Cost optimization of composite aircraft structures including variable laminate qualities

Fibre Reinforced Plastics (FRPs) have been used in many practical structural applications due to their excellent strength and weight characteristics as well as the ability for their properties to be tailored to the requirements of a given application. Thus, designing with FRPs can be extremely challenging, particularly when the number of design variables contained in the design space is large. For example, to determine the ply orientations and the material properties optimally is typically difficult without a considered approach. Optimization of composite structures with respect to the ply angles is necessary to realize the full potential of fibre-reinforced materials. Evaluating the fitness of each candidate in the design space, and selecting the most efficient can be very time consuming and costly. Structures composed of composite materials often contain components which may be modelled as rectangular plates or cylindrical shells, for example. Modelling of components such as plates can be useful as it is a means of simplifying elements of structures, and this can save time and thus cost. Variations in manufacturing processes and user environment may affect the quality and performance of a product. It is usually beneficial to account for such variances or tolerances in the design process, and in fact, sometimes it may be crucial, particularly when the effect is of consequence. The work conducted within this project focused on methodologies for optimally designing fibre-reinforced laminated composite structures with the effects of manufacturing tolerances included. For this study it is assumed that the probability of any tolerance value occurring within the tolerance band, compared with any other, is equal, and thus the techniques are aimed at designing for the worst-case scenario. This thesis thus discusses four new procedures for the optimization of composite structures with the effects of manufacturing uncertainties included.

The demand for more efficient commercial aircraft with lower maintenance and operation cost has promoted more extensive use of composite materials for a significant reduction in weight and fuel consumption. Making a structure as light as possible suggests solving an optimization problem with the goal of mass minimization. To avoid failure modes related to composite structures, such as delamination and transverse matrix crack propagation, some design guidelines are commonly suggested by aircraft industries. These guidelines are: symmetry, covering ply, disorientation, percentage rule, balance, and contiguity of the layup. The manufacturability of the final design with available technologies must be guaranteed. Different regions may be subject to different loads in a large-scale structure. Laminate thickness may vary throughout the structure depending on the distributed loads in an optimized design. Additionally, for large-scale composite structures, such as an aircraft wing or fuselage, stiffeners are added to enhance structural performance in carrying compressive and tensile loads. The stiffeners divide the structure into smaller panels. To ensure manufacturability of a composite structure, it is crucial for the plies to be continuous among adjacent panels while the laminate thickness varies. Continuity of plies in adjacent panels, which is commonly referred to as blending, is a particularly difficult constraint to deal with. A feasible composite structure has to have sufficient stiffness and strength while complying with the aforementioned design and manufacturability guidelines. The feasibility of a composite structure is evaluated through the constraints added to the optimization problem. The variables in the design of a composite structure include, but are not limited to, ply stacking sequences and thickness distribution. Depending on the application, it may also be desirable to design the shape of a structure. Holes may be required in the design of parts such as ribs of aircraft wings to reduce the weight. Therefore, the shape and the topology may be additional variables in the design of a composite structure. Optimization of a composite structure subject to the design, manufacturing, and strength related guidelines is a very challenging problem. Fulfilling the manufacturability guideline in particular has been a major goal in recent studies. This confirms the interest of the experts in the field in the applicability of their proposed methods to real-life engineering problems. However, an investigation among the existing research, revealed that these methods require a significantly large number of computations and their inherent complexity makes them inadmissible for application to real structures. This motivated performing the present research with the goal of developing a design tool that can optimize complex fiber-reinforced composite structures in practice. The present research consists of a design tool for the optimization of variable stiffness composite structures (where fibers are not steered), and a method which is developed mainly for the optimization of an aircraft wing. To optimize a variable stiffness composite structure, the proposed method separates the optimization of stacking sequences from the optimization of the thickness distribution. A set of laminates with optimized stacking sequences with respect to the problem at hand is generated and stored in a reference table known as the Stacking Sequence Table (SST). The laminates in an SST must satisfy the desired laminate design guidelines. As long as the ply stacks in a design are selected from the SST, manufacturability of the final design is guaranteed. Next, a novel level-set gradient based method is introduced for the global optimization of ply drop locations. The proposed method aims at turning the discrete optimization problem associated with the integer number of plies into a continuous problem. This is done through the way the problem is parametrized; the design variables are never rounded in this approach. The level-set function gives the optimum thickness distribution over the structure for a specific SST. The developed method is verified by its application to the well-known horseshoe panel optimization problem. To investigate the performance of the method in dealing with a real problem, the proposed method is then applied to the layup optimization of a composite skin and rib structure of a wing. Local buckling and allowable strain are considered as the constraint of the problem and a standard finite element package is used to calculate buckling factors. The structural optimization of an aircraft wing is a highly complex problem. This is due to the large number of variables as well as structural and aerodynamics constraints influencing the design of skins and stiffeners. To make it computationally more efficient, a large problem can be decomposed into several smaller subproblems (subsystems) while preserving the couplings among these subproblems. A special method is subsequently introduced for the optimization of interacting skins and ribs of an aircraft wing box.

Genetic Algorithms (GAs) are a powerful technique for search and optimization problems, and are particularly useful in the optimization of composite structures. The search space for an optimal composite structure is generally discontinuous and strongly multimodal, with the possibility for many local sub-optimal solutions or even singular extrema. These facts severely limit gradient-type approaches to optimization, bringing this broad class of problems under scrutiny for application of GAs. Examples described here of the successful use of parallel GAs to design composite structures by the authors include energy-absorbing laminated beams [1], airfoils with tailored bending-twisting coupling [2], and flywheel structures [3]. Optimal design of laminated composite beams was performed using a GA with a specialized finite element model to design material stacking sequences to maximize the mechanical energy absorbed before fracture. An initial GA approach to the optimal design of a specialized, idealized composite airfoil is now being refined for a practical application. The optimum stacking sequence to produce a desired twisting response while minimizing weight, maximizing in-plane stiffness and maintaining acceptable stress levels is determined. The GA has also been used to maximize the Specific Energy Density (SED) of composite flywheels. SED is defined as the amount of rotational energy stored per unit mass. Optimization of SED was achieved by allowing the GA to search for various flywheel shapes and allowing the GA to pick material sequences along the radius of the flywheel.KeywordsAutomated DesignGenetic AlgorithmsComposite MaterialLaminated StructureOptimization

This paper describes the development of an instrumented tap test and imaging system for the inspection of defects and damages in composite aircraft structures, especially honeycomb sandwich constructions. The resulting tap test images based on the impact duration and displayed in a C- scan format can readily reveal the shape and extent of honeycomb core damages. The images also reveal considerable detail of internal substructures such as core splices, ply build-up, and changes of foam core density. Based on a grounded-spring mechanical model, the tap scan images can be converted into images that show the quantitative changes of the local stiffness. A manual tap scan imaging system was assembled for laboratory experiments. Using actual composites parts for airplanes, impact damages and substructures were imaged. In addition, effects of the tapper mass, impactor radius, tap velocity and operator dependence were studied. A compact fieldable system is being built and an automated tapping and imaging system is also under development. The technique was compared to other similar devices currently available.

Concurrent multi-material and multi-scale design optimization of fiber-reinforced composite material and structures for minimum structural compliance

Impact analysis of fiber reinforced polymer honeycomb composite sandwich beams

PurposeThis paper provides a general review of automated processing methods currently being used to fabricate aircraft composite structure.Design/methodology/approachPresents a description of the Automated Tape Layer (ATL) process and the Fiber Placement (FP) process. These processes are the most “automated” of all processes being used to fabricate composite aircraft structure. Fiber Placement machines and Automated Tape Layers are composites machine tools and they are the closest comparison the composites industry has to metals machining equipment.FindingsThere is a need for more variety of composites automation and more affordable machines in the aerospace composites industry. The limited variety of automation and the cost of equipment tend to limit the spread of automation throughout the aerospace composites industry. ATL and FP are composites laminating technologies that could be adapted to a wide range of machine sizes, configurations, and price ranges.Originality/valueMore widespread use of automated processes in composites would tend to lower the cost of composite aircraft structure on a global basis.

Cost Analysis of Spinal Versus General Anesthesia for Lumbar Diskectomy and Laminectomy Spine Surgery

In order to explore the optimization of composite wing structure, this paper reviews the past research works of optimization of composite wing structures worldwide and in China. It concludes that most optimization methods and calculations are useful but only applied with specified theoretic model. Then relevant composite optimization theories and strength constraints with composite lamination failure criteria, stability, aeroelasticity are sorts out. Optimizing algorithm and applied solution rounding strategy are also listed. Concentrating on weight-reduction analysis on the thickness, sequence, and angle of the composite stacking under multiple constraints, a practical model of a composite wing structure example with a large aspect ratio is completed, and the calculation is based on the finite element analysis tool Nastran and the genetic algorithm Matlab program. Two optimal solutions are given with detailed differences and the comparison of the composite structure parameters before and after calculation is also recorded. The application example indicates the feasibility of the optimization method and quadratic rounding process, the detail structure performance improvements with skin lamination change under the calculation are illustrated at the same time.

As composite materials have been used for marine structures such as boats, ships, offshore structures, etc., the effect of FSI should be understood.In particular, FSI effect is expected to be significant for polymer composites because the water density is very comparable to the composites' densities.For example, the density of a carbon composite is approximately 50% greater than the water density.Sandwich composites consisting of very low densities of core materials have lower densities than that of water.As a result, the hydrodynamic mass associated with FSI would be very critical to composite structures under water.The growing use of composites in ship masts, superstructures, deck grates, piping, ducting, rudders, propellers, stacks, and various submarine structures requires extensive modeling and testing to help designers, builders and operators better understand composite response [Mouritz, et al., 2001].These materials are subjected to a wide spectrum of loads during manufacturing and service life.Dynamic loadings, in particular, impact type event, represent a serious design concern for use of composite.Composite structures are more susceptible to impact damage than similar metallic structures which are more ductile in nature and can absorb typically large amounts of energy without failure.Furthermore, the damage in composites from impact can go undetected even when the mechanical properties may be drastically reduced from an impact.For these reasons, numerous experimental and analytical studies have been conducted to study the dynamic response of composites subjected to impact loading [Abrate, 1994;Aslan, et al., 2003;Kwon & Wojcik, 1998].According to the review of past works, most of the research effort has been focused on low velocity impact damage, specifically, the damage predictions, and the evaluation and prediction of residual properties of damaged laminates.All of the research completed thus far has focused on damage in composites under impact loading in dry environments to support development of composites in aircraft structures.As far as dynamic response of structures under water is concerned, a great deal amount of analytical and experimental studies have been conducted on the effect of fluid force on the natural frequencies, damping ratios and mode shapes of vibrating structures in contact with fluid.This is commonly known as the Fluid Structure Interaction (FSI) problem.FSI investigations have supported many problems in submarine signaling, offshore oil structure stability, and ship structure vibrations.Through these studies, many numerical and analytical methods have been developed in order to predict the added mass and the www.intechopen.com Advances in Composites Materials -Ecodesign and Analysis 338resulting changes in natural frequency of a structure in contact with fluid.It has been determined and widely proven that the effect of fluid surrounding a structure decreases the natural frequency of a structure due to the increase in total kinetic energy of the vibrating structure and fluid from the addition of kinetic energy of the fluid.This effect can be interpreted as an added mass to the vibrating structure in the analysis of the dynamic response.Essentially as the structure vibrates, its mass is increased by the mass of the vibrating fluid with which it is in contact, consequently decreasing its natural frequency.Studies of fluid structure interaction and the added mass effect, also known as virtual mass effect, hydrodynamic mass, and hydroelastic vibration of structures, started with Lamb [Lamb, 1921] who calculated the first bending mode of a submerged circular plate.In response to a problem of submarine signaling, Lamb investigated the vibrations of a thin elastic circular plate in contact with water.In his investigation he discovered that the natural frequencies for structures in contact with fluid are lower than the frequencies in air, based on the assumption that the modes shapes are virtually the same in water as in a vacuum.The resonant frequency was determined using Rayleigh's method.Lamb's theortical results were verified experimentally [Powell & Roberts, 1923].Much later, more research was conducted for measuing and calculating natural frequencies of free vibartions of beams and plates under water [Lindholm, et al., 1965; Fu & price, 1979;Kwak, 1996].Another branch of FSI studies is underwater explosion.Both experimental and numerical studies have been conducted for metallic structures [

Cost/weight optimization of composite prepreg structures for best draping strategy

Composite aircraft structures commonly use bonded joints to adhere the wing and fuselage skins to the wing spars, stringers, ribs, and frames. Unfortunately the bond out-ofplane normal (peel) strength is many orders of magnitude less than the fiber strength of the adjacent laminates. Moreover, the adhesive failure is brittle in nature so that it is difficult to monitor an impending overload or fatigue failure. In order to avoid failures due to interlaminar damage without using high margins of safety, a Structural Health Monitoring (SHM) system can be used. While extensive literature on SHM systems for fiber-driven damage exists, there is no mention of an SHM system that specifically addresses these types of matrix-driven damage. The proposed SHM system spans the acquisition, location, reconstruction, and damage type decision of Acoustic Emission (AE) damage signals. An algorithm is presented to calculate the location of AE damage signals in a laminated composite structure using an over-determined approach to minimize location errors. Also, a novel procedure is presented to simultaneously reconstruct the loading source time series for a recorded damage signal and identify the source type. Parametric studies are presented for both algorithms to gauge their effectiveness under varying levels of noise, number of sensors, material properties and different types of uncertainty. Two experimental studies were undertaken to exercise the SHM algorithm presented in this study. The first experiment models a predominantly single type quasi-static damage case, and the second experiment studies a mixed damage impact case.

A technical and design evaluation was carried out to meet DOE hydrogen fuel targets for 2010. These targets consisted of a system gravimetric capacity of 2.0 kWh/kg, a system volumetric capacity of 1.5 kWh/L and a system cost of $4/kWh. In compressed hydrogen storage systems, the vast majority of the weight and volume is associated with the hydrogen storage tank. In order to meet gravimetric targets for compressed hydrogen tanks, 10,000 psi carbon resin composites were used to provide the high strength required as well as low weight. For the 10,000 psi tanks, carbon fiber is the largest portion of their cost. Quantum Technologies is a tier one hydrogen system supplier for automotive companies around the world. Over the course of the program Quantum focused on development of technology to allow the compressed hydrogen storage tank to meet DOE goals. At the start of the program in 2004 Quantum was supplying systems with a specific energy of 1.1-1.6 kWh/kg, a volumetric capacity of 1.3 kWh/L and a cost of $73/kWh. Based on the inequities between DOE targets and Quantum’s then current capabilities, focus was placed first on cost reduction and second on weight reduction. Both of these were to be accomplished without reduction of the fuel system’s performance or reliability. Three distinct areas were investigated; optimization of composite structures, development of “smart tanks” that could monitor health of tank thus allowing for lower design safety factor, and the development of “Cool Fuel” technology to allow higher density gas to be stored, thus allowing smaller/lower pressure tanks that would hold the required fuel supply. The second phase of the project deals with three additional distinct tasks focusing on composite structure optimization, liner optimization, and metal.

The determination of lamination parameters feasible region is extremely important for design and optimization of composite structures. While the feasible region for unconstrained laminates is well-established, applying mandatory design rules alters both its size and shape. This is especially relevant in aeronautical composite structures, where design rules are quite often mandatory. However, existing studies do not specify the exact boundaries of the feasible region, which is significant since the optimum for critical loading conditions, such as buckling, often lies on or near the boundary. This work proposes an innovative approach for finding the exact feasible region with several design rules. The initial step is to generate laminate databases containing all feasible stacking sequences with the prescribed design rules, up to 36 layers. These data provide a rigorous insight into the distribution patterns. Based on this information, a simple and computationally efficient procedure, based on the exact determination of a few key boundary points, allows for the exact determination of the feasible region for any number of layers.

The curing distortion is a critical determinant of the quality of integrally manufactured composite structures, playing a pivotal role in the design and fabrication of composite. This paper presents two simplified methods in predicting the curing distortion for large-scale composite aircraft structures manufactured through the autoclave process. Firstly, the refined finite element models of the two simplified methods were developed. Then, it was utilized to predict the curing distortion of Ω-shaped composite laminates. The comparative study between the experimental data and numerical results shows that the proposed second simplified method balanced the prediction accuracy and efficiency, which is urgently needed in practice. Finally, using the second simplified method, predictions were conducted for the curing distortion of practical large-scale composite skin structures. The results were in good agreement with the corresponding experiments. This study provides a new solution for the rapid iterative design of large-scale composite structures, as well as the efficient design of frame molds for their manufacturing.