A simple, computationally efficient analysis approach is developed to predict the buckling of geodesically stiffened composite panels under in-plane loads. This procedure accounts for the discrete flexural contribution of each stiffener through the use of Lagrange multipliers in an energy method solution. An analysis is also implemented for the buckling of simply supported anisotropic rhombic plates. Examples are presented to verify results of the stability analyses and to demonstrate their convergence behavior. Analysis routines are coupled with a versatile numerical optimizer to create a package for the design of minimum-mass stiffened panels, subject to constraints on buckling of the panel assembly, local buckling of the stiffeners, and material strength failure. The design code is used to conduct a preliminary design study of structurally efficient stiffened aircraft wing rib panels. Design variables include thickness of the skin laminate, stiffener thickness, and stiffener height. Applied loads are uniaxial compression, pure shear, and combined compression-shear. Two different geodesically stiffened wing nib configurations with increasing numbers of stiffeners are considered. Results are presented in the form of structural efficiency curves and are compared with those for minimum-weight longitudinally stiffened panels and unstiffened flat plates. Trends in design parameters, including skin thickness and stiffener height, stiffener thickness, stiffener aspect ratio, stiffener load fraction, and stiffener mass fraction, are also examined for the geodesic panels under compression and shear. The effects of skin laminate geometry and anisotropy on the local buckling behavior of cross-stiffened geodesic panels are examined using the rhombic plate analysis. / Master of Science

This investigation was aimed at understanding the global and local strain and deflection responses of stiffened skins. Global deformations of the stiffened skins, under load, produce high local stresses in the interface region between the stiffener and skin. Test panels were designed to study the stiffener and skin interactions using parameters typical of stiffened skins for aircraft fuselages. A total of six panels were tested. Two skin laminates, both 0.04 in. thick, and three stiffener configurations were studied. The panels, having clamped edge boundary conditions, were subjected to pressure loads of up to 14.5-14.8 psi. Out-of-plane deflections and longitudinal and transverse strains were measured in several locations. The deflection responses showed a strongly nonlinear behavior at pressure loads of less than 5 psi. In addition relatively severe gradients of both longitudinal and transverse strains were measured in the interface region of the stiffener and skin. Finite element models incorporating geometric nonlinearities were made of four of the panels. Results of these models substantiated the overall findings of the experimental measurements. / Master of Science / incomplete_metadata

Linear, small deflection plate theory is used to study the stress at the contact axis and the deflection of an infinite plate caused by the impact of an axisymmetric cometary meteoroid. The analysis assumes that momentum exchange is the primary mechanism, that the time of exchange is instantaneous, and that the momentum of the meteoroid is negligible after impact. The stress at the origin is reduced to a single definite integral and the deflection to the Hankel inversion integral, both requiring definition of the particular projectile before further evaluation. A particular cometary meteoroid is mathematically represented in the analysis by its projected momentum per unit area onto the plate. The three specific shapes studies are the usual projectile shapes used in hypervelocity laboratories - cylinder, cone, and sphere - even though the analysis is not intended for the high-strength, high-density laboratory projectiles. Projectile comparisons based on equal mass, diameter, and total momentum indicate that frangible, low-strength cone projectiles cause significantly higher stresses and larger displacements of the plate at short times after impact than similar sphere and cylinder projectiles. / Master of Science

The increasing use of laminated composites in moving structures such as aircraft has led to a need for an efficient and accurate procedure for performing transient bending analysis of laminated composite plates. Classical theory is inadequate because it neglects transverse shear deformation, rotatory inertia, and geometric nonlinearities. In this thesis, a theory to account for transverse shear deformation and rotatory inertia is combined with the von Karman theory of geometric nonlinearities to develop the nonlinear governing equations of laminated composite plate bending. A finite element program is developed to solve these equations, using the Newmark direct integration technique to integrate the equations in time. Apparently, this constitutes the first transient finite-element analysis of laminated composite plate bending which accounts for transverse shear deformation, rotatory inertia, and geometric nonlinearities. The program accuracy is verified by comparison with results previously reported in the literature. Finally, results of a study of various material and plate geometry parameters are presented. The results of the parametric study show that transverse shear deformation, rotatory inertia, and geometric nonlinearity may all have a profound effect on the predicted bending response. In addition, the effects of material orthotropy, plate aspect ratio, plate thickness, lamination scheme, and load magnitude are shown to be significant. Computational constants such as the Newmark coefficients, the time-step size, and the element mesh are also investigated, and appropriate observations are made on the computational aspects of the program. / Master of Science

Results of an experimental and analytical study of the nonlinear response and failure characteristics of internally pressurized 4- to 16-ply-thick graphite-epoxy cylindrical panels are presented. Specimens with clamped boundaries simulating the skin between two frames and two stringers of a typical transport fuselage were tested to failure. Failure results of aluminum specimens are compared with the graphite-epoxy test results. The specimens failed at their edges where the local bending gradients and interlaminar stresses are maximum. STAGS nonlinear two-dimensional shell analysis computer code results are used to identify regions of the panels where the response is independent of the axial coordinate. A geometrically nonlinear one-dimensional cylindrical panel analysis was derived and used to determine panel response and interlaminar stresses. Inclusion of the geometric nonlinearity was essential for accurate prediction of panel response. Measurements of panel radius and edge circumferential displacements associated with specimen slipping were also required in the one-dimensional analysis for good correlation between analytical and experimental results. Some panels failed with significant damage in the form of tensile fiber breaks and ply delaminations preceding the ultimate pressure. Other panels failed suddenly without any apparent damage preceding the ultimate pressure. The failure usually occurred along one edge of the panel leaving the other edge intact. The damage on the panel surfaces and through-the-thickness were examined to determine the failure characteristics of the panels. Various failure criteria were applied to the stresses predicted from the one-dimensional analysis. The maximum stress failure criterion applied to the predicted tensile stress in the fiber direction agreed best with the experimentally determined first damage pressures. Results indicate that all panels tested would support applied internal pressures well above fuselage proof pressures. / Ph. D.

"June 2002" Includes bibliographical references (leaves 349-374)

"June 2002" / Includes bibliographical references (leaves 349-374) / xxxix, 416 leaves : ill., plates ; 30 cm. / Title page, contents and abstract only. The complete thesis in print form is available from the University Library. / Thesis (Ph.D.)--University of Adelaide, Dept. of Civil and Environmental Engineering, 2002

Indiana University-Purdue University Indianapolis (IUPUI) / This thesis proposes an automated grid-based alignment system utilizing lasers and an array of light-detecting photodiodes. The intent is to create an inexpensive and scalable alignment system for pick-and-place robotic systems. The system utilizes the transformation matrix, geometry, and trigonometry to determine the movements to align the robot with a grid-based array of photodiodes. The alignment system consists of a sending unit utilizing lasers, a receiving module consisting of photodiodes, a data acquisition unit, a computer-based control system, and the robot being aligned. The control system computes the robot movements needed to position the lasers based on the laser positions detected by the photodiodes. A transformation matrix converts movements from the coordinate system of the grid formed by the photodiodes to the coordinate system of the robot. The photodiode grid can detect a single laser spot and move it to any part of the grid, or it can detect up to four laser spots and use their relative positions to determine rotational misalignment of the robot. Testing the alignment consists of detecting the position of a single laser at individual points in a distinct pattern on the grid array of photodiodes, and running the entire alignment process multiple times starting with different misalignment cases. The first test provides a measure of the position detection accuracy of the system, while the second test demonstrates the alignment accuracy and repeatability of the system. The system detects the position of a single laser or multiple lasers by using a method similar to a center-of-gravity calculation. The intensity of each photodiode is multiplied by the X-position of that photodiode. The summed result from each photodiode intensity and position product is divided by the summed value of all of the photodiode intensities to get the X-position of the laser. The same thing is done with the Y-values to get the Y-position of the laser. Results show that with this method the system can read a single laser position value with a resolution of 0.1mm, and with a maximum X-error of 2.9mm and Y-error of 2.0mm. It takes approximately 1.5 seconds to process the reading. The alignment procedure calculates the initial misalignment between the robot and the grid of photodiodes by moving the robot to two distinct points along the robot’s X-axis so that only one laser is over the grid. Using these two detected points, a movement trajectory is generated to move that laser to the X = 0, Y = 0 position on the grid. In the process, this moves the other three lasers over the grid, allowing the system to detect the positions of four lasers and uses the positions to determine the rotational and translational offset needed to align the lasers to the grid of photodiodes. This step is run in a feedback loop to update the adjustment until it is within a permissible error value. The desired result for the complete alignment is a robot manipulator positioning within ±0.5mm along the X and Y-axes. The system shows a maximum error of 0.2mm in the X-direction and 0.5mm in the Y-direction with a run-time of approximately 4 to 5 minutes per alignment. If the permissible error value of the final alignment is tripled the alignment time goes down to 1 to 1.5 minutes and the maximum error goes up to 1.4mm in both the X and Y-directions. The run time of the alignment decreases because the system runs fewer alignment iterations.