WRINKLE AND DELAMINATION TYPE DEFECTS GRADUATION PROJECT. Muhammet Emre ÖZARPACI. Department of Aerospace Engineering. Programı : Herhangi Program

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1 ISTANBUL TECHNICAL UNIVERSITY FACULTY OF AERONAUTICS AND ASTRONAUTICS WRINKLE AND DELAMINATION TYPE DEFECTS GRADUATION PROJECT Muhammet Emre ÖZARPACI Department of Aerospace Engineering Thesis Anabilim Advisor: Dalı Prof. : Dr. Herhangi Vedat Ziya Mühendislik, DOĞAN Bilim Programı : Herhangi Program AUGUST, 018 i

2 ISTANBUL TECHNICAL UNIVERSITY FACULTY OF AERONAUTICS AND ASTRONAUTICS WRINKLE AND DELAMINATION TYPE DEFECTS GRADUATION PROJECT Muhammet Emre ÖZARPACI Department of Aerospace Engineering Thesis Advisor: Prof. Dr. Vedat Ziya DOĞAN Anabilim Dalı : Herhangi Mühendislik, Bilim Programı : Herhangi Program AUGUST,018 ii

3 Muhammet Emre ÖZARPACI,student of ITU Faculty of Aeronautics and Astronauticsstudent ID , successfully defended the graduation entitled WRINKLE AND DELAMINATION TYPE DEFECTS, which he prepared after fulfilling the requirements specified in the associated legislations, before the ury whose signatures are below. Thesis Advisor : Prof. Dr. Vedat Ziya DOĞAN... İstanbul Technical University Jury Members : Prof.Dr. Zahit MECİTOĞLU... İstanbul Technical University Dr.Öğr.Üye. Özge ÖZDEMİR... İstanbul Technical University Date of Submission : 03 August 018 Date of Defense : 06 August 018 iii

4 To my family, iv

5 FOREWORD I would like to thank to my advisor Prof. Dr. Vedat Ziya DOĞAN for his helpful lead,patience and time.also I want to thank to my dear mother Suna ÖZARPACI for her endless support and encouragements. August 018 Muhammet Emre ÖZARPACI v

6 vi

7 Content SUMMARY... x History Wrinkle Problem description and formulation Layer-by-layer approach Single layer large deflection, SLLD approach Delamination in Composites Causes of Delamination Manufacturing Environmental effects Drilling Geometrical configuration Low velocity impact Delamination Modeling Fracture mechanics based models Damage mechanics based models Standardization of Delamination Testing Historical development Mode I delamination testing Mode II delamination testing... 3 REFERENCES vii

8 ABBREVIATIONS BC CTP LbL SLLD M E G FRP DCB ASTM ESIS JIS ISO ENF SENF 4ENF ELS : Before Christian : Glass fiber reinforced polyester : Layer-by-Layer : Single-Layer Large-Deflection : Moment : Elastic Modulus : Shear Modulus : Fibre reinforced polymers : Double Cantilever Beam : American Society for Testing and Materials : European Structural Integrity Society : Japanese Industrial Standards : International Standards Organization : End Notch Flexure : Stabilized End Notch Flexure : Four Point Bend End Notch Flexure : End Loaded Split viii

9 List of Figures Page Figure 1.1 Formation of wrinkles as a result of deformation rate Figure 1.. Schematic of layers and interfaces for an n-layer laminate under moment loading Figure 1.3. An n-layer composite under moment loading where bending stresses have been transformed into the inplane compressive force Figure 1.4. Deformed geometry of the interface with adacent plies used in calculating interfacial strains Figure 1.5. Free body diagram of th layer used in calculating interactions from adacent layers Figure 1.6. th layer in the form of a beam Figure 1.7. Free body diagram of th layer used for deriving governing equations Figure.1 Delamination caused by drilling... 3 Figure.. Three fracture modes... 5 Figure.3. Crack in a homogenous isotropic linear elastic infinite plate... 6 Figure.4. Traction-displacement curve for the tip of the delamination... 9 Figure.5. DCB specimen... 3 ix

10 THESIS TITLE IN ENGLISH HERE SUMMARY x

11 History Since the early ages, human beings have tried to remove fragility by putting vegetable or animal fibers into fragile materials. The best example of this is the mudbrick. In the production of mudbrick, stalks and fibers such as straw, ivy branches which are introduced into clayey mud increase the strength of the mudbrick. Another example is the arrow springs made by stacking features and fiber directions on top of different wood plates. Moreover, the production of thin glass fibers in Egypt in 1600 BC., XVIII. from the age of the dynasty, is understood from the various items of glass fibers of different colors. For this reason, composite materials can be said to be technically used from past to present. There are various uses ranging from mudbrick to asphalt and concrete. There are some patents taken in the early 19th century on the method of manufacturing artificial stone plates using hydraulic binders and fiber materials. The first record concerning the use of glass fibers in the industry is dated Synthetic resins reinforced with fibers have started to be used in the industry since 1950's. The most well-known group of this material is "glass fiber reinforced polyester (CTP)". These materials, known as "fiber-glass" in our country, have been used since 1960 for the manufacture of components such as liquid tanks, roofing plates, small-sized marine vessels. "Anadol", the first domestic car produced in our country in series, is also the body of this material. Composite material can be simply described as "new materials created by combining multiple materials in such a way that they do not dissolve into one another." The aim here is to develop some properties (lightness, strength, flexibility, etc.) that are not present when the components are stand-alone and to bring them together. The composites are generally formed from a more basic material given a matrix name and a reinforcing element fiber name given as a base material. From these two material groups, reinforcing material plays a role of preventing the strength and load bearing property of composite material and the matrix material plays a role in crack propagation which may occur in the transition to plastic deformation, and delays the breaking of composite material. Along with improvements in composite material technology, composites have begun to be used increasingly in industry and technology applications. It is used in a wide variety of areas thanks to its composite materials properties and properties. Since each industry has different 11

12 needs and expectations, product flexibility of composite materials is a maor advantage. Scientific studies on composites, which are used in aerospace industry and in many other sectors, are continuing intensively today. Composite materials with their characteristic features and many advantages over other materials are preferred due to their long life, lightness, high chemical and mechanical strength. Despite all of these, composite materials can undergo deformation, such as other materials, exposed to external influences, collisions and stresses under pressure during their service life. These effects can lead to the formation of layers of composite materials in layered structures, the formation of wrinkles and delamination zones. The presence of such damages reduces the strength and rigidity of the material by making an important preventive effect on the load carrying capacity. Due to these effects, the material may come into contact with unexpected structural deterioration. Delaminations that are obviously unfavorable in composite materials may not only be caused by the effects of the exposure, but may also result from manufacturing mistakes. 1.Wrinkle Geometry, stacking sequence, material stiffness and especially in-plane loading can cause wrinkles to form. During the processing of continuous-fiber polymeric composites have observed two species of wrinkle formation. Firstly, pre-wrinkled state of the wrinkling layer have flat structure and wrinkle length is too limited. For this reason, with the global bending of laminate, every layer is thought to be under in-plane load in the laminate. It was thought that the upper layer could be wrinkled under various loads. The so-called critical load is the minimum amount of this load, which is directly related to the length of the wrinkle and this load. If this critical value is exceeded, wrinkling will occur. Secondly, the reason for the wrinkling is that a layer subected to maximum buckling causes it to wrinkle.the reason for the wrinkling is that a layer subected to maximum buckling is caused to wrinkle. The unwrinkled part of the wrinkling layer being supported by the remaining layers. For predicting the formation of wrinkles is benefited from large-deflection theory and a variation in the potential energy of the wrinkle layer as a function of its perturbed states. 1

13 Rosen says compressive strength of a composite laminate can be estimated using the fiber micro-buckling approach. It is thought here that the laminate is very thick, while at the same time the applied in-plane strain is equal on the cross-section. It has been determined that the buckling load obtained under these conditions is much higher than experimental observations. Sun and Jun presented a model to explain the cause of this situation (experimental and analytic pressure differential). As Morris and Sun stated, it has been determined that the wrinkles that occur during the processing of composite materials have appeared on certain production conditions. During processing, the resin-rich interfacial layers become weaker than the other layers. Resin-rich plates differ from others in the stifnesses. This causes the plies to slide over each other at high temperatures. As a function depending on time, temperature and process ratio, stress occurs during deformation in the plies. An example of sufficient ply stress in figure 1.1 is that it causes wrinkling. Some or all of the stress generated during deformation by interlaminar slip has been discharged. A Layer-by-Layer, LbL Approach (1), and a Single- Layer Large-Deflection, SLLD Approach () are two types of wrinkles that are formed during the formation of composite laminates. With the deformation model created by Pandey and Sun, enough stress can be calculated to analyze stability/wrinkling in any time. Figure 1.1 Formation of wrinkles as a result of deformation rate Firstly, throughout the places where wrinkles are present, the laminate is assumed to be a flat structure.pandey and Sun explain this situation in the following way, "The bending moment in the laminate developed as a result of the deformation is discretized as in-plane loads in the individual layers. A set of governing equations is derived to describe the deformation of individual layers under the in-plane loads. The interactions between the adacent layers occur 13

14 through the interfaces. A global system matrix is formed from the derived governing equations. The relationship between the applied load and the wrinkle length is obtained from the eigenvalues of the above- mentioned system matrix. The minimum of the load and corresponding wrinkle length are called, respectively, the critical load and the wrinkle length." Secondly, the first layer in the direction of the compression portion of the laminate causes a wrinkle. The full or partial length of the inner layer (as mentioned above) is supported by the remaining layers. It is seen that there is no connection between the wrinkled part and the other strata at the beginning of the wrinkling. The wrinkled portion are supported by the end points of the laminate that contact the portion. The total potential energy of the laminate is found for any curvature of laminate and plane load. Utilizing numerous perturbation displacements, the potential energy change and the beginning of wrinkle formation are estimated for the laminate deformation state. The potential energy states of the different transverse deformed configurations are found for the wrinkle layer. The externally applied perturbation load and the two curved parts of the wrinkle layer are considered to have the same potential energy when trying to learn the wrinkle start. 1.. Problem description and formulation There are two different types of wrinkle formation analysis, LbL and SLLD. İnitially, the approach that transforms interfaces into arc sets is LbL. Interfaces make it possible for all layers to deform simultaneously with their own degrees of freedom. Secondly, SSLD considers the main cause of wrinkle formation as only the innermost layer. The wrinkle layer is supported by the other layers. Between the wrinkled layer and the supporting layers over the length of the wrinkle don't have material continuity at the beginning of wrinkling Layer-by-layer approach When considering the length of the wrinkle length, a small portion of the laminate is taken. This small part is corrected before the wrinkle takes place. Inter laminar thin resin-rich layer of thickness and plies of thickness are named respectively ' h and h as in figure

15 Figure 1.. Schematic of layers and interfaces for an n-layer laminate under moment loading Global moment M acts laminae along the cross-sectional area during deformation.instead of moment (M), P is divided into load parts (in-plane load) and applied to th layer as in figure 1.3. Besides the total bending moment, it is assumed that the "bending moment of the individual layers" is negligible. A subscript and a superscript `prime' define the interface between the th and (+1)th layer.the elastic modulus and shear modulus are defined as E and G, respectively. Figure 1.3. An n-layer composite under moment loading where bending stresses have been transformed into the inplane compressive force. The relative movements of adacent layers are limited along the interfaces. This interaction between the two can be explained as the kinematic of the deformation. If the layers are examined separately, the free body diagram (on the upper and lower surface) will contain 15

16 normal stress and additional shear strain. The th interface and its adacent layers in a deformed configuration are showed in figure 1.4. Figure 1.4. Deformed geometry of the interface with adacent plies used in calculating interfacial strains. The upper and lower face of the interface (with the benefit of identification) are subect to the shear strains shown below: ' t h w w w ' h x x x (1.1) ' b h w w w h x x x 1 ' (1.) 16

17 The average of the upper and lower shear strain gives the mean shear strain of the interfaces. Interfaces remain thin when compared to plies ' ( h h), the average strain in the th interface shown below: ' h h w w 1 1 ' 4h x x (1.3) The normal strain moving at the th interface (with transverse displacement of the layers) shown below: '( ) yy w w 1 (1.4) ' h The stresses in the th interface layer in term of the kinematic variable can be write with eqs. (3) and (4): G ' ' h h w w 1 1 ' 4h x x (1.5) q ' e w w ' 1 E (1.6) ' h For the change of interface material, " K 'e and 's K " are equivalent spring constant. Then K ' e E (1.7) h ' ' K ' s G (1.8) h ' ' 17

18 Normal stresses on the upper and lower surfaces of the interface compensate the shear stress in Eq. (6) as shown in figure 1.5. The relationship between normal and shear movements can be found in the way in figure 1.5 as q ' s ' ' h x (1.6.1) Figure 1.5. Free body diagram of th layer used in calculating interactions from adacent layers. and, the relative axial and transverse motions of the layer cause the total normal traction as ' s ' K ' h ( h h 1) w w 1 ' e q K ( w 1 w ) 8 x x (1.9) Figure 1.6 and figure 1.7 are free body diagram for the th ply as shown, the resultant distributed moments and tractions are represented as h ' ' m 1 (1.10) 18

19 q ' s ' K 1h 1( h 1 h ) w 1 w 8 x x ' s ' K h ( h h 1) w w 1 8 x x K ( w w ) K ( w w ) ' e ' e (1.11) Figure 1.6. th layer in the form of a beam. The deformation equation of the th ply can be written like at figure 7 for governing: w w m E I P q x x x 4 4 (1.1) Figure 1.7. Free body diagram of th layer used for deriving governing equations. Deformation of plates with governing equation can be expressed as (all layers are considered to be of the same thickness and if resin-rich places are overlooked): 19

20 4 w1 w1 ' e E1I P 4 1 K ( w 1 w ) x x h ' s w1 w K 1 4 x x 4 w w ' e E I P ( 4 K w w 1) x x h w ' s w K 1 1 n 4 x x 4 wn wn ' e EnI P ( 4 n K w n wn 1) x x h ' s wn 1 w K n n 4 x x (1.13) (1.14) (1.15) Let's consider the transverse deformation of the th layer as follows: w x Wsin l (1.16) where, half of the wavelenth of buckled shape is "l". In Eqs. (13) and (15) the sinusoidal shape is modified, we get: h h E I P K W K K W ' e ' s ' e h h 4 4 ' s ' e 4 ' e K K W 1 E I P K W h ' s ' e K K W h h 4 4 ' s ' e 4 ' e K K Wn 1 E I Pn K Wn 0 (1.17) (1.18) (1.19) Define: 0

21 4 h ' s ' s d1 E1I P1 K K, 4 4 h ' s ' s d E I P K K, 4 h ' s ' s dn EnI Pn K K, 4 h k K K 4 l ' s ' e, Equations 17 and 19 are homogeneous (expressing the deformation under in-plane loading). For this reason, the determinant of the coefficients must be zero for a nontrivial solution: i.e., d1 k k d k k d3 k k dn k k dn 1 k k d n The determinant value given by the relational relation can be found. The determinant of the th order upper left sub-matrix think D. Then: D D d k D (1.0) n n1 n n where, D 1, D0 1. For the set of in-plane load values for which the determinant can be zero, the nontrival solution " Dn 0 " should be taken. As a result, the critical load and the length corresponding to this critical load must be found by finding the minimum load Single layer large deflection, SLLD approach 1

22 . Delamination in Composites.1. Causes of Delamination High interlaminar stress conditions affect the initiation or spreading of delamination. Permanent high interlaminar stress; residual stresses can arise due to the production process, temperature and humidity conditions, or geometric configuration. These can affect material for long periods of time. Mechanical loads cause the "interlaminar strase" to act instantaneously and increase its effect immediately Manufacturing Resin is used in the production of FRP composites. The resin must be cured to benefit the composite material, to achieve this, the resin is heated to certain temperatures and cooled to certain temperatures. This causes the laminate to be exposed to a temperature difference of degrees. This difference in temperature causes residual stress on the laminate, due to different expansion coefficients. Incompatible thermal expansions are observed, due to the orthotropic nature of the material. The altered laminate placement angles are subected to residual stress. And over time it increases the likelihood of laminate delamination..1.. Environmental effects Laminates are environmentally affected by temperature and humidity. Changes in humidity and temperature cause residual stress to develop over time. Anisotropic differentiation occurs in the properties of the material. These conditions increase the likelihood that the material will experience delamination..1.3 Drilling Composite materials need to undergo drilling in some cases, for example; for fixing, wiring or weight saving. However, this process causes high interlaminar stress on the laminate. Delamination effects are seen in the area where the drilling process is. The top layers are folded down and up by the drill. After this step, the back surface of the material is more damaged than the front surface, as shown in figure.1. This resulting delamination can

23 become important in case of overloading of the layers. For this reason, the mechanisms of drilling must be known and precautions should be taken accordingly. Figure.1 Delamination caused by drilling.1.4. Geometrical configuration The discontinuities in the material and configurations cause singularities in interlaminar stress and are the main cause of geometric dependent delaminations. Various geometric situations related to this delamination have been tried to be explained below. Taper: Tension and bending on the internal or external ply drops reason the formation of critical stress densities. Inclusions: Stress singularity in the zone being affected can be formed by bolts, holes and notches. It is likely that delamination will occur under bearing conditions. Skin-stringer debonding: Joining techniques are preferred during curing of FRPs compared to older methods. During this process, some oin elements are used. The parts that are separated from the laminate under load may cause delamination. 3

24 Free-edge: Free edge effect is a common occurrence in the causes of delamination. The difference in the characteristics of together stored lamina is due to this. This causes excessive interlaminar stress on the edge of the laminates and causes delamination over time Low velocity impact During periods of manufacturing, maintenance or service life, the materials can be reduced or like ways and this can cause low velocity impacts. Composite crusts can cause matrix cracks after a certain speed threshold. These matrix cracks can cause delamination. This effect may occur suddenly or in time.. Delamination Modeling Delamination has a significant effect on the materials, however this condition can be tolerated. These delaminations may not be vital in any case. For this, the critical limits to which construction can tolerate must be well defined. First, the conditions that cause delamination must be eliminated or prevented from spreading. FRP composites may be more useful and preferable by developing analytical models, numerical methods and standard tests. In FRP composites, Fracture mechanics based models and damage mechanics based models are two main ways for delamination modeling...1 Fracture mechanics based models Interlaminar cracks in anisotropic media conditions can be seen as delimination of FRP composites. For this reason, analytical and numerical models used in the concept of fracture mechanics of isotropic materials are used to predict the occurrence of delaminations. 4

25 Fracture modes The triple combination of fracture modes can sometimes be seen as deformation and crack propagation in fracture mechanics, as shown in figure.. Figure.. Three fracture modes The deformation caused by normal stresses applied to the crack plane in the normal direction is defined as mode I or opening mode. The in-plane shear stress inducing crack propagation perpendicular to the front edge results in sliding mode expressed by Mode II. Out-of-plane shear stress causes deformation parallel to the front edge, which is defined as tearing mode, mode III Stress and deformation fields In the late 1950s, George R. Irwin identified the elastic stress near the crack tip. In the fracture space, he proposed the concept of the stress intensity factor. Here it is easy to calculate the crack-tip stress and calculate the strain areas subected to tensile and shear loads. A lot of research has been done about the elastic stress area near a crack tip. These studies have detected that a line of discontinuity with zero thickness and a length in a homogeneous isotropic linear elastic infinite plate is as a crack. 5

26 Figure.3. Crack in a homogenous isotropic linear elastic infinite plate If the crack at infinity affected by a tensile stress ( ), for Mode I, stresses of an element (dxdy) at distance (r) and angle ( ) to the crack tip is found as below that: a 3 x cos 1 sin sin r a 3 y cos 1 sin sin r a 3 xy cos sin cos r 0 plane stress z ( ) plane strain z x y (.1) Generalized form can be written as: i K I r f ( ) where K a I i (.) 6

27 Stress intensity factor (Kı) for mode I can be written in Equation.1 and Equation.. At the same time, this expression can be defined for mode II and mode III, and the stress field around a crack type can be defined as below that: KI 3 KII 3 x cos 1 sin sin sin cos cos r r KI 3 KII 3 y cos 1 sin sin sin cos cos r r xy yz KI 3 KII 3 cos sin cos cos 1 sin sin r r KIII cos r KIII zx sin r 0 plane stress z ( ) plane strain z x y (.3) Also, deformations can be written as: KI r KII r u cos K 1 sin sin K 1 cos KI r KII r sin K 1 cos cos K 1 sin KIII r w sin 3 K plane stress 1 K 3 4 plane strain (.4) where is shear modulus and is Poisson s ratio Irwin s crack closure concept and strain energy release rate 7

28 In a fragile material, the energy enough to close the new cracked surfaces is equal to the energy that causes crack formation, according to Irwin. This is why it requires the same energy as expanding a crack from a to a + a and closing the crack from a + a to a. The energy required to close the crack from a + a to a can be written using the equations of stress and deformation: a 1 W y ( a r) ( r) dr (.5) 0 And strain energy release rate G is obtained as: a W 1 G lim lim ( a r) ( r) dr a a (.6) a0 a0 y 0 Substituting G in stress and deformation field equations and integrating: KI KII KIII G GI GII GIII (1 ) (.7) ' ' E E E Respectively, total strain energy release rate and energy release rate are called G and mode I, for mode I, GII for mode II, G I for G III for mode III. Furthermore, stress intensity factors is called K I K II for mode II, K III for mode III. ' E E plane stress ' E E / (1 ) plane strain (.8) where E is the Young s modulus and is the Poisson s ratio of the material. To predict the onset or spread of delamination, the strain energy release rate is defined as a failure criterion. So, when this amount reaches the critical value, the crack extension appears. This value is variable in all three modes, and is defined the critical energy release rate. 8

29 .. Damage mechanics based models With the framework of damage mechanics, attempts have been made to define another approach to delamination failure. Cohesive zone or interface models are these. It is often found in finite element models that follow the growth of the declaration because it does not require modification for the representation of material softening, which provides advantages over other fracture mechanics models. Another advantage is that delamination initiation and propagation estimation is possible without the need for initial cracks. Accordingly, it is thought that there is a cohesive or interface layer between the delamination surfaces. This is used to model material softening, and it is possible that there may be linear or nonlinear reactions under normal and shear stress conditions. The Traction displacement curve forms the core of the model. If traction reaches the interfacial strength or maximum traction, damage initiation happen as show figure.4. Traction reaches zero level when the area under the curve is equal to the interlaminar fracture toughness (Gc). This allows us to understand that the fracture energy is free and the crack formation is over. The area with delamination is expressed by a traction-free geometrical discontinuity, after a new crack is formed. Figure.4. Traction-displacement curve for the tip of the delamination...1 Pure mode loading criterion 9

30 For pure mode loading, delamination initiation and propagation criteria are simple. If the traction of mode I, mode II or mode III, interfacial strength or maximum traction in mode I, mode II or mode III is respectively reached, the onset of delamination can be estimated. Thus, for pure mode loading, the general initialization criterion can be expressed as: (.9) i i If the energy release rate calculated from the area under the traction displacement curve of mode I ( G ), mode II ( G ) or mode III ( G I ( G ), II ( G ) or mode III ( G I c II c II III c III ) is greater than the interface force or the mode I ), the delamination is spreaded.... Mixed-mode loading criterion The coupling effects between modes cause the mixed mode loading criterion to be more complex. According to the Ye criterion, the cohesive model is predicted from the beginning of the delamination and the interaction of modes oins the account. f initiation f 3 1 ( i) (.10) where initiation f, f and. is called respectively the failure criterion, the norm of tractions and the MacAuley bracket, describe as: i 1 x ( x x ) (.11) The power law expression is one of the commonly used error criteria to estimate delamination propagation. f propagation G G G f Gi G G G I II III Ic IIc IIIc (.1) 30

31 , and are found as a result of experimental studies. In cases where experimental data can not be obtained, the variables for the linear are chosen as 1 and, for the quadratic failure criterion is selected as..3 Standardization of Delamination Testing According to research, delamination damage is a serious effect that affects fiber reinforced plastic. In order to prevent these situations, the conditions that cause this situation must be known. And at the same time, the interlaminar strength of the materials must be learned in order to achieve this. For quality and safety, some standard tests are needed to determine the interlaminar fracture toughness of FRP composites..3.1 Historical development Fracture testing methods have been used for a long time in composite materials. Test methods with mica were started to be tried in the 1930s, followed by materials such as crystals in the 1950s and timber in the 1970s. Beginning in the 1980s, studies were started to measure the delamination resistance of composite materials. In 198, the first test procedure for delamination by NASA was introduced with respect to the mode I delamination test using the DCB (Double Cantilever Beam) example. Standardization work continues for test methods for pure and mixed fracture modes by the American Society for Testing and Materials (ASTM), European Structural Integrity Society (ESIS) and Japanese Industrial Standards (JIS) organizations. There are 3 mode types, Mode I, Mode II and Mode III..3. Mode I delamination testing In 1994, ASTM, ESIS and JIS concluded that the mode I testing method was the ASTM standard, using the DCB specimen, as show figure.5. In the following period, this test method was adopted as an ISO (International Standards Organization standard in 001). The way to follow the Mode I delamination test is by traction with a constant displacement rate from the edge of the delaminated area on either side. As the crack extends from the edge of 31

32 the initial delamination, the delamination propagation and the load displacement curve are saw together and saved. Using numerous data analysis techniques, mode I interlaminar fracture toughness G is found and the delamination resistance of the material is achieved. I c Figure.5. DCB specimen.3.3 Mode II delamination testing There is currently no internationally recognized mode II delamination test method available. However, there are many suggestion test methods. The four most preferred ones are: ENF (end notch flexure) specimen, SENF (stabilized end notch flexure) specimen, 4ENF (four point bend end notch flexure) specimen, ELS (end loaded split) specimen. 3

33 REFERENCES ASTM Standard test method for mode I interlaminar fracture toughness of unidirectional fibre-reinforced polymer matrix composites, D558. American Society for Testing and Materials International, ASTM Babu, P. R. & Pradhan, B Effect of damage levels and curing stresses on delamination growth behaviour emanating from circular holes in laminated FRP composites. Composites: Part A, 38, Broek, D Elementary engineering fracture mechanics (3rd ed.). Dordrecht: Martinus Nihoff Publishers, 8-9. Broek, D Elementary engineering fracture mechanics (3rd ed.). Dordrecht: Martinus Nihoff Publishers, Hocheng, H. & Tsao, C. C Effects of special drill bits on drillinginduced delamination of composite materials. Machine Tools & Manufacture, 46, Irwin, G. R Relation of Stresses near a Crack to the Crack Extension Force. Proceedings of the 9th International Congress on Applied Mechanics, Brussels, ISO 001. Fibre-reinforced plastic composites - determination of mode I interlaminar fracture toughness, G, for unidirectionally reinforced materials, International I c Organization for Standardization, ISO 001. Kaw, A. K Mechanics of composite materials (nd ed.). Boca Raton, FL: CRC Press, Krueger, R. 00. The virtual crack closure technique - history, approach and applications. Applied Mechanics Reviews, 57 (),

34 Landes, J. D The Contributions of George Irwin to Elastic-Plastic Fracture Mechanics Development, in Fatigue and Fracture Mechanics, 31st Volume, STP1389. American Society for Testing and Materials, MSC Software 010. MD/MSC Nastran 010 Online Documentation [User s guide]. Santa Ana, CA: MSC Software Corp. 85 Moore, D. R., Pavan, A., Williams, J. G. (Eds.) Fracture mechanics testing methods for polymers, adhesives and composites. Amsterdam: ESIS Publication, NASA 198. Standard test method for mode I interlaminar fracture toughness of unidirectional fibre-reinforced polymer matrix composites, NASA RP-109. National Aeronautics and Space Administration, NASA 198. Newman, J. C., Jr Irwin's stress-intensity factor - a historical perspective, in Fatigue and Fracture Mechanics, 31st Volume, STP1389. American Society for Testing and Materials, Paris, P. C. & Sih, G. C Stress analysis of cracks, in Fracture Toughness Testing and Its Applications, STP381. American Society for Testing and Materials, SIMULIA 010. Abaqus 6.10 Online Documentation [User s guide]. Providence, RI: Dassault Systèmes Simulia Corp. Sridharan, S. (Ed.) Delamination behaviour of composites. Boca Raton, FL: CRC Press, Tay, T. E Characterization and analysis of delamination fracture in composites: An overview of developments from 1990 to 001. Applied Mechanics Reviews, 56 (1). Turon, A., Camanho, P. & Costa, J Delamination in composites: simulation of delamination in composites under static and fatigue loading using cohesive zone models. Saarbrücken: VDM Verlag Dr. Müller,

35 Ye, L Role of matrix resin in delamination onset and growth in composite laminates. Composite Science and Technology, 33, Wu, E. M. & Reuter, R. C. Jr Crack Extension in Fiberglass Reinforced Plastics. T&AM Report No. 75. Department of Theoretical and Applied Mechanics, University of Illinois, Urbana, IL, February

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