An experimental investigation using grids with a frequency of 125 lines/in. (5 lines/mm) was performed on inert propellant and pure binder at two different global head rates of 0.1 in./min (2.5 mm/min) and 1.0 in./min (25.4 mm/min). From the extracted data, displacements, strains, and dominant eigenvalue for displacement were calculated. An idealized model was used to explain the high strain zone ahead of inert propellant that caused severe blunting at the crack tip. Using the available algorithms and three dimensional photoelasticity, the dominant stress singularity order values were calculated in a four point single edged cracked bend specimen with both straight front and thumbnailed cracks. The free surface values are the same as for the inert propellant and in good agreement with analytical values. A boundary layer is observed in the singularity order which extends towards the mid-plane of the specimen. This region is about twenty percent of the distance from the free surface to mid-depth of the fractured body. The slow and fast head rates alter the global behavior of the specimen as well as the density of the displacement and strain contours. However, the near tip mechanisms are not altered. / Ph. D.

In this study, an analytical procedure to predict the strength and failure of laminated composite structures under monotonically increasing static loads is presented. A degenerated 3-D shell finite element that includes linear elastic and plastic material behavior with full geometric nonlinearity is used to determine stresses at selected points (Gauss quadrature points in each element) of the structure. Material stiffness (constitutive) matrices are evaluated at each Gauss point, in each lamina and in each element, and when the computed stress state violates a user selected failure criterion, the material stiffness matrix at the failed Gauss point is reduced. The reduction procedure involves setting the material stiffnesses to unity. Examples of isotropic, orthotropic, anisotropic and composite laminates are presented to illustrate the validity of the procedure developed and to evaluate various failure theories. Maximum stress, modified Hills (Mathers), Tsai-Wu (F₁₂ = 0), and Hashin's failure criteria are included. The results indicate that for large length-to-thickness ratios, the geometric nonlinear effect should be incorporated for both isotropic and anisotropic structures. The nonlinear material model influences the behavior of isotropic structures with small length-to-thickness ratios, while having nearly no effect at all on laminated anisotropic structures. Of the four failure theories compared, each predicts failure at nearly the same load levels and locations. Hashin's criterion is particularly noteworthy in that the mode is also predicted. / Ph. D.

Composite laminates are prone to delamination at free edges, straight edges or at holes, due to the mismatch at interfaces where two adjacent plies have different fiber orientations and/or different material properties. The linear analysis of the mismatch at the edge results in a mathematical singularity. That phenomenon occurs in a boundary layer and has to be treated mathematically and physically as such. In the literature it is called the "Boundary Layer Effect" or simply the "Edge Effect". It is of great importance to recognize and be able to predict delamination locations at edges prone to such events. The goal of this research was to create a model capable of providing such a prediction. In an effort to generalize the model, the more complicated case of a free edge at holes in the composite laminate was chosen rather than the case of a straight free edge. A sequel of three major efforts was completed: 1) Development of the analysis of the free-edge effect at a hole in a composite laminate, 2) Performance of an extensive experimental program to provide data for the creation of the prediction model, and 3) On the basis of the analysis, establishment of the model, and comparison with the experimental results. The prediction model consists of two major products of the analysis, the order of the singularity and the strain energy release rate. Both are useful in locating the interface most prone to delaminate and the point along the hole circumference where it initiates. Two material systems (AS4/3501-6 and AS4/1808) and two stacking sequences [(0/45/0/-45)_s)₄] and [ (0/45/90/-45)_s)₄]s , quasi-orthotropic and quasi-isotropic respectively, were quasi-statically tested under tension and compression. The specimens were X-rayed after each loading stage in order to locate the initiation of delaminations. The fact that both materials consisted of the same type of fibers, was an excellent opportunity to examine the performance of the matrix and its influence on the process of delamination. Matrix dependent behavior was successfully examined and studied through the experiments and the prediction model. Results showed good correlation and high sensitivity to the type of matrix material involved. / Ph. D.

Compressive loads can cause local buckling in composite laminates that have a near-surface delamination. This buckling causes load redistribution and secondary loads, which in turn cause interlaminer stresses and delamination growth. The goal of this research effort was to enhance the understanding of this instability-related delamination growth in laminates containing either an embedded or an edge delamination. There were three primary tasks: 1) development of a geometrically nonlinear finite element analysis named NONLIN3D; 2) performance of a parametric analytical study to determine the effects of strain, delamination shape, and delamination size on the distribution of the strain energy release rate components along the delamination front; and 3) performance of a combined experimental and analytical study of instability-related delamination growth (IRDG). Two material systems (AS4/PEEK and IM7/8551-7) and two stacking sequences (0/90/90/0)₆ and (90/0/0/90)₆ were examined. The laminates were fabricated with Kapton inserts between the fourth and fifth plies from the top surface to give an initial delamination. The analysis predicted a large variation of G_I and G_II along the delamination front. The G_III component was always small. The location of maximum G_I and G_II depended on the delamination shape and applied strain. In general, the strain-energy release rates were small except in a small region. Hence, delamination growth was expected to occur over only a small portion of the delamination front. Experiments corroborated this prediction. The laminate stacking sequence had a large effect on the shape of the deformed region, the direction of delamination growth, and the strain at which delamination growth occurred. These effects were predicted by the analysis. The G_I component appeared to govern initial delamination growth in the IM7/8551-7 laminates. Matrix ply cracking generally accompanied delamination growth. In some cases fiber micro-buckling also occurred shortly after delamination growth occurred. / Ph. D.

A basic physical explanation, a model, and comments on NDE application of the acoustoultrasonic method for composite materials are presented. The basis of this work is a set of experiments where a sending and a receiving piezoelectric transducer were both oriented normal to the surface, at different points, on aluminum plates, various composite plates, and a tapered aluminum plate. Chapter one introduces the purpose and basic idea of the dissertation, while supporting its need. Also, general comments on the AU method are offered. The second chapter offers a literature review of areas pertinent to the dissertation, such as composite materials, wave propagation, ultrasonics, and the AU method. Special emphasis is given to theory which is used later on and past experimental results that are important to the physical understanding of the AU method. The third chapter describes the experimental set-up, procedure, and the ensuing analysis. In the fourth chapter, the experimental results are presented in both a quantitative and qualitative manner. Chapter five furnishes a physical understanding of experimental results based on elasticity solutions, Lamb wave theory, and through-the-thickness-transverse·resonance (TTTR). Computer results are presented for sake of comparison. The sixth chapter discusses modeling and applications of the AU method for composite materials and the seventh chapter states general conclusions. The unique offering of this work is the physical model of the AU method for composite materials, something which has been much needed and sorely lacking. This physical understanding is possible due to the extensive set of experimental measurements, also reported in this dissertation. / Ph. D.

Failure processes in unidirectional composite materials subjected to quasi-static tensile load along the fiber direction are investigated. The emphasis in this investigation is to identify the physical processes taking place during the evolution of failure in these materials. An extensive literature review is conducted and the information relevant to the present topic is summarized. The nature of damage growth in five different commercially available composite systems are studied. In-situ scanning electron microscopy is employed for identifying the failure events taking place at the microscopic level. Acoustic emission monitoring is used for estimating the rate of damage growth on a global scale and determining the size of individual failure events. The results of this study have shown the important roles of the matrix material and the interphase in determining the tensile strength of unidirectional composite materials. Several failure modes occurring at the microscopic scale are revealed for the first time. Further, the results indicate that dynamic fracture participates to a significant extent in determining the failure process in these materials. Based on the results of this study the influence of various parameters in determining the composite strength is described. / Ph. D.

Damage states in laminated composites were studied by considering the model problem of a laminated beam subjected to three-point bending. A combination of experimental and theoretical research techniques was used to correlate the experimental results with the analytical stress distributions. The analytical solution procedure was based on the stress formulation approach of the mathematical theory of elasticity. The solution procedure is capable of calculating the ply-level stresses and beam displacements for any laminated beam of finite length using the generalized plane deformation or plane stress state assumption. The beam lamination can be any arbitrary combination of monoclinic, orthotropic, transversely-isotropic, and isotropic layers. Prior to conducting the experimental phase of the study, the results from preliminary analyses were examined. Significant effects in the ply-level stress distributions were seen depending on the fiber orientation, aspect ratio, and whether or not a grouped or interspersed stacking sequence was used. The experimental investigation was conducted to determine the different damage modes in laminated three-point bend specimens. The test matrix consisted of three-point bend specimens of 0° unidirectional, cross-ply, and quasi-isotropic stacking sequences. The dependence of the damage initiation loads and ultimate failure loads were studied, and their relation to damage susceptibility and damage tolerance of the beam configuration was discussed. Damage modes were identified by visual inspection of the damaged specimens using an optical microscope. The four fundamental damage mechanisms identified were delaminations, matrix cracking, fiber breakage, and crushing. The correlation study between the experimental results and the analytical results was performed for the midspan deflection, indentation, damage modes, and damage susceptibility. The correlation was primarily based on the distributions of the in-plane component of shear stress, t_xz. The exceptions were for the case of a very small aspect ratio (less than 1.0) where the crushing mode of damage was predicted based on the maximum contact pressure, and for very large aspect ratios (greater than 12.0) where a maximum tensile bending stress criterion was used for predicting the damage initiation loads. / Ph. D.

Free vibration, buckling and postbuckling analyses of laminated composite plates with multiple delaminations are presented. A fInite element method based on a layer-wise laminated composite plate theory is developed to formulate the problem. Geometric nonlinearity in the sense of von Karman and the imperfection in the plate in the form of initial global deflection and initial delamination openings are included. A simple contact algorithm which precludes the physically inadmissible overlapping between delaminated surfaces is proposed and incorporated in the analysis. A sublaminate concept is adopted in the analysis to reduce the computational efforts, and found to be efficient. Numerical results are obtained for through-the-width, circular and rectangular delaminations addressing the effects of the number of delaminations, their lengths and through-the-thickness and axial locations on the critical buckling load and buckling mode shapes as well as free vibration frequency and modes. Postbuckling responses are investigated with respect to different magnitudes and directions of initial imperfections. The effects of material anisotropy and contact condition between delaminated surfaces are also considered. It is found that the proposed approach is very efficient and powerful for solving the above mentioned problems. / Ph. D.

The objective of this study was to observe and characterize the out-of-plane fracture of a 2-D carbon-carbon composite and to gain an understanding of the factors influencing the stress distribution in such a laminate. The experimental portion of this study consisted of performing an out-of-plane tensile test in a scanning electron microscope and determining the modes of failure. Failure was found to be interlaminar, with cracks propagating along the fiber-matrix interface. Finite element analyses of a two-ply carbon-carbon composite under in-plane, out-of-plane, and thermal loading were performed. Stress distributions were studied as a function of stacking sequence, undulation aspect ratio, and undulation offset ratio. The results indicated that under out-of-plane loading σ_x and τ_xz were strongly dependent on the geometric parameters studied, but σ_z and σ_y were relatively independent of geometry. Under in-plane loading all components of stress were strong functions of the geometry, and large interlaminar stresses were predicted in regions of undulation. The thermal analysis predicted the presence of large in-plane normal stresses throughout the laminate and large interlaminar stresses in regions of undulation. An elasticity solution was utilized to analyze an orthotropic fiber in an isotropic matrix under uniform thermal load. The analysis reveals that the stress distributions in the fiber are singular when the radial stiffness C_rr is greater than the hoop stiffness C₀₀. Conversely, if C_rr < C₀₀ the maximum stress in the composite is finite and occurs at the fiber-matrix interface. In both cases the stress distributions are radically different than those predicted assuming the fiber to be transversely isotropic (C_rr = C₀₀). / Ph. D.

Thesis (M.Tech.: Mechanical Engineering)-Dept. of Mechanical Engineering, Durban University of Technology, 2007 xvi, 123 leaves / Interlaminar cracking or delamination is an inherent disadvantage of composite materials. In this study the fracture properties of nano and fibre-reinforced polypropylene and epoxy composite structures are examined. These structures were subjected to various tests including Single Edge Notched Bend (SENB) and Mixed Mode Bending (MMB) tests. Polypropylene nanocomposites infused with 0.5, 1, 2, 3 and 5 weight % nanoclays showed correspondingly increasing fracture properties. The 5 weight % specimen exhibited 161 % improvement in critical stress intensity factor (KIC) over virgin polypropylene. XRD and TEM studies show an increase in the intercalated morphology and the presence of agglomerated clay sites with an increase in clay loading. The improvement in KIC values may be attributed to the change in structure. Tests on the fibre-reinforced polypropylene composites reveal that the woven fibre structure carries 100 % greater load and exhibits 275 % lower crack propagation rate than the chopped fibre specimen. Under MMB conditions, the woven fibre structure exhibited a delamination propagation rate of 1.5 mm/min which suggests delamination growth propagates slower under Mode I dominant conditions. The woven fibre / epoxy structure shows 147 % greater tensile modulus, 63 % greater critical stress intensity factor (KIC), and 184 % lower crack propagation rate than the chopped fibre-reinforced epoxy composite. MMB tests reveal that the load carrying capability of the specimens increased as the mode-mix ratio decreased, corresponding to an increase in the Mode II component. Delamination was through fibre–matrix interface with no penetration of fibre layers. A failure envelope was developed and tested and may be used to determine the critical applied load for any mode-mix ratio. The 5 weight % nanocomposite specimen exhibited a greater load carrying capability and attained a critical stress intensity factor that was 10 % less than that of the fibre-reinforced polypropylene structure, which had three times the reinforcement weight. Further, the nanocomposite exhibited superior strain energy release rates to a material with ten times the reinforcement weight. The hybrid structure exhibited 27 % increase in tensile modulus over the conventional fibre-reinforced structure. Under MMB conditions, no significant increase in load carrying capability or strain energy release rate over the conventional composite was observed. However, the hybrid structure was able to resist delamination initiation for a longer period, and it also exhibited lower delamination propagation rates.