In this study an automated multidisciplinary design optimization code is developed for the minimum weight design of a composite wing box. The multidisciplinary static strength, aeroelastic stability, and manufacturing requirements are simultaneously addressed in a global optimization environment through a genetic search algorithm. The static strength requirements include obtaining positive margins of safety for all the structural parts. The modified engineering bending theory together with the coarse finite element model methodology is utilized to determine the stress distribution. The nonlinear effects, stemming from load redistribution in the structure after buckling occurs, are also taken into account. The buckling analysis is based on the Rayleigh-Ritz method and the Gerard method is used for the crippling analysis. The aeroelastic stability requirements include obtaining a flutter/divergence free wing box with a prescribed damping level. The root locus method is used for aeroelastic stability analysis. The unsteady aerodynamic loads in the Laplace domain are obtained from their counterparts in the frequency domain by using Rogers rational function approximations. The outer geometry of the wing is assumed fixed and the design variables included physical properties like thicknesses, cross sectional dimensions, the number of plies and their corresponding orientation angles. The developed code, which utilizes MSC/NASTRAN&reg / as a finite element solver, is used to design a single cell, wing box with internal metallic substructure and composite skins.

This thesis aims to establish a greater understanding of the effects of the single-sided robotic tufting process on component manufacturing costs and weight, and subsequently determine whether potential economic benefits can be realised through its implementation. To complete this assessment, the study uses the novel composite rib post application of the Next Generation Composite Wing (NGCW) project as an example case, investigating the component’s manufacturing cost, performance and weight relative to those of a non-tufted rib post with the same initial fabric layup, and a prepreg rib post with the same fibre orientation and fibre volume fraction. Detailed analyses of the tufted composite rib post manufacturing process have established the relationships between its manufacturing activities and the consumed resources, facilitating the creation of a novel cost model incorporating algorithms defining the robotic tufting process, which may be adapted for different geometries and tuft parameters. In addition, novel permeability data for the considered tufted preform have been obtained experimentally allowing the simulation of the liquid composite manufacturing process. The insertion of tufting in the studied arrangement was shown to decrease the in-plane permeability of the preform, reducing the characteristic by a factor of 3.5 in the direction perpendicular to the tuft seams, and 1.8 in the direction parallel. Completing structural analyses and further cost modelling, the study considers the manufacturing costs and weight of each of the studied components when performance is normalised. Implementing weight penalties including lifetime fuel consumption and lifetime revenue, appreciation is given to the potential lifetime costs generated by each of the structural solutions. Comparing the amalgamated costs of each component, relative economy has been evaluated both when considering equipment amortisation and also the use of pre-amortised equipment. Considering lifetime fuel cost/kg as the weight penalty, the novel composite rib post has greater economy within the pure pull-off load case without tufting due to its lower manufacturing costs. However, with the consideration of a more significant weight penalty, such as potential increased lifetime revenue, the use of the novel composite rib post results in superior economy, its lifetime cost/part is estimated to be 14% less than the component without tufting and 61% less than the lifetime cost/part of the prepreg component. Critical evaluations of the project’s methodologies have highlighted the various assumptions made due to the limitations of available information and means to perform certain validations. Under these circumstances the assumptions are believed to be the best representations of what is experienced in reality. As far as possible, errors believed to be the result of assumptions have been accounted for in rigorous error analyses to establish their propagation through the various calculations performed. These analyses have provided confidence in the results of the study, ensuring the significance of the determined economical differences between the investigated components.

The main aim of the this diploma thesis is the wing design of the UAV aircraft, including the appropriate material choice, calculation of the wing load and also strength analysis. Other goals include the design of the location and volume of fuel tanks, as well as the design of wingspan reduction after landing.

In this study, the structural optimization of a cruise missile wing is accomplished for the aerodynamic loads for four different flight conditions. The flight conditions correspond to the corner points of the V-n diagram. The structural analysis and optimization is performed using the ANSYS finite element program. In order to construct the flight envelope and to find the pressure distribution in each flight condition, FASTRAN Computational Fluid Dynamics program is used. The structural optimization is performed for two different wing configurations. In the first wing configuration all the structural members are made up of aluminum material. In the second wing configuration, the skin panels are all composite material and the other members are made up of aluminum material. The minimum weight design which satisfies the strength and buckling constraints are found for both wings after the optimization analyses.

This diploma thesis deals with the calculation of composite wing loading and its stressanalysis. Furthermore work contains design of static strength test wing.

An analytical model of cracked composite beams vibrating in coupled bending-torsion is developed. The beam is made of fiber-reinforced composite with fiber angles in each ply aligned in the same direction. The crack is assumed open. The local flexibility concept is implemented to model the open crack and the associated compliance matrix is derived. The crack introduces additional boundary conditions at the crack location and these effects in conjunction with those of material properties are investigated. Free vibration analysis of the cracked composite beam is presented. The results indicate that variation of natural frequencies in the presence of a crack is affected by the crack ratio and location, as well as the fiber orientation. In particular, the variation pattern is different as the magnitude of bending-torsion coupling changes due to different fiber angles. When bending and torsional modes are essentially decoupled at a certain fiber angle if there is no crack, the crack introduces coupling to the initially uncoupled bending and torsion. Based on the crack model, aeroelastic characteristics of an unswept composite wing with an edge crack are investigated. The cracked composite wing is modeled by a cracked composite cantilever and the inertia coupling terms are included in the model. An approximate solution on critical flutter and divergence speeds is obtained by Galerkin's method in which the fundamental mode shapes of the cracked wing model in free vibration are used. It is shown that the critical divergence/flutter speed is affected by the elastic axis location, the inertia axis location, fiber angles, and the crack ratio and location. Moreover, model-based crack detection (size and location) by changes in natural frequencies is addressed. The Cawley-Adams criterion is implemented and a new strategy in grouping frequencies is proposed to reduce the probability of measurement errors. Finally, sensitivity of natural frequencies to model parameter uncertainties is investigated. Uncertainties are modeled by information-gap theory and represented with a collection of nested sets. Five model parameters that may have larger uncertainties are selected in the analysis, and the frequency sensitivities to uncertainties in the five model parameters are compared in terms of two immunity functions. / Ph. D.

Indiana University-Purdue University Indianapolis (IUPUI) / In a multifidelity structural design process, depending on the required analysis, different levels of structural models are needed. Within the aerospace design, analysis and optimization community, there is an increasing demand for automatic generation of parametric feature tree (build recipe) attributed multidisciplinary models. Currently, this is mainly done by creating separate models for different disciplines such as mid-surface model for aeroelasticity, outer-mold line for aerodynamics and CFD, and built-up element model for structural analysis. Since all of these models are built independently, any changes in design parameters require updates on all the models which is inefficient, time-consuming and prone to deficiencies. In this research, Engineering Sketch Pad (ESP) is used to create attribution and maintain consistency between structural models with different fidelity levels. It provides the user with the ability to interact with a configuration by building and/or modifying the design parameters and feature tree that define the configuration. ESP is based an open-source constructive solid modeler, named OpenCSM, which is built upon the OpenCASCADE geometry kernel and the EGADS geometry generation system. The use of OpenCSM as part of the AFRL’s CAPS project on Computational Aircraft Prototype Syntheses for automatic commercial and fighter jet models is demonstrated. The rapid generation of parametric aircraft structural models proposed and developed in this work will benefit the aerospace industry with coming up with efficient, fast and robust multidisciplinary design standardization of aircraft structures. Metallic aircraft wings are usually not optimized to their fullest potential due to shortage of development time. With roughly \$1000 worth of potential fuel savings per pound of weight reduction over the operational life of an aircraft, airlines are trying to minimize the weight of aircraft structures. A stiffness based strategy is used to map the nodal data of the lower-order fidelity structural models onto the higher-order ones. A simple multi-fidelity analysis process for a parametric wing is used to demonstrate the advantage of the approach. The loads on the wing are applied from a stick model as is done in the industry. C program is created to connect the parametric design software ESP, analysis software Nastran, load file and design configuration file in CSV format. This problem gets compounded when it comes to optimization of composite wings. In this study, a multi-level optimization strategy to optimize the weight of a composite transport aircraft wing is proposed. The part is assumed to initially have some arbitrary number of composite super plies. Super plies are a concept consisting of a set of plies all arranged in the same direction. The thickness and orientation angles of the super plies are optimized. Then, each ply undergoes topometry optimization to obtain the areas of each super ply taking the least load so that it could be cut and removed. Each of the super plies are then optimized for the thickness and orientation angles of the sub plies. The work presented on this paper is part of a project done for Air Force Research Laboratory (AFRL) connecting the parametric geometry modeler (ESP) with the finite element solver (Nastran).

O σχεδιασμός και η ανάπτυξη μιας σύγχρονης αεροναυπηγικής κατασκευής περιλαμβάνει ως επιμέρους φάσεις (μεταξύ άλλων) τον αρχικό και τον προκαταρκτικό σχεδιασμό. Οι φάσεις αυτές έχουν ιδιαίτερη σημασία διότι εκεί δίνεται η αρχική μορφή και οι διαστάσεις της κατασκευής. Είναι γεγονός ότι η συμβατική σχεδίαση των βασικών δομικών στοιχείων των αεροσκαφών έχει φτάσει σε πολύ υψηλό επίπεδο βελτιστοποίησης που επιδέχεται πλέον μόνο μικρά περιθώρια περαιτέρω βελτίωσης. Οι σύγχρονες όμως απαιτήσεις των ελαφρών κατασκευών, όπως δραστική μείωση του βάρους, αύξηση του ωφέλιμου φορτίου κτλ. ωθεί τις αεροναυπηγικές βιομηχανίες στη δημιουργία δομών που ξεφεύγουν από τις παραδοσιακές (μη-συμβατικές δομές). Παράλληλα με τα παραπάνω γίνεται προσπάθεια για μερική αντικατάσταση μεταλλικών υλικών από σύνθετα υλικά στις πρωτεύουσες δομές αεροναυπηγικών κατασκευών. Για να σχεδιαστούν και να εξελιχθούν μη-συμβατικές αεροναυπηγικές δομές χωρίς να καταφύγει κάποιος σε εκτενείς πειραματικές δοκιμές, η σύγχρονη τάση είναι η ανάπτυξη και ο συνδυασμός προτύπων συμπεριφοράς στη λογική της εξομοίωσης των πειραματικών δοκιμών. Η εξομοίωση αυτή επιτυγχάνεται με τη βοήθεια ηλεκτρονικών υπολογιστών και κατάλληλων μεθόδων βασισμένων στη θεωρία των πινάκων (Πεπερασμένα Στοιχεία, Συνοριακά Στοιχεία κλπ.). Στη φάση του αρχικού και προκαταρκτικού σχεδιασμού η εφαρμογή των μεθοδολογιών προσομοίωσης δεν είναι πάντοτε εύκολη και απλή, λόγω των πολλαπλών αλλαγών στη γεωμετρία, το υλικό και τις κατασκευαστικές λεπτομέρειες που πραγματοποιούνται στη δομή κατά την επαναληπτική διαδικασία του σχεδιασμού. Επομένως, η αποκλειστική χρήση αριθμητικών μεθόδων ανάλυσης καθίσταται αναποτελεσματική από άποψη χρονικών απαιτήσεων, αν δεν συνοδεύεται από αναλυτικές ή ημιαναλυτικές προσεγγίσεις επιμέρους προβλημάτων του σχεδιασμού. Βασικό μέρος του προκαταρκτικού σχεδιασμού μιας πτέρυγας μη συμβατικής δομής αποτελεί η αποφυγή της αστοχίας του άνω τμήματός της, διότι οι λεπτότοιχες ενισχυμένες με δοκούς πλάκες που χρησιμοποιούνται στην κατασκευή υφίστανται λυγισμό λόγω των θλιπτικών φορτίσεων που κυρίως παραλαμβάνουν. Η διαστασιολόγηση των σύνθετων πλακών που φέρουν δοκούς ενίσχυσης στις κατασκευές αυτές απαιτούν συνήθως πλήθος επαναληπτικών υπολογισμών για διαφορετικές γεωμετρίες, φορτίσεις, οριακές συνθήκες κλπ. Η εξέταση της κάθε περίπτωσης μεμονωμένα με τη χρήση αριθμητικών μεθόδων καθιστά την επίλυση ολόκληρης της κατασκευής εξαιρετικά χρονοβόρα. Για το λόγο αυτό, στη φάση της αρχικής θεωρητικής μελέτης και της αρχικής διαστασιολόγησης η χρησιμοποίηση αναλυτικών μεθόδων για την εύρεση του κρίσιμου φορτίου λυγισμού πλακών με δοκούς ενίσχυσης οδηγεί στην εξοικονόμηση υπολογιστικού κόστους. Επομένως, η ανάπτυξη αναλυτικών ή ημιαναλυτικών μεθόδων προσδιορισμού των φορτίων λυγισμού ενισχυμένων με δοκούς συνθέτων πλακών και κελυφών θεωρείται πολύ σημαντική. Για τον σκοπό αυτό, στο πλαίσιο αυτής της διατριβής, αναπτύσσονται αναλυτικές και ημιαναλυτικές λύσεις για το λυγισμό πολύστρωτων πλακών ενισχυμένων με ενισχυτικές διαμήκεις δοκούς, οι οποίες ενσωματώνονται σαν κριτήρια στη μέθοδο διαστασιολόγησης της δομής. Η μεθοδολογία συμπληρώνεται με πλήθος άλλων κατάλληλων κριτηρίων για τον έλεγχο της αντοχής των δομικών στοιχείων της πτέρυγας καθώς και με κριτήρια για την επαναδιαστασιολόγηση των στοιχείων κατά την επαναληπτική διαδικασία της βελτιστοποίησης. Με τη μεθοδολογία που αναπτύσσεται διερευνούνται διατάξεις δομής πτερύγων από σύνθετα υλικά με πολυάριθμες κύριες δοκούς. Πιο συγκεκριμένα, αναπτύσσονται αναλυτικές/ημιαναλυτικές λύσεις ολικού και τοπικού λυγισμού πλακών που φέρουν δοκούς ενίσχυσης. Όσον αφορά τον ολικό λυγισμό αναπτύσσεται μια μεθοδολογία που βασίζεται στη μαθηματική μετατροπή μιας πλάκας που φέρει δοκούς ενίσχυσης σε μια ισοδύναμη ομογενή πλάκα. Η αναπτυχθείσα μεθοδολογία ομογενοποίησης των ενισχυμένων πλακών εμφανίζει σημαντικά πλεονεκτήματα σε σύγκριση με τις αντίστοιχες ήδη υπάρχουσες. Παράλληλα, η ενεργειακή μέθοδος Rayleigh-Ritz εφαρμόζεται για τη λύση προβλημάτων λυγισμού μερικώς ανισότροπων πλακών με ενισχυτικές δοκούς από σύνθετα υλικά, λαμβάνοντας διακριτά υπόψη τις ενισχυτικές δοκούς. Όσον αφορά το πρόβλημα του τοπικού λυγισμού, αναπτύσσεται μια νέα μεθοδολογία για την εύρεση των κρίσιμων φορτίων τοπικού λυγισμού λεπτότοιχης πλάκας με χρήση ενεργειακών μεθόδων. Το μαθηματικό μοντέλο που χρησιμοποιείται για την περίπτωση του τοπικού λυγισμού της επικάλυψης είναι η απομόνωση του τμήματος της επικάλυψης μεταξύ δυο ενισχυτικών δοκών και η αντικατάσταση της δυσκαμψίας της υπόλοιπης πλάκας με ελατήρια μεταβλητής δυσκαμψίας. Η μεθοδολογία αυτή επεκτείνεται και στον προσδιορισμό της μεταλυγισμικής συμπεριφοράς μιας πλάκας ενισχυμένης με διαμήκεις δοκούς. Οι παραπάνω μεθοδολογίες υπολογισμού του κρίσιμου φορτίου λυγισμού που αναπτύσσονται, στα πλαίσια αυτής της διατριβής, εφαρμόζονται στη διαστασιολόγηση πτέρυγας μη συμβατικής δομής από σύνθετα υλικά με πολυάριθμες κύριες δοκούς, σε αντίθεση με τις συμβατικές πτέρυγες (με δύο κύριες δοκούς). Η ανάλυση τάσεων της πτέρυγας πραγματοποιείται με τη βοήθεια της μεθόδου των πεπερασμένων στοιχείων. Η τελική διαστασιολόγηση επιτυγχάνεται με επαναληπτική διαδικασία βελτιστοποίησης βασισμένη σε αναλυτικές και ημιαναλυτικές σχέσεις. Με τον τρόπο αυτό, συγκρίνεται λεπτομερώς η συμβατική δομή πτέρυγας με 2 κύριες δοκούς και οι αντίστοιχες πτέρυγες με 4, 5 και 6 κύριες δοκούς. Για την περαιτέρω βελτιστοποίηση της συμπεριφοράς της πτέρυγας, διερευνάται η επίδραση που έχει η αλλαγή των μηχανικών ιδιοτήτων του υλικού και των επιτρεπόμενων ορίων παραμόρφωσης στη δυνατότητα ελαχιστοποίησης της μάζας της πτέρυγας. Υπολογίστηκε ότι κάτω από συγκεκριμένες συνθήκες η χρήση της μη συμβατικής πτέρυγας μπορεί να οδηγήσει σε μείωση μάζας μέχρι και 12%. / The design and development of a modern aerospace structure consists of many design stages. The most important stages are the conceptual and the preliminary where the initial sizing of the structure is obtained. It is known that the conventional design of the aircraft’s main structural members has reached a high optimization level, where margins for further improvement are small. The current demands of the lightweight structures such as weight reduction, payload increase etc. have led the aerospace industries develop unconventional structures and partially substitute the metallic materials of the primary structures with composites. The current trend of designing and evolving unconventional aerospace structures, without performing extended experimental tests, leads to the development of behavior models. The simulation of the experimental tests (through the behavior models) is achieved using high performance computers and numerical methods (Finite Element Method, Boundary Element Method etc). To apply simulation methods during the conceptual and preliminary stage is not an easy task. Most of the difficulties are the numerous geometrical, material parameters and the structural details that alter during the iterative process of the design. So, the exclusive usage of numerical analysis methods becomes very time consuming, if it is not accompanied by analytical or semi analytical methods of the sub-problems of the design. Part of the preliminary design of an unconventional wing structure is to prevent upper skin from failure. The stiffened panels that comprise the upper skin of the wing suffer from buckling due to the applied compressive loads. The sizing of the composite stiffened panels usually requires numerous of iterative calculations for various geometries, loading and boundary conditions etc. The examination of each case separately, with the use of numerical methods, results to time consuming analyses of the entire structure. Therefore, the development of appropriate analytical or semi analytical methods for estimating stiffened panels’ critical buckling load is of great importance. For this purpose, in the present thesis, analytical and semi analytical methodologies are developed for estimating the critical buckling load of stiffened panels. The developed methodologies are incorporated as design criteria in the sizing routine of the entire structure. The sizing routine comprises additional sizing criteria for checking the strength of wing’s structural members at each phase of the iterative process. Applying the developed sizing routine in various wing configurations made of composite materials, multispar wing designs are studied. Specifically, analytical and semi analytical methods for global and local buckling problems of stiffened panels are developed. The methodology of global buckling problems is based on the mathematical conversion of a stiffened panel to an equivalent homogeneous panel. The developed method of homogenization of stiffened panels appears to have significant advantages over the already existed homogenization methods. Additionally, the energy method Rayleigh-Ritz is applied for solving global buckling problems of stiffened panels with partial anisotropy considering discrete stiffeners. Regarding local buckling problems of stiffened panels, a new methodology is developed for estimating the critical local buckling load with the use of energy methods. The approach considers the stiffened panel segment located between two stiffeners, while the remaining panel is replaced by equivalent transverse and rotational springs of varying stiffness, which act as elastic edge supports. The buckling analysis of the segment provides an accurate and conservative prediction of the panel local buckling behavior. Consequently, the developed methodology is extended in the prediction of post-buckling response of stiffened panels where skin has undergone local buckling. The developed methodologies for calculating the critical buckling load are applied for sizing the wing members of an unconventional wing (multispar configuration) from composite materials. An efficient methodology based on fast Finite Element (FE) stress analysis combined to analytically formulated design criteria is presented for the initial sizing of a large scale composite component. A detailed comparison between optimized designs of conventional (2-spar) and three alternative wing configurations which comprise 4-, 5-, and 6-spars for the wing construction is performed. In order to understand the effect of different material properties, as well as the variation of maximum strain level allowed in the total wing mass, parametric analyses are performed for all wing configurations considered. It arises that under certain conditions the multispar configuration demonstrates significant advantages over the conventional design. This would lead to a mass reduction of 12%.