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Hybrid composite architectures employing traditional advanced composites and carbon nanotubes offer significant potential mechanical and multifunctional performance benefits. The architecture investigated here is composed of aligned fibers with carbon nanotubes grown radially on their surface. A novel process for rapidly growing dense, long, high-quality, aligned carbon-nanotube forests is employed. Two fundamental issues related to realizing hybrid composite architectures are investigated experimentally: wetting of the carbon nanotubes by thermoset polymers and retention of mechanical (stiffness and strength) properties of the fibers after the carbon-nanotube growth process. Wetting of carbon-nanotube forests by two commercial polymers (including a highly viscous epoxy) is demonstrated at rates conducive to creating a fully dispersed carbon-nanotube/matrix region around the fibers in a typical composite. Single-fiber tension tests indicate no mechanical degradation for alumina fibers undergoing the carbon-nanotube growth process. Results indicate that hybrid carbon-nanotube/composite architectures are feasible, and future work focuses on mechanical and multifunctional property characterization of other hybrid architectures and scaling to a continuous carbon-nanotube growth process.

A benchmark, geometrically nonlinear, shell-buckling problem is reviewed and the correct solution is identified and discussed vis-a-vis the incorrect solution from the open literature. The problem is that of a hinged, thin, isotropic cylindrical shell section, point-loaded transversely at the center, undergoing large deformations (5x thickness). This benchmark problem and solution were introduced in 1972 and have been reproduced in at least 30 journal articles and books in the more than three decades since that time. The benchmark problem is of interest because it demonstrates many features possible in shell buckling, including a load-limit point (snap-through buckling) and a deflection-limit point (snap-down, or snap-back). The benchmark problem is typically used to demonstrate the capability of commercial/private finite element codes to traverse such complicated nonlinear load paths. The existing incorrect benchmark solution involves the nonlinear growth of only symmetric deformation modes. The correct solution to this benchmark problem involves bifurcation from the nonlinear load path into a simple asymmetric deformation mode. Experimental data do not exist for verification of the calculated benchmark response, and so a similar problem using a laminated composite shell is suggested as an alternative benchmark problem for shell stability analysis.

Experiments have been performed to verify power-optimized modal models of piezoelectric vibration harvesters for microelectromechanical systems. Such harvesters can power a variety of sensors, and there have been recent national workshops dedicated to harvesting. Detailed experimental results, including displacement histories and electrical output, are provided over a range of frequencies and electrical loadings to compare with (optimized) modal models. The harvester geometry considered is that of a symmetric bimorph macroscale cantilever. Although some experimental work for cantilevered bimorph harvesters has been published, key testing and/or device parameters needed for model verification are missing and/or data at and near power optima (the most interesting operating points) are not provided. Therefore, a detailed set of experiments was performed using power-optimized modeling results to guide the test matrix. Over the broad range of parameters tested, the models accurately predicted all trends and device performance away from device resonances (resonance and antiresonance frequencies). Near the resonance frequencies, the model consistently underpredicts electrical performance, which is satisfactorily attributed (and experimentally supported) to the well-known piezoelectric coupling nonlinearity in the large-strain region. The data presented herein can serve as benchmark data to verify other modeling efforts. The verified models have been used to optimally design microelectromechanical system harvesters for commercial aircraft and microfabrication is ongoing. DOI: 10.2514/1.25047

The response of transversely point-loaded composite shells is investigated via the finite element method. The thin, circular cylindrical shells are hinged at a fixed distance apart on the straight edges and are free on the curved edges. Stability issues are considered, including nonlinear prebuckling, limit points, bifurcation, and post buckling. In the finite element formulation, a novel technique is utilized to identify and traverse bifurcation points efficiently in the nonlinear elastic response. A classic shell buckling benchmark problem is solved here correctly for the first lime; previous studies have round the isotropic shell to have a limit point, when, in fact, the shell bifurcates before reaching a limit point. Predicted results for graphite/epoxy laminated shells are in excellent agreement with measured load-deflection and mode-shape evolutions as previously reported. Two cases of extraordinarily good agreement between the data and model, particularly large-deflection asymmetric deformations modes through a bifurcation point and far into the postbuckling regime, clearly establish validated new benchmarks. Asymmetries in both directions of the shell plane, relative to the centrally applied point load, are discussed with regard to bifurcation and material couplings in the composite laminate. Asymmetric stress distributions are discussed in the context of composite damage resistance and previous experimental findings on atypical shell damage mode and extent. The influence of test fixture compliance is evaluated and briefly discussed.

The transverse loading response of convex composite shell structures typical of aircraft fuselage sections was investigated, and the experimental work is reported. Important mechanisms in the response, particularly instabilities (buckling), were studied by investigation of the force-deflection response and the evolution of full-field deformation shapes. Quasi-static tests were conducted to simulate impact of convex shells. The specimens were laminated, cylindrical shell sections in [[[+ or -][45.sub.n]/[0.sub.n]].sub.s] configurations, where n takes on values of 1, 2, and 3. The three structural parameters of radius, span, and thickness were varied according to a scaling relation and were chosen to represent approximate fuselage dimensions of general aviation and commercial transport aircraft. All specimens were evaluated for damage with dye-penetrant enhanced x radiography and sectioning after mechanical testing. The structural response changes both quantitatively and qualitatively for the different shell geometries and was categorized into three types based on the existence of the instability transition characteristics such as deformation shapes. The presence of the instability becomes more likely for deeper, thinner specimens where the ratio of membrane stiffness to bending stiffness is higher. This can be characterized by a structural parameter involving geometric and material factors. The majority of the specimens showed no damage, but the limited experimental evidence did show plate-like damage occurring in the shell specimens before any instability. Suggestions for further work to extend these findings are made.

A study was conducted to analyze the impact and quasi-static response of different composite shell structures with a wide range of structural configurations to transverse loading. Scaling considerations were characterized by variations in the specimen structural parameters while the shells were restrained in a test fixture with boundary conditions of pinned/no in-plane sliding on the axial edges. Results showed that damage resistance was correlated with variations in the structural parameters.

The unstable loading response of convex shells is responsible for the differences in the plate and convex shell responses to transverse loading. The plate and the convex shell responses differ in terms of trends in the response parameters, and the damage extent and distribution. The instability helps the convex shells in dissipating impact energy through structural deformation, thus minimizing the impact damage. This concept of instability has implications for the designing of tolerant aerospace components.

Nanoparticulate zirconia (Zr[O.sub.2]) has catalyzed both growth of single-wall and multiwall carbon nanotubes (CNTs) by thermal chemical vapor deposition (CVD) and graphitization of solid amorphous carbon. The studies have shown that a nonmetallic catalyst has catalyzed the CNT growth by thermal CVD while remaining in an oxidized state and has provided new insights into the interactions between nanoparticulate metal oxides and carbon at elevated temperatures.

The transverse-loading response of laminated composite shell structures is studied experimentally and numerically. Monolithic graphite/epoxy shell structures having layups of [[[+ or -][45.sub.n]/[0.sub.n]].sub.s] (n = 1, 2, and 3) closely represent commercial fuselage structures in both geometry and boundary conditions. A combined experimental and numerical approach is used to assess shell response to centered transverse loading. Experimentally, load-deflection response and mode-shape evolutions are measured and damage resistance characterized via dye-penetrant enhanced x radiography and sectioning. Nonlinear finite element analyses including buckling and dynamic collapse are conducted for comparison to the experimental data. Modeling results allow a more refined interpretation of observed bifurcation phenomena, particularly premature transition to a secondary equilibrium path attributed to geometric imperfections. A novel finite element technique introduced in previous work is found to be superior to traditional methods for identifying and traversing bifurcation points in this work. A simply supported axial boundary condition is found to give a much more complex buckling response (bifurcation and limit-point buckling, as well as dynamic collapse) than specimens with a free axial edge (bifurcation or limit-point buckling). Experimental and numerical comparisons for the range of thicknesses considered indicate that elasticity of the in-plane boundary condition and transverse shear effects need further consideration. Observed shell damage is typical of that observed for composite plates. Results of this work give new insight into the response of composite fuselage panels to damaging transverse events, particularly in regard to instability/buckling behavior.