Optimal design of laminated composites with focus on aircraft structures

Farshad Farzan Nasab

    Research output: ThesisPhD Thesis - Research UT, graduation UT

    72 Downloads (Pure)

    Abstract

    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.
    Original languageEnglish
    QualificationDoctor of Philosophy
    Awarding Institution
    • University of Twente
    Supervisors/Advisors
    • Geijselaers, Bert, Supervisor
    • Akkerman, R., Supervisor
    • Baran, Ismet , Co-Supervisor
    • de Boer, Andries , Advisor
    Award date29 Mar 2019
    Place of PublicationEnschede
    Publisher
    Print ISBNs978-90-365-4736-9
    DOIs
    Publication statusPublished - 29 Mar 2019

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