A Comparison Study of Composite Laminated Plates with Holes under Tension A study was conducted to quantify the accuracy of numerical approximations to deem sufficiency in validating structural composite design, thus minimizing, or even eliminating the need for experimental test. Error values for stress and strain were compared between Finite Element Analysis (FEA) and analytical (Classical Laminated Plate Theory), and FEA and experimental tensile test for two composite plate designs under tension: a cross-ply composite plate design of [(0/90)4]s, and a quasi-isotropic layup design of [02/+45/-45/902]s, each with a single, centered hole of 1/8” diameter, and 1/4" diameter (four sets total). The intent of adding variability to the ply sequences and hole configurations was to gauge the sensitivity and confidence of the FEA results and to study whether introducing enough variability would, indeed, produce greater discrepancies between numerical and experimental results, thus necessitating a physical test. A shell element numerical approximation method through ABAQUS was used for the FEA. Mitsubishi Rayon Carbon Fiber and Composites (formerly Newport Composites) unidirectional pre-preg NCT301-2G150/108 was utilized for manufacturing—which was conducted and tested to conform to ASTM D3039/D3039M standards. A global seed size of 0.020, or a node count on the order of magnitude of 30,000 nodes per substrate, was utilized for its sub-3% error with efficiency in run-time. The average error rate for FEA strain from analytical strain at a point load of 1000lbf was 2%, while the FEA-to-experimental strains averaged an error of 4%; FEA-to-analytical and FEA-to-tensile test stress values at 1000lbf point load both averaged an error value of 6%. Suffice to say, many of these strain values were accurate up to ten-thousandths and hundred-thousandths of an in/in, and the larger stress/strain errors between FEA and test may have been attributed to the natural variables introduced from conducting a tensile test: strain gauge application methods, tolerance stacks from load cells and strain gauge readings. Despite the variables, it was determined that numerical analysis could, indeed, replace experimental testing. It was observed through this thesis that a denser, more intricate mesh design could provide a greater level of accuracy for numerical solutions, which proves the notion that if lower error rates were necessitated, continued research with a more powerful processor should be able to provide the granularity and accuracy in output that would further minimize error rates between FEA and experimental. Additionally, design margins and factors of safety would generally cover the error rates expected from numerical analysis. Future work may involve utilizing different types of pre-preg and further varied hole dimensions to better understand how the FEA correlates with analytical and tensile test results. Other load types, such as bending, may also provide insight into how these materials behave under loading, thus furthering the conversation of whether numerical approximations may one day replace testing all together.

The Cal Poly SLO Turbo-Prop Racer project aims to design a world speed record-breaking aircraft, capable of flying more than 550 miles per hour on a 3-kilometer closed course. To further this endeavor, this thesis presents the calculations of load distributions across the aircraft’s wing and tail and preliminary structural estimates of primary structural components for verification of the loads calculations and for use in a future finite element model. The aircraft’s fundamental design characteristics effect on the structure of the aircraft, namely the unique Y-tail design, are first examined. Then, loads are calculated in accordance with the regulation dictated by CS-23. Maneuvering loads, gust loads, ground loads, and engine loads are calculated through the Vortex-Lattice Method and CS-23 to provide input for detailed structural analysis. Structural thickness estimates are found using simplified analytical stress analysis. The wing and tail’s primary spars’ spar-caps and shear-webs, the wing and tail skins, and the rear fuselage are all calculated. The loads and thicknesses found are shown to be within order of reason and to support the fundamental design characteristics of the aircraft, pushing the project to continue toward its goal.

The demand for an insert on composite sandwich structures to aid in the arrestment of face-core delamination is of great need. This research studies the use of a damage arrestment device (DAD) that connects the carbon fiber face sheets to the foam core to find whether an increase in the structural integrity of the sandwich beam results. Experimental analysis was employed to test the samples and was verified by a theoretical and finite element approach. The mechanical properties of LTM45/CF1803 pre-impregnated carbon fiber and Last-A-foam FR 6710 polyvinylchloride foam were experimentally analyzed using ASTM D3039 and ASTM D1621 standards respectively to verify the manufacturer’s data for the given material. With all the mechanical data, the effects of adding DAD keys to a delaminated composite sandwich beam were studied under a four-point bending test using ASTM standard D6272 and compared with non-delaminated beams to see if an increase in ultimate strength could be achieved. The initial delamination in the beams under consideration was one inch in length and located in between the loaded span of the beam. Two control beams were utilized for comparison: one with no defects, and another with a one inch delamination introduced at the face-core interface. The DAD keys were added in two different configurations to potentially stop the delamination propagation and increase the ultimate strength. In the first configuration DAD keys were added 0.25 inches on either side of the initial delamination in the transverse direction and provided a significant increase in strength over the delaminated control beam. The second configuration had a DAD key running along the longitudinal axis of the sandwich beam and resulted in a significant increase in ultimate strength over the delaminated control beam. After testing ten successful samples for each of the six different configurations, it was concluded that the addition of DAD keys in both configurations significantly increased the structural integrity of both the delaminated and non-delaminated control beams. With all the experimental data acquired, finite element models were created in COSMOS. The purpose of the finite element analysis was to validate the experimental results by comparing the deflections of the beam subjected to four-point bending during the experiment to the deflections found numerically. The deflections for the various DAD key configurations found in the experimental work were in agreement with the finite element results.

This research studies the effects of a damage arrestment device embedded between a carbon fiber facesheet and foam core to find whether there is an increase in the structural integrity of the sandwich composites. Experimental and theoretical finite element analyses are implemented for two different composite sandwich geometries; plates and beams. Each structure consisted of the same loading criteria and was restricted to the same vibration fixture during the experiment. An accelerometer was placed on the composite plate to record the amplitude and the natural frequencies of the composite structure. Each composite specimen is then fixed to the surface of the Cal Poly Shake Table by two aluminum block fixtures. The mechanical properties of LTM45/CF1803 pre-impregnated carbon fiber and Last-A-foam FR 6710 polyvinylchloride foam were experimentally analyzed using ASTM D3039 and ASTM D1621 standards respectively to determine the material’s mechanical properties. By using the finite element program COSMOS with the pre-software GeoStar, accuracy representation were created to compare numerical, analytical, and theoretical results.

The structural design of flight vehicles is a balancing act between maximizing loading capability while minimizing weight. An engineer must consider not only the classical static structural yielding failure of a vehicle, but a variety of ways in which structural deformations can in turn, affect the loading conditions driving those deformations. Lift redistribution, divergence, and flutter are exactly such dynamic aeroelastic phenomena that must be properly characterized during the design of a vehicle; to do otherwise is to risk catastrophe. Relevant within the university context is the design of small-scale aircraft for student projects and of particular consideration, the DBF competition hosted by AIAA. This work implements a variety of aeroelastic analysis methods: K and PK with Theodorsen aerodynamics via Matlab, NASA EZASE, and the FEMAP NX NASTRAN Aeroelasticity Package. These techniques are applied to a number of baseline test cases in addition to two representative DBF wings. Both wings considered ultimately indicated stability within reasonable flight conditions, although each for a different reason. Analysis results for the Cal Poly 2020 wing, a spar-rib construction emblematic of the collocation design approach, showed that the wing was stable within expected flight regions. The USC 2020 wing model, a composite top spar construction, exhibited unstable behavior, however this was well outside the scope of expected flight conditions. The codebase developed as a part of this work will serve as a foundation for future student teams to perform aeroelastic analyses of their own and support continued aeroelastic research at Cal Poly - SLO.

Accurate mechanical analysis of composite structures is necessary for the prediction of laminate behavior. Cylindrical composite tubes are a mainstay in many structural applications. The fundamental design of circular composite cross-sections necessitates the development of a comprehensive composite lamination theory. A new analytical method is developed to characterize the behavior of thin-walled composite cylindrical tubes using a modified plate theory. A generated numerical solver can predict properties such as axial stiffness, bending stiffness, layer stresses, and layer strains in composite tubes subjected to combined mechanical loading and thermal effects. The model accounts for the curvature by transforming and translating the material in-plane lamina properties over a global reference coordinate system. A MATLAB-based solver is used to determine the lamina stiffness and stress outcomes with adjustable parameters, including elastic material properties, thermal coefficients, tubing radius, the orientation of fibers, and the ply stacking sequence. The results are then validated using a FE model developed in ABAQUS using a simple quadrature S4R element type. Parametric case studies confirm the validity of the analytical model by accounting for different ply orientations, stacking sequence, and thermal, mechanical loading.

The radiation environment in space is a threat that engineers and astronauts need to mitigate as exploration into the solar system expands. Passive shielding involves placing as much material between critical components and the radiation environment as possible. However, with mass and size budgets, it is important to select efficient materials to provide shielding. Currently, NASA and other space agencies plan on using water as a shield against radiation since it is already necessary for human missions. Water has been tested thoroughly and has been proven to be effective. Liquid propellants are needed for every mission and also share similar characteristics to water such as their density and hydrogenous composition. This thesis explores the shielding potentials of various liquid propellants as they compare to water for the purpose of providing an additional parameter when choosing propellants for any given mission. Testing propellants is done by first creating an experimental setup involving radioisotope sources Cs-137 and Co-60, a column of liquid with variable depths, and a Geiger-Mueller tube. Water and three other liquids: acetone, 70% isopropyl alcohol, and 12% hydrogen peroxide are physically tested and their linear attenuation coefficients are calculated. Then, the test setup is replicated in CERN’s Monte Carlo base radiation transport code, FLUKA. Although the calculated linear attenuation outputs from FLUKA are discrepant from experimental results by an average of 34%, they produce the same trends. FLUKA is used to expand upon experimental results by simulating a multitude of liquid propellants and comparing them all to water. FLUKA has the ability to simulate all propellants including hydrogen, oxygen, hydrazine, and dinitrogen tetroxide. Most of the tested propellants are found to have similar, to within 35%, gamma radiation linear attenuation coefficients as compared to water. The gamma radiation in this thesis’s experiment and simulations comes from Cs-137 and Co-60 radioisotope sources. For gamma radiation from the Co-60 source, liquid hydrogen provides 90% less attenuation than water and nitric acid and AF-M315E provide 35% and 38% more attenuation than water respectively. For gamma radiation emitted by Cs-137, liquid hydrogen, isopropyl alcohol, and methane have 90%, 35%, and 29% less attenuation than water respectively. Dinitrogen tetroxide, hydrogen peroxide, AF-M315E, and nitric acid have 34%, 41%, 46%, and 52% greater attenuation coefficients than water respectively. The liquids that are similar to water for the Cs-137 source have linear attenuation coefficients within 20% of water’s. Ultimately, most of the tested liquid propellants are shown to shield against radiation at a similar rate to water. Thus, an additional parameter for choosing liquid propellants on any given mission should be their radiation shielding capabilities.

The use of composite sandwich structures is rapidly increasing in the aerospace industry because of their increased strength-to-weight and stiffness-to-weight characteristics. The effects of low velocity impacts on these structures, however, are the main weakness that hinders further use of them in the industry because the damages from these loadings can often be catastrophic. Impact behavior of composite materials in general is a crucial consideration for a designer but can be difficult to describe theoretically. Because of this, experimental analysis is typically used to attempt to describe the behavior of composite sandwiches under impact loads. Experimental testing can still be unpredictable, however, because low velocity impacts can cause undetectable damage within the composites that weaken their structural integrity. This is an important issue with composite sandwich structures because interlaminar damage within the composite facesheets is typical with composites but the addition of a core material results in added failure modes. Because the core is typically a weaker material than the surrounding facesheet material, the core is easily damaged by the impact loads. The adhesion between the composite facesheets and the core material can also be a major region of concern for sandwich structures. Delamination of the facesheet from the core is a major issue when these structures are subjected to impact loads. This study investigated, through experimental and numerical analysis, how varying the core and facesheet material combination affected the flexural strength of a composite sandwich subjected to low velocity impact. Carbon, hemp, aramid, and glass fiber materials as facesheets combined with honeycomb and foam as core materials were considered. Three layers of the same composite material were laid on the top and bottom of the core material to form each sandwich structure. This resulted in eight different sandwich designs. The carbon fiber/honeycomb sandwiches were then combined with the aramid fiber facesheets, keeping the same three layer facesheet design, to form two hybrid sandwich designs. This was done to attempt to improve the impact resistance and post-impact strength characteristics of the carbon fiber sandwiches. The two and one layer aramid fiber laminates on these hybrid sandwiches were always laid up on the outside of the structure. The sandwiches were cured using a composite press set to the recommended curing cycle for the composite facesheet material. The hybrid sandwiches were cured twice for the two different facesheet materials. The cured specimens were then cut into 3 inch by 10 inch sandwiches and 2/3 of them were subjected to an impact from a 7.56 lbf crosshead which was dropped from a height of 38.15 inches above the bottom of the specimen using a Dynatup 8250 drop weight machine. The impacted specimen and the control specimen (1/3 of the specimens not subjected to an impact) were loaded in a four-point bend test per ASTM D7250 to determine the non-impacted and post-impact flexural strengths of these structures. Each sandwich was tested under two four-point bend loading conditions which resulted in two different extension values at the same 100 lbf loading value. The span between the two supports on the bottom of the sandwich was always 8 inches but the span between the two loading pins on the top of the sandwich changed between the two loading conditions. The 2/3 of the sandwiches that were tested after being impacted were subjected to bending loads in two different ways. Half of the specimens were subjected to four-point bending loads with the impact damage on the top facesheet (compressive surface) in between the loading pins; the other half were subjected to bending loads with the damage on the bottom facesheet (tensile surface). Theoretical failure mode analysis was done for each sandwich to understand the comparisons between predicted and experimental failures. A numerical investigation was, also, completed using Abaqus to verify the results of the experimental tests. Non-impacted and impacted four-point bending models were constructed and mid-span deflection values were collected for comparison with the experimental testing results. Experimental and numerical results showed that carbon fiber sandwiches were the best sandwich design for overall composite sandwich bending strength; however, post-impact strengths could greatly improve. The hybrid sandwich designs improved post-impact behavior but more than three facesheet layers are necessary for significant improvement. Hemp facesheet sandwiches showed the best post-impact bending characteristics of any sandwich despite having the largest impact damage sizes. Glass and aramid fiber facesheet sandwiches resisted impact the best but this resulted in premature delamination failures that limited the potential of these structures. Honeycomb core materials outperformed foam in terms of ultimate bending loads but post-impact strengths were better for foam cores. Decent agreement between numerical and experimental results was found but poor material quality and high error in material properties testing results brought about larger disagreements for some sandwich designs.

Hybrid carbon fiber reinforced composites are a new breed of materials that are currently being explored and characterized for next generation aerospace applications. Through the introduction of secondary reinforcements, such as alumina nanoparticles, hybrid properties including improved mechanical properties and stress sensing capabilities can be achieved. In order to maximize these properties, it is necessary to achieve a homogeneous dispersion of particulate filler. Utilizing the photoluminescent properties of alumina, it is possible to compare local levels of particle concentration through emission intensities as a way to determine the effectiveness of the tested manufacturing parameters in increasing material homogeneity. Parameters of these photoluminescence emissions have been established to be stress dependent, which allows for in situ residual stress measurements. It is shown here that the application of silane coupling agents as particle surface treatments improves particle dispersion when compared to untreated samples. Reactive silane coupling agent (RSCA) treatments were found to provide for greater dispersion improvements when compared to non-reactive silane coupling agents (NRSCA). Higher resolution investigations into these samples found that treatment with a reactive coupling agent altered the stress state of particles concentrated around the fiber from a tensile stress state to a compressive stress state. This is proposed to result from bonding of the reactive groups on the coupling agent to the organic groups on the carbon fibers which adjusts the stress state of the particle. Future mechanical tests will verify the effects of the particle surface functionalization treatments on mechanical properties of the composites.

In sub-millimeter sheets of wrought H-X750 Nickel-base superalloy, aluminum-rich coatings are bonded to matrix with a vapor phase aluminization process. If an appropriate amount of aluminum is bonded to matrix with homogenization treatment, the resulting diffusion couple will diffuse into coherent (g/g’) heterogeneous phases creating matrix that is both precipitation and solid solution strengthened. The diffusional mechanisms for solid solution mass transport involved with the growth and dispersion of bonded aluminum-rich coatings in the aluminization process only differ from the no external mass flow homogenization process with annealing treatment in that the boundary conditions are different. In each case these forces that activate diffusion at the macroscopic level are connected to the activation energies of random walks of atoms on a wide scale at the angstrom level. An overview of wrought Nickel-base superalloy is presented. Starting with thin sheets the alloy will be aluminized and homogenized. The research from this study will determine the parameters for the movement of the phase boundaries, mass transport, and the time variant concentration fields for both the aluminization and homogenization processes. This is predictable for both single dimension fluxes assuming the interdiffusivities and fluxes at the phase boundaries are known. Because mass-transport is related to the movement of the phase boundaries through density, an investigation into the less dense aluminum-rich coatings and resultant matrix is also included.