The organization of behavior: A neuropsychological theory

Aircraft wings are a compromise that allows the aircraft to fly at a range of flight conditions, but the performance at each condition is sub-optimal. The ability of a wing surface to change its geometry during flight has interested researchers and designers over the years as this reduces the design compromises required. Morphing is the short form for metamorphose; however, there is neither an exact definition nor an agreement between the researchers about the type or the extent of the geometrical changes necessary to qualify an aircraft for the title ‘shape morphing.’ Geometrical parameters that can be affected by morphing solutions can be categorized into: planform alteration (span, sweep, and chord), out-of-plane transformation (twist, dihedral/gull, and span-wise bending), and airfoil adjustment (camber and thickness). Changing the wing shape or geometry is not new. Historically, morphing solutions always led to penalties in terms of cost, complexity, or weight, although in certain circumstances, thes...

A Review of Morphing Aircraft

Dynamic fracture mechanics

Probability, Reliability and Statistical Methods in Engineering Design

Past, present and future of nonlinear system identification in structural dynamics

Stiffened thin-walled pressure vessels are widely used, especially in the aerospace, marine, automobile, and chemical industries. Accurate structural analysis of such complex structures under external and internal pressure loads is key to making critical design decisions during the conceptual and preliminary stages. With the advancements in adaptive optimization frameworks, multi-fidelity finite element (FE) models (shell-solid coupled) are increasingly sought in the early stages of design development for accurate and efficient analysis. However, the time and effort required to create such mixed finite element models with acceptable meshes for highly complex structures is still significant and is a major bottleneck in the FE modelling. Thus, automation of FE model creation and mesh generation is critical for advancements in technology. A major hurdle in automating highly complex structural FE models is the generation of robust meshing. This paper presents a framework for the automation of multi-fidelity model development from the solid and shell models using parameterization and scripting within the commercial FE packages. Additionally, issues related to the automated meshing of highly complex assemblies is discussed, and a volume decomposition scheme is proposed based on face partitions. A comparison of the all-solid, all-shell and different solid/shell multi-fidelity models of a highly complex stiffened thin-walled pressure vessel under external pressure and internal tank pressure is presented. Results reveal that automated generation of mixed models in an integrated manner including the geometry creation, meshing, and post-processing is capable of significant computational savings and reducing considerable interactive effort involved in analyzing complex structures. 

Automated FE Analysis of a Stiffened Tank Pressure Vessel using Shell-Solid Multi-Fidelity Modeling

I. Abstract Aerospace structures utilize honeycomb panels because the panels offer high stiffness to weight ratios. In these structural designs, members often require a combination of mechanical joining and adhesive bonding. Adhesive bonding is preferred because traditional mechanical attachments increase weight and introduce local stress concentration areas at the joint. In the space industry, where low weight is critical and high thermal and structural loads are the norm, structural members are often bonded sandwich composite materials that require adequate strength prediction at the structural joints. Insufficient adhesive joint strength can be catastrophic and costly, and as such there is intense focus in proper design for manufacturing and analysis of adhesively bonded structures. Although structural testing is necessary to verify space structure integrity, it cannot address the structural integrity of all the spacecraft joints and at 100% of predicted flight loading. That is why it is very important to be able to accurately and expeditiously predict the structural capability of bonded joints in a space structure. A typical method for analytical strength predictions is the finite element method. Finite element modeling of these structures is challenging because one often must address the modeling of the interface of thick sandwich composite panels to relatively thin bonded joints and properly address the discontinuity at the materials interface. In this paper, three different approaches using the finite element method are compared based on their ability to predict failure of various joints under different loading conditions, their ease of implementation and their computational rigor. All methods attempt to eliminate the effects of the discontinuity through a variety of modeling and postprocessing techniques.

Adhesively Bonded Joint Modeling Approaches using Linear Finite Element Analysis

Traveling to space requires a great deal of energy. This then limits the size of spacecraft accessible to transport to space. An optimization of a flexible tube that could be used as a satellite deployable structure was conducted by varying the cross section of the tube and its composite material properties. The material properties manipulated include the selection of a fiber, matrix, filler volume ratio, and orientation. HEEDS, a commercially available software, conducts the optimization process using the SHERPA algorithm. In the optimization, the finite element code, ABAQUS, iteratively performs two simulations. First, ABAQUS determines the stress distribution along the tube when wrapping the tube in its stored configuration. Second, ABAQUS finds the first natural frequency of the deployed structure. The objective function driving the optimization process is minimizing the weight and strain energy of the tube to create a light but highly flexible tube. This provides benefits of avoiding a violent deployment and lowering the dynamic response of the spacecraft during deployment. Three optimizations were performed with 1000 iterations each, using different initial geometries. While all three produce very similar results, one design converges to a clear best result. Using the best design, a series of deployment simulations are performed, using different boundary conditions to represent various scenarios. These boundary conditions include a free body dynamic response to deployment, a restricted response to only allow for rotation about the direction of deployment, and an increased damping deployment. Energy is dissipated differently comparing the results, showing that the most realistic case, being a free body deployment, has the lowest effect on the system. The spacecraft can dissipate energy by oscillating in the other axis. While damping does reduce the settling time for the deployed tube, there is notable oscillation in the middle of the tube seen in the transient state. 

/pdf/optimization-of-geometric-parameters-and-material-properties-2snnfwxz.pdf

Optimization of Geometric Parameters and Material Properties for a Deployable Space Structure

Honeycomb Core Buckling Due to Transverse Compression

Torsional vibrations and buckling of elastically supported thin-walled beams using orthogonal polynomials

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