Manufacturing considerations have been incorporated into the design optimization of a blade-stiffened composite panel. For the manufacturing analysis, a one-dimensional resin film infusion model is developed to compute the infiltration time of the resin into a fabric preform of the panel. Results are presented showing the effects of structurally important design variables, such as cross-sectional geometry and material properties, on the manufacturing performance of the panel. In addition, the effects of manufacturing process variables, such as pressure and temperature, on the structural performance are studied. The structural problem is formulated to minimize the panel mass subject to buckling constraints. A simplified buckling analysis model for the panel is used to compute the critical buckling loads. The objective of the manufacturing problem is to minimize the resin infiltration time. Optimum panel designs for the manufacturing and structures problems alone, as well as for the combined problem, are generated using a genetic algorithm. These results indicate a strong connection between the structures and manufacturing design variables and trade-offs between the responses, illustrating that a multidisciplinary approach to the problem is essential to incorporating manufacturing into the preliminary design stage. / Master of Science

Diese Forschung zielte darauf ab, die vorgespannten und versteiften Membranstrukturen unter Verwendung eines experimentellen Ansatzes und einer Computersimulation konzeptionell zu entwickeln. Die physikalische Methode der Formfindung beinhaltete das vorgespannte Gewebe mit dem verleimten Gitter aus den Holzstäben. Die Relaxation der belasteten Membran trug zur Bildung der spezifischen antiklastischen hyparischen Oberfläche durch Energiefreisetzung bei. Der Einfluss der starren Elemente Muster, Intensität und Richtung der Vorspannung auf die Endform wurde untersucht. Auch die Tensegrity-Strukturen wurden nach dem gleichen Formfindungsweg gebaut. Diese Experimente führten zur Modellierung der resultierenden Proben mit parametrischen Entwurfswerkzeugen, nämlich Rhino und Grasshopper. Die Optimierung der endgültigen Form erfolgte durch Änderung von Parametern wie der Versteifungskonfiguration und der Membranfestigkeit. Dieser digitale Ansatz demonstrierte die erfolgreiche Simulation und Rationalisierung der betrachteten Strukturen. Darüber hinaus können die endgültigen Modelle für weitere Statik und BIM verwendet werden. Berücksichtigte Membranstrukturen weisen ein sehr effizientes Tragverhalten auf. Sie zeichnen sich durch geringes Gewicht, hohe Lichtdurchlässigkeit und die Möglichkeit aus, große, säulenfreie Nutzräume zu schaffen. Die gefährlichsten Belastungen für Membrankonstruktionen sind Wind und Seewasser. In der Praxis sind PTFE-beschichtete Glasfasergewebe und PVC-beschichtete Polyestergewebe für vorgespannte und versteifte Membrankonstruktionen am besten geeignet. Die Rolle steifer Elemente können Stahlprofile oder Metallrohre spielen. Die durchschnittliche Zeit für den Bau einer Membrankonstruktion beträgt 6-15 Monate. Die resultierenden vorgespannten und versteiften Membrankonstruktionen können als Pavillons, Dächer und Markisen verwendet werden. Sie zeichnen sich durch eine spektakuläre architektonische Aussicht und ein sehr effektives Struktursystem aus. Darüber hinaus zeichnen sich Membranzugtragwerke durch eine hohe Ökoeffizienz und Nachhaltigkeit im Vergleich zu anderen Bauweisen aus. / This research aimed to develop conceptually the pretensioned and stiffened membrane structures, using an experimental approach and computer simulation. The physical method of form finding included the pretensioned fabric with the glued grid made of the wooden sticks. Relaxation of the stressed membrane contributed to forming the specific anticlastic hyparic surface by energy release. The influence of the rigid elements pattern, intensity and direction of pretensioning on the final shape was investigated. The tensegrity structures were also built applying the same form finding way. These experiments led to the modelling of resulting samples with parametric design tools, namely Rhino and Grasshopper. Optimization of the final shape was carried out by changing parameters such as stiffenings configuration and membrane strength. This digital approach demonstrated successful simulation and rationalization of considered structures. Moreover, the final models can be used for further structural analysis and BIM. Considered membrane structures have very efficient load-bearing behavior. They are characterized by small weight, high light transmission and the ability to create large usable spaces free from columns. The most dangerous loads for membrane structures are wind and ponding. In practice, PTFE coated glass-fibre fabric and PVC coated polyester fabric are most suitable for pretensioned and stiffened membrane structures. The role of stiff elements can be played by steel profiles or metal tubes. The average time for the construction of a membrane structure is 6-15 months. Resulted pretensioned and stiffened membrane structures can be used as pavilions, roofs and awnings. They are distinguished by spectacular architectural view and very effective structural system. In addition, membrane tensile structures are characterized by high eco-efficiency and sustainability compared to other types of construction.

No description available.

Large-aspect-ratio wings and composite structures both have been considered for the next-generation civil transport aircraft to achieve improved aerodynamic efficiency and to save aircraft structural weight. The use of the large-aspect-ratio and the light-weight composite wing can lead to an enhanced flexibility of the aircraft wing, which may cause many aeroelastic problems such as large deflections, increased drag, onset of flutter, loss of control authority, etc. Aeroelastic tailoring, internal structural layout design and aerodynamic wing shape morphing are all considered to address these aeroelastic problems through multidisciplinary design, analysis and optimization (MDAO) studies in this work. Performance Adaptive Aeroelastic Wing (PAAW) program was initiated by NASA to leverage the flexibility associated with the use of the large-aspect-ratio wings and light-weight composite structures in a beneficial way for civil transport aircraft wing design. The biologically inspired SpaRibs concept is used for aircraft wing box internal structural layout design to achieve the optimal stiffness distribution to improve the aircraft performance. Along with the use of the active aeroelastic wing concept through morphing wing shape including the wing jig-shape, the control surface rotations and the aeroelastic tailoring scheme using composite laminates with ply-drop for wing skin design, a MDAO framework, which has the capabilities in total structural weight minimization, total drag minimization during cruise, ground roll distance minimization in takeoff and load alleviation in various maneuver loads by morphing its shape, is developed for designing models used in the PAAW program. A bilevel programming (BLP) multidisciplinary design optimization (MDO) architecture is developed for the MDAO framework. The upper-level optimization problem entails minimization of weight, drag and ground roll distance, all subjected to both static constraints and the global dynamic requirements including flutter mode and free vibration modes due to the specified control law design for body freedom flutter suppression and static margin constraint. The lower-level optimization is conducted to minimize the total drag by morphing wing shape, to minimize wing root bending moment by scheduling flap rotations (a surrogate for weight reduction), and to minimize the takeoff ground roll distance. Particle swarm optimization and gradient-based optimization are used, respectively, in the upper-level and the lower-level optimization problems. Optimization results show that the wing box with SpaRibs can further improve the aircraft performances, especially in a large weight saving, as compared to the wing with traditional spars and ribs. Additionally, the nonuniform chord control surface associated with the wing with SpaRibs achieve further reductions in structural weight, total drag and takeoff ground roll distance for an improved aircraft performance. For a further improvement of the global wing skin panel design, an efficient finite element approach is developed in designing stiffened composite panels with arbitrarily shaped stiffeners for buckling and vibration analyses. The developed approach allows the finite element nodes for the stiffeners and panels not to coincide at the panel-stiffeners interfaces. The stiffness, mass and geometric stiffness matrices for the stiffeners can be transformed to those for the panel through the displacement compatibility at their interfaces. The method improves the feasible model used in shape optimizing by avoiding repeated meshing for stiffened plate. Also, it reduces the order of the finite element model, a fine mesh typically associated with the skin panel stiffened by many stiffeners, for an efficient structural analysis. Several benchmark cases have been studied to verify the accuracy of the developed approach for stiffened composite panel structural analyses. Several parametric studies are conducted to show the influence of stiffener shape/placement/depth-ratio on panel's buckling and vibration responses. The developed approach shows a potential benefit of using gradient-based optimization for stiffener shape design. / Ph. D.

A global/local framework for multidisciplinary optimization of generalized aircraft wing structure has been developed. The concept of curvilinear stiffening members (spars, ribs and stiffeners) has been applied in the optimization of a wing structure. A global wing optimization framework EBF3WingOpt, which integrates the static aeroelastic, flutter and buckling analysis, has been implemented for exploiting the optimal design at the wing level. The wing internal structure is optimized using curvilinear spars and ribs (SpaRibs). A two-step optimization approach, which consists of topology optimization with shape design variables and size optimization with thickness design variables, is implemented in EBF3WingOpt. A local panel optimization EBF3PanelOpt, which includes stress and buckling evaluation criteria, is performed to optimize the local panels bordered by spars and ribs for further structural weight saving. The local panel model is extracted from the global finite element model. The boundary conditions are defined on the edges of local panels using the displacement fields obtained from the global model analysis. The local panels are optimized to satisfy stress and buckling constraints. Stiffened panel with curvilinear stiffeners is implemented in EBF3PanelOpt to improve the buckling resistance of the local panels. The optimization of stiffened panels has been studied and integrated in the local panel optimization. EBF3WingOpt has been applied for the optimization of the wing structure of the Boeing N+2 supersonic transport wing and NASA common research model (CRM). The optimization results have shown the advantage of curvilinear spars and ribs concept. The local panel optimization EBF3PanelOpt is performed for the NASA CRM wing. The global-local optimization framework EBF3GLWingOpt, which incorporates global wing optimization module EBF3WingOpt and local panel optimization module EBF3PanelOpt, is developed using MATLAB and Python programming to integrate several commercial software: MSC.PATRAN for pre and post processing, MSC.NASTRAN for finite element analysis. An approximate optimization method is developed for the stiffened panel optimization so as to reduce the computational cost. The integrated global-local optimization approach has been applied to subsonic NASA common research model (CRM) wing which proves the methodology's application scaling with medium fidelity FEM analysis. Both the global wing design variables and local panel design variables are optimized to minimize the wing weight at an acceptable computational cost. / Ph. D.

Unitized aircraft structures have the potential to be more efficient than current aircraft structures. The Electron Beam Freeform Fabrication (EBF3) process can be used to manufacture unitized aircraft structures. The structural efficiency of blade stiffened panels made with EBF3 was compared to panels made by integrally machining from thick plate. The panels were tested under two load cases in a combined compression-shear load test fixture. One load case tested the panels' responses to a higher compressive load than the shear load. The second load case tested the panels' responses to an equal compressive and shear load. Finite element analysis was performed to compare with the experimental results. The EBF3 panels failed at a 18.5% lower buckling load than the machined panels when loaded mostly in compression but at an almost two times higher buckling load than the machined panels when the shear matched the compressive load. The finite element analysis was in good agreement with the experimental results prior to buckling. The results demonstrate that the EBF3 process has the capabilities of manufacturing stiffened panels that behave similarly to machined panels prior to buckling. Once the EBF3 panels buckled, the buckled shape of the EBF3 panels was different from the machined panels, generally buckling in the opposite direction of what was observed with the machined panels. This was also expected based on the finite element analysis. The different post-buckling response between the two manufacturing techniques was attributed to the residual stress and associated distortion induced during the EBF3 manufacturing process. / Master of Science

This thesis presents the results of 67 ABAQUS elasto-plastic Riks ultimate strength analyses of cross-stiffened panels. These panels cover a wide range of typical geometries. Uniaxial compression is applied to the panels, and in some cases combined with lateral pressure. For eight of the panels full-scale experimental results are available, and these verified the accuracy of the ABAQUS results. The 67 ABAQUS results were then compared to the ultimate strength predictions from the computer program ULSAP. In all but 10 cases the ULSAP predicted strength is within 30% of the ABAQUS value, and in all but 4 cases the predicted failure mode also agrees with that of ABAQUS. In one case the ULSAP predicted ultimate strength is 51% below the experimental value, and so this case is studied in detail. The discrepancy is found to be caused by the method which ULSAP uses for panels that experience overall collapse initiated by beam-column-type failure. The beam-column method program ULTBEAM is used to predict the ultimate strength of the 61 panels that ULSAP predicts to fail due to overall collapse of the stiffeners and plating which may or may not be triggered by yielding of the plate-stiffener combination at the midspan (Mode III or III-1). ULTBEAM is found to give more accurate results than ULSAP for Mode III or III-1 failure. Future work is recommended to incorporate ULTBEAM into ULSAP to predict the ultimate strength of panels that fail in Mode III or III-1. / Master of Science

In this study, the &ldquo / Free Flexural (or Bending) Vibrations of Stiffened Plates or Panels&rdquo / are investigated in detail. Two different Groups of &ldquo / Stiffened Plates&rdquo / will be considered. In the first group, the &ldquo / Type 4&rdquo / and the &ldquo / Type 6&rdquo / of &ldquo / Group I&rdquo / of the &ldquo / Integrally-Stiffened and/or Stepped-Thickness Plate or Panel Systems&rdquo / are theoretically analyzed and numerically solved by making use of the &ldquo / Mindlin Plate Theory&rdquo / . Here, the natural frequencies and the corresponding mode shapes, up to the sixth mode, are obtained for each &ldquo / Dynamic System&rdquo / . Some important parametric studies are also presented for each case. In the second group, the &ldquo / Class 2&rdquo / and the &ldquo / Class 3&rdquo / of the &ldquo / Bonded and Stiffened Plate or Panel Systems&rdquo / are also analyzed and solved in terms of the natural frequencies with their corresponding mode shapes. In this case, the &ldquo / Plate Assembly&rdquo / is constructed by bonding &ldquo / Stiffening Plate Strips&rdquo / to a &ldquo / Base Plate or Panel&rdquo / by dissimilar relatively thin adhesive layers. This is done with the purpose of reinforcing the &ldquo / Base Plate or Panel&rdquo / by these &ldquo / Stiffening Strips&rdquo / in the appropriate locations, so that the &ldquo / Base Plate or Panel&rdquo / will exhibit satisfactory dynamic response. The forementioned &ldquo / Bonded and Stiffened Systems&rdquo / may also be used to repair a damaged (or rather cracked) &ldquo / Base Plate or Panel&rdquo / . Here in the analysis, the &ldquo / Base Plate or Panel&rdquo / , the &ldquo / Stiffening Plate Strips&rdquo / as well as the in- between &ldquo / adhesive layers&rdquo / are assumed to be linearly elastic continua. They are assumed to be dissimilar &ldquo / Orthotropic Mindlin Plates&rdquo / . Therefore, the effects of shear deformations and rotary moments of inertia are considered in the theoretical formulation. In each case of the &ldquo / Group I&rdquo / and &ldquo / Group II&rdquo / problems, the &ldquo / Governing System of Dynamic Equations&rdquo / for every problem is reduced to the &ldquo / First Order Ordinary Differential Equations&rdquo / . In other words the &ldquo / Free Vibrations Problem&rdquo / , in both cases, is an &ldquo / Initial and Boundary Value Problem&rdquo / is reduced to a &ldquo / Two- Point or Multi-Point Boundary Value Problem&rdquo / by using the present &ldquo / Solution Technique&rdquo / . For this purpose, these &ldquo / Governing Equations&rdquo / are expressed in &ldquo / compact forms&rdquo / or &ldquo / state vector&rdquo / forms. These equations are numerically integrated by the so-called &ldquo / Modified Transfer Matrix Method (MTMM) (with Interpolation Polynomials)&rdquo / . In the numerical results, the mode shapes together with their corresponding non-dimensional natural frequencies are presented up to the sixth mode and for various sets of &ldquo / Boundary Conditions&rdquo / for each structural &ldquo / System&rdquo / . The effects of several important parameters on the natural frequencies of the aforementioned &ldquo / Systems&rdquo / are also investigated and are graphically presented for each &ldquo / Stiffened and Stiffened and Bonded Plate or Panel System&rdquo / . Additionally, in the case of the &ldquo / Bonded and Stiffened System&rdquo / , the significant effects of the &ldquo / adhesive material properties&rdquo / (i.e. the &ldquo / Hard&rdquo / adhesive and the &ldquo / Soft&rdquo / adhesive cases) on the dynamic response of the &ldquo / plate assembly&rdquo / are also presented.

No description available.

This thesis presents a study about the determination of the stress intensity factors in cracked sheets with riveted stiffeners. Stress intensity factors are determined with both analytical method and finite element method for different combination of rivet/stringer spacing and stringer to sheet stiffness ratio. Analytical part of the thesis is a replication of the original study of Poe which assumes rigid rivet connections with no stringer offset. In the analytical part, the whole systems of equations of Poe are re-derived, and it is shown that there are two typographical errors in the expressions for the calculation of the influence coefficients of the cracked sheet and the stringer. Major objective of the analytical part is to develop a computer code which calculates the variation of the normalized stress intensity factor with the crack length for any combination of rivet/stringer spacing and stringer to sheet stiffness ratio. Analytical part of the study also covers the effect of broken stiffener on the stress intensity factor of the cracked sheet. The stress intensity factors of stiffened cracked sheets are calculated by the finite element method by incorporating fastener flexibility and stringer offset. Finite element solutions are performed by Franc2D/L and Abaqus, and comparisons are made. The effect of geometry, fastener flexibility, and stringer offset on the stress intensity factors are studied by presenting normalized stress intensity factor versus crack length curves. Finally, as a case study a sample damage tolerant stiffened panel is designed according to FAR 25 safety criteria. Experiments are performed for determining mechanical and crack growth properties of Al 2124 which is used as the material in the case study. Present study showed that the most significant effect on the stress intensity factor is seen when stringer-cracked sheet offset is included in the analysis model.