The in-plane shear responses of continuous fiber reinforced composite materials under high strain rates were characterized experimentally. The V-notch rail shear configuration was used for characterizing the in-plane shear behavior of Newport NB321/3k70 plain weave carbon fabric/epoxy, NCT/321/G150 Carbon Fiber Unitape/epoxy, NB321/7781 fiberglass/epoxy, Cytec PWC/T300/3KNT plain weave carbon fabric/epoxy and Fibercote 3KPW/E365 plain weave carbon fabric/epoxy systems. The testing was conducted using a servo hydraulic testing machine at nominal stroke rates ranging between 0.00083in/sec to 500in/sec. A maximum average shear strain rate of 654rad/sec was achieved up to shear strain levels of 0.08 radians, during the tests. The stress-strain behavior of all material systems exhibited contrasting behavior with increasing stroke rates. The stress-strain curves for all materials exhibited an asymptotic behavior for stroke rates approaching 10in/sec and were material dependent for stroke rates exceeding 250in/sec. For Newport NB321/3k70 and NB321/7781 systems at the highest test rate, the shear strengths increased by a factor of three relative to that of the quasi-static rate, and were independent of the reinforcement type. The shear strength for Cytec PWC/T300/3KNT, Fibercote 3KPW/E365 and Newport NCT/321/G150 increased until a stroke rate of 100in/sec and later decreased with increasing stroke rate. The failure modes for Cytec PWC/T300/3KNT, Fibercote 3KPW/E365 and Newport NCT/321/G150 were at the minimum section of the shear coupons, while for Newport NB321/3k70 and NB321/7781, failure was observed to change from a shear mode across the minimum section to a complex failure mode away from the minimum section. / "December 2006."

Predicting failure of fabric composite laminates in the presence of flaws such as cracks and stress raisers has been an important research problem for the last two decades. Most of the existing models for predicating notched strength are of a ‘cure fit’ nature, wherein the model parameters e.g. characteristic distance or damage zone sizes are chosen so as to fit the experimental data. Such parameters have been shown to depend on notch size and laminate orientation, and as such, cannot be considered material constants for the composite system. The energy dissipation is a physical phenomenon that captures the collective behavior of the failure mechanisms without requiring an explicit knowledge of the mechanisms, and it can also be related to local stiffness changes, leading to a form of nonlinear structural behavior. An approach to characterize failure behavior and degree of load induced internal damage in single lamina of satin weave fabric with off-axis loading and having a central notch is proposed. From the experimental results it was found that energy absorption for notched and un notched woven fabric composite is a function of fiber orientation and notch size. Energy absorption increases as the fiber orientation increases but decreases with increase in the notch size. The failure loads for hysteresis loading were about the same as that of one time loading, indicating that no damage growth occurred due to unloading and subsequent reloading. The energy density is directly proportional to fiber orientation indicating that higher energy densities for higher off-axis angles. The failure modes and the fracture propagating direction for the off-axis notched specimen were very much dependent on the fiber orientation. Recommendations were made for future work for the determination energy absorption characteristics, from lamina level to laminate level and for different weave types. / Thesis (M.S.)--Wichita State University, College of Engineering, Dept. of Aerospace Engineering

The transverse compressive behavior of honeycomb cores has been widely exploited for energy absorption applications in the automotive and aerospace industry. The energy absorption in honeycomb cores with thin reinforced thermoplastic/thermoset cell walls is achieved by progressive folding/fracture of cell walls. The energy absorption by honeycomb cores has been assumed to be primarily due to loading along thickness direction. However, inclined loads cannot be precluded and their influence on energy absorption process must be measured. In this investigation, the effects of inclined/oblique loads on the compressive behavior and energy absorption of Plascore Nomex honeycomb core has been investigated experimentally under static and dynamic loading. The load inclination to the cell wall direction was varied by using off-axis core specimens. The effect of loading angle, core thickness and test speed on the resulting energy absorption has been studied. Test specimens with varying thicknesses of 0.5˝, 1˝, 1.5˝ and 2˝, with cell walls oriented at 0º, 5º, 15º, 30º and 45º to the loading axis were used in the study. These specimens are tested using a servo hydraulic testing machine at rates of 0.008333, 10-1, 1 and 10 in/s. It was observed that the initial load peak in the compression response, the sustained crushing load and energy absorption was influenced by the cell-wall orientation, core thickness and the test speed. / Thesis (M.S)--Wichita State University, College of Engineering, Dept. of Aerospace Engineering / "December 2007."

The compressive residual strength and behavior of impact damaged sandwich panels are governed by the amount of non-visible core damage underneath the facesheets. The failure modes under in-plane compressive loading are either due to crack precipitation from the damage zone or unstable dimple propagation across the width. In both cases, the local bending of facesheet (dimple) triggers the final failure sequence. Thus, containment of this local bending by reinforcing the core cell(s) in the damage region will mitigate the failure initiation and thus increase residual properties. A repair technique involving the containment of local bending of sandwich facesheet by strategically filling honeycomb core cells has been explored to improve the damage tolerance of impact damaged sandwich panels. The experimental results indicate that filling honeycomb cells at the center of the damage region in addition to those at the edge of the damage region produces the maximum benefit. The test data and final failure mode of the repaired specimen under in-plane compression loading indicated that the present repair technique can help recover the undamaged strength of the sandwich panels. / Thesis [M.S.] - Wichita State University, College of Engineering, Dept. of Aerospace Engineering

Structural elements are subjected to variable amplitude/spectrum fatigue loading where tensile overloads and compressive underloads may occur randomly. Depending on the magnitude of these loads, localized plastic deformation may occur at notches leading to residual stresses. These residual stresses have been known to significantly alter the fatigue life. In a spectrum, overloads may seldom occur in blocks and the effects of single overload cycles on the fatigue life must be investigated. In the present investigation, 2024-T3 clad aluminum sheet specimens with a gage width of 0.75” and circular hole diameter of 0.161” were subjected to constant amplitude fatigue loading with periodic overloads. The theoretical net stress concentration factor based on net section stress for each of the specimens is 2.48. Constant amplitude fatigue tests were conducted to determine SN curves at mean remote stress levels of 3ksi and 6ksi. Periodic tensile overloads of 44ksi, 35ksi and 20ksi were applied to constant amplitude loading blocks of 40ksi, 30ksi, 20ksi and 15ksi, each with a mean stress of 6ksi. The overload cycles reached the compressive load level prior to transitioning to the constant amplitude portion of the loading. The period between the overloads was varied at 10 cycles, 50 cycles, 100 cycles, 1,000 cycles and 5,000 cycles of constant amplitude loading. The experimental results showed an increase in the number of cycles to failure when the periodic tensile overloads were applied after every 100 cycles, 1,000 cycles and 5,000 cycles, and if frequent reverse yielding did not occur, as compared to the number of cycles to failure in the case of constant amplitude loading without overloads. In the analytical approach, Miner‟s rule was first used to estimate the cycles to failure. The study incorporated notch residual stresses into the Miner‟s rule in an effort to improve the life estimation. Neuber‟s rule was used to estimate residual stresses due to overloads. Two ideal conditions of strain hardening were considered – isotropic hardening and kinematic hardening – to establish a stress-strain curve for one cycle of loading. The results showed improved estimation of cycles to failure, when residual stresses were included in Miner‟s rule and if no reverse yielding occurred at the notch while fatigue loading. The loading that showed such a trend was maximum constant amplitude stress of 15ksi and a mean stress of 6ksi on which an overload of 20ksi was periodically applied. An improvement of 17% - 19% was obtained for a periodicity of 50 cycles or greater. In other cases of loading, if reverse yielding occurred at every cycle or frequently between cycles (periodicity of 10 cycles to 100 cycles) during fatigue then this caused detrimental damage and the specimen experienced a reduction in life. / Thesis (M.S.)--Wichita State University, College of Engineering, Dept. of Aerospace Engineering.

The rate sensitivities of mode-I and mode-II fracture toughness, in the form of GIc and GIIc of Toray T800S/3900 carbon unitape/epoxy, T700G/3900 Plain weave carbon fabric/epoxy, and Newport NB321/7781 fiberglass/epoxy materials were investigated experimentally. Static and dynamic tests, with stroke rates ranging from 8.33×10-4in/s to 100in/s, were conducted on double cantilever beam and end notch flexure specimens. Both [0°] N and [±45°] N laminates were studied in this investigation. The rate sensitivities were characterized in terms of the initiation fracture toughness and crack growth resistance. The average mode-I initiation fracture toughness, GIc of NB321/7781 [0°] laminate was observed to decrease with stroke rate from 6.5lbf-in/in2 to 4lbf-in/in2 and further increased to 6lbf-in/in2 over a range of 5 decades of crack opening displacement rates. The resistance tended to be constant with crack length for the low rate tests. At higher rates, the resistance decreased gradually. The tests data indicated ductile stable and brittle unstable behavior with a transition stick-slip behavior. The average mode-II fracture toughness of the NB321/7781 [0°] laminate was observed to increase from 12lbf-in/in2 to 17lbf-in/in2 over the entire range of 5 decades of shear displacement rates. The fracture behavior was observed to be brittle stable at all rates. The NB321/7781 [±45°] laminate’s opening mode initiation toughness was observed to decrease from 9lbf-in/in2 to 3lbf-in/in2 and further increased to 4lbf-in/in2 over a range of 5 decades of crack opening displacement rates. The delamination resistance increased with crack length for the low rate tests. At higher rates, the resistance decreased gradually. The fracture behavior transitioned from ductile stable to brittle unstable. The opening mode initiation toughness of T700G/3900 [0°] laminate decreased from 3.5lbf-in/in2 to 2lbf-in/in2 and further increased to 3.2lbf-in/in2 over a range of 5 decades of viii crack opening displacement rates. The delamination resistance indicated constant pattern irrespective of the stroke rates. The fracture behavior at low rates indicated transition stick-slip behavior, whereas brittle unstable behavior was evident at higher rates. The mode-II fracture toughness tended to be constant at 4.5lbf-in/in2 over a range of 5 decades of shear displacement rates. Brittle stable fracture behavior was observed at all rates. The opening mode initiation toughness of T700G/3900 [±45°] laminate was observed to decrease from 2.75lbf-in/in2 to 1.8lbf-in/in2 and further increased to 2.8lbf-in/in2 over a range of 5 decades of crack opening displacement rates. The delamination resistance tended to increase with crack length at lower rates, whereas it tended to decrease at higher rates. The fracture behavior indicated transition stick-slip behavior at lower rates, whereas brittle unstable behavior was evident at the highest rate tested. The T800S/3900 [0°] laminate’s opening mode initiation toughness was observed to steadily decrease from 5lbf-in/in2 to 2.5lbf-in/in2 over a range of 5 decades of crack opening displacement rates. The delamination resistance was observed to increase with crack length at lower rates, whereas it tended to decrease gradually at higher rates. Brittle stable fracture behavior was observed at all rates. Fiber bridging was prominent with an increase in test rates. The mode-II fracture toughness increased steadily from 9lbf-in/in2 to 20lbf-in/in2 over a range of 5 decades of shear displacement rates. Brittle stable fracture behavior was observed at all rates. / Thesis (M.S.)--Wichita State University, College of Engineering, Dept. of Aerospace Engineering.

The adhesively bonded layered aluminum is used in aircraft structures to avoid knife edge situations when flush head fasteners are used with minimum gage skins. Due to the countersunk hole and adhesive bonding, stress flow becomes more complicated. Extensive knowledge of the different parameters that affect the behavior of the bonded joints with countersunk holes is essential for dependable and effective design. A 3-D finite element model was used to estimate the location and magnitude of stress concentration under remote tension for the aforementioned problem. The influence of the various parameters on stress concentration was investigated for a counter sunk angle of 100º. Different parameters such as ratio of young‟s modulus of adhesive to aluminum, position of adhesive layer, countersunk sunk depths, ratio of thickness to radius and ratio of width to radius have been addressed in this study. The stress flow varies significantly when the plates are filled with fasteners of different pre-tension loads. Also the effects of pre-tension loading were compared for the cases of open hole and fastener filled hole without pre-tension for bonded, monolithic and straight shank hole. The results obtained from the finite element analysis for the monolithic cases have been validated against those reported in literature. / Thesis (M.S.)--Wichita State University, College of Engineering, Dept. of Aerospace Engineering.

A three dimensional solid finite element model was assembled to investigate the influence of parameters such as friction coefficient, fastener preload and material plasticity on the structural response of Hi-Lok bolted single-lap joints under remote tension/compression loading. Three different specimen types which produce nominal load transfer levels of 6%, 30% and 50% have been investigated. The effects of these parameters where characterized in terms of load transfer, fastener rotation, and stress concentration factor. Friction coefficient has been varied from 0 to 1.25 and the fastener preloads of 1.35lbs, 13.5lbs and 135lbs were assigned for different simulations. Furthermore, constant amplitude load cycles of 15, 20, 30 and 40Ksi were applied in different FE analysis. Results of elastic and isotropic hardening based elastic-plastic material models have been compared and effects of plasticity have been explored. The analysis results indicate that friction coefficient is the most important parameter and friction has a great influence on load transfer, stress concentration and fastener rotation, for all loading conditions, and for both elastic and elastic-plastic material models. Friction, which is affected by both friction coefficient and clamp-up force, has the most influence on load transfer and stress concentration factor at lower remote loads and smaller amount of load transfer. Plasticity which is more prevalent at high remote loads reduces load transfer and stress concentration factors, especially during tension. / Thesis (M.S.)--Wichita State University, College of Engineering, Dept. of Aerospace Engineering.

The stress-strain behavior and failure of composite materials are strain rate sensitive, and influenced by the dimensions of the structure. To elucidate the combined effects of scaling and strain rate on the strength of unnotched continuous fiber reinforced composites, an experimental investigation has been conducted on Newport NB321/7781 fiberglass/epoxy and Toray T800/3900-2B unitape/epoxy materials. The experimental results have been characterized in terms of failure strength, failure modes and the Weibull modulus m. A 2D-scaling approach has been followed and composite coupons were fabricated with [0]4 and [±45]s stacking sequences. The experimentation has been conducted at strain rates ranging from quasi-static (0.0002 s^-1) to high strain rate (50 s^-1), to study the mechanical responses and associated failure modes. Subsequently, the Weibull statistical model was utilized to characterize the scaling behavior at different strain rates. The average failure stress of [0]4 carbon, [0]4 fiberglass and [±45]s fiberglass specimens were observed to decrease with increasing specimen size at each strain rate. However, at high strain rate, the percentage of strength reduction was observed to be lower in comparison to the quasi-static strain rate. Owing to the free edge effects, the scaling effect was maximum for [+45/-45]s carbon unitape specimens. But unlike the other stacking sequences, the percentage of strength reduction at higher strain rates was higher compared to quasi-static strain rate, indicating increased scaling effects with strain rate. Weibull modulus m for the specimens tended to increase with increasing strain rate indicating diminishing scaling effects, while [+45/-45]s carbon specimens exhibited opposite trend. Failure at multiple locations was observed in larger coupons at high strain rate, which results in size and strain rate dependent fracture behavior. / Thesis (M.S.)--Wichita State University, College of Engineering, Dept. of Aerospace Engineering.

The Split Hopkinson Pressure Bar (SHPB) technique is widely used to dynamically characterize metals and is increasingly being used to characterize non-metallic materials such as fiber reinforced polymeric composite materials. However, the tensile version of the SHPB apparatus requires specimen gripping devices and/or complex loading mechanisms that distort and attenuate the loading pulse. Strain estimations in the test specimen based on the one-dimensional wave propagation theory are found to differ from direct measurements of the same. The discrepancy between the measured and estimated strains along with a general lack of guidance for tensile load generation limit the broader application of the testing technique. The current investigation addresses tensile load generation in a tensile SHPB apparatus, establishes a reliable loading methodology, develops a correction methodology for one dimensional theory strain estimation, and identifies the factors that contribute to wave modulation. A correction methodology for one-dimensional wave propagation analysis is presented to address the discrepancy between strains estimated by one dimensional wave propagation theory and strains measured directly over the test specimen. A method for correcting the strains in the frequency domains using Fourier analysis is presented. The correction methodology is applied to virtual strain measurements from simulations to establish its applicability to experimental results. Subsequently, the methodology is validated with experimental data from carbon fabric laminated composite specimens. / Thesis (Ph.D.)--Wichita State University, College of Engineering, Dept. of Aerospace Engineering