carbon Fiber Reinforced Plastics) application in aircraft has turned from secondary structures to primary load bearing structures. High strength and middle modulus CFRP promote composites application to load bearing structures, such as T700, T800, T1100 and MJ series. Damage of aircraft load bearing level structures always occurred in impacting and maintaining. As composites have low interlaminar toughness, low velocity impact leads to an invisible inner damage. For the upgraded composites, the mechanical characteristics would change. Hence, the problems of composite repair exist in load bearing structures and are complicated and important [1,2]. At present, composite repair technique has become one of key factors for further composite applications. The existing aircraft structures design considered only the static strength, stiffness recovering level, but didn’t include the second impact in the service of repaired structures. When stepped and scarf repaired structures are impacted, the innerlaminar, interlaminar and adhesive will suffer damage in some extent. The competed failure mecha⇑ Corresponding author. E-mail address: [email protected] (B. Liu). http://dx.doi.org/10.1016/j.compstruct.2017.08.087 0263-8223/Ó 2017 Elsevier Ltd. All rights reserved.
nisms of different materials must be revealed clearly to obtain the impact tolerance and resistance for composite repairs. Low velocity impact has strong influence to composite stepped and scarf repairs [3–5]. In 2006, U.K. Vaidya [6] found that outplane load leads to higher peel stress and concentration than inplane load case, and adhesive crack initiates as mixture mode and transforms to mode II. In 2007, I. Takahashi [7] applied health monitoring technique to detect scarf adhesive damage. A.B. Harman [8] discovered that the impact damage tolerance for composite scarf repair structure reduced comparing with laminate plates. H.C.H. Li adopted quasi-static out-plane load to substitute impact dynamic load and found that in-plane prestrain affects bending stiffness. In 2012, M.K. Kim [9] studied the impact damage of laminate plates and composite scarf repair under combined in-plane load and out-plane impact, and gave the conclusion of that impact resistance increase obviously as the prestrain increases. Berrin Gunaydın [10] investigated the effects of composite repair patches and number of patch layers on the fatigue behavior of surfacenotched composite pipes. In 2015, C.H. Wang [11] experimentally studied CAI (Compression after Impact) mechanical performance for 2 mm thickness stepped repair. C.H. Wang [12] proposed that the existed design methodologies consider only loading capacity of integrated repair structures without including adhesive damage
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and delamination. J.J. Andrew [13] conducted experiments of multiple low velocity impact and CAI for GFRP bonding lap joints. Effects of number of patch layers on the burst pressure of low velocity impact damaged tubes that have been repaired with composite patches were investigated [14]. In 2016, G. Balaganesan [15] found that the patch of GFRP stepped repair structures could absorb 50% to 80% impact energy. In 2017, O. Balcı [16] studied the honey comb GFRP structures repaired by lap joint and concluded 3 J impact energy produce surface damage and 8 J is the critical penetrated energy. S.R.M. Coelho [17] concerned the mechanical performance and the damage for single patch and double patches with multiple impacts. In summary, impact problems of composite scarfed and stepped repair were studied by few researchers, but impact response, damage mechanisms and damage modes were not researched systematically. The conjunction studies of macroscopic damage phenomenon and microscopic failure cracks [18–20] were conducted much less and needed to be carried out. 2. Specimens and tests setup 2.1. Specimens As shown in Fig. 1, the specimen length, width and thickness are 300 mm, 50 mm and 4 mm [5]. Scarfed surfaces with the angle of 5° were cut and polished by machines. The scarf bond length is 45.7 mm and the bond surface thickness is 0.2 mm. The supplements of both ends are clamped area for tensile test machine
and the size is 50 50 mm. To avoid the stress concentration, the clamped supplements were machined with the angle of 15°. The composite scarf adherends are hard patch repair which were pre-cured. The composite adherends were fabricated using the layup consisting of 32 plies of high performance T700/LT03A carbon/epoxy prepregs. The stack sequence of composites is [45/ 0/-45/90/0/45/0/-45/0/0/90/0/-45/0/45/0]S. The percentage of 0°, 45° and 90° is [50/37.5/12.5]%. This kind of layup being composed of 0° of 50% is a representative for primary-load bearing composite structures [1]. The prepreg material has a nominal ply with thickness of 0.125 mm and curing temperature of 120 °C. The adhesive used to joint two composite adherends which were made from epoxy of Cytec FM73M with thickness of approximately 0.2 mm and curing temperature of 120 °C. FM73M is an aerospace high-performance film adhesive. 2.2. Tests setup On account of that the existing impact standards for composites didn’t include relative narrow specimen, we designed the narrow impact setup for composite scarf repair by referring to A.B. Harman and A.N. Rider [5]. Fig. 2(a) shows the schematic of impact fixture and specimen. Fig. 2(b) shows the experimental setup coinciding to Fig. 2(a). The drop-testing machine Instron-9250HV was used and the tip diameter was 16 mm. The quality of the punch is 5.067 kg. During impact process, the responses of impact load, deflection, velocity, absorbed energy and time can be obtained by the testing machine. To investigate the effect of impact energy on composite
Fig. 1. Specimen schematic of the composite scarf repair.
Fig. 2. The impact fixture and specimen: (a) schematic; (b) setup.
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scarf repair, 4 J, 6 J, 8 J, 12 J, 16 J, 20 J, 23 J, 25 J and 30 J were chosen. The impact initiation velocities were 1.194 m/s, 1.462 m/s, 1.689 m/s, 2.069 m/s, 2.389 m/s, 2.671 m/s, 2.986 m/s, and 3.271 m/s for the nine impact energy levels respectively. Tension and tension after impact tests were operated on the static testing machine with the displacement control of 0.5 mm/min by ASTMD3039/D3039M-008. 3. Experimental results and discussion 3.1. Impact response On the basis of the experimental results with different impact energy, the graphs of impact load vs. time show double peaks. We gave the impact load and deflection with time for different impact energy levels of 8 J and 25 J respectively as
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shown in Fig. 3(a) and Fig. 3(b). From load-time graphs, we provided the schematic of load-time in Fig. 4(a). The first peak of force denoted by Fmax1 keeps constant in 5000 N–6000 N, and the second peak of force denoted by Fmax2 rises from 4000 N to 12000 N as the impact energy increases as shown in Fig. 4(b). The regularity research of impact load and deflection were conducted as follows. The typical 4 phases can be summarized from the studies. Firstly, the example of 8 J impact energy was provided as shown in Fig. 5(a). Phase I, linear elastic phase: the impact load varied from 0 to Fmax1. In this phase, the repair structure suffered an elastic deformation. Phase II, impact load dropping: the impact load dropped suddenly after Fmax1. The damage of repair structure could occur. In this short time, the partial strain energy of structure was released
Fig. 3. Impact load and deflection with time of (a) 8 J and (b) 25 J.
Fig. 4. (a) Schematic of impact load with time and (b) variation of load peak Fmax1 and Fmax2 with impact energy.
Fig. 5. Impact load-deflection curves (a) 8 J (b) 25 J.
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by the damage. Thus the impact load dropped suddenly and the stiffness of structure also decreased. Phase III, impact load rising again: the graph of load-deflection vibrated and tended to being stable in this phase till the impact load reached to Fmax2. Phase IV, the impactor rebounding: the strain energy was released completely and translated to impactor kinetic energy. After Fmax2, the impact load decreased non-linearly to zero. However, as the impact energy rose to being greater than 23 J, the second peak of impact load Fmax2 dropped suddenly. Here we
Fig. 6. The front impact dent and the back adhesive crack with 25 J impact energy.
gave the relative distinct graph of load-deflection with 25 J impact energy as shown in Fig. 5(b). 3.2. Impact damage In order to reveal the damage mechanism of impact and tension after impact, we studied three kinds of damage which respectively were visual visible damage, microscope visible damage and ultrasonic inspecting damage. Firstly, the visual visible damage showed that when impact energy rose greater than 23 J, the dent depth of front surface became clearly and the scarfed tip of back surface appeared obvious adhesive fracture. The front surface and back surface of 25 J impacted specimen are illustrated in Fig. 6. In order to get the inner damage details, the specimens were cut along the longitudinal central line. To avoid producing new damage for impacted specimens, we chose diamond cutting blade with rotate speed of 1500 r/min and feed speed of 300 mm/min. From the central cut section, by using microscope, the damage details of 25 J were indicated in Fig. 7. The damage modes of 25 J included adhesive damage, matrix crack of 90° plies, and the interlaminar delamination 0°–90° and 0°–±45°. Among all the damage modes, adhesive damage was most obvious and located from the back scarfed tip to the tenth ply counting from the back. The adhesive damage amount was totally 31% of the scarf bonding length. The microscope visible damage of adhesive under 20 J and 23 J was not found in the cut section. As the impact energy increases to 30 J, not only the adhesive had more obvious damage (37.5%) than 25 J (31%), but also the
Fig. 7. The microscopic damage image captured from central cut section for the specimen with 25 J impact energy.
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Fig. 8. The microscopic damage image captured from central cut section for the specimen with 30 J impact energy.
fiber on back side existed fracture. Meanwhile, a plastic deformation appeared in composite materials and adhesive under the impact point, as shown in Fig. 8. Summarizing Fig. 7 and Fig. 8, Fig. 9 and Fig. 10 showed the visualized schematics of the central cut section damage found by microscope for 25 J and 30 J impact energy respectively. In the schematics, the white color denotes 0° plies, gray color ±45°, black color 90°, and red lines illustrate the crack path. Fig. 11 shows the ultrasonic C-scan images of impacted damage for different impact energy. The different colors denote the damage
Fig. 11. The ultrasonic C-scan images of impacted damage for different impact energy.
depth and the blue color denotes no damage exists. The length and width can be also red from Fig. 11. Due to the shortage of C-scan, we couldn’t get the inner adhesive damage information. However, the damage size gotten by C-scan revealed the composite materials damage variation with impact energy. Damage length and width
Fig. 9. The schematic of microscopic visible damage from central section with 25 J impact energy.
Fig. 10. The schematic of microscopic visible damage from central section with 30 J impact energy.
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The two parts of adherends of specimens impacted were separated from each other under tension load. The two parts firstly were combined together, then, by means of microscope, were observed to reveal the mechanisms for tension after impact. Here, the microscope images of 25 J impact energy was provided as illustrated in Fig. 14. The damage details in Fig. 14 can be shown as the schematic type in Fig. 15. The damage modes included compositeadhesive interface fracture, 90°–0° and ±45°–0° delamination, adhesive cohesive damage, 90° plies and 45° plies matrix crack. Compared with specimens impacted without tension, the delamination and matrix crack length and width of tension after impact became larger. It explains that tension load made the impacted damage severe and adhesive damage was the major reason for load capacity reducing.
Fig. 12. The variation of length and width of damage size with impact energy.
varied nonlinearly as the impact energy increased. We adopted the least square method and obtained the fitted quadratic polynomial to descript damage size variation with impact energy. The damage of the composite scarf repair propagates easily along length direction compared to width (Fig. 12). 3.3. Tension after impact After studying the impact damage mechanisms for composite scarf repair, we tried to combine impact damage mechanisms with load capacity of tension after impact. Below the threshold impact energy, there was no obvious decreasing of load capacity of tension after impact as shown in Fig. 13. When impact energy increased to 25 J, load capacity of tension after impact decreased obviously. As impact energy equaled to 30 J, load capacity of tension after impact continued to decline. Thus 25 J can be considered as the critical energy or threshold level of energy. The original average strength T0 of no impact (0 J) repairs is 69.9 kN. TE delegates TAI strength under impact energy E. So the TAI capability reduction can be calculated by equation (T0 TE)/ T0 100%. Compared with no impact specimens, the load capacity T25 of TAI with 25 J dropped 23% and 41% relatively for two specimens while T30 with 30 J dropped 29% and 40%. As what was mentioned before, adhesive damage percentage were 31% and 37.5% for 25 J and 30 J. We can clearly see that TAI load capacity decreasing has nearly the same amount as the adhesive damage percentage. Therefore, we could draw a conclusion that TAI load capacity is closely related to adhesive damage for composite scarf repair.
Fig. 13. Load capacity of TAI with discrete impact energies.
4. Simulation On account of that impact procedure time is very short, and C-scan and microscope tools cannot capture the sequence of different materials mechanisms, simulation is needed to recruit with experimental studies. The mechanical properties of the composite material and the adhesive, as provided by the manufactures, are presented in Tables 1–3. The whole FEM model mainly includes impactor, support, press head, and specimen as shown in Fig. 16. From Fig. 16, the FEM model of composite scarf repair includes composite innerlayer 3D elements, composite interlamina cohesive zone elements, and scarfed adhesive cohesive zone elements. The composite innerlayer elements are 8 nodes hexahedral reduced integration elements and 6 nodes pentahedral reduced integration elements. The composite interlamina elements and scarfed adhesive elements are 8 nodes hexahedral cohesive zone elements. Support, press, and impactor include 4 nodes tetrahedral, 6 nodes pentahedral and 8 nodes hexahedral reduced integration elements. The typical 4 phases of impact load vs. displacement was found from the experimental studies. In the 4 phases, the reason of load dropping of phase II and phase IV and the competed damage sequence of materials cannot be obtained in experiments. We tried to explain and reveal them by FEM simulation of two kinds of impact energy 8 J and 25 J. As shown in Fig. 17 and Fig. 18, impact load and time relationship of simulation coincides well with the experimental results. The simulation indicated that the first impact load dropping of 8 J and 25 J was caused by composite delamination and the second impact load dropping of 25 J was caused by scarfed adhesive damage. In the whole impact procedure, composite material damage certainly contained matrix crack, but the strain energy was mainly released by composite delamination. For low impact energy such as 8 J below 23 J, materials damage sequence can be divided into three phases of delamination initiation, delamination propagation and no delamination propagation. For relative high impact energy such as 25 J, materials damage sequence can be divided, by the time t1, t2, t3, and t4, into five phases of delamination initiation, delamination propagation, no delamination propagation, scarfed adhesive damage, and no adhesive propagation. The phases were divided by the time t1, t2, t3, and t4. In phase I, linear process had no materials damage till the first force peak at the time of t1. In phase II, impact load dropped suddenly for the reason of that composite delamination initiating and propagating fleetly leaded the structure strain energy releasing and shear stiffness declining in short time t1-t2. In phase III, the impact load climbed again with the composite delamination propagation. When the impact load reached to the second force peak, for the impact energy above 23 J, the scarfed adhesive appearing damage leaded to the second impact load dropping.
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Fig. 14. The section microscope damage for specimen of tension after 25 J impact.
The critical points of t1, t2, t3 and t4 separated the impact procedure into different materials damage included that the composites delamination dominated the first force dropping, and that the
adhesive damage dominated the second force dropping. The delamination of initiation and propagation got by simulations was presented from two views: top view, xoy plane; and side view,
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Fig. 15. The schematic of crack path for specimen of tension after 25 J impact.
Table 1 Mechanical properties of T700/LT03A and FM73M. Material
E11/GPa
E22 = E33/GPa
m
G12 = G23 = G13/GPa
q
T700/LT03A FM73M
128 2.27
8.46
0.322 0.35
3.89 0.84
1560 kg/m3 1200 kg/m3
Table 2 Damage related properties of T700/LT03A and FM73 (Unit: MPa). Material
XT
XC
YT
YC
ZT
S
sf
T700/LT03A FM73M
2372
1234
50
178
50 7.3
107
80.7 46.1
Table 3 Fracture related properties of T700/LT03A and FM73 (Unit: mJ/mm2). Material T700/LT03A FM73M
G1c 0.9
G2c 1.8
G3c
Fig. 17. Impact load and time relationship of simulation and tests with 8 J.
Gft
Gfc
Gmt
Gmt
91.6
79.9
0.22
1.1
1.8
xoz plane (enlargement). Figs. 19–21 are the simulation results of 8 J. Figs. 22–26 illustrate the simulation results of 25 J. At time of t1 and t2, the model possesses the same size of composites delamination. From view of xoy plane, it could be found that delamination initiated at the impact location and then propagated around. From view of xoz plane, detail delamination location along the thickness
direction z and length direction x were implied. The largest of delamination occurred between 10th layer 0° and 11th layer 90°, and was labeled in red slash. The final impact delamination size for 8 J and 25 J were 34 mm(x direction) 30 mm(y direction) and 64 mm 50 mm by c-scan respectively. The delamination of simulation coinciding with the c-scan results explains that the damages detected by c-scan are the main delamination. The difference of 25 J from relative low energy 8 J was that delamination continued propagating in the process of t2-t3 till t4. For 25 J case, scarfed adhesive damage appeared at the time t4. Fig. 26 shows
Fig. 16. The impact model and FEM mesh illustration.
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that the scarfed adhesive damage occurred at the scarf feathered tip of back side. On the contrary, the adhesive under the impact point wasn’t damaged. It reveals that the most easily broken location is the feathered tip at the back of impact point. 5. Conclusion The impact damage mechanisms, responses and load-carrying capacity of TAI for the composite scarf repair were studied in this paper.
Fig. 18. Impact load and time relationship of simulation and tests with 25 J.
(1). In the impact procedure, it has double force peaks phenomenon and four typical phases. The first force peak keeps constant in 5000 N–6000 N, and the second force peak rises from 4000 N to 12000 N as the impact energy increases.
Fig. 19. Delamination initiation with 8 J at the time of t1: (a) from xoy plane, and (b) from xoz plane (enlargement).
Fig. 20. Delamination propagation with 8 J at the time of t2: (a) from xoy plane, and (b) from xoz plane (enlargement).
Fig. 21. The final delamination with 8 J at the time of t3: (a) from xoy plane, (b) c-scan, and (c) from xoz plane (enlargement).
Fig. 22. Delamination initiation with 25 J at the time of t1: (a) from xoy plane, and (b) from xoz plane (enlargement).
Fig. 23. Delamination propagation with 25 J at the time of t2: (a) from xoy plane, and (b) from xoz plane (enlargement).
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Fig. 24. Delamination propagation with 25 J at the time of t2-t3: (a) from xoy plane, and (b) from xoz plane (enlargement).
Fig. 25. The final delamination with 25 J at the time of t3: (a) from xoy plane, (b) c-scan, and (c) from xoz plane (enlargement).
(3). When the impact energy is below the critical 23 J, composite delamination and matrix cracks can be found in the process of first force dropping. When the impact energy is higher than the critical 23 J, except for the composites damage, the adhesive damage can be observed at the second force dropping. (4). The scarfed adhesive damage occurs at the scarf feathered tip of back side. It reveals that the most easily broken location is the feathered tip on the back of impact point. The adhesive damage has strong influence over tension after impact capability.
Acknowledgements
Fig. 26. The final scarfed adhesive damage with 25 J at the time of t4 from xoy plane.
Phase Ⅰ, linear elastic phase: the impact load varied from 0 to Fmax1. In this phase, the repair structure suffered the elastic deformation. When the impact load reached to Fmax1, large amount of delamination occurred in this short time. Phase II, impact load dropping: the impact load dropped suddenly after Fmax1. The damage of repair structure could occur. In this short time, the partial strain energy of structure was released by the damage. In the force dropping procedure, the materials, which were being damaged, were not elastic status. But other part of materials still suffered the elastic deformation. Phase III, impact load rising again: the graph of load-deflection vibrated and tended to be stable in this phase till the impact load reached to Fmax2.Phase Ⅳ, the impactor rebounding: the strain energy was released completely and translated to impactor kinetic energy after Fmax2. (2). The critical impact energy exists for the composite scarf repair, which would be 23 J for this size of specimen. For relative impact energy below 8 J, phase I was no damage linear phase; phase II of force dropping was the composites damage phase. When impact rises greater than 8 J, phase III started to make the composites damage arising.
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