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Dissertations / Theses on the topic 'Aerospace structures'

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Published: 4 June 2021
Last updated: 23 February 2023

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1

Jenett, Benjamin (Benjamin Eric). "Digital material aerospace structures." Thesis, Massachusetts Institute of Technology, 2015. http://hdl.handle.net/1721.1/101837.

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Thesis: S.M., Massachusetts Institute of Technology, Department of Civil and Environmental Engineering, 2015.
Cataloged from PDF version of thesis.
Includes bibliographical references (pages 71-76).
This thesis explores the design, fabrication, and performance of digital materials in aerospace structures in three areas: (1) a morphing wing design that adjusts its form to respond to different behavioral requirements; (2) an automated assembly method for truss column structures; and (3) an analysis of the payload and structural performance requirements of space structure elements made from digital materials. Aerospace structures are among the most difficult to design, engineer, and manufacture. Digital materials are discrete building block parts, reversibly joined, with a discrete set of positions and orientations. Aerospace structures built from digital materials have high performance characteristics that can surpass current technology, while also offering potential for analysis simplification and assembly automation. First, this thesis presents a novel approach for the design, analysis, and manufacturing of composite aerostructures through the use of digital materials. This approach can be used to create morphing wing structures with customizable structural properties, and the simplified composite fabrication strategy results in rapid manufacturing time with future potential for automation. The presented approach combines aircraft structure with morphing technology to accomplish tuned global deformation with a single degree of freedom actuator. Guidelines are proposed to design a digital material morphing wing, a prototype is manufactured and assembled, and preliminary experimental wind tunnel testing is conducted. Seconds, automatic deployment of structures has been a focus of much academic and industrial work on infrastructure applications and robotics in general. This thesis presents a robotic truss assembler designed for space applications - the Space Robot Universal Truss System (SpRoUTS) - that reversibly assembles a truss column from a feedstock of flat-packed components, by folding the sides of each component up and locking onto the assembled structure. The thesis describes the design and implementation of the robot and shows that an assembled truss compares favorably with prior truss deployment systems. Thirds, space structures are limited by launch shroud mass and volume constraints. Digital material space structures can be reversibly assembled on orbit by autonomous relative robots using discrete, incremental parts. This will enable the on-orbit assembly of larger space structures than currently possible. The engineering of these structures, from macro scale to discrete part scale, is presented. Comparison with traditional structural elements is shown and favorable mechanical performance as well as the ability to efficiently transport the material in a medium to heavy launch vehicle. In summary, this thesis contributes the methodology and evaluation of novel applications of digital materials in aerospace structures.
by Benjamin Jenett.
S.M.
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2

Spendley, Paul R. "Design allowables for composite aerospace structures." Thesis, University of Surrey, 2012. http://epubs.surrey.ac.uk/810072/.

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Recent developments in aircraft design have seen the Airbus A380 and the Boeing Dreamliner employ significant amounts of advanced composite materials. There is some thought however, und the motivation for this current work, that these materials continue to suffer a weight penalty. In this work tests required to generate design allowables which accommodate environmental effects and holes arc performed on Carbon/epoxy quasi-isotropic laminatcs. The test data is treated statistically to provide B-basis allowables for each specimen type and condition. It was seen that the notched specimens (coupons containing a centrally placed through hole) displayed significantly less scatter in strength than unnotched specimens. This is significant when considering the widespread use of deterministic knock-down factors as an alternative route to obtain design allowables which accommodate environmental effects and/or holes. This results in an over-conservative design allowable being employed in subsequent structural design calculations. The possibility for using notched coupons to determine design allowables was explored using the COG (Critical Damage Growth) model. This showed that. given two of the three parameters. the unnotched and notched strength, and fracture toughness the variation in strengths could be reasonable predicted. This leads to a more representative design allowable by maintaining the statistical nature of the B-basis allowable. During the statistical treatment of the test data it was also seen that although current aerospace guidelines recommend a particular distribution model (i.e. the Wcibull distribution) this can also leads to an artificially reduced design allowable. These findings suggest that the use of notched specimens can lead to a reduced development test programme and reduced structural weight.
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3

Hanuska, Alexander Robert Jr. "Thermal Characterization of Complex Aerospace Structures." Thesis, Virginia Tech, 1998. http://hdl.handle.net/10919/36617.

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Predicting the performance of complex structures exposed to harsh thermal environments is a crucial issue in many of today's aerospace and space designs. To predict the thermal stresses a structure might be exposed to, the thermal properties of the independent materials used in the design of the structure need to be known. Therefore, a noninvasive estimation procedure involving Genetic Algorithms was developed to determine the various thermal properties needed to adequately model the Outer Wing Subcomponent (OWS), a structure located at the trailing edge of the High Speed Civil Transport's (HSCT) wing tip. Due to the nature of the nonlinear least-squares estimation method used in this study, both theoretical and experimental temperature histories were required. Several one-dimensional and two-dimensional finite element models of the OWS were developed to compute the transient theoretical temperature histories. The experimental data were obtained from optimized experiments that were run at various surrounding temperature settings to investigate the temperature dependence of the estimated properties. An experimental optimization was performed to provide the most accurate estimates and reduce the confidence intervals. The simultaneous estimation of eight thermal properties, including the volumetric heat capacities and out-of-plane thermal conductivities of the facesheets, the honeycomb, the skins, and the torque tubes, was successfully completed with the one-dimensional model and the results used to evaluate the remaining in-plane thermal conductivities of the facesheets, the honeycomb, the skins, and the torque tubes with the two-dimensional model. Although experimental optimization did not eliminate all correlation between the parameters, the minimization procedure based on the Genetic Algorithm performed extremely well, despite the high degree of correlation and low sensitivity of many of the parameters.
Master of Science
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4

White, Caleb, and caleb white@rmit edu au. "Health Monitoring of Bonded Composite Aerospace Structures." RMIT University. Aerospace, Mechanical and Manufacturing Engineering, 2009. http://adt.lib.rmit.edu.au/adt/public/adt-VIT20090602.142122.

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Airframe assemblers have long recognised that for a new aircraft to be successful it must use less fuel, have lower maintenance requirements, and be more affordable. One common tactic is the use of innovative materials, such as advanced composites. Composite materials are suited to structural connection by adhesive bonding, which minimises the need for inefficient mechanical fastening. The aim of this PhD project was to investigate the application of existing, yet immature Structural Health Monitoring (SHM) techniques to adhesively bonded composite aerospace structures. The PhD study focused on two emerging SHM technologies - frequency response and comparative vacuum monitoring (CVM). This project aimed to provide missing critical information for each technique. This included determining sensitivity to damage, repeatability of results, and operating limitations for the frequency response method. Study of the CVM technique aimed to address effectiveness of damage detection, manufacture of sensor cavities, and the influence of sensor integration on mechanical performance of bonded structures. Experimental research work is presented examining the potential of frequency response techniques for the detection of debonding in composite-to-composite external patch repairs. Natural frequencies were found to decrease over a discrete frequency range as the debond size increased; confirming that such features could be used to both detect and characterise damage. The effectiveness of the frequency response technique was then confirmed for composite patch and scarf repair specimens for free-free and fixed-fixed boundary conditions. Finally, the viability of the frequency response technique was assessed for a scarf repair of a real aircraft component, where it was found that structural damping limited the maximum useable frequency. The feasibility of CVM technique for the inspection of co-cured stiffener-skin aircraft structures was explored. The creation of sensor cavities with tapered mandrels was found to significantly alter the microstructure of the stiffener, including crimping and waviness of fibres and resin-rich zones between plies. Representative stiffened-skin structure with two sensor cavity configurations (parallel and perpendicular to the stiffener direction) was tested to failure in tension and compression. While tensile failure strength was significantly reduced for both configurations (up to 25%), no appreciable differences in compression properties were found. Two potential sensor cavity configurations were investigated for the extension of the CVM technique to pre-cured and co-bonded scarf repair schemes. The creation of radial and circumferential CVM sensor cavities was found to significantly alter the microstructure of the adhesive bond-line and the architecture of the repair material in the case of the co-bonded repair. These alterations changed the failure mode and reduced the tensile failure strength of the repair. A fibre straightening mechanism responsible for progressive failure (specific to co-bonded repairs with circumferential cavities) was identified, and subsequently supported with acoustic emission testing and numerical analysis. While fatigue performance was generally reduced by the presence of CVM cavities, the circumferential cavities appeared to retard crack progression, reducing sensitivity to the accumulation of fatigue damage. These outcomes have brought forward the implementation of SHM in bonded composite structures, which has great potential to improve the operating efficiency of next generation aircraft.
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5

Zhang, Haochuan. "Nonlinear aeroelastic effects in damaged composite aerospace structures." Thesis, Georgia Institute of Technology, 1996. http://hdl.handle.net/1853/12150.

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6

Navarro, Zafra Joaquin. "Computational mechanics of fracture on advanced aerospace structures." Thesis, University of Sheffield, 2016. http://etheses.whiterose.ac.uk/16883/.

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In this thesis, the computational simulation of cracks in advanced composite structures subjected to biaxial loading is studied. A structural integrity analysis using the eXtended Finite Element Method (XFEM) is considered for simulating the crack behaviour of a chopped fibre-glass-reinforced polyester (CGRP) cruciform specimen subjected to a quasi-static tensile biaxial loading [99]. This is the first time this problem is accomplished for computing the stress intensity factors (SIFs) produced in the biaxially loaded area of the cruciform specimen. SIFs are calculated for infinite plates under biaxial loading as well as for the CGRP cruciform specimens in order to review the possible edge effects. A new ratio relating the side of the central zone of the cruciform and the crack length is proposed. Additionally, the initiation and evolution of a three-dimensional crack are successfully simulated. Specific challenges such as the 3D crack initiation, based on a principal stress criterion, and its front propagation, in perpendicular to the principal stress direction, are conveniently addressed. No initial crack location is pre-defined and an unique crack is developed. A three-dimensional progressive damage model (PDM) is implemented within a CGRP cruciform structure for modelling its damage under loading [100]. In order to simulate the computational behaviour of the composite, the constitutive model considers an initial elastic behaviour followed by strain-softening. The initiation criterion defined is based on the maximum principal stress of the composite and once this criterion is satisfied, stiffness degradation starts. For the computation of damage, the influence of the fibre and the matrix are taken into account within the damage rule. This is the first time a three-dimensional PDM is implemented into a composite cruciform structure subjected to biaxial loading. A new approach for dynamic analysis of stationary cracks using XFEM is derived. This approach is capable of addressing dynamic and static fracture mechanics problems. Additionally, by means of this relatively simple approach, it is possible to address correctly the crack pattern of the 10 degrees off-axis laminate manufactured solving the limitation observed with progressive damage modelling. During the whole thesis, the computational outcomes have been validated by means of comparison with theoretical and experimental results.
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7

Lam, Daniel F. "STRAIN CONCENTRATION AND TENSION DOMINATED STIFFENED AEROSPACE STRUCTURES." University of Akron / OhioLINK, 2006. http://rave.ohiolink.edu/etdc/view?acc_num=akron1145393262.

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8

Vishwanathan, Aditya. "Uncertainty Quantification for Topology Optimisation of Aerospace Structures." Thesis, University of Sydney, 2020. https://hdl.handle.net/2123/23922.

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The design and optimisation of aerospace structures is non-trivial. There are several reasons for this including, but not limited to, (1) complex problem instances (multiple objectives, constraints, loads, and boundary conditions), (2) the use of high fidelity meshes which impose significant computational burden, and (3) dealing with uncertainties in the engineering modelling. The last few decades have seen a considerable increase in research output dedicated to solving these problems, and yet the majority of papers neglect the effect of uncertainties and assume deterministic conditions. This is particularly the case for topology optimisation - a promising method for aerospace design that has seen relatively little practical application to date. This thesis will address notable gaps in the topology optimisation under uncertainty literature. Firstly, an observation underpinning the field of uncertainty quantification (UQ) is the lack of experimental studies and dealing with non-parametric variability (e.g. model unknowns, experimental and human errors etc.). Random Matrix Theory (RMT) is a method explored heavily in this thesis for the purpose of numerical and experimental UQ of aerospace structures for both parametric and non-parametric uncertainties. Next, a novel algorithm is developed using RMT to increase the efficiency of Reliability-Based topology optimisation, a formulation which has historically been limited by computational runtime. This thesis also provides contributions to Robust Topology optimisation (RTO) by integrating uncertain boundary conditions and providing experimental validation of the results. The final chapter of this thesis addresses uncertainties in multi-objective topology optimisation (MOTO), and also considers treating a single objective RTO problem as a MOTO to provide a more consistent distribution of solutions.
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9

Pozegic, Thomas R. "Nano-modified carbon-epoxy composite structures for aerospace applications." Thesis, University of Surrey, 2016. http://epubs.surrey.ac.uk/809603/.

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Carbon fibre reinforced plastics (CFRP) have revolutionised industries that demand high specific strength materials. With current advancements in nanotechnology there exists an opportunity to not only improve the mechanical performance of CFRP, but to also impart other functionalities, such as thermal and electrical conductivity, with the aim of reducing the reliance on metals, making CFRP attractive to many other industries. This thesis provides a comprehensive analysis of the nano-phase modification to CFRP by growing carbon nanotubes (CNTs) on carbon fibre (CF) and performing mechanical, electrical and thermal conductivity tests, with comparisons made against standard CFRP. Typical CFs are coated with a polymer sizing that plays a vital role in the mechanical performance of the composite, but as a consequence of CNT growth, it is removed. Therefore, in addition, an ‘intermediate’ composite was fabricated – based on CFs without a polymer sizing – which enabled a greater understanding of how the mechanical properties and processability of the material responds to the CNT modification. A water-cooled chemical vapour deposition system was employed for CNT growth and infused into a composite structure with an industrially relevant vacuum-assisted resin transfer moulding (VARTM) process. High quality CNTs were grown on the CF, resulting in properties not reported to date, such as strong intra-tow binding, leading to the possibility of a polymer sizing-free CFRP. A diverse set of spectroscopic, microscopic and thermal measurements were carried out to aid understanding for this CNT modification. Subsequent electrical conductivity tests performed in three directions showed 300%, 230% and 450% improvements in the ‘surface’, ‘through-thickness’ and ‘through-volume’ directions, for the CNT modified CFRP, respectively. In addition, thermal conductivity measurements performed in the through-thickness direction also gave improvements in excess of 98%, boding well for multifunctional applications of this hybrid material concept. A range of mechanical tests were performed to monitor the effect of the CNT modification, including: single fibre tensile tests, tow pull-out tests (from the polymer matrix), composite tensile tests, in-plane shear tests and interlaminar toughness tests. Single fibre tensile tests demonstrated a performance reduction of only 9.7% after subjecting the fibre to the low temperature CNT growth process, which is significantly smaller than previous reports. A reduction in tensile performance was observed in the composite tensile test however, with a reduction of 33% reduction in the ultimate tensile strength, but a 146% increase in the Young’s modulus suggests that the CNTs may have improved the interfacial interactions between the fibre and the polymer matrix. To support this, improvements of 20% in the in-plane shear stress and 74% and the shear chord modulus, were recorded. Negligible differences were observed using a pull-out test to directly measure the interfacial strength as a consequence of the inherently difficult mechanical test procedure. The fracture toughness was tested under mode-I loading of a double cantilever beam configuration and improvements of 83% for CNT modified composite alluded to CNT pull-out fracture mechanism and crack propagation amongst the microstructures. The changes in the physical properties are correlated to the microstructure modifications ensured by the low temperature CNT growth on the CF substrates used in the CFRP composites. This allows for a new generation of modified multifunctional CFRPs to be produced.
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10

Sebastian, Christopher. "Towards the validation of thermoacoustic modelling in aerospace structures." Thesis, University of Liverpool, 2015. http://livrepository.liverpool.ac.uk/2012079/.

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The research presented in this thesis has been performed over the course of three years under funding from the European Office of the United States Air Force (EAORD) as a part of a long-term project to collect high quality data for the validation of computational mechanics models of thermoacoustic loading. The focus is on the adaptation of stereoscopic (3D) Digital Image Correlation for use in a combined thermal and high temperature measurements. To that end, a background is provided which highlights the current state of the art in high temperature, vibration experiments and data acquisition. A system is described in which a pulsed laser of duration 4 nanoseconds is used to capture high-quality displacement and strain data from vibrating components (PL- DIC). Based on this a novel method of capturing data from a component subjected to random excitation was developed. A laser vibrometer was used along with a custom LabVIEW program to trigger the pulsed laser relative to points of maximum velocity in the components vibration cycle. A dynamic calibration procedure was performed of both a high speed DIC system and the Pulsed-Laser DIC system to assess and compare the measurement uncertainty from the respective systems. It is crucial to know the uncertainty in experimental data when using it for the validation of computational models. A new way to validate computational models of vibration behavior using full-field DIC data and image decomposition is described. This is a phasic approach in which data from the entire cycle of vibration is used. The validation assessment is performed using the expanded uncertainty calculated and a concordance correlation coefficient. An example is provided using an aerospace component to validate four different simulation conditions of a modal frequency response model. An apparatus was designed and built which uses a 10 kW array of quartz lamps to reproduce some aspects of the heating provided by the Air Force test chambers. Experiments were performed in collaboration with the University of Illinois using induction heating and a small Hastelloy plate. A thermal buckling phenomena was observed using the PL-DIC system, the first full-field results of such.
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11

Zhou, Jin. "The energy-absorbing behaviour of novel aerospace composite structures." Thesis, University of Liverpool, 2015. http://livrepository.liverpool.ac.uk/2014139/.

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The aim of this research is to investigate the structural response of PVC foam based sandwich structures, composite reinforced foam cores and fibre metal laminates (FMLs) subjected to quasi-static and dynamic loading conditions. It also includes the investigation of the mechanical properties and energy-absorbing characteristics of the novel hybrid materials and structures for their potential use in aerospace and a wide range of engineering applications. Firstly，a series of experimental tests have been undertaken to obtain the mechanical properties of all constituent materials and structural behavior of the composite structures, which are used to develop and validate numerical models. The material tests carried out include (1) tension properties of composite laminates and aluminium alloys, (2) compression of PVC foams, carbon and glass fibre rods and tubes, and fibre metal laminates in the edge wise and flat wise, (3) shear and bending of PVC foams, (4) Hopkinson Bar, (5) quasi-static and dynamic crushing of composite reinforced foams, and (6) projectile impact on fibre reinforced laminates, aluminium alloy panels, PVC foam based sandwich panels and fibre metal laminates. The corresponding failure modes are obtained to validate the numerical predictions. In addition, perforation energy and specific energy absorptions of various composite structures investigated are evaluated. Moreover, the rate-sensitivity of FMLs based on glass fibre reinforced epoxy and three aluminium alloys has been investigated though a series of quasi-static and impact perforation tests on multilayer configurations ranging from a simple 2/1 lay-up to a 5/4 stacking sequence. FMLs based on a combination of the composite and metal constituents exhibit a low degree of rate-sensitivity, with the impact perforation energy increasing slightly in passing from quasi-static to dynamic rates of loading. Then, finite element (FE) models are developed using the commercial code Abaqus/Explicit to simulate the impact response of PVC foam sandwich structures. The agreement between the numerical predictions and the experimental results is very good across the range of the structures and configurations investigated. The FE models have produced accurate predictions of the impact load-displacement responses, the perforation energies and the failure characteristics recorded. The analyses are used to estimate the energy absorbed by the skins and the core during the perforation process. The validated FE models are also used to investigate the effect of oblique loading and to study the impact response of sandwich panels on an aqueous environment and subjected to a pressure differential (equivalent to flying at an altitude of 10000 m). The modelling has been further undertaken on the low velocity impact response of the sandwich structures based on graded or composite reinforced PVC foam cores, with reasonably good correlation to the corresponding experimental results. Consequently, a series of finite element analyses have been conducted to investigate the influence of varying foam density, rod diameter, rod length and fibre type on the energy-absorbing characteristics of the reinforced foams. Perforation energies, impact resistance performance and unit cost of the structures have been evaluated. Furthermore, the low velocity impact response of fibre metal laminates has been studied numerically. Here, the composite layer in FMLs is modelled using the modified 3D Hashin’s failure criteria, which are implemented into the main programme through a user-defined subroutine, whilst aluminium alloys are modelled using Johnson-Cook plasticity and the corresponding damage criterion. A large number of simulations have been undertaken to cover FMLs with all stacking sequences and alloy types studied, which are compared with the experimental results in terms of the load-displacement trace and failure modes, with very good correlation. Similar modelling work has been carried out on the aluminium layer and composite layer individually. The energy to perforate the various FMLs is plotted and fitted on a single curve that can be used to predict the perforation energies of other configurations. The dynamic characteristics of the composite structures through a series experimental tests and numerical predictions investigated in this project can be used in the design of lightweight composite structures for energy-absorbing applications.
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12

McCrory, John. "Advanced Acoustic Emission (AE) monitoring techniques for aerospace structures." Thesis, Cardiff University, 2016. http://orca.cf.ac.uk/89212/.

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This thesis contains the development of advanced Acoustic Emission (AE) monitoring techniques for aerospace structures. The techniques developed in this work explore AE’s ability to detect, locate and characterise signals. Experimental studies were conducted on a range of structures made from typical aerospace materials, including carbon fibre composite, GLARE and high grade steel; and the data collected from these studies was processed using the newly developed AE techniques, in order to determine their effectiveness. The work was divided into three main areas of research: 1. AE Source Location A location test was conducted on a GLARE fuselage panel specimen with complex geometric features in order to test the effect that altering the training grid resolution has on the accuracy of the delta-T mapping location technique. Delta-T mapping yielded more accurate results than the conventional Time of Arrival (TOA) method and the development of this technique formed the basis from which AE signals could be confidently located. 2. Damage Identification A fatigue test was conducted on a pre-notched, 300M grade steel, cantilevered beam which was monitored using both AE and Digital Image Correlation (DIC) during loading. The work considered the detection and tracking of fatigue crack growth. A novel form of data acquisition and analysis called an Additive Hits Analysis (AHA) was proposed and developed. The AHA provided a similar result to a conventional wavestreaming approach which was also used, though it did so in a much more streamlined manner. Specific delta-T mapping located signals were used to determine the frequency bands of interest for the cracking process to be tracked. DIC was noted as being a useful tool for validation of AE testing. 3. Characterisation on Large Scale Specimen A buckling test was conducted on a large-scale carbon fibre composite specimen which was monitored using AE and DIC. The work focused on the detection, location and characterisation of signals occurring in the specimen due to the applied loading. An ultrasonic C-scanner was used to quantify the damage which occurred in the specimen and this was found to be a useful tool for validation. A novel form of the modal analysis technique Measured Amplitude Ratio (MAR) called Automated Corrected MAR was developed. The new method was found to be able to successfully distinguish between in and out-of-plane signals arising in the specimen during the test whilst also providing time saving benefits over conventional methods. The combination of delta-T mapping with the Automated Corrected MAR results proved useful to the analysis.
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Jefferson, Thomas G. "Reconfigurable assembly system design methodology for aerospace wing structures." Thesis, University of Nottingham, 2017. http://eprints.nottingham.ac.uk/42778/.

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The aerospace industry is facing new challenges to meet burgeoning customer demand. An unprecedented number of orders for commercial aircraft has placed great urgency on aerospace manufacturers to make gains in production efficiency. Wing assembly is one such area where cycle times are in the order of hundreds of hours and commissioning cells can take several years which has led to a significant order backlog. In light of these challenges, new techniques are required to bring about greater agility to respond to market changes. Aerospace manufacturers must seize the opportunity to innovate and readdress approaches to ensure their prosperity. Recent research advocates Reconfigurable Assembly Systems (RAS) as a viable solution. A RAS is designed at the outset to change in structure to modify production capacity and functionality to meet new requirements. Yet, adding reconfigurability further increases design complexity. Despite the increased complexity, few formal methodologies exist to support RAS design for aerostructures. A novel RAS design methodology is presented to address the design complexity and the specific needs of Airbus. The methodology is a systems design approach consisting of reconfigurability principles, Axiomatic Design and Design Structure Matrices. Customer needs and existing knowledge are used to systematically specify scalable and customisable functionalities from the outset. These requirements and constraints are then translated into physical system designs modelled using CATIA, a 3D modelling software suite. The design methodology is applied in two case studies for wing assembly to produce full-scale factory designs. The designs are compared with current Airbus baselines in production ramp-up scenarios. The RAS demonstrate capability to change in structure for rapid increase in capacity at comparable cost to fixed systems. Greater capacity and shortened ramp-up time evidently reduces backlog compared to current systems. The first case study focused on technical development of a RAS and found potential ramp-up reduction of 88% and 10% less Capital Expenditure (CapEx) over 10 years. The second case study for a current wing scenario found reductions of 50% to ramp-up and 41% less tooling CapEx compared to a pulse line for a 12-year production cycle. The designs and scenarios were validated in formal Airbus design reviews. The case studies present the first instances of production-scale RAS for aerostructures. The RAS designs are made possible by designing from the outset using a novel design methodology which sets a precedent for the future of aerostructure assembly and opens up new possibilities for future research.
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El-Nounu, AbdulRahman. "Redesign methodology for cost effective assembly of aerospace structures." Thesis, University of Nottingham, 2018. http://eprints.nottingham.ac.uk/51634/.

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The research addresses the topic of design for assembly from an aerospace structures perspective. Aerospace assembly has traditionally taken secondary important to aircraft performance. This approach has been validated through healthy sales, most recently demonstrated in the Single Aisle market. However, in recent times, design for assembly has become more important. There are two main drivers behind this shift in focus. The first is a desire from aircraft manufacturers to maximise profits on existing aircraft orders through redesign. The second is the future outlook of aircraft sales, estimated to be in the trillions of dollars 2035. Aircraft manufacturers have therefore recognised that optimising their manufacturing system is critical lest market share is lost to emerging aircraft manufacturers through an inability to meet rising demand. Three methods are then developed to provide design for assembly indicators for development decisions. The underpinning methodology behind these methods is a data driven approach. This is that cost saving decisions can be made using the mass of existing available data from production systems at early stages provided that key indicators are identified. The methods allow engineers to make informed decisions on design for assembly and technology development. The first method addresses the issue of redesign. A tool is presented that relies on available data of assembly processes to make recommendations on redesign projects. The method is populated with real data and its output is compared against real business decisions. The results show that the method provides positive direction and is beneficial when filtering between costly redesign projects. The second method addresses design for assembly at early product development. A complexity metric is developed using a combination of historical data and known data at a particular development stage to produce a complexity metric that carries out an analysis of a full assembly system. It provides the engineer with a macro view enabling the identification of potential bottlenecks. Data from a previous product was used to demonstrate this method. The results shown were able to highlight real issues and make recommendations about technology strategy. The final method developed in this research recognised that design for assembly and assembly technology were synergetic and should be developed together. It proposed an assembly process characterisation technique to enable future technology strategy planning at design for assembly stage. The tool was demonstrated using existing data and proposed several concepts for a future product to enable higher levels of automation and more cost effective future technology implementation. The research concluded that there was a definite advantage in using the demonstrated methods in providing direction to an aircraft manufacturing business. In the redesign method and the complexity analysis method this was validated through comparison against real business decisions. The two methods were in line with business thinking. Also, where the redesign method was different in its advice compared with business direction, it was shown that following the advice of the method would have been beneficial to the business. It was more difficult to validate the shared platform approach method due to its results providing indicators for future decisions. Early analysis into its potential validity through technology benchmarking looked promising.
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Lim, Sang Seok. "Modeling and control of large space structures." Thesis, University of Ottawa (Canada), 1990. http://hdl.handle.net/10393/5822.

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In this thesis a preliminary formulation of large space structures and their stabilization is considered. The system consists of a (rigid) massive body and flexible configurations which consist of several beams, forming the space structure. The rigid body is located at the center of the space structure and may play the role of experimental modules. A complete dynamics of the system has been developed using Hamilton's principle. Euler-Bernoulli, Rayleigh and Timoshenko beam theories are utilized to derive the dynamic equations governing the vibration of the beams. The equations that govern the motion of the complete system consist of six ordinary differential equations and several partial differential equations together with appropriate boundary conditions. The partial differential equations govern the vibration of flexible components. The ordinary differential equations describe the rotational and translational motion of the central body. The dynamics indicate very strong interaction among rigid body translation, rigid body rotation and vibrations of flexible members through nonlinear couplings. Hence any rotation of the rigid body induces vibration in the beams and vice-versa. Also any disturbance in the orbit induces vibration in the beams and wobbles in the body rotation and vice-versa. This makes the system performance unsatisfactory for many practical applications. In this thesis stabilization of the above mentioned system subject to external disturbances is considered. The asymptotic stability of the perturbed system by applying several types of stabilizing controls such as proportional controls, deadzone controls or saturation controls is proved using Lyapunov's method. Numerical simulations are carried out in order to illustrate the impact of dynamic coupling or interaction among several members of the system and the effectiveness of the suggested feedback controls for stabilization. Stability of a spacecraft and a space station under the influence of slew maneuvering is numerically investigated.
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Orifici, Adrian Cirino, and adrian orifici@student rmit edu au. "Degradation Models for the Collapse Analysis of Composite Aerospace Structures." RMIT University. Aerospace, Mechanical and Manufacturing Engineering, 2007. http://adt.lib.rmit.edu.au/adt/public/adt-VIT20080619.090039.

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17

Qu, Shuang. "Multilevel optimisation of aerospace and lightweight structures incorporating postbuckling effects." Thesis, Cardiff University, 2011. http://orca.cf.ac.uk/55080/.

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The optimisation of aerospace structures is a very complex problem, due to the hundreds of design variables a multidisciplinary optimisation may contain, so that multilevel optimisation is required. This thesis presents the recent developments to the multilevel optimisation software VICONOPT MLO, which is a multilevel optimisation interface between the well established analysis and design software packages VICONOPT and MSC/NASTRAN. The software developed is called VICONOPT MLOP (Multilevel Optimisation with Postbuckling), and allows for postbuckling behaviour, using analysis based on the Wittrick-Williams algorithm. The objective of this research is to enable a more detailed insight into the multilevel optimisation and postbuckling behaviour of a complex structure. In VICONOPT MLOP optimisation problems, individual panels of the structural model are allowed to buckle before the design load is reached. These panels continue to carry load with differing levels of reduced stiffness. VICONOPT MLOP creates new MSC/NASTRAN data files based on this reduced stiffness data and iterates through analysis cycles to converge on an appropriate load re-distribution. Once load convergence has been obtained with an appropriate criterion, the converged load distribution is used as a starting point in the optimisation of the constituent panels, i.e. a new design cycle is started, in which the updated ply thicknesses for each panel are calculated by VICONOPT and returned to MSC/NASTRAN through VICONOPT MLOP. Further finite element analysis of the whole structure is then carried out to determine the new stress distributions in each panel. The whole process is repeated until a mass convergence criterion is met. A detailed overview of the functionality of VICONOPT MLOP is presented in the thesis. A case study is conducted into the multilevel optimisation of a composite aircraft wing, to demonstrate the capabilities of VICONOPT MLOP and identify areas for future studies. The results of the case study show substantial mass savings, proving the software's capabilities when dealing with such problems. The time taken for this multilevel optimisation also proves the efficiency of the software.
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18

Halbert, Keith. "Estimation of probability of failure for damage-tolerant aerospace structures." Thesis, Temple University, 2014. http://pqdtopen.proquest.com/#viewpdf?dispub=3623167.

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The majority of aircraft structures are designed to be damage-tolerant such that safe operation can continue in the presence of minor damage. It is necessary to schedule inspections so that minor damage can be found and repaired. It is generally not possible to perform structural inspections prior to every flight. The scheduling is traditionally accomplished through a deterministic set of methods referred to as Damage Tolerance Analysis (DTA). DTA has proven to produce safe aircraft but does not provide estimates of the probability of failure of future flights or the probability of repair of future inspections. Without these estimates maintenance costs cannot be accurately predicted. Also, estimation of failure probabilities is now a regulatory requirement for some aircraft.

The set of methods concerned with the probabilistic formulation of this problem are collectively referred to as Probabilistic Damage Tolerance Analysis (PDTA). The goal of PDTA is to control the failure probability while holding maintenance costs to a reasonable level. This work focuses specifically on PDTA for fatigue cracking of metallic aircraft structures. The growth of a crack (or cracks) must be modeled using all available data and engineering knowledge. The length of a crack can be assessed only indirectly through evidence such as non-destructive inspection results, failures or lack of failures, and the observed severity of usage of the structure.

The current set of industry PDTA tools are lacking in several ways: they may in some cases yield poor estimates of failure probabilities, they cannot realistically represent the variety of possible failure and maintenance scenarios, and they do not allow for model updates which incorporate observed evidence. A PDTA modeling methodology must be flexible enough to estimate accurately the failure and repair probabilities under a variety of maintenance scenarios, and be capable of incorporating observed evidence as it becomes available.

This dissertation describes and develops new PDTA methodologies that directly address the deficiencies of the currently used tools. The new methods are implemented as a free, publicly licensed and open source R software package that can be downloaded from the Comprehensive R Archive Network. The tools consist of two main components. First, an explicit (and expensive) Monte Carlo approach is presented which simulates the life of an aircraft structural component flight-by-flight. This straightforward MC routine can be used to provide defensible estimates of the failure probabilities for future flights and repair probabilities for future inspections under a variety of failure and maintenance scenarios. This routine is intended to provide baseline estimates against which to compare the results of other, more efficient approaches.

Second, an original approach is described which models the fatigue process and future scheduled inspections as a hidden Markov model. This model is solved using a particle-based approximation and the sequential importance sampling algorithm, which provides an efficient solution to the PDTA problem. Sequential importance sampling is an extension of importance sampling to a Markov process, allowing for efficient Bayesian updating of model parameters. This model updating capability, the benefit of which is demonstrated, is lacking in other PDTA approaches. The results of this approach are shown to agree with the results of the explicit Monte Carlo routine for a number of PDTA problems.

Extensions to the typical PDTA problem, which cannot be solved using currently available tools, are presented and solved in this work. These extensions include incorporating observed evidence (such as non-destructive inspection results), more realistic treatment of possible future repairs, and the modeling of failure involving more than one crack (the so-called continuing damage problem).

The described hidden Markov model / sequential importance sampling approach to PDTA has the potential to improve aerospace structural safety and reduce maintenance costs by providing a more accurate assessment of the risk of failure and the likelihood of repairs throughout the life of an aircraft.

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19

Shengnan, Geng, Wang Xinglai, and Feng Hui. "FIBER BRAGG GRATING SENSOR SYSTEM FOR MONITORING COMPOSITE AEROSPACE STRUCTURES." International Foundation for Telemetering, 2016. http://hdl.handle.net/10150/624242.

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To investigate strain-sensitive characteristics of fiber Bragg grating (FBG) sensors, a minimal sensing system consisting of multiplex FBG sensors and signal demodulating and processing instruments was constructed. FBG sensors were designed with different package structures for respectively sensing strain or temperature parameters, and they returned measurand-dependent wavelengths back to the interrogation system for measurement with high resolution. In this paper, tests were performed on structure samples with step-wise increase of deformations. Both FBG sensing system and strain gages were tested and compared. Experimental work proved that the FBG sensing system had a good level of accuracy in measuring the static response of the tested composite structure. Moreover the additional advantages such as damp proofing, high sampling rates and real-time inspection make the novel system especially appropriate for load monitoring and damage detection of aerospace structures.
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Footdale, Joseph N. "Multi-axis real-time hybrid testing for precision aerospace structures." Connect to online resource, 2008. http://gateway.proquest.com/openurl?url_ver=Z39.88-2004&rft_val_fmt=info:ofi/fmt:kev:mtx:dissertation&res_dat=xri:pqdiss&rft_dat=xri:pqdiss:3337052.

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21

XIE, QIULIN. "PROBABILISTIC DESIGN OPTIMIZATION OF BUILT-UP AIRCRAFT STRUCTURES WITH APPLICATION." MSSTATE, 2003. http://sun.library.msstate.edu/ETD-db/theses/available/etd-07292003-211728/.

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This thesis discusses a methodology for probabilistic design optimization of aircraft structures subject to a multidisciplinary set of requirements originating from the desire to minimize structural weight while fulfilling the demands for quality, safety, producibility, and affordability. With this design methodology as the framework, a software is developed, which is capable of performing design optimization of metallic built-up beam structures where the material properties, external load, as well as the structural dimensions are treated as probabilistic random variables. The structural and failure analyses are based on analytical and semi-empirical methods whereas the component reliability analysis is based on advanced first-order second moment method. Metrics-based analytical models are used for the manufacturability analysis of individual parts with the total manufacturing cost estimated using models derived from the manufacturing cost / design guide developed by the Battelle¡¯s Columbus Laboratories. The resulting optimization problem is solved using the method of sequential quadratic programming. A wing spar design optimization problem is used as a demonstrative example including a comparison between non-buckling and buckling web design concepts. A sensitivity analysis is performed and the optimization results are used to highlight the tradeoffs among weight, reliability, and manufacturing cost.
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22

Han, Yong. "Aeroelastic oscillations of damaged wing structures with bonded piezoelectric strips." Thesis, McGill University, 2013. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=116892.

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This thesis examines a new method of detecting the presence of structural cracks in wing-like structures at an incipient stage. This method is based on the analysis of the dynamics of damaged structures with bonded piezoelectric strips executing flexural oscillations. Such oscillations can be generated by mechanical loads, piezoelectric actuators, or unsteady aerodynamic loads in certain flight conditions of the aircraft. The proposed method of crack detection uses pairs of piezoelectric strip sensors bonded on the opposite sides of the structure and is based on the fact that the presence of a crack causes a difference between the strains measured by the two sensors of a given pair. The structural analysis presented in this thesis uses a nonlinear model for the cracks and a finite element formulation for the piezoelectric strips coupled with the structure. A 3D panel method is used to determine the unsteady aerodynamic loads acting on the oscillating wing. This study includes the dynamic analysis in time domain of cracked wing-like structures undergoing forced flexural vibrations in a range of frequencies generated by a pair of piezoelectric actuators, as well as the analysis of the oscillating wings with piezoelectric strips subjected to unsteady aerodynamic loads. The numerical simulations have shown that the presence of a crack in wing-like structures can be efficiently detected at an early stage by monitoring the response of the piezoelectric sensor pairs.
Cette thèse étudie une nouvelle méthode de détection de la présence de fissures structurelles à un stade précoce dans une structure de type aile. Cette méthode est basée sur l'analyse des oscillations en flexion des structures endommagées munies de bandes piézoélectriques collées. Ces oscillations peuvent être générées par des charges mécaniques, des actionneurs piézoélectriques, ou des charges aérodynamiques instationnaires dans certaines conditions de vol de l'avion. La méthode de détection des fissures proposée utilise des paires de capteurs piézoélectriques collés sur les côtés opposés de la structure et est basée sur le fait que la présence d'une fissure entraîne une différence entre les déformations mesurées par les deux capteurs d'une paire donnée. L'analyse structurale présentée dans cette thèse utilise un modèle non linéaire pour les fissures et une formulation par éléments finis pour les bandes piézoélectriques couplées avec la structure. Une méthode de panneau tridimensionnelle est utilisée pour déterminer les charges aérodynamiques instationnaires agissant sur l'aile oscillante. Cette étude comprend l'analyse dynamique dans le domaine temporel de structure de type aile fissurée subissant des vibrations en flexion forcées dans une gamme de fréquences générées par une paire d'actionneurs piézoélectriques, ainsi que l'analyse des ailes oscillantes équipées de bandes piézoélectriques soumises à des charges aérodynamiques instationnaires. Les simulations numériques ont montré que la présence d'une fissure dans ces structures peut être efficacement détectée à un stade précoce en surveillant la réponse des capteurs piézoélectriques.
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23

Paget, Christophe. "Active Health Monitoring of Aerospace Composite Structures by Embedded Piezoceramic Transducers." Doctoral thesis, KTH, Aeronautical Engineering, 2001. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-3277.

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The objectives of the thesis work were to study theinteraction between embedded piezoceramic transducers andcomposite structures as well as determine techniques tosimplify the Lamb waves analysis. Firstly, this studyconsidered the design of the embedded piezoceramic transducers.Secondly, the effect of the embedded transducer on thecomposite strength as well as the influence of the mechanicallyloaded composite on the characteristics of the embeddedtransducer were investigated. Finally, to simplify the analysisof such complex Lamb wave responses, two techniques weredeveloped. They were based on the wavelet technique and amodelling technique, respectively.

The design of the embedded piezoceramic transducers wasimproved by reducing the stress concentrations in the compositeas well as in all components constituting the piezoceramictransducer, that is, the piezoceramic element, interconnectorand conductive adhesive. The numerical analysis showed that thethickness of the interconnector had no significant influence onthe stress state of the piezoceramic transducer. It was alsofound that a compliant conductive adhesive reduced the stressconcentration located at the edge of the piezoceramic element.The structural integrity of composites embedded with theimproved piezoceramic transducer was investigated. Theexperiments, performed in tensile and compressive staticloading, indicated that the strength of the composite was notsignificantly reduced by the embedded piezoceramic transducer.Further investigations were conducted to evaluate theperformance of the improved piezoceramic transducer used as aLamb wave generator embedded in composites subjected tomechanical loading. The tests were conducted in tensile andcompressive static loading as well as fatigue loading. Thestudy showed a large working range of the embedded piezoceramictransducer. A post processing technique based on the waveletswas further assessed in the detection of damage and in thedamage size evaluation. A new wavelet basis was developedspecially for processing the Lamb wave response. This method,focused on the wavelet coefficients from the decomposition Lambwave response, showed promising results in evaluating thedamage size. The wavelets offered a sensitive tool to detectsmall damage, compared to other detection methods, improvingthe damage detection capabilities. The other technique wasdevoted to the simplification of the generated Lamb waves bythe use of multi-element transducers. The transducers weredesigned using both a normal-mode expansion and a FE-method.This technique allowed reducing the effect of a Lamb wave modetowards another. This technique was successfully implemented ina damage detection system in composites.

Keywords:Embedded piezoceramic, transducer, composite,structural integrity, health monitoring, damage detection, Lambwaves, wavelets, normal-mode expansion, FE-method

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24

Gunel, Murat. "Linear And Nonlinear Progressive Failure Analysis Of Laminated Composite Aerospace Structures." Master's thesis, METU, 2011. http://etd.lib.metu.edu.tr/upload/12614033/index.pdf.

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This thesis presents a finite element method based comparative study of linear and geometrically non-linear progressive failure analysis of thin walled composite aerospace structures, which are typically subjected to combined in-plane and out-of-plane loadings. Different ply and constituent based failure criteria and material property degradation schemes have been included in a PCL code to be executed in MSC Nastran. As case studies, progressive failure analyses of sample composite laminates with cut-outs under combined loading are executed to study the effect of geometric non-linearity on the first ply failure and progression of failure. Ply and constituent based failure criteria and different material property degradation schemes are also compared in terms of predicting the first ply failure and failure progression. For mode independent failure criteria, a method is proposed for the determination of separate material property degradation factors for fiber and matrix failures which are assumed to occur simultaneously. The results of the present study show that under combined out-of-plane and in-plane loading, linear analysis can significantly underestimate or overestimate the failure progression compared to geometrically non-linear analysis even at low levels of out-of-plane loading.
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25

Elvin, Niell Glen. "Damage detection in civil and aerospace structures with fiber optic sensors." Thesis, Massachusetts Institute of Technology, 1995. http://hdl.handle.net/1721.1/37018.

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Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Civil and Environmental Engineering, 1995, and Thesis (M.S.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1995.
Includes bibliographical references.
by Niell Glen Elvin.
M.S.
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26

Neri, Luca. "Negative Stiffness Structures: an additively manufactured design solution for aerospace applications." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2021. http://amslaurea.unibo.it/24997/.

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The aim of this project is to investigate two damping structures based on negative stiffness behavior and realized by Additive Manufacturing in the rubber like material Tango Gray. The first structure is based on the circular geometry by Wang et al. and Corsi et al. This structure is obtained by a circular repetition of an unit cell based on a curved beam plane geometry. When the curved beam is loaded the buckling instability phenomenon appears and the collapse of the structure reduces the force reaction: the fast change of configuration is called snap-through behaviour. If after buckling the force applied is removed the structure can remain in a different stable position: this is the bistability phenomenon. The second structure is a toroidal design. It is obtained by a rotation about a vertical axis parallel to the symmetry axis of the unit cell and at a distance "R" from the unit cell symmetry axis. The first step of this project is to determine which parameters might influence the behaviour of both structures. The high number of possible configuration to study are analized with the DOE (Design of Experiment) method. In this way is possible to select a reduced number of samples to be analized and find which variation of the characteristic parameters is favourable. The structures selected are verified with the numerical software ANSYS with a FEA (Finite Elements Analysis): the 3-dimensional model takes into consideration both material properties and geometrical dimensions. The obtained results are fundamental to choose which structures are the best to be prototyped and experimentally tested in a quasi-static compression test. At this point the experimental results are compared with the numerical ones to validate the numerical model. In this way it is possible to use this generical model for any tipes of material and geometry. A possible application of this innovative structure in the aerospace sector could be the possibility to dissipate impact energy.
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Martin, Oliver. "Metrology enabled tooling for the assembly of aero-structures." Thesis, University of Bath, 2016. https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.690722.

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Metrology and Tooling are considered as discrete disciplines within Manufacturing Engineering, however, assembly tooling often acts as a checking mechanism. Assembly tooling has the primary function of controlling part location during assembly; with a secondary requirement as a quality gate. In-tool checks are manual mechanical checks of the assembly, these gauging checks assume the tooling has the correct, nominal geometry. Tooling conformance is certified periodically; however these intervals can be up to three years. Further examination of the metrology requirements within the aerospace industry with respect to large scale assembly tooling identify a requirement to: reduce manual metrology checks, reduce tooling recertification time, and enable greater automation. Currently, there is a lack of integration between metrology and Wing-box assembly tooling. This research investigates how to increase manufacturing confidence with respect to tooling conformance; and, ultimately improve the manufacturing process for aero-structures, through the increased and enhanced use of metrology in the assembly tooling environment. The Metrology Enhanced Tooling for Aerospace (META) framework has been created to provide a robust framework for deploying metrology in the tooling environment. The major elements of the framework are subsequently detailed and demonstrated in three chapters: i) large volume metrology networks, for the measurement of tooling structures; testing instrument performance, quantifying and improving the uncertainty estimation, and ultimately, establishing a rapid measurement process for assembly tooling; ii) embedded metrology systems demonstrates how local measurement systems can be utilised to replace and improve on, traditional in-tool checks; and iii) metrology feedback presents an example of an automated tooling pick-up that manipulates the assembly to achieve the design intent. The contributions can be summarised as: firstly, the creation of the META framework for the deployment of metrology in assembly tooling environment, accommodating and facilitating a number of the future tooling and assembly requirements. Secondly, the establishment of a generic commissioning methodology and measurement strategy for the rapid measurement of assembly tooling to increase tooling confidence. The research output was demonstrated in a case study, through a combination of physical measurement and digital automation simulation to prove the process time was greatly decreased from current methods.
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Fulcher, Jared T. "MECHANICAL CHARACTERIZATIONS OF ENVIRONMENTALLY CONDITIONED SHAPE MEMORY POLYMERS FOR RECONFIGURABLE AEROSPACE STRUCTURES." UKnowledge, 2011. http://uknowledge.uky.edu/gradschool_theses/81.

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Shape memory polymers (SMPs) have been candidate materials for morphing applications. However, the SMPs have not been fully tested to work in relevant environments required for Air Force missions. In this study, an epoxy-based SMP was separately exposed to moisture, lubricating oil and UV radiation, which are simulated service environments designed to be reflective of anticipated performance requirements. The thermomechanical properties and shape memory effects were studied by using novel high-temperature nanoindentation technique. Results show that environmental conditions have affected the glass transition temperature and mechanical properties of the SMPs. In most cases, the conditioned SMPs exhibited higher elastic moduli than the unconditioned SMP. The shape recovery ability of the SMP was assessed by creating an indent and then observing the corresponding recovery according to the standard shape memory cycle. It was found that the deformation was mostly recovered for both conditioned and unconditioned SMP samples on heating the material above its glass transition temperature.
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Ginzburg, Dmitri. "Damage propagation and detection using nonlinear elastic wave spectroscopy in aerospace structures." Thesis, University of Bath, 2016. https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.698994.

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The sustainable future of aerospace industry in large part relies on two factors: (i) development of advanced damage tolerant materials and (ii) the ability to detect and evaluate defects at very early stages of component service life. The use of laminated composite materials, such as carbon fibre reinforced plastics (CFRP), has had a significant contribution to reducing airframe weight while improving passenger safety and comfort. However, it is well known that these materials exhibit poor resistance to impact damage caused by foreign objects. Inspired by the naturally occurring impact resistant structures, the first part of this work has shown that enhanced damage tolerance can be achieved with standard CFRP layers by creatively arranging them into bio-inspired configurations. Through an extensive numerical modelling study supported by the experimental results, a further insight into the possibilities that these structures can offer in terms of damage resistance was attained. The second part of this PhD work focused on developing a range of nonlinear nondestructive evaluation techniques that are sensitive to the early signs of material degradation. A range of defect types in metallic and composite structures has been considered, such as fatigue cracks, impact damage and disbonds in adhesively bonded components typical in aerospace industry. Furthermore, throughout this work, an advanced explicit finite element analysis (FEA) software code LS-DYNA® has been used for modelling the nonlinear effects associated with the propagation of elastic waves in damaged solid media.
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Tomko, Jason Robert. "Fluid-loaded vibration of thin structures due to turbulent excitation." Thesis, University of Notre Dame, 2014. http://pqdtopen.proquest.com/#viewpdf?dispub=3583070.

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Flow-induced structural acoustics involves the study of the vibration of a structure induced by a fluid flow as well as the resulting sound generated and radiated by the motion of the system. The thesis examines several aspects of flow-induced structural vibration for fluid-loaded systems. A new method, termed Magnitude-Phase Identification, is derived to experimentally obtain a modal decomposition of the vibration of a structure using two-point measurements. MPI was used to measure the auto-spectral density of various modes for a non-fluid-loaded, rectangular, clamped plate excited by a spatially-homogeneous turbulent boundary layer. These results agreed well with theory. Using MPI, it was shown that when both fluid-loading and a spatially non-homogeneous wall pressure field is applied to a structure that the mode shapes become dependent on the forcing field, an effect which does not occur when either characteristic is applied individually. Furthermore, the resulting mode shapes are potentially highly asymmetric. It was shown through a discretized string model that these results can be attributed to the increased damping induced by fluid loading. Internal acoustic wall pressure fields due to a ducted rotor were measured, and it was shown that the acoustic effects of the rotor can be approximated by replacing the rotor with a continuous ring of dipoles located at the blade tip. The finite length of the duct was accounted for through use of a method of images. The theoretical results from this model match well with the measured values. Lastly, the vibration of a fluid-loaded duct excited by an internal rotor is measured through use of MPI. The resulting vibration field appears similar to the field examined earlier due to fluid loading, with a decrease in the coherent vibration magnitude for increasing spatial separation from the reference location.

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31

Ekren, Mustafa. "Structural Optimization Strategies Via Different Optimization And Solver Codes And Aerospace Applications." Master's thesis, METU, 2008. http://etd.lib.metu.edu.tr/upload/12610250/index.pdf.

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In this thesis, structural optimization study is performed by using three different methods. In the first method, optimization is performed using MSC.NASTRAN Optimization Module, a commercial structural analysis program. In the second method, optimization is performed using the optimization code prepared in MATLAB and MSC.NASTRAN as the solver. As the third method, optimization is performed by using the optimization code prepared in MATLAB and analytical equations as the solver. All three methods provide certain advantages in the solution of optimization problems. Therefore, within the context of the thesis these methods are demonstrated and the interface codes specific to the programs used in this thesis are explained in detail. In order to compare the results obtained by the methods, the verification study has been performed on a cantilever beam with rectangular cross-section. In the verification study, the height and width of the cross-section of the beam are taken as the two design parameters. This way it has been possible to show the design space on the two dimensional graph, and it becomes easier to trace the progress of the optimization methods during each step. In the last section structural optimization of a multi-element wing torque box has been performed by the MSC.NASTRAN optimization module. In this section geometric property optimization has been performed for constant tip loading and variable loading along the wing span. In addition, within the context of shape optimization optimum rib placement problem has also been solved.
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32

Meshreki, Mouhab. "Dynamics of thin-walled aerospace structures for fixture design in multi-axis milling." Thesis, McGill University, 2009. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=32614.

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Milling of thin-walled aerospace structures is a critical process due to the high flexibility of the workpiece. Available models for the prediction of the effect of the fixture on the dynamic response of the workpiece are computationally demanding and fail to represent practical cases for milling of thin-walled structures. Based on the analysis of typical structural components encountered in the aerospace industry, a generalized unit-element, with the shape of an asymmetric pocket, was identified to represent the dynamic response of these components. Accordingly, two computationally efficient dynamic models were developed to predict the dynamic response of typical thin-walled aerospace structures. These models were formulated using Rayleigh's energy and the Rayleigh-Ritz methods. In the first model, the dynamics of multi-pocket thin-walled structures is represented by a plate with torsional and translational springs. A methodology was proposed and implemented for an off-line calibration of the stiffness of the springs using Genetic Algorithms. In the second model, the dynamics of a 3D pocket is represented by an equivalent 2D multi-span plate. Through a careful examination of the milling of thin-walled structures, a new formulation was developed to represent the continuous change of thickness of the workpiece due to the material removal action. Two formulations, based on holonomic constraints and springs with finite stiffness, were also developed and implemented to take into account the effect of perfectly rigid and deformable fixture supports. All the developed models and formulations were validated numerically and experimentally for different workpiece geometries and
Le fraisage des structures aérospatiales à parois minces est un processus critique dû à la flexibilité élevée de la pièce. Les modèles disponibles pour la prévision de l'effet du système de fixation sur la réponse dynamique de la pièce sont basés sur des méthodes numériques très lentes et n'arrivent pas à représenter les cas pratiques du fraisage des structures à parois minces. Basé sur une analyse des composants structurels typiques produits dans l'industrie aérospatiale, un élément généralisé de base avec la forme d'une poche asymétrique, a été identifié pour représenter la réponse dynamique de ces composants. En conséquence, deux modèles dynamiques efficaces ont été développés pour prévoir la réponse dynamique des structures aérospatiales types à parois minces. Ces modèles ont été formulés en utilisant les méthodes de Rayleigh et Rayleigh-Ritz. Dans le premier modèle, les réponses dynamiques des structures de poches multiples à parois minces sont représentées par des plaques avec des ressorts de torsion et de translation. Une méthodologie a été proposée et mise en application pour calibrer la rigidité des ressorts en utilisant les algorithmes génétiques. Dans le deuxième modèle, la réponse dynamique d'une poche en 3D est représentée par une plaque équivalente de multi-travées en 2D. À travers une étude approfondie du fraisage des structures à parois minces, une nouvelle formulation a été développée pour représenter le changement continu de l'épaisseur de la pièce durant l'usinage. Deux formulations, basées sur des contraintes holonomes et des ressorts avec des rigidités finies, ont été$
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Njuguna, James A. K. "Micro- and macro-mechanical properties of aerospace composite structures and their dynamic behaviour." Thesis, City University London, 2006. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.440734.

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34

Christian, William J. R. "The development of a strain-based defect assessment technique for composite aerospace structures." Thesis, University of Liverpool, 2017. http://livrepository.liverpool.ac.uk/3010051/.

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This thesis details the work conducted over three years on the development of strain-based defect assessment techniques for carbon-fibre reinforced composites. This material, whilst exhibiting a high specific strength, is sensitive to defects and thus there is an industrial need for assessment techniques that are capable of characterising defects and obtaining predictions of residual strength or life. The most commonly applied techniques are currently ultrasonic and thermographic non-destructive evaluation. A strain-based defect assessment could lead to more accurate predictions of residual strength, resulting in a reduction of the costs associated with operating composite aerospace structures. The aim of this project is to increase the quality and confidence in residual strength information gained from the non-destructive evaluation of composite defects using strain-based assessments, in addition to currently applied ultrasonic practices for composite structures. A literature review on composite defects and existing techniques for assessing defects was conducted. Knowledge gaps were then identified that if filled, could improve residual strength predictions. Initially, a statistical framework was developed that used Bayesian regression to predict the residual strength of impacted composites, based on ultrasonic non-destructive measurements, that is robust to data outliers. As part of this framework a performance metric for quantifying the accuracy of residual strength predictions was introduced, allowing currently applied assessment techniques to be compared with the novel strain-based assessment. Then, a novel technique for performing strain-based defect assessments was developed that utilised image decomposition and the statistical framework to make residual strength predictions. Digital image correlation was used to measure strain fields which were then dimensionally reduced to feature vectors using image decomposition. The difference between feature vectors representing virgin and defective laminates were quantified, resulting in a strain-based defect severity measure. Bayesian regression was used to fit an empirical model capable of predicting the residual strength of an impacted laminate based on the strain-based defect severity. The accuracy of the strain-based predictions were compared to the accuracy of ultrasound-based predictions and found to outperform the currently applied ultrasonic technique. Strain-based assessment of in-plane fibre-waviness was also explored, as minimal research had been conducted studying waviness defects with full-field techniques. This required the development of a procedure for creating controlled levels of local waviness in laminates. The same strain-based assessment used for assessing impact damage was applied to the fibre-waviness specimens, but for this defect the accuracy of predictions were found to be comparable to the ultrasound-based predictions. However, residual strain measurements were found to be effective for predicting the strength of laminates, indicating that knowledge of the residual strains around a waviness defect may be important when predicting a laminates residual strength.
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Horton, Brandon Alexander. "Comprehensive Multi-Scale Progressive Failure Analysis for Damage Arresting Advanced Aerospace Hybrid Structures." Diss., Virginia Tech, 2017. http://hdl.handle.net/10919/93961.

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In recent years, the prevalence and application of composite materials has exploded. Due to the demands of commercial transportation, the aviation industry has taken a leading role in the integration of composite structures. Among the leading concepts to develop lighter, more fuel-efficient commercial transport is the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept. The highly integrated structure of PRSEUS allows pressurized, non-circular fuselage designs to be implemented, enabling the feasibility of Hybrid Wing Body (HWB) aircraft. In addition to its unique fabrication process, the through-thickness stitching utilized by PRSEUS overcomes the low post-damage strength present in typical composites. Although many proof-of-concept tests have been performed that demonstrate the potential for PRSEUS, efficient computational tools must be developed before the concept can be commercially certified and implemented. In an attempt to address this need, a comprehensive modeling approach is developed that investigates PRSEUS at multiple scales. The majority of available experiments for comparison have been conducted at the coupon level. Therefore, a computational methodology is progressively developed based on physically realistic concepts without the use of tuning parameters. A thorough verification study is performed to identify the most effective approach to model PRSEUS, including the effect of element type, boundary conditions, bonding properties, and model fidelity. Using the results of this baseline study, a high fidelity stringer model is created at the component scale and validated against the existing experiments. Finally, the validated model is extended to larger scales to compare PRSEUS to the current state-of-the-art. Throughout the current work, the developed methodology is demonstrated to make accurate predictions that are well beyond the capability of existing predictive models. While using commercially available predictive tools, the methodology developed herein can accurately predict local behavior up to and beyond failure for stitched structures such as PRSEUS for the first time. Additionally, by extending the methodology to a large scale fuselage section drop scenario, the dynamic behavior of PRSEUS was investigated for the first time. With the predictive capabilities and unique insight provided, the work herein may serve to benefit future iteration of PRSEUS as well as certification by analysis efforts for future airframe development.
PHD
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36

Kapidzic, Zlatan. "Strength analysis and modeling of hybrid composite-aluminum aircraft structures." Licentiate thesis, Linköpings universitet, Hållfasthetslära, 2013. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-91894.

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The current trend in aircraft design is to increase the proportion of fiber composites in the structures. Since many primary parts also are constructed using metals, the number of hybrid metal-composite structures is increasing. Such structures have traditionally often been avoided as an option because of the lack of methodology to handle the mismatch between the material properties. Composite and metal properties differ with respect to: thermal expansion, failure mechanisms, plasticity, sensitivity to load type, fatigue accumulation and scatter, impact resistance and residual strength, anisotropy, environmental sensitivity, density etc. Based on these differences, the materials are subject to different design and certification requirements. The issues that arise in certification of hybrid structures are: thermally induced loads, multiplicity of failure modes, damage tolerance, buckling and permanent deformations, material property scatter, significant load states etc. From the design point of view, it is a challenge to construct a weight optimal hybrid structure with the right material in the right place. With a growing number of hybrid structures, these problems need to be addressed. The purpose of the current research is to assess the strength, durability and thermo-mechanical behavior of a hybrid composite-aluminum wing structure by testing and analysis. The work performed in this thesis focuses on the analysis part of the research and is divided into two parts. In the first part, the theoretical framework and the background are outlined.Significant material properties, aircraft certification aspects and the modeling framework are discussed.In the second part, two papers are appended. In the first paper, the interaction of composite and aluminum, and their requirements profiles,is examined in conceptual studies of the wing structure. The influence of the hybrid structure constitution and requirement profiles on the mass, strength, fatigue durability, stability and thermo-mechanical behavior is considered. Based on the conceptual studies, a hybrid concept to be used in the subsequent structural testing is chosen. The second paper focuses on the virtual testing of the wing structure. In particular, the local behavior of hybrid fastener joints is modeled in detail usingthe finite element method, and the result is then incorporated into a global model using line elements. Damage accumulation and failure behavior of the composite material are given special attention. Computations of progressive fastener failure in the experimental setup are performed. The analysis results indicate the critical features of the hybrid wing structure from static, fatigue, damage tolerance and thermo-mechanical points of view.
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Deaton, Joshua D. "Design of Thermal Structures using Topology Optimization." Wright State University / OhioLINK, 2014. http://rave.ohiolink.edu/etdc/view?acc_num=wright1401302982.

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38

Chatla, Priyanjali. "LS-Dyna for Crashworthiness of Composite Structures." University of Cincinnati / OhioLINK, 2012. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1352993298.

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39

Lenahan, Kristie M. "Thermoelastic control of adaptive composites for aerospace applications using embedded nitinol actuators." Thesis, Virginia Tech, 1996. http://hdl.handle.net/10919/44955.

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Aerospace structures have stringent pointing and shape control requirements during long-term exposure to a hostile environment with no scheduled maintenance. This makes them excellent candidates for a smart structures approach as current passive techniques prove insufficient. This study investigates the feasibility of providing autonomous dimensional control to aerospace structures by embedding shape memory alloy elements inside composite structures. Increasing volume fractions of nitinol wire were embedded in cross-ply graphite/ epoxy composite panels. The potential of this approach was evaluated by measuring the change in longitudinal strain with increasing temperature and volume fraction. Reduction of thermal expansion is demonstrated and related to embedded volume fraction.

Classical lamination theory is used to formulate a two-dimensional model which included the adaptive properties of the embedded nitinol. The model was used to predict the increased modulus and reduction of thermal strain in the modified plates which was verified by the experimental data.
Master of Science

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40

Kececi, Erkan. "Highly durable hydrophobic thin films for moisture prevention of composite structures for aerospace applications." Diss., Wichita State University, 2012. http://hdl.handle.net/10057/6096.

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41

Chronopoulos, Dimitrios. "Prediction of the vibroacoustic response of aerospace composite structures in a broadband frequency range." Phd thesis, Ecole Centrale de Lyon, 2012. http://tel.archives-ouvertes.fr/tel-00787864.

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During its mission, a launch vehicle is subject to broadband, severe, aeroacoustic and structure-borne excitations of various provenances, which can endanger the survivability of the payload and the vehicles electronic equipment, and consequently the success of the mission. Aerospace structures are generally characterized by the use of exotic composite materials of various configurations and thicknesses, as well as by their extensively complex geometries and connections between different subsystems. It is therefore of crucial importance for the modern aerospace industry, the development of analytical and numerical tools that can accurately predict the vibroacoustic response of large, composite structures of various geometries and subject to a combination of aeroacoustic excitations. Recently, a lot of research has been conducted on the modelling of wave propagation characteristics within composite structures. In this study, the Wave Finite Element Method (WFEM) is used in order to predict the wave dispersion characteristics within orthotropic composite structures of various geometries, namely flat panels, singly curved panels, doubly curved panels and cylindrical shells. These characteristics are initially used for predicting the modal density and the coupling loss factor of the structures connected to the acoustic medium. Subsequently the broad-band Transmission Loss (TL) of the modelled structures within a Statistical Energy Analysis (SEA) wave-context approach is calculated. Mainly due to the extensive geometric complexity of structures, the use of Finite Element(FE) modelling within the aerospace industry is frequently inevitable. The use of such models is limited mainly because of the large computation time demanded even for calculations in the low frequency range. During the last years, a lot of researchers focus on the model reduction of large FE models, in order to make their application feasible. In this study, the Second Order ARnoldi (SOAR) reduction approach is adopted, in order to minimize the computation time for a fully coupled composite structural-acoustic system, while at the same time retaining a satisfactory accuracy of the prediction in a broadband sense. The system is modelled under various aeroacoustic excitations, namely a diffused acoustic field and a Turbulent Boundary Layer (TBL) excitation. Experimental validation of the developed tools is conducted on a set of orthotropic sandwich composite structures. Initially, the wave propagation characteristics of a flat panel are measured and the experimental results are compared to the WFEM predictions. The later are used in order to formulate an Equivalent Single Layer (ESL) approach for the modelling of the spatial response of the panel within a dynamic stiffness matrix approach. The effect of the temperature of the structure as well as of the acoustic medium on the vibroacoustic response of the system is examined and analyzed. Subsequently, a model of the SYLDA structure, also made of an orthotropic sandwich material, is tested mainly in order to investigate the coupling nature between its various subsystems. The developed ESL modelling is used for an efficient calculation of the response of the structure in the lower frequency range, while for higher frequencies a hybrid WFEM/FEM formulation for modelling discontinuous structures is used.
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Mullan, Matthew Noel. "Understanding cost drivers within an aerospace manufacturing supply chain for fibre reinforced plastic structures." Thesis, Queen's University Belfast, 2013. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.602691.

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Increasingly stringent emission targets has encouraged prime airframe assemblers to reduce aircraft weight, in doing so reducing CO2 emissions associated with fuel burn. All three major aircraft manufacturers have chosen to achieve this by incorporating greater proportions of advanced composite material within aircraft structural design. This research presents the development of a novel approach to simultaneously model continuous and discontinuous production processes, enabling new understanding of composite material production cost drivers within an aerospace manufacturing supply chain. The sources of composite component cost drivers are traced through a complex manufacturing supply chain to understand specific driver impact. A hybrid mass and energy balance methodology is presented to simulate composite material production processes within a framework using physics inputs. Verification of the simulation results has been demonstrated through comparison with real aerospace material pricing. A production cost mark-up factor has been identified. as the greatest contributor toward material price, owing to its recurring manifestation at each tier interface. Dominant production costs at discrete supply chain tiers have been shown to develop from raw process materials, through energy dominant production techniques, finally resulting in large labour commitments for material handling and equipment preservation. The methodology investigates two non-deterministic aspects of production simulation; cost drivers changing with time and production processes changing with technology. This research has identified the cost drivers associated with composite component production are resultant from a multi-tier production supply chain, introducing non-value added costs at each ~tier interface. Through an analytical modelling method, the understanding of material production costs is encouraged, which may prompt a shift in composite material procurement strategies. The model has demonstrated that with ever increasing material production costs, product ion technology can provide a quantifiable reduction in component weight and composite material cost.
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Mohammed, Mohammed Abdelaziz Elamin. "IMPACT AND POST IMPACT RESPONSE OF COMPOSITE SANDWICH STRUCTURES IN ARCTIC CONDITION." University of Akron / OhioLINK, 2018. http://rave.ohiolink.edu/etdc/view?acc_num=akron1518520473027006.

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44

Wild, Graham. "Distributed optical fibre smart sensors for acoustic sensing in the structural health monitoring of robust aerospace vehicles." Thesis, Edith Cowan University, Research Online, Perth, Western Australia, 2010. https://ro.ecu.edu.au/theses/1873.

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The use of distributed optical fibre smart sensors for the detection of acoustic signals in the Structural Health Monitoring (SHM) of robust aerospace vehicles has been demonstrated. Current distributed optical fibre sensors are multiplexed along a single fibre. Inherent problems exist with a multiplexed architecture. Two significant issues are; the possibility of fibre breakage, and the possibility of failure of the single transmitter, the single receiver, or the single processor. In a ‘smart’ architecture, the intelligence, as well as the sensors, is distributed. Hence, if destructive damage occurs, then the SHM system can continue to operate in all other locations on the vehicle, making the system robust. Work on the optical fibre sensors was limited to acoustic signals. This included acoustic emissions, acousto-ultrasonics, acoustic transmissions and other dynamic strain signals. Fibre Bragg Gratings (FBGs) were chosen as the optical fibre sensor for the detection of the acoustic signals. FBGs offer significant advantages over other types of optical fibre sensors. The most significant of these is the ease of multiplexing and their versatility, i.e. the ability of FBGs to detect a significant number of measurands. In the work on optical fibre sensing, we showed the implementation of an innovative detection system. This Transmit Reflect Detection System (TRDS) made use of both the transmitted and reflected signals from the FBG. The TRDS is an improvement on conventional power detection where either the transmitted or reflected component is used. The TRDS was used to successfully detect all types of dynamic and static signals, the most significant being the acoustic emission from a lead pencil break test. The use of the FBG sensor as a receiver for acoustic communications was also shown. Acoustic communications have been proposed for use in the SHM of robust aerospace vehicles with the use of autonomous agents, e.g. inspection or repair robots. The FBG receivers were compared with PZT receivers. When communicating through aluminium, the FBG performance was not as good as the PZT receiver, specifically due to the properties of the FBG which limit the frequency response. However, in Carbon Fibre Composites (CFC), the FBG outperformed the PZT due to the properties of the CFC. We also note that when contained within the thermal packaging the FBG had a very interesting frequency response, likely due to the suspended beam nature of the structure. This type of packaging could be used to tune the response of the FBG sensor. The work on the distributed optical fibre smart sensors showed the implementation of a Smart Transducer Interface Module (STIM), which used the TRDS with a Digital Signal Processor (DSP). The output of the TRDS was differentially amplified with a high speed amplifier, and the output was passed to the ADC onboard the DSP. The DSP was also used to toggle on and off output, including closed loop actuation, and controlling a 1550nm laser, which would represent the source used in the implemented system. The use of STIM to form a distributed optical fibre sensor network was also shown in principle.
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45

McEwan, Matthew Ian. "A combined modal/finite element technique for the non-linear dynamic simulation of aerospace structures." Thesis, University of Manchester, 2001. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.504961.

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The surface panels of modern high-speed aircraft are subjected to high-intensity acoustic loading from sources such as jet efflux and turbulent fluid flow. The vibration induced by this loading may result in fatigue crack growth near stress concentrations, and may eventually cause in-service failure of the structure. Central to the acoustic fatigue design problem is knowledge of the dynamic response of the structure in questions, so that some fatigue damage model may be applied in order to determine the fatigue life of the structure. In this thesis, a combined normal mode /finite element technique is developed for modelling non-linear beams, plates and stiffened panels, undergoing large amplitude vibrations. The loads and displacements from a number of static non-linear finite element test cases are transformed into modal co-ordinates using the normal modes of the underlying linear system. Regression analysis is then used to find the unknown coupled non-linear modal stiffness coefficients. The inclusion of finite element derived modal mass, and an arbitrary damping model completes the governing non-linear equations of motion. Time domain numerical integration is then used to simulate the response to excitation with a wide variety of possible spatial and temporal components. The particular benefits of this approach are that a significant time-saving mnay be achieved in comparison to conventional finite element methods, and that almost am commercial finite element package may be employed without modification. The proposed method is applied to homogeneous isotropic beams. Fully simply supported and fully clamped boundary conditions are considered. For the free vibration case, results are compared to those of previous researchers. For the case of steady state harmonic, and random excitation, results are compared with the direct integration non-linear finite element method. For plates and stiffened panels, the stress response of the structure is determined by, identifying the non-linear relationship between stress and modal displacements, using a regression analysis. Also a formulation is proposed to allow the simulation of travelling planar pressure waves. These waves may impinge upon the structure at any angle. In order to demonstrate the proposed method, random excitation is applied to flat rectangular plates, and orthogonally stiffened panels. The Autopower Spectral Density estimates of the displacement and stress are obtained for a number of different excitation cases. In all of these problems, the proposed method demonstrates good agreement with the direct integration finite element method.
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46

Doyle, Keith Brian. "An optimization method for the design of structures for maximum fundamental frequency." Diss., The University of Arizona, 1993. http://hdl.handle.net/10150/186321.

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An optimization method to maximize the fundamental frequency of a structure is developed. The procedure uses the stresses due to the mechanical loading and the free-vibration mode shapes to determine design coefficients for the elements. Each element of the structure is assigned a design coefficient rated on a scale of zero to ten. The design coefficients are used to modify an initial design following an iterative procedure. This method of optimal structural design, referred to as the Maximum Stiffness Design (MSD), may be classified as an intuitive optimality criteria method. The MSD method is demonstrated by increasing the fundamental frequency of simple beam structures, truss structures, and complex structures. These examples include a support structure for a telescope, a support structure for a beam collapser, an airplane wing, and a truss railroad bridge. The MSD optimization method is compared to NASTRAN's Design Sensitivity Analysis to provide a benchmark comparison. It is shown that the MSD method compares well to NASTRAN's optimization method. Furthermore, the optimization technique is used to develop optimum contour shapes for single arch, double arch, and edge-supported mirrors.
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47

Khalili, Ashkan. "Spectrally formulated user-defined element in Abaqus for wave motion analysis and health monitoring of composite structures." Thesis, Mississippi State University, 2017. http://pqdtopen.proquest.com/#viewpdf?dispub=10269016.

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Wave propagation analysis in 1-D and 2-D composite structures is performed efficiently and accurately through the formulation of a User-Defined Element (UEL) based on the wavelet spectral finite element (WSFE) method. The WSFE method is based on the first order shear deformation theory which yields accurate results for wave motion at high frequencies. The wave equations are reduced to ordinary differential equations using Daubechies compactly supported, orthonormal, wavelet scaling functions for approximations in time and one spatial dimension. The 1-D and 2-D WSFE models are highly efficient computationally and provide a direct relationship between system input and output in the frequency domain. The UEL is formulated and implemented in Abaqus for wave propagation analysis in composite structures with complexities. Frequency domain formulation of WSFE leads to complex valued parameters, which are decoupled into real and imaginary parts and presented to Abaqus as real values. The final solution is obtained by forming a complex value using the real number solutions given by Abaqus. Several numerical examples are presented here for 1-D and 2-D composite waveguides. Wave motions predicted by the developed UEL correlate very well with Abaqus simulations using shear flexible elements. The results also show that the UEL largely retains computational efficiency of the WSFE method and extends its ability to model complex features.

An enhanced cross-correlation method (ECCM) is developed in order to accurately predict damage location in plates. Three major modifications are proposed to the widely used cross-correlation method (CCM) to improve damage localization capabilities, namely actuator-sensor configuration, signal pre-processing method, and signal post-processing method. The ECCM is investigated numerically (FEM simulation) and experimentally. Experimental investigations for damage detection employ a PZT transducer as actuator and laser Doppler vibrometer as sensor. Both numerical and experimental results show that the developed method is capable of damage localization with high precision. Further, ECCM is used to detect and localize debonding in a composite material skin-stiffener joint. The UEL is used to represent the healthy case whereas the damaged case is simulated using Abaqus. It is shown that the ECCM successfully detects the location of the debond in the skin-stiffener joint.

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Moses, Mychal-Drew. "A Study on the Micro Electro-Discharge Machining of Aerospace Materials." TopSCHOLAR®, 2015. http://digitalcommons.wku.edu/theses/1448.

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Electrical Discharge Machining (EDM) is a non-traditional machining process that uses hundreds of thousands of minute electrical sparks per second to machine any electrically conductive material, no matter the hardness or how delicate it is. EDM allows a much greater range of design possibilities, unconstrained from the traditional machining processes, in which material is removed mechanically by either rotating the cutting tool or the work piece. Shapes that were impossible to machine by any other method, such as deep, precision, square holes and slots with sharp inside corners, are readily produced. It provides accurate geometries in high- aspect ratio holes and slots, blind undercuts, small holes adjacent to deep sidewalls, and complex cuts in thin, fragile parts. Micro-EDM is a growing form of manufacturing and will continue to expand within various production fields. Micro-EDM is especially attractive for the applications where the cutting time is minimal, but precision and accuracy are maximized. Micro- EDM is a non-traditional cutting process, which consistently produces ultra-precise holes with fine surface finishes and better roundness, while holding extremely close diameter tolerances. The process could be an excellent problem-solving tool for configurations that are difficult or impossible to produce using conventional machining processes. This study presents a comparative experimental investigation on the micro-EDM machinability of difficult-to-cut Ti-6Al-4V and soft brass materials. As both materials are electrically conductive, they were machinable using the micro-EDM process irrespective of their hardness. The machining performance of the two materials was evaluated based on the quality of the micro-features produced by the micro-EDM process. Both blind and through micro-holes and micro-slots were machined on brass and Ti-6Al-4V materials. The quality of micro-features was assessed based on the shape accuracy, surface finish and profile accuracy of the features. Finally, the arrays of micro-features were machined on both materials to compare the mass production capability of micro-EDM process on those materials.
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Culler, Adam John. "Coupled Fluid-Thermal-Structural Modeling and Analysis of Hypersonic Flight Vehicle Structures." The Ohio State University, 2010. http://rave.ohiolink.edu/etdc/view?acc_num=osu1280930589.

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Kral, Zachary Tyler. "Development of a decentralized artificial intelligence system for damage detection in composite laminates for aerospace structures." Diss., Wichita State University, 2013. http://hdl.handle.net/10057/10612.

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Because of economic impact that results from downtime, aircraft maintenance is an important issue in the aerospace industry. In-service structures will decay over time. Compared to low-cycle loading structures, aerospace structures experience extreme loading conditions, resulting in rapid crack propagation. The research involved in this dissertation concerns development of the initial stages of structural health monitoring (SHM) system that includes a network of ultrasonic testing sensors with artificial intelligence capable of detecting damage before structure failure. A series of experiments examining the feasibility of ultrasonic sensors to detect the initial onset of damage on a composite laminate, similar in structure to that used in aerospace components, was conducted. An artificial neural network (ANN) with the best accuracy was found to be a hybrid of a self-organizing map (SOM) with a feed-forward hidden and output layer. This was used for the single actuator-to-sensor scans on a composite laminate with simulated damage. It was concluded that a decentralized network of sensors was appropriate for such a system. The small four-sensor system was proven to be capable of predicting the presence of damage within a scanning area on a composite laminate, as well as predict the location once damage was detected. The main experimentation for this dissertation involved four ultrasonic sensors operated in a pitch-catch configuration. Simulated damage, verified through experimentation, was placed at various locations in the scanning area of interest. Signals obtained from the ultrasonic sensors were analyzed by a multi-agent system in which each agent describes an ANN. The system was trained to determine damage size. A second multi-agent system was constructed to determine the location of the detected damage. The architecture was similar to the damage-sizing system. Results demonstrated that with the artificial intelligence post-processing of ultrasonic sensors, 95% confidence can be obtained for detecting and locating damage that is 0.375 in. in diameter, which was verified through a bootstrap method. This dissertation validated the initial stages of constructing such a network of ultrasonic sensors. Future research in this area could involve combining the four-sensor network into a larger network of sensors by means of multi-agent processing (i.e., developing scanning regions). The novel method presented here provides the basis for the development of the SHM system for typical aerospace structures.
Thesis (Ph.D.)--Wichita State University, College of Engineering, Dept. of Aerospace Engineering.
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