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1
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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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.
3
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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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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Liu, Albert Darien. "THE EFFECT OF SENSOR MASS, SENSOR LOCATION, AND DELAMINATION LOCATION OF DIFFERENT COMPOSITE STRUCTURES UNDER DYNAMIC LOADING." DigitalCommons@CalPoly, 2013. https://digitalcommons.calpoly.edu/theses/917.
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This study investigated the effect of sensor mass, sensor location, and delamination location of different composite structures under dynamic loading. The study pertains to research of the use of accelerometers and dynamic response as a cost-effective and reliable method of structural health monitoring in composite structures. The composite structures in this research included carbon fiber plates, carbon fiber-foam sandwich panels, and carbon-fiber honeycomb sandwich panels. The composite structures were manufactured with the use of a Tetrahedron MTP-8 heat press. All work was conducted in the Cal Poly Aerospace Structures/Composites Laboratory. Initial delaminations were placed at several locations along the specimen, including the bending mode node line locations. The free vibration of the composite structure was forced through a harmonic horizontal vibration test using an Unholtz-Dickie shake system. A sinusoidal sweep input was considered for the test. The dynamic response of the composite test specimens were measured using piezoelectric accelerometers. Measurements were taken along horizontal and vertical locations on the surfaces of the composite structures. Square inch grids were marked on the surfaces to create a meshed grid system. Accelerometer measurements were taken at the center of the grids. The VIP Sensors 1011A piezoelectric accelerometer was used to measure vibration response. The measurements were then compared to response measurements taken from a PCB Piezotronics 353B04 single access accelerometer to determine the effects of sensor mass. Deviations in bending mode natural frequency and differences in mode shape amplitude became the criteria for evaluating the effect of sensor mass, sensor location, and delamination location. Changes in damping of the time response were also studied. The experimental results were compared to numerical models created using a finite element method. The experimental results and numerical values were shown to be in good agreement. The sensor mass greatly affected the accuracy and precision of vibration response measurements in the composites structures. The smaller weight and area of the VIP accelerometer helped to minimize the decrease in accuracy and precision due to sensor mass. The effect of sensor location was found to be coupled with the effect of sensor mass and the bending mode shape. The sensor location did not affect the vibration response measurements when the sensor mass was minimized. Off-center horizontal sensor placement showed the possibility of measuring vibration torsion modes. The effect of delamination changed the bending mode shape of the composite structure, which corresponded to a change in natural frequency. The greatest effect of the delamination was seen at the bending mode node lines, where the bending mode shape was most significantly affected. The effect of delamination was also dependent on the location of the delamination and the composite structure type. The results of this study provided considerations for future research of an active structural health monitoring system of composite structures using dynamic response measurements. The considerations included sensor mass reduction, sensor placement at constraints and bond areas and the presence of damping material.
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Davis, Richard Anthony. "The Effects of a Damage Arrestment Device on the Mechanical Behavior of Sandwich Composite Beams Under Four-Point Bending." DigitalCommons@CalPoly, 2011. https://digitalcommons.calpoly.edu/theses/506.
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The demand for an insert on composite sandwich structures to aid in the arrestment of face-core delamination is of great need. This research studies the use of a damage arrestment device (DAD) that connects the carbon fiber face sheets to the foam core to find whether an increase in the structural integrity of the sandwich beam results. Experimental analysis was employed to test the samples and was verified by a theoretical and finite element approach.
The mechanical properties of LTM45/CF1803 pre-impregnated carbon fiber and Last-A-foam FR 6710 polyvinylchloride foam were experimentally analyzed using ASTM D3039 and ASTM D1621 standards respectively to verify the manufacturer’s data for the given material. With all the mechanical data, the effects of adding DAD keys to a delaminated composite sandwich beam were studied under a four-point bending test using ASTM standard D6272 and compared with non-delaminated beams to see if an increase in ultimate strength could be achieved. The initial delamination in the beams under consideration was one inch in length and located in between the loaded span of the beam. Two control beams were utilized for comparison: one with no defects, and another with a one inch delamination introduced at the face-core interface. The DAD keys were added in two different configurations to potentially stop the delamination propagation and increase the ultimate strength. In the first configuration DAD keys were added 0.25 inches on either side of the initial delamination in the transverse direction and provided a significant increase in strength over the delaminated control beam. The second configuration had a DAD key running along the longitudinal axis of the sandwich beam and resulted in a significant increase in ultimate strength over the delaminated control beam. After testing ten successful samples for each of the six different configurations, it was concluded that the addition of DAD keys in both configurations significantly increased the structural integrity of both the delaminated and non-delaminated control beams.
With all the experimental data acquired, finite element models were created in COSMOS. The purpose of the finite element analysis was to validate the experimental results by comparing the deflections of the beam subjected to four-point bending during the experiment to the deflections found numerically. The deflections for the various DAD key configurations found in the experimental work were in agreement with the finite element results.
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Elmushyakhi, Abraham. "In-Plane Fatigue Characterization of Core Joints in Sandwich Composite Structures." University of Dayton / OhioLINK, 2017. http://rave.ohiolink.edu/etdc/view?acc_num=dayton1510678155755824.
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Balatbat, Richard Vincent S. "An Investigation of Damage Arrestment Devices Application with Fastener/Hole Interaction." DigitalCommons@CalPoly, 2010. https://digitalcommons.calpoly.edu/theses/402.
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This thesis presents a parametric study on the effects of how damage arrestment devices application interacts with a fastener in a composite sandwich panel. The primary objective of the damage arrestment device was to prevent the failure of the composite face sheet, such as crack propagation, around the hole/fastener joint. The damage arrestment devices are made of composite strips that are inserted under the face sheet to increase the overall structural strength of the panel and to prevent the propagation of failure along the hole. This was supposed to be a quicker and stronger alternative to potted inserts for composite sandwich panels for designer. The manufacturing curing cycle of the composite sandwich specimens has been carried out by using a Tetrahedron Composite Air Press. The press has been used to fabricate composite sandwich panels by applying constant pressure and variable heat to create panels with dimensions of 5” x 2” x .552”. The panels were stacked using a polyurethane foam, Last-A-Foam FR-6710 with two layers of a carbon-fiber/epoxy weave, LTM45, on both sides of the foam. The specimens were loaded under a compressive strain of 0.5 mm/min. The damage arrestment devices’ thickness was varied and tested under both monotonic and fatigue loading. The experimental results indicate that as the thickness of the device increased the overall strength of the part increased at a parabolic curve with it topping at a thickness of 0.065” and a strength increase of 109%. Under fatigue loading, a control group test case and damage arrestment device configuration case was tested. The experimental results indicate that both cases have similar fatigue trends but shows that the damage arrestment specimens are stronger due to the increase of structural strength. The experimental results were compared with numerical results or Finite Element Model. The results showed that numerical results can capture the linear or elastic portion of the experimental results having identical Elastic Modulus values. The models do differ in the maximum displacement of the specimen and the failure mode around the hole of the composite sandwich panel. The discrepancy in displacement and the failure mode was attributed to inaccurate loading on the hole of the composite sandwich panel and non-linear modeling of the solution. The correlation between the FEM and the experimental data was good enough in predicting the trends of the composite sandwich panels.
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Carter, Jeffrey Scott. "Effects of Low Velocity Impact on the Flexural Strength of Composite Sandwich Structures." DigitalCommons@CalPoly, 2014. https://digitalcommons.calpoly.edu/theses/1327.
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The use of composite sandwich structures is rapidly increasing in the aerospace industry because of their increased strength-to-weight and stiffness-to-weight characteristics. The effects of low velocity impacts on these structures, however, are the main weakness that hinders further use of them in the industry because the damages from these loadings can often be catastrophic. Impact behavior of composite materials in general is a crucial consideration for a designer but can be difficult to describe theoretically. Because of this, experimental analysis is typically used to attempt to describe the behavior of composite sandwiches under impact loads. Experimental testing can still be unpredictable, however, because low velocity impacts can cause undetectable damage within the composites that weaken their structural integrity. This is an important issue with composite sandwich structures because interlaminar damage within the composite facesheets is typical with composites but the addition of a core material results in added failure modes. Because the core is typically a weaker material than the surrounding facesheet material, the core is easily damaged by the impact loads. The adhesion between the composite facesheets and the core material can also be a major region of concern for sandwich structures. Delamination of the facesheet from the core is a major issue when these structures are subjected to impact loads.
This study investigated, through experimental and numerical analysis, how varying the core and facesheet material combination affected the flexural strength of a composite sandwich subjected to low velocity impact. Carbon, hemp, aramid, and glass fiber materials as facesheets combined with honeycomb and foam as core materials were considered. Three layers of the same composite material were laid on the top and bottom of the core material to form each sandwich structure. This resulted in eight different sandwich designs. The carbon fiber/honeycomb sandwiches were then combined with the aramid fiber facesheets, keeping the same three layer facesheet design, to form two hybrid sandwich designs. This was done to attempt to improve the impact resistance and post-impact strength characteristics of the carbon fiber sandwiches. The two and one layer aramid fiber laminates on these hybrid sandwiches were always laid up on the outside of the structure. The sandwiches were cured using a composite press set to the recommended curing cycle for the composite facesheet material. The hybrid sandwiches were cured twice for the two different facesheet materials. The cured specimens were then cut into 3 inch by 10 inch sandwiches and 2/3 of them were subjected to an impact from a 7.56 lbf crosshead which was dropped from a height of 38.15 inches above the bottom of the specimen using a Dynatup 8250 drop weight machine.
The impacted specimen and the control specimen (1/3 of the specimens not subjected to an impact) were loaded in a four-point bend test per ASTM D7250 to determine the non-impacted and post-impact flexural strengths of these structures. Each sandwich was tested under two four-point bend loading conditions which resulted in two different extension values at the same 100 lbf loading value. The span between the two supports on the bottom of the sandwich was always 8 inches but the span between the two loading pins on the top of the sandwich changed between the two loading conditions. The 2/3 of the sandwiches that were tested after being impacted were subjected to bending loads in two different ways. Half of the specimens were subjected to four-point bending loads with the impact damage on the top facesheet (compressive surface) in between the loading pins; the other half were subjected to bending loads with the damage on the bottom facesheet (tensile surface).
Theoretical failure mode analysis was done for each sandwich to understand the comparisons between predicted and experimental failures. A numerical investigation was, also, completed using Abaqus to verify the results of the experimental tests. Non-impacted and impacted four-point bending models were constructed and mid-span deflection values were collected for comparison with the experimental testing results. Experimental and numerical results showed that carbon fiber sandwiches were the best sandwich design for overall composite sandwich bending strength; however, post-impact strengths could greatly improve. The hybrid sandwich designs improved post-impact behavior but more than three facesheet layers are necessary for significant improvement. Hemp facesheet sandwiches showed the best post-impact bending characteristics of any sandwich despite having the largest impact damage sizes. Glass and aramid fiber facesheet sandwiches resisted impact the best but this resulted in premature delamination failures that limited the potential of these structures. Honeycomb core materials outperformed foam in terms of ultimate bending loads but post-impact strengths were better for foam cores. Decent agreement between numerical and experimental results was found but poor material quality and high error in material properties testing results brought about larger disagreements for some sandwich designs.
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Renova, Elvia Paola. "Simulated annealing algorithms for the optimization of particulate composite structures analyzed by X-FEM." To access this resource online via ProQuest Dissertations and Theses @ UTEP, 2008. http://0-proquest.umi.com.lib.utep.edu/login?COPT=REJTPTU0YmImSU5UPTAmVkVSPTI=&clientId=2515.
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Thummalapalli, Vimal Kumar. "Biomimetic Composite T-Joints." University of Dayton / OhioLINK, 2011. http://rave.ohiolink.edu/etdc/view?acc_num=dayton1323547304.
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Richard, Brandon Demar. "Thermal Infrared Reflective Metal Oxide Sol-Gel Coatings for Carbon Fiber Reinforced Composite Structures." Scholar Commons, 2013. http://scholarcommons.usf.edu/etd/4569.
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Recent trends in composite research include the development of structural materials with multiple functionalities. In new studies, novel materials are being designed, developed, modified, and implemented into composite designs. Typically, an increase in functionality requires additional material phases within one system. The presence of excessive phases can result in deterioration of individual or overall properties. True multi-functional materials must maintain all properties at or above the minimum operating limit. In this project, samples of antimony and cobalt-doped tin oxide (ATO(Co2O3)) sol-gel solutions are used to coat carbon fibers and are heat treated at a temperature range of 200 - 500 °C. Results from this research are used to model the implementation of sol-gel coatings into carbon fiber reinforced multifunctional composite systems. This research presents a novel thermo-responsive sol-gel/ (dopant) combination and evaluation of the actuating responses (reflectivity and surface heat dissipation) due to various heat treatment temperatures. While ATO is a well-known transparent conductive material, the implementation of ATO on carbon fibers for infrared thermal reflectivity has not been examined. These coatings serve as actuators capable of reflecting thermal infrared radiation in the near infrared wavelengths of 0.7-1.2 μm. By altering the level of Co2O3 and heat treatment temperatures, optimal optical properties are obtained. While scanning electron microscopy (SEM) is used for imaging, electron diffraction
spectroscopy (EDS) is used to verify the compounds present in the coatings. Fourier transform infrared (FT-IR) spectroscopy was performed to analyze the chemical bonds and reflectivity in the infrared spectra after the heat treatments. Total reflection and angle-dependent reflectivity measurements were performed on the coatings in the wavelengths of 0.7-2 μm. Laser induced damage threshold testing was done to investigate the dielectric breakdown and used to calculate surface temperatures.
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Puttmann, John Paul. "Spatially Targeted Activation of a SMP." University of Dayton / OhioLINK, 2018. http://rave.ohiolink.edu/etdc/view?acc_num=dayton1525166147319011.
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Heil, Joshua W. "Methods of Processing Kenaf Chopped Strand Mats for Manufacturing Test Specimens and Composite Structures." DigitalCommons@USU, 2015. https://digitalcommons.usu.edu/etd/4376.
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Bio-composites are increasing in demand due to governmental incentives across the globe for both environmental and human health reasons. The ideal bio-composite is renewable, recyclable, available, and non-toxic. To properly engineer bio-composite products, the physical properties of the fiber as well as fiber/matrix interactions must be known. The problem lies in the fact that many suitable natural fibers are not currently available in a material form that may be easily worked with. This research investigates methods to process raw kenaf (hibiscus cannabinus) on a scale that allows researchers to make more consistent samples for testing. Though kenaf is highlighted, these processing methods may be applied to any natural fiber. The raw fibers are processed into kenaf chopped strand mats (KCSM) by adapting basic paper-making techniques. KCSM exhibit paper-like qualities and mechanical properties and provide a material of uniform thickness for use in composite parts. Also presented are a basic understanding of natural fiber constituents and effects of mechanical and co-mechanical treatments on those constituents. To test KCSM, samples are made for the ASTM D3039 tensile testing and for testing in a dynamic material analyzer (DMA). Both mechanically and chemo-mechanically processed samples are made for the purpose of comparison. Also, I-beam bridges are built with KCSM to demonstrate how KCSM may be used to create a structure. This is spurred on by the annual SAMPE bridge competition that includes special categories for natural fiber beams. The lay-up procedure is shown in detail to provide a framework that future competitors may use to build quality I-beams for this competition.
The properties obtained by using the KCSM are competitive with other reported properties for kenaf-based composites. A kenaf I-beam demonstrates a strength-to-weight ratio that is 65% of a berglass I-beam built to the same dimensions. Trade-os of using KCSM are the random 2d-fiber orientation and brittle failure, which are not usually desirable in composite components. The chemically treated samples indicate a higher degree of crystallinity but demonstrate inferior mechanical properties when compared to the untreated samples.
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Ruggeri, Charles R. "High Strain Rate Data Acquisition of 2D Braided Composite Substructures." University of Akron / OhioLINK, 2009. http://rave.ohiolink.edu/etdc/view?acc_num=akron1255968114.
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Chillara, Venkata Siva Chaithanya. "Multifunctional Laminated Composites for Morphing Structures." The Ohio State University, 2018. http://rave.ohiolink.edu/etdc/view?acc_num=osu1524104865278235.
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Sweeney, Benjamin Andrew. "The Effect of Biocomposite Material In A Composite Structure Under Compression Loading." DigitalCommons@CalPoly, 2017. https://digitalcommons.calpoly.edu/theses/1932.
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While composite structures exhibit exceptional strength and weight saving possibilities for engineering applications, sometimes their overall cost and/or material performance can limit their usage when compared to conventional structural materials. Meanwhile ‘biocomposites’, composite structures consisting of natural fibers (i.e. bamboo fibers), display higher cost efficiency and unique structural benefits such as ‘sustainability’. This analysis will determine if the integration of these two different types of composites are beneficial to the overall structure. Specifically, the structure will consist of a one internal bamboo veneer biocomposite ply; and two external carbon fiber weave composite plies surrounding the bamboo biocomposite. To acquire results of this study, the hypothesized composite structure will consist of varied trapezoidal corrugated specimens and tested in uniaxial compression loading. Thereafter, this test data will be used to ultimately design, manufacture, and test a structural biocomposite/composite box, intended to carry extremely high compressive loads; relative to its own weight. A finite element analysis of this test will be used to validate experimental data. After running the experiment, the carbon fiber with bamboo test sample results were compared to that of only carbon fiber test sample. The carbon fiber samples resulted in a maximum compressive load difference of only 23% higher loads when compared to the carbon fiber with bamboo, on average. These findings are discussed throughout.
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Subramaniam, C. "Chemorhelogical Modeling Of Amine-Cured Multifunctional Epoxy Resin Systems Used As Matrices In Aerospace Composites." Thesis, Indian Institute of Science, 1994. http://hdl.handle.net/2005/127.
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High performance multifunctional epoxy resin systems are becoming increasingly important as matrix materials for the advanced composites used in aerospace, electronics, automotive and other industries. In a composite based on epoxy resin systems, a three-dimensional network of the matrix is formed around the reinforcing fibre as a result of the chemical reaction between the resin and the curing agent. This chemical process, known as curing, is an important event to he considered in the production of composite components made up of these resin systems. Two process parameters namely viscosity and chemical conversion are of paramount significance in the production of composite materials Curing studies of the resin systems based on these two parameters, would therefore assume great importance in deciding the performance reliability of the end product.
The objectives of the present investigation are
1. to study the cure kinetics of three thermoset resin systems, viz.,
i) epoxy novolac (EPIT)/ diamino diphenyl methane{DDM), ii) trigylcidyl para- ammo phenol (TGPAP)/toluene diamine (TDA) and iii) tetraglycidyl diamino diphenyl methane (TGDDM)/pyridine diamine(PDA) using the cure kinetic models based on chemical conversion (α), Theological conversion (β) and viscosity.
2.to develop a correlation between a and viscosity (η) and modify an existing autocatalytic model based on α, to the viscosity domain and
3.to investigate the cure behaviour of these systems in terms of the TTT cure diagram and its associated models.
EPN/DDM, TGPAP/PDA and TGDDM/PDA resin systems were chosen for the studies to represent a range of functionalities, The cure was monitored using differential scanning calorimetry (DSC), fourier transform infrared (FTIR) and dynamic mechanical analysis (DMA) techniques by following the changes in enthalpy, functional groups and rheology, respectively.
The kinetic parameters namely, order of reaction and activation energy were estimated from dynamic DSC data using the methods of Freeman-Carroll and Ellerstein using nth rate expression. Barton, Kissinger and Osawa methods were employed to find out the activation energy from the peak/equal conversion at different heating rates. Isothermal DSC data were also analyzed using nth order model and it was observed that the data could be fitted satisfactorily only for higher temperatures The results obtained from the analysis of both dynamic and isothermal DSC data using nth order model clearly indicate that this model is inadequate for describing the cure behavior. The isothermal DSC data was analyzed by the autocatalytic models of Hone and Kamal Good correlation was observed with Hum and Kamal models up to 60-70%, 25% and 45% conversions for EPN/DDM, TGPAP/TDA and TCDDM/PDA systems respectively. However, the parameters m and n in Kamal model were found to be temperature dependent for EPN/DDM and TCPAP/TDA systems. The limited applicability of the autocatalytic models IK attributed to the counter-effect offered by the intra-molecular bonding taking place.
The primary amine and epoxy groups conversions obtained from FTIR were analyzed using autocatalytic model and the kinetic parameters were calculated. The reactivity ratio of the primary amine and the secondary amine with epoxy was found to be dependent on temperature in agreement with the recent findings reported m the literature.
The existing models that relate the cure kinetics and the rheological changes, are dual Arrhenius nth order model and autocatalytic model The nth order kinetic model was used to evaluate the kinetic parameters using the viscosity data at different cure temperatures under isothermal conditions As the storage modulus, G' is proportional to the chemical cross links and becomes significant only after the g<4 point, it was used to follow the changes in conversion known as rheoconversion after the gel point The rheoconversion was found by normalizing the G' data with G1^, the storage modulus of the fully cured resin It was used to study the cuie kinetics using an autocatalytic model The kinetic parameters such as rate constant, acceptation and retardation parameters were evaluated and that temperature dependence was established.
While the existing models relate viscosity and conversion only up to gel point the new proposed model, termed VISCON model takes into account the changes up to vitrification. The relation so developed is used to modify the autocatalytic cure model based on chemical conversion. The parameters appearing in this model were evaluated using Levenberg-Marquardt error minimization algorithm. The kinetic parameters obtained are comparable with the values estimated using the DSC data.
All the models cited above represent the microkinetic aspects. The models based on the information of TTT cure diagrams, however, represent the macrokinetic aspects of the cure, as they are based on the cure stages such as gelation and vitrification TTT diagram relates the cure characteristics like cure temperature, cure time, Ta and, indirectly, chemical conversion Hence the ultimate properties of the composite could he predicted and established with the help of the models based on TTT cure diagrams The changes in the storage modulus, G1 and loss modulus, G", were followed to identify the gel and vitrification points of the resin systems at different cure temperatures Gel point and vitrification point were used to generate gelation and vitrification hues in the construction of TTT cure diagrams for EPN/DDM, TGPAP/TDA and TGDDM/PDA resin systems Theoretical TTT diagrams were generated and IBO-T, contours were established using the TTT diagram-based models The cure schedule for the resin systems investigated could be determined from the TTT diagram and the respective rheological data.
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Andrejic, Mateja. "Effects of Curing Cycle and Loading Rates on the Bearing Stress of Double Shear Composite Joints." DigitalCommons@CalPoly, 2016. https://digitalcommons.calpoly.edu/theses/1549.
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In the last few decades, there has been a shift to using more lightweight materials for the potential of fuel consumption reduction. In the Aerospace Industry, conventional metal structures are being replaced by advanced composite structures. The major advantage of an advanced composite structure is the huge reduction in the number of parts and joints required. Also composite materials provide better resistance to creep, corrosion, and fatigue. However, one cannot eliminate all the joints and attachments in an aircraft’s structure. Eliminating structural joints is impractical in present-day aircraft because of the requirements for inspection, manufacturing breaks, assembly and equipment access, and replacement of damaged structures. Currently, composite joints are overdesigned which leads to weight penalties. Understanding how to optimize the ultimate bearing strength of a composite joint by altering the cure cycle might be beneficial to the composite joint design process.
This study investigates, through numerical and experimental analysis, the mechanical behavior of double shear joints. The first task is to test Aluminum double shear joint specimens inside the double shear joint fixture at a loading rate of 0.05 in./min. (quasi-static). The second task is to numerically model and validate the aluminum double shear joint specimen. The third task is to test the Unidirectional MTM 49 carbon fiber pre-preg double shear composite joint specimens with two different cure cycles and five different loading rates (0.05 in./min., 0.1 in./min., 1 in./min., 2 in./min. and 6 in./min.). The double shear composite joint specimens are made, using a heat press, with a quasi-isotropic laminate orientation of [0 0 +45 -45 +45 -45 90 90]s. The first cure cycle used is called the alternate cure cycle, which is Cytec’s MTM 49 Unidirectional Carbon Fiber pre-preg material cure cycle, and the second cure cycle used is called the datasheet cure cycle, which is Umeco's MTM 49 Unidirectional Carbon Fiber pre-preg material cure cycle. The recommended datasheet cure cycle and an alternate cure cycle are both compared to see how they affect the mechanical characteristics of the matrix along with the bearing stress. The fourth task is to adjust the Aluminum double shear joint numerical model for the double shear composite joint specimen. The numerical results for both the Aluminum and the composite specimens are in agreement with the experimental results. The theoretical in-plane material properties of the quasi-isotropic laminate were in agreement with the experimental results. One can see that at 0.05 in./min. and 0.1 in./min. (for both cure cycles) the composite double shear specimens carried more load compared to the higher loading rates of 1 in./min., 2 in./min. and 6 in./min. The tensile modulus of elasticity of an Aluminum sample is measured using a crosshead displacement, a strain gage and an extensometer. The crosshead displacement yielded very inaccurate results when compared to the strain gage and the extensometer.
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Carter, Justin B. "Vibration and Aeroelastic Prediction of Multi-Material Structures based on 3D-Printed Viscoelastic Polymers." Miami University / OhioLINK, 2021. http://rave.ohiolink.edu/etdc/view?acc_num=miami1627048967306654.
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Terzak, John Charles. "Modeling of Microvascular Shape Memory Composites." Youngstown State University / OhioLINK, 2013. http://rave.ohiolink.edu/etdc/view?acc_num=ysu1389719238.
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Ahn, Junghyun. "Integrated analysis procedure of aerospace composite structure." Thesis, Massachusetts Institute of Technology, 2008. http://hdl.handle.net/1721.1/43106.
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Thesis (S.M.)--Massachusetts Institute of Technology, System Design and Management Program, 2008.Includes bibliographical references (p. 50).
The emergence of composite material application in major commercial aircraft design, represented by the Boeing 787 and Airbus A350-XWB, signals a new era in the aerospace industry. The high stiffness to weight ratio of continuous fiber composites (CFC) makes CFCs one of the most important materials to be introduced in modern aircraft industry. In addition to inherent strength (per given weight) of CFCs, they also offer the unusual opportunity to design the structure and material concurrently. The directional properties (and the ability to change these properties through the design process) of composite materials can be used in aeroelastically tailored wings, the fuselage and other critical areas. Due to the longer lifecycle (25-30 years) of a commercial airliner and the tools and processes developed for the airplane of previous product development cycles, new technology often ends up being deployed less effectively because of the mismatch in the technical potential (what can be done) vs. design tools and processes (what was done before). Tools and processes need to be current to take advantage of latest technology, and this thesis will describe one possible approach in primary composite structural design area using integrated structural analysis
by Junghyun Ahn.
S.M.
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Key, Ross A. "Automated manufacturing processes for secondary structure aerospace composites." Thesis, University of Nottingham, 2016. http://eprints.nottingham.ac.uk/33572/.
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As projected manufacturing rates for commercial aircraft increase to levels of multiple ship sets per day from individual manufacturing facilities, GE Aviation have expressed the need for a shift in composites secondary structure manufacturing philosophy. Traditional manufacturing processes tend to be touch labour intensive and hence costly. The manual placement of large numbers of individual ply profiles, lengthy debulking operations and complex cure cycles, result in excessive component lead times and manufacturing costs. As a result, direct labour cost is a major factor in the total economies of production processes. The implementation of industrial robotics has proved highly successful in automotive manufacturing, and various methods for automating individual aspects of the composites manufacturing process have been suggested. Technical cost modelling has been used to anticipate the production costs of a prototype secondary structure component, as supplied by GE Aviation, through direct simulation of the existing manufacturing process. This work has clearly highlighted the potential for cost and cycle time reductions if process automation can be successfully introduced. Observation of the existing manufacturing process has allowed three alternative manufacturing scenarios to be considered with respect to cost-effectiveness and feasibility, whilst highlighting long term cost benefits. Investigations have been undertaken to identify and evaluate alternative material and processing methodologies ranging from resin infused woven dry fabrics to UD prepreg tape and tow. In addition, candidate processing routes have been systematically evaluated using design of experiments techniques, which focussed on assessing the feasibility and technology readiness of robotic deposition and consolidation methodologies, including pick and place and debulking. Process automation in these areas has the potential for total component cost and cycle time reductions in the order of 2.8 to 21.6 and 0.6 to 63.4 per cent respectively. The quasi-static mechanical testing of a range of face sheet materials has provided a performance assessment based on tensile, compressive and shear properties and laminate Vf. Findings suggest that materials offering increased suitability for automation typically have reduced mechanical performance when compared to candidate prepregs; tensile modulus and strength reductions of 5 and 34 per cent were reported when comparing a 6k woven 2X2 twill fabric and equivalent prepreg respectively. Furthermore, 26 and 4 per cent reductions in tensile modulus and 38 and 40 per cent reductions in tensile strength were observed for 179 and 318gsm UD NCF, when compared with a candidate UD prepreg. Data has also been presented on the effect of varying the traditional consolidation frequency and methodology. While earlier findings suggest that debulking has little effect on the laminate tensile modulus; ply compaction level varies considerably. Furthermore, it has been shown that on-the-fly consolidation, using a robotically mounted, roller-based end effector has the advantages of mechanical performance retention, cycle time reduction and repeatable laminate post cure thickness. In addition, when compared with candidate woven and UD prepreg laminates manufactured using the traditional vacuum bagging approach; equivalent tensile modulus, strength and fibre volume fraction have been observed and with less variability. Handling characteristics inherent to vacuum and needle grippers, including pickup performance, defined as the pickup or holding force required to overcome fabric weight, shear force performance; the maximum force that can be exerted on the fabric before the onset of slip, and the accuracy with which non-rigid-materials (NRMs) can be handled, have also been considered. The achievable positional accuracy of robotically pick and placed prepreg plies greatly exceeds that of dry fabrics in all cases and with less variability, irrespective of the gripping mechanism used. Vacuum grippers exhibit more uniform positional error and increased positional accuracy when handling dry fabrics, whilst needle grippers outperformed the vacuum alternative when handling prepregs, irrespective of form. Robotic pick and place solutions offer low variability in ply positional error with a guaranteed placement accuracy of ±0.8mm and ±2.3mm for prepregs and dry fabrics respectively. Characterisation of the gap type defect and butt and overlapping joining methodologies has provided a performance trend based on ply positional error. Quasi-static mechanical testing has revealed that laminates with equivalent tensile modulus to an un-spliced control could be achieved. However, significant reductions in the tensile strength and an increase in overall laminate thickness and thickness variation highlighted the negative effect of ply splicing on laminate performance. However, it has been shown that a robotic placement accuracy of ±0.8mm gives rise to acceptable tensile strength reductions in candidate prepreg laminates. The up-scaling of laminate level robotic manipulators has been discussed and addressed in conjunction with the commissioning of a flexible robotic manufacturing cell, facilitating the manufacture of full-scale secondary structure aerospace components. Comparisons have been made between a benchmark prepreg panel, manufactured using traditional manual methods and alternative dry fabric and prepreg panels manufactured using increased levels of process automation. In each case, manufacturing feasibility, mechanical performance and component geometric accuracy have been assessed. It has been shown that there are significant advantages to be gained from the implementation of robotic automation within the traditional manufacturing process. Component cost and cycle time reductions, coupled with the processing and performance advantages and increased suitability to automation of woven dry fibre materials are clear. Findings which support a key driver of this project, which seeks to justify alternative dry fabrics as a viable alternative to traditional prepreg broadgoods for the manufacture of secondary structure aerospace components.
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Lee, Jin Woo. "Multi-level Decoupled Optimization of Wind Turbine Structures Using Coefficients of Approximating Functions as Design Variables." University of Toledo / OhioLINK, 2017. http://rave.ohiolink.edu/etdc/view?acc_num=toledo1501003238831086.
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Beji, Faycel Ben Hedi. "Buckling Analysis of Composite Stiffened Panels and Shells in Aerospace Structure." Thesis, Virginia Tech, 2018. http://hdl.handle.net/10919/81620.
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Stiffeners attached to composite panels and shells may significantly increase the overall buckling load of the resultant stiffened structure. Initially, an extensive literature review was conducted over the past ten years of published work wherein research was conducted on grid stiffened composite structures and stiffened panels, due to their applications in weight sensitive structures. Failure modes identified in the literature had been addressed and divided into a few categories including: buckling of the skin between stiffeners, stiffener crippling and overall buckling. Different methods have been used to predict those failures. These different methods can be divided into two main categories, the smeared stiffener method and the discrete stiffener method. Both of these methods were used and compared in this thesis. First, a buckling analysis was conducted for the case of a grid stiffened composite pressure vessel. Second, a buckling analysis was conducted under the compressive load on the composite stiffened panels for the case of one, two and three longitudinal stiffeners and then, using different parameters, stiffened panels under combined compressive and shear load for the case of one longitudinal centric stiffener and one longitudinal eccentric stiffener, two stiffeners and three stiffeners.Master of Science
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Palsule, Sanjay. "Structure and properties of aerospace molecular composites : third generation polymers." Thesis, Heriot-Watt University, 1994. http://hdl.handle.net/10399/1388.
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Hoffer, Jacob. "Development of a Draping Algorithm for Non-Structural Aerospace Composites." Thesis, Université d'Ottawa / University of Ottawa, 2020. http://hdl.handle.net/10393/40650.
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Fibre reinforced polymer matrix composites are used frequently in aerospace applications. Manufacturers of aerospace components favour composites over traditional metallic alloys due to their light weight, high modulus, corrosion resistance and fatigue resistance. Advantages of composites for non-structural interior components over metallic include: ease of manufacturing for single parts of complex geometry as opposed to assemblies, cheaper manufacturing of a limited series of parts and composites greatly reduced noise, vibration and harshness. However, manufacturing interior composite components requires critical attention to detail during the preforming stages and handling of dry fabric textiles. Since these components are handmade they often yield lower profits and therefore efficient preforming is critical. Designing draping strategies for industrial liquid composite moulding processes requires a significant amount of time and testing, in simulation and also working on physical moulds. Mould and part surfaces are often defined by a number of geometric features, labelled base surfaces in the context of this thesis, which can be used to quickly probe multiple draping strategies and identify the best one. Traditionally, trial and error work is performed over a full mould surface until a working or acceptable draping strategy is found, rarely identifying the best strategy. The work in this thesis presents the initial development stages for a draping predictive tool aimed at quickly probing multiple draping scenarios in simulation prior to receiving moulds and identifying the best draping strategy for industrial non-structural aerospace composites. A multi-parameter
remodelling tool – the conical frustum – was developed for uniformly identifying base surfaces through 12 geometric parameters linked into a database of in-plane shear and yarn orientations results. The development of the database is discussed, detailing Taguchi methods of experimental design used for developing linear functions from the database results, which allow interpolation of results on base surfaces that do not directly exist within the database. This thesis also includes major developments for the core draping algorithm used for linking individual base surface results together when probing draping strategies. Further investigations were performed on unique elements of in-plane shear behaviour that are encountered during draping, so that these could ultimately be considered during the development of this version of the draping algorithm whilst others may be included in future developments.
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Simsiriwong, Jutima. "STRUCTURAL TESTING OF AN ULTRALIGHT UAV COMPOSITE WING AND FUSELAGE." MSSTATE, 2009. http://sun.library.msstate.edu/ETD-db/theses/available/etd-04022009-160649/.
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The details of an experimental investigation focusing on obtaining the static and vibration characteristics of a full-scale carbon composite wing and fuselage structural assemblies of an ultralight unmanned aerial vehicle (UAV) are presented. The UAV has a total empty weight of 155-lb and an overall length of approximately 20.6-ft. A three-tier whiffletree system and the tail fixture were designed and used to load the wing and the fuselage in a manner consistent with a high-g flight condition. A shaker-table approach was used for the wing vibration testing, whereas the modal characteristics of the fuselage structure were determined for a free-free configuration. The static responses of the both structures under simulated loading conditions as well as their dynamic properties such as the natural frequency, damping coefficient and associated mode shapes were obtained. The design and implementation of the static and vibration tests along with the experimental results are presented in this thesis.
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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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Lokits, Jeremy Craig. "Structural Reinforcement Layout and Sizing Optimization of a Composite Advanced Sail." MSSTATE, 2006. http://sun.library.msstate.edu/ETD-db/theses/available/etd-04162006-162337/.
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Structural reinforcement layout optimization can be a very useful tool in the preliminary stages of design. In this research, sizing optimization techniques are used to generate results very similar to traditional layout optimization techniques with advantages in composite modeling and available strength and stability responses. Both linear and nonlinear sizing-to-design variable relationships are applied to a composite advanced sail design problem with high and low-complexity finite element models. An alternate methodology based on fractional-factorial-design and response surface modeling is also presented with promising results for finding the globally optimum reinforcement layout design. The stiffener layouts obtained from the different approaches are used to define an improved stiffener layout for sizing optimization for minimum weight. A weight savings of more than 19% is obtained over a baseline model using these methodologies.
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Meink, Troy Edward. "Structural Analysis and Design of Composite Isogrid Panels for Increased Buckling Efficiency." The Ohio State University, 1995. http://rave.ohiolink.edu/etdc/view?acc_num=osu1392917065.
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Roy, Steven. "Mechanical modeling and testing of a composite helicopter structure made by resin transfer moulding." Thesis, McGill University, 2009. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=40829.
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The design and mechanical performance of a helicopter horizontal stabilizer slat made by resin transfer moulding (RTM) can be evaluated with finite element analysis (FEA). To verify the validity of the assumptions used in the modelling of the slat structure, static mechanical tests were performed on prototype slats which were half of the full-size length. The slat complex boundary conditions were simplified to make static mechanical testing possible. Two fixtures were designed and built to introduce simplified loads in specimens with two different bracket configurations: a full and a half bracket. A finite element (FE) model of the specimens was made with shell elements and the finite element solution was compared with the experimental results. In most cases, comparison between the finite element analysis solution and experimental results showed good agreement in terms of structure stiffness, strength, strain and damage location. It is believed that out-of-plane stresses should be considered to improve the finite element solution accuracy.Le design et la performance mécanique d’un bec de bord d’attaque de stabilisateur horizontal d’hélicoptère fabriqué par moulage par injection sur renfort (« resin transfer moulding ») peuvent être évalués par des analyses par éléments finis. Pour vérifier la validité des hypothèses utilisées dans la modélisation du bec de bord d’attaque, des essais mécaniques ont été effectués sur des prototypes demi-longueurs. Les conditions frontières complexes ont été simplifiées pour rendre les essais mécaniques possibles. Deux gabarits ont été conçus et construits pour introduire les chargements simplifiés dans les pièces d’essai possédant deux configurations de support: un support complet et un demi-support. Un modèle par éléments finis des pièces d’essai a été réalisé avec des éléments de type membrane et la solution est comparée avec les résultats expérimentaux. Dans la plupart des cas, la comparaison entre la solution par éléments finis et les résultats expérimentaux coïncide concernant la rigidité de la structure, la résistance, l’allongement et la localisation de l’endommagement. Les contraintes hors du plan devraient être considérées pour améliorer la précision de la solution par éléments finis.
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Kousourakis, Asimenia, and asimeniak@hotmail com. "Mechanical Properties and Damage Tolerance of Aerospace Composite Materials Containing CVM Sensors." RMIT University. Aerospace, Mechanical and Manufacturing Engineering, 2009. http://adt.lib.rmit.edu.au/adt/public/adt-VIT20090506.095922.
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The PhD thesis evaluates the mechanical properties and damage tolerance of aerospace carbon/epoxy laminates containing long, narrow interlaminar galleries. The term 'galleries' refers to thin and long holes in a laminate used for the installation of small measuring devices, such as structural health monitoring (SHM) sensors. The galleries considered in this study are similar to those used in a novel SHM system known as 'Comparative Vacuum Monitoring (CVM)'. CVM was developed by the Australian company - Structural Monitoring Systems (SMS) - for damage detection in aircraft structures. CVM is a SHM system that utilises pressure differentials between a parallel series of galleries at atmospheric or low pressure to detect damage initiation and propagation. Thus far, CVM has been used for the monitoring of surface cracks in metallic structures using surface mounted sensors. Recent research has also demonstrated that it may be possible to monitor damage along the bond- line of both metallic and composite joints using CVM. The ability of CVM sensors to detect delamination damage inside composite structures is less well understood. It is envisaged that CVM can be used for the through-life health monitoring of composite aircraft structures prone to delamination damage. However, a major concern with applying CVM to composite laminates is the open-hole design of the galleries that may initiate damage growth under external loading. Material property data, structural tests, and models for predicting the properties of laminates containing galleries is needed before CVM technology can be certified for use in aircraft composite structures. The primary objectives of this PhD thesis are the development of an optimum process method for introducing multiple interlaminar CVM galleries in composite laminates; the development of a validated model for calculating changes to the mechanical properties of laminates containing CVM galleries; and the determination of optimum CVM gallery shape, size and orientation combinations for minimising the effect of the galleries on the mechanical properties of laminates. The effects of the shape, size and orientation of CVM galleries on the mechanical properties of carbon/epoxy laminates are evaluated by an extensive experimental research program, and the results are presented in the thesis. The properties investigated include the in-plane tensile and compressive properties, tensile and compressive fatigue life, through-thickness tensile strength, interlaminar shear strength, mode I and mode II interlaminar fracture toughness, and impact damage resistance. The results from tensile tests on lap-joints and T-joints containing CVM galleries are also presented.
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Nicolas, Matthew James. "Structural analysis and testing of a carbon-composite wing using fiber Bragg gratings." Thesis, Mississippi State University, 2013. http://pqdtopen.proquest.com/#viewpdf?dispub=1536133.
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The objective of this study was to determine the deflected wing shape and the out-of-plane loads of a large-scale carbon-composite wing of an ultralight aerial vehicle using Fiber Bragg Grating (FBG) technology. The composite wing was instrumented with an optical fiber on its top and bottom surfaces positioned over the main spar, resulting in approximately 780 strain sensors bonded to the wings. The strain data from the FBGs was compared to that obtained from four conventional strain gages, and was used to obtain the out-of-plane loads as well as the wing shape at various load levels using NASA-developed real-time load and displacement algorithms. The composite wing measured 5.5 meters and was fabricated from laminated carbon uniaxial and biaxial prepreg fabric with varying laminate ply patterns and wall thickness dimensions. A three-tier whiffletree system was used to load the wing in a manner consistent with an in-flight loading condition.
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Bail, Justin L. "Non-desctructive investigation & FEA correlation on an aircraft sandwich composite structure." Akron, OH : University of Akron, 2007. http://rave.ohiolink.edu/etdc/view?acc%5Fnum=akron1196702586.
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Thesis (M.S.)--University of Akron, Dept. of Civil Engineering, 2007."December, 2007." Title from electronic thesis title page (viewed 02/25/2008) Advisor, Wieslaw Binienda; Faculty readers, Craig Menzemer, Robert Goldbert; Department Chair, Wieslaw Binienda; Dean of the College, George K. Haritos; Dean of the Graduate School, George R. Newkome. Includes bibliographical references.