Mohamed Jrad

Global/Local Optimization of Aircraft Wing Using Parallel Processing

The structural optimization of a cantilever aircraft wing with stiffeners and curvilinear spars and ribs is described. The decomposition technique for the wing structure is widely used for the optimization of a complex wing. The optimization procedure is divided into two subsystems: the global wing optimization, which determines the geometry and location of spars and ribs, and local panel optimization to further reduce wing weight. Because the design variables that are changed in the global step have an impact on those changed in the local step and vice versa, an iterative process that iterates between the global and local optimizations is employed. Particle swarm optimization and gradient-based optimization are used to perform integrated global/local optimization. Parallel computing is used, implemented using Python, to reduce the central processing unit time. The license cycle-check method and memory self-adjustment method are developed and applied in the parallel processing framework to optimize the us...

Global–Local Analysis of Composite Plate with Thin Notch

Laminated composites are becoming very popular in the aerospace industry because of their high strength-to-weight ratio and the fact that their properties can be tailored by changing different parameters like fiber and matrix materials, ratio of fiber to matrix volumes, fiber orientations, sequence of the plies, etc. However, because of their heterogeneity, an accurate analysis of composite laminates requires very high time and storage space. In the presence of structural discontinuities like cracks, delaminations, cutouts, etc., the computational complexity increases significantly. A possible alternative to reduce the computational complexity is the global–local analysis, which involves an approximate analysis of the whole structure followed by a detailed analysis of a significantly smaller region of interest. We investigate here the performance of the global–local scheme based on the finite-element method by comparing it to the traditional finite-element method. To do so, we conduct a two-dimensional st...

Integrated Global Wing and Local Panel Optimization of Aircraft Wing

The structural optimization of a cantilever aircraft wing with curvilinear spars and ribs and stiffeners is described. The design concept of reinforcing the wing structure using curvilinear stiffening members has been explored due to the development of novel manufacturing technologies like electron-beam-free-form-fabrication (EBF 3 ). For the optimization of a complex wing, a common strategy is to divide the optimization procedure into two subsystems: the global wing optimization which optimizes the geometry of spars, ribs and wing skins; and the local panel optimization which optimizes the design variables of local panels bordered by spars and ribs. The stiffeners are placed on the local panels to increase the stiffness and buckling resistance. The panel thickness, size and shape of stiffeners are optimized to minimize the structural weight. The geometry of spars and ribs greatly influences the design of stiffened panels. The interaction between the global wing optimization and the local panel optimization is usually computationally expensive. An approximate approach is implemented for the stiffened panel optimization to obtain approximate optimal panels using the polynomial curve fitting of a series of optimized panel thicknesses, so as to reduce the computational cost. The aircraft design is characterized by multiple disciplines: structures, aeroelasticity and buckling. Particle swarm optimization is used in the integration of global/local optimization to optimize the SpaRibs. A parallel computing technology has been developed in Python programming to reduce the CPU time. The license cycle-check method and memory self-adjustment method are two approaches that have been applied in the parallel framework in order to optimize the use of the resources by reducing the license and memory limitations and making the code robust. 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. The structural weight of the wing has been, therefore, reduced by 40% and the parallel implementation allowed a reduction in the CPU time by 89%.

A Framework for Damage Tolerance and Optimization of Stiffened Panels

A damage tolerance framework, EBF3PanelOpt, has been developed to design and analyze curvilinearly stiffened panels. The framework is written with the scripting language Python and it interacts with the commercial software MSC. Patran (for geometry and mesh creation), MSC. Nastran (for finite element analysis), and MSC. Marc (for damage tolerance analysis). The crack location is set to the location of the maximum value of the major principal stress while its orientation is set normal to the major principal axis direction. The effective stress intensity factor is calculated using the Virtual Crack Closure Technique and compared to the fracture toughness of the material in order to decide whether the crack will grow or not. The ratio of these two quantities is used as a constraint, along with the buckling constraint, Kreisselmeier and Steinhauser criteria, and crippling constraint. The EBF3PanelOpt framework is integrated within a two-step Particle Swarm Optimization in order to minimize the weight of the panel while satisfying the aforementioned constraints and using all the shape and thickness parameters as design variables. The result of the PSO is used then as an initial guess for the Gradient Based Optimization using only the thickness parameters as design variables and employing VisualDOC. Stiffened panel with two curvilinear stiffeners is optimized and significant reduction has been made for the panels weight. doi: 10.12783/SHM2015/3

SpaRibs Geometry Parameterization for Wings with Multiple Sections using Single Design

The SpaRibs topology of an aircraft wing has a significant effect on its structural behavior and stability as well as the flutter performance. The development of additive manufacturing techniques like Electron Beam Free Form Fabrication (EBF3) has made it feasible to manufacture aircraft wings with curvilinear spars, ribs (SpaRibs) and stiffeners. In this article a new global-local optimization framework for wing with multiple sections using curvilinear SpaRibs is described. A single design space is used to parameterize the SpaRibs geometry. This method has been implemented using MSC-PATRAN to create a broad range of SpaRibs topologies using limited number of parameters. It ensures C0 and C1 continuities in SpaRibs geometry at the junction of two wing sections with airfoil thickness gradient discontinuity as well as mesh continuity between all structural components. This method is advantageous in complex multi-disciplinary optimization due to its potential to reduce the number of design variables. For the global-local optimization the local panels are generated by an algorithm which is totally based on a set algebra on the connectivity matrix data. The great advantage of this method is that it is completely independent of the coordinates of the nodes of the finite element model. It is also independent of the order in which the elements are distributed in the FEM. The code is verified by optimizing of the CRM Baseline model at trim condition at Mach number equal to 0.85 for five different angle of attack (-2deg, 0deg,2deg,4deg and 6deg). The final weight of the wing is 19,090.61 lb. This value is comparable to that obtained by Qiang et al. 6 (19,269 lb).

Buckling Analysis of Curvilinearly Stiffened Composite Panels with Cracks

STIFFENERS attached to composite panels may significantly increase the overall buckling load of the resultant stiffened structure. First, buckling analysis of a composite panel with attached longitudinal stiffeners under compressive loads is performed using Ritz method with trigonometric functions. Results are then compared to those from ABAQUS FEA for different shell elements. The case of composite panel with one, two, and three stiffeners is investigated. The effect of the distance between the stiffeners on the buckling load is also studied. The variation of the buckling load and buckling modes with the stiffeners’ height is investigated. It is shown that there is an optimum value of stiffeners’ height beyond which the structural response of the stiffened panel is not improved and the buckling load does not increase. Furthermore, there exist different critical values of stiffener’s height at which the buckling mode of the structure changes. Next, buckling analysis of a composite panel with two straight stiffeners and a crack at the center is performed. Finally, buckling analysis of a composite panel with curvilinear stiffeners and a crack at the center is also conducted. ABAQUS is used for these two examples and results show that panels with a larger crack have a reduced buckling load. It is shown also that the buckling load decreases slightly when using higher order 2D shell FEM elements.