Design Of Aircraft Structures Research Articles

Application of Multidisciplinary Design Optimization on Advanced Configuration Aircraft

An optimization strategy is constructed to solve the aerodynamic and structural optimization problems in the conceptual design of double-swept flying wing aircraft. Aircraft preliminary aerodynamic and structural design optimization is typically based on the application of a deterministic approach of optimizing aerodynamic performance and structural weight. In aerodynamic optimization, the objective is to minimize induced drag coefficient, and the structural optimization aims to find the minimization of the structural weight. In order to deal with the multiple objective optimization problems, an optimization strategy based on collaborative optimization is adopted. Based on the optimization strategy, the optimization process is divided into system level optimization and subsystem level optimization. The system level optimization aims to obtain the optimized design which meets the constraints of all disciplines. In subsystem optimization, the optimization process for different disciplines can be executed simultaneously to search for the consistent schemes. A double-swept configuration of flying wing aircraft is optimized through the suggested optimization strategy, and the optimization results demonstrate the effectiveness of the method.

Experimental and numerical investigations of compression stability of stiffened composite panel with ply interleaving

A novel composite layup technique (ply interleaving) is proposed to satisfy the extraordinary needs in aircraft structure design. Buckling experiments and corresponding simulations with different configurations are carried out to study the mechanical properties of the stiffened panel with this ply layout and get a suitable numerical prediction method. A ply-by-ply-based modeling technique is applied to characterize the real discontinuity in each ply. Additionally, the relationship between the macroscopic buckling and damage evolution is investigated by the comparative analyses of experiments and simulations. The investigations show the following results: the panel with ply interleaving has a considerable anti-buckling capability besides its flexible layout; both adhesive layer damage and lamina damage should be included in the simulation for obtaining reliable solution; the panel would buckle while adhesive layer damage occurs between stiffeners and skin, which initiates firstly in the ply interleaving zone and then expands to the surrounding area; internal damage of each ply also originates around discontinuities then develops rapidly, leading to the panel failure. The prediction results of buckling and failure show good agreements with the experimental results, so the suggested simulation technique is verified to be an effective method to study the mechanical behavior of the composite plate with preset discontinuities.

Multiscale virtual testing: the roadmap to efficient design of composites for damage resistance and tolerance

This paper proposes a novel approach to the determination of the mechanical behaviour of composite materials up to failure using numerical and experimental techniques in parallel. A bottom-up multiscale virtual testing strategy is presented to take into account the physical mechanisms of deformation at different length scales on the behaviour of the composite. Starting from the microscale, the contributions of the basic constituents, microstructure and loading conditions to the mechanical response are considered in a rigorous way. This hierarchical multiscale approach describes systematically the material behaviour at different length scales from ply to laminate to component level, allowing the determination of ply properties, laminate characteristics and structural response. Additionally, this approach easily allows consideration of changes in properties of the constituents (fibre, matrices), fibre architecture or laminate lay-up and provides fast predictions of their influences on the macroscopic behaviour of composite structures. Hence, this approach constitutes a promising tool to provide significant efficiency gains in the design, testing and certification of composite aircraft structures.

Safety envelope for load tolerance of structural element design based on multi-stage testing

Structural elements, such as stiffened panels and lap joints, are basic components of aircraft structures. For aircraft structural design, designers select predesigned elements satisfying the design load requirement based on their load-carrying capabilities. Therefore, estimation of safety envelope of structural elements for load tolerances would be a good investment for design purpose. In this article, a method of estimating safety envelope is presented using probabilistic classification, which can estimate a specific level of failure probability under both aleatory and epistemic uncertainties. An important contribution of this article is that the calculation uncertainty is reflected in building a safety envelope using Gaussian process, and the effect of element test data on reducing the calculation uncertainty is incorporated by updating the Gaussian process model with the element test data. It is shown that even one element test can significantly reduce the calculation uncertainty due to lacking knowledge of actual physics, so that conservativeness in a safety envelope is significantly reduced. The proposed approach was demonstrated with a cantilever beam example, which represents a structural element. The example shows that calculation uncertainty provides about 93% conservativeness against the uncertainty due to a few element tests. It is shown that even a single element test can increase the load tolerance modeled with the safety envelope by 20%.

Probabilistic Method to predict Remaining Usage Life of Aircraft Structures

Current trends in the design of new aircraft components and high level innovations in the older type of aircraft are clearly pointing to automation and integration of all aircraft systems to give increased cost efficiency in the aircraft operation. The same trend can be observed in the design, operation and maintenance of aircraft structures where advanced Structure Health Monitoring (SHM) systems are about to enter into service. These advanced systems are designed to provide both diagnostic and prognostic information enabling application of Condition Based Maintenance (CBM) or even Prognostics Maintenance (PM) concepts into the maintenance of the aircraft structures. To make this CBM/PM concept a reality, an accurate, reliable and robust method to predict the Remaining Usage Life (RUL) of the structure is the foremost step.This paper presents a probabilistic method for RUL prediction. A hybrid approach is used, comprising of two different algorithms. The first algorithm adopts fracture- mechanics based fatigue crack growth model. This approach uses physics of failure to predict the crack growth curve and underlying degradation process. It calculates the accurate value of Stress Intensity Factor (SIF) to calculate the crack growth curve. The second algorithm is a mathematical model which quantify various sources of uncertainty such as future loading, crack length, model parameters etc. The process described in this paper results in enhanced remaining usage life estimation by compensating for the aforementioned modeling uncertainties. The model results were verified and validated on a typical aero structure with experimental, FEA simulation and fractography data.

Optimization design of unitized panels with stiffeners in different formats using the evolutionary strategy with covariance matrix adaptation

Curvilinear stiffener concept has been introduced to aircraft panel structures most recently for possible further weight reduction during optimization design. However, due to enlarged design space and high design complexity, more computational time is needed to optimize curvilinearly stiffened panels. Considering the requirement for both lighter structure and less design time, optimization designs of a unitized panel with stiffeners in three different formats, i.e. curvilinear, oblique, and evenly distributed straight, are conducted under various loading conditions to figure out which stiffener format should be selected in a real design environment. Four single loading cases with different compression-shear ratios and a set of multiple load cases are considered. Evolution strategy with covariance matrix adaption is combined with sequential quadratic programming to seek out the optimum structure with minimum mass taking into account buckling and strength constraints. Comparative analysis of the optimization results indicates that evenly distributed straight format is suitable for compression-dominant loading conditions whereas curvilinear format become superior in case that shear loading is considerable or multiple biaxial compression and shear loads are applied. Besides, curvilinear format adds more design flexibilities to the stiffeners, which enables them to play a more important role in the optimized structures. Furthermore, evolution strategy with covariance matrix adaption is found to be more efficient than particle swarm optimization for this optimization problem. This study can provide a useful guidance for future optimization design of aircraft structures.

Rapid Design Method and System Development for Aircraft Wing Structure

This work is mainly done by too many manual operations in the aircraft structure design process resulting in heavy workload, low efficiency and quality, non-standardized processes and procedures. A top-down associated design method employing the template parametric technology is proposed here in order to improve the quality of design and efficiency of aircraft wing structure at the preliminary design stage. The appropriate parametric tool is chosen and the rapid design system of knowledge-driven aircraft wing structure is developed. First, a skeleton model of aircraft wing structure is rapidly built up through the template encapsulated design knowledge. Associated design is then introduced to realize the association between the typical structural part and skeleton model. Finally, the related elements are referenced from skeleton model, and a typical structural part reflecting an automatic response for design changes of the upstream skeleton model is quickly constructed within the template. The rapid design system proposed and developed in this paper is able to formalize the design standardization of aircraft wing structure and thus the rapid generation of different aircraft wing structure programs and achieve the structural design knowledge reuse as well.

Automated sizing of a composite wing for the usage within a multidisciplinary design process

Purpose The purpose of this paper is to present a structural design and optimization module for aircraft structures that can be used stand-alone or in a high-fidelity multidisciplinary design optimization (MDO) process. The module is capable of dealing with different design concepts and novel materials properly. The functionality of the module is also demonstrated.

A Data-Driven Approach to Detect Faults in the Airbus Flight Control System

This paper presents a data-driven strategy for the detection of failures impacting the flight control system. Early and robust detection of Oscillatory Failure Case (OFC) allows the aircraft structural design to be optimized, which in turn helps improve the aircraft environmental footprint thanks to weight saving. Compared to existing model-based techniques already used on in-service Airbus aircraft, this paper studies a novel signal processing approach based on distance and correlation. It is shown that a mixed similarity index between Euclidean distance and logarithmic invariant divergence gives promising detection results. This paper details the proposed approach by insisting on practical constraints due to implementation in embedded real-time systems such as the flight control computer. Preliminary results obtained from a Verification & Validation (V&V) on-going campaign are presented.

On the monitoring and implications of growing damages caused by manufacturing defects in composite structures

Damage tolerance is a classical safety concept for the design of aircraft structures. Basically, this approach considers possible damages in the structure, predicts the damage growth under applied loading conditions and predicts the following decrease of the structural strength. As a fundamental result the damage tolerance approach yields the maximum inspection interval, which is the time a damage grows from a detectable to a critical level. The above formulation of the damage tolerance safety concept targets on metallic structures where the damage is typically a simple fatigue crack. Fiber-reinforced polymers show a much more complex damage behavior, such as delaminationsin laminated composites. Moreover, progressive damage in composites is often initiated by manufacturing defects. The complex manufacturing processes for composite structures almost certainly yield parts with defects, e.g. pores in the matrix or undulations of fibers. From such defects growing damages may start after a certain time of operation.The demand to simplify or even avoid the inspection of composite structures has therefore led to a comeback of the traditional safe-life safety concept. The aim of the so-called safe-life flaw tolerance concept is a structure that is capable of carrying the static loads during operation, despite significant damages and after a representative fatigue load spectrum. A structure with this property does not need to be inspected, respectively monitored at all during its service life. However, its load carrying capability is thereby not fully utilized.This article presents the possible refinement of the state-of-the-art safe-life flaw tolerance concept for composite structures towards a damage tolerance approach considering also the influence of manufacturing defects on damage initiation and growth. Based on fundamental physical relations and experimental observations the challenges when developing damage growth and residual strength curves are discussed.

Dynamic tests of composite panels of an aircraft wing

The paper describes the analysis of aerospace composite structures under dynamic loading. Today, it is common to use design procedures based on assumption of static loading only, and dynamic loading is rarely assumed and applied in design and certification of aerospace structures. The paper describes the application of dynamic loading for the design of aircraft structures, and the validation of the procedure on a selected structure. The goal is to verify the possibility of reducing the weight through improved design/modelling processes using dynamic loading instead of static loading. The research activity focuses on the modelling and testing of a composite panel representing a local segment of an aircraft wing section, investigating in particular the buckling behavior under dynamic loading. Finite Elements simulation tools are discussed, as well as the advantages of using a digital optical measurement system for the evaluation of the tests. The comparison of the finite element simulations with the results of the tests is presented.

Topology Optimization in Aircraft and Aerospace Structures Design

Topology optimization has become an effective tool for least-weight and performance design, especially in aeronautics and aerospace engineering. The purpose of this paper is to survey recent advances of topology optimization techniques applied in aircraft and aerospace structures design. This paper firstly reviews several existing applications: (1) standard material layout design for airframe structures, (2) layout design of stiffener ribs for aircraft panels, (3) multi-component layout design for aerospace structural systems, (4) multi-fasteners design for assembled aircraft structures. Secondly, potential applications of topology optimization in dynamic responses design, shape preserving design, smart structures design, structural features design and additive manufacturing are introduced to provide a forward-looking perspective.

The Application of Multiple Vibration Neutralizers for Vibration Control in Aircraft

A current challenge for researchers is the design and implementation of an effective vibration control method that reduces vibration transmission from vehicle structures such as aircraft. This challenge has arisen due to the modern trend of utilizing lightweight thin panels in aircraft structural design, which have the potential to contribute towards significant vibration in the structures. In order to reduce structural vibration, one of the common approaches is considering vibration neutralizer system attached to the structure. In this study, a vibration neutralizer is developed in a small scale size. The effectiveness of attached vibration neutralizers on a thin plate are investigated through experimental study. Prior to the experiment, a finite element analysis of Solidworks® and analytical modelling of Matlab® are produced in order to determine the structural dynamic response of the thin plate such as the natural frequency and mode shapes. The preliminary results of finite element analysis demonstrate that the first four natural frequency of clamped plate are 48Hz, 121Hz, 194Hz and 242Hz, and these results are in agreement with the plate’s analytical equations. However, there are slight discrepancies in the experiment result due to noise and error occurred during the set up. In the later stage, the experimental works of thin plate are performed with attached vibration neutralizer. Result shows that the attachment of vibration neutralizer produces better outcome, which is about 41% vibration reduction. It is expected that by adding more vibration neutralizer to the structure, the vibration attenuation of thin plate can be significant.

Design and Analysis of an Aircraft Composite Hinge Bracket Using Finite Element Approach

The use of composite materials in aircraft structures have been increasing for the past decade. The anisotropic and heterogeneous nature of composites remains a major challenge to the design and analysis of composite aircraft structures. Composite structures require a different design approach compared to the design of metallic structures. This paper aims to provide a step by step definitive guide to design and analyze composite structures using finite element approach. A simplified design model for the composite structural design was used to analyze an aircraft composite hinge bracket. The composite hinge bracket which is made of IM7/8552 laminated composite plates was successfully designed with a margin of safety of 0.216 and a weight savings of 43.77 percent was estimated.

Reexamined Structural Design Procedures for Very Flexible Aircraft

This paper reviews the applicability of some conventional structural design practices to the analysis and design of very flexible aircraft. The effect of large structural deformations and the coupling between aeroelasticity and flight dynamics is investigated in different aspects of the aircraft structural design process, including aeroelastic stability, loads, and flight dynamics and control. This is illustrated with a numerical example of the static and dynamic responses of a representative high-altitude long-endurance vehicle. Suggestions are presented for the development of appropriate frameworks to design and analyze very flexible aircraft.

Rubber Ball Impacts on Composite Plates Investigation by 3D Correlation

Aircraft structures and wings should endure tyre debris impacts. Therefore, numerical simulations which are able to predict such phenomena are highly appreciated for the aircraft structures design. However these simulations have to be developed, tuned and validated by means of experimental data. Nowadays, full-scale tests are widely used. However, they include too complex mechanisms, which are difficult to consider in a process of model development. Besides, these experiments are extremely expensive. In this work, we are dealing with a simplified experimental set-up with elementary boundary conditions. Namely, we are experimentally investigating rubber balls’ impact on composites panels. The main novelty is the use of 3D-stereo correlation technique.

Fatigue, Fatigue Crack Propagation and Mechanical Fracture Behaviour of Laser Beam-Welded AZ31 Magnesium Sheets

Weight reduction is the main driving force in automotive and aircraft structural design. As a result, magnesium alloys, with their high potential for lightweight construction, have attracted a considerable amount of industrial attention. The determining criterion for the structural applications of magnesium alloys is the availability of efficient joining technologies for the construction of lightweight structures and the availability of reliable data for the assessment of their damage tolerance behaviour. Laser beam welding (LBW), as a high-speed and easily controllable process, allows the welding of complex geometric forms that are optimised in terms of mechanical stiffness, strength, production velocity and visual quality. The work accomplished in this study addresses the challenges of the LBW process for typical joint configurations using the magnesium alloy AZ31HP: butt joints, T joints and overlap joints. LBW processes were developed for use with a 3.3-kW Nd:YAG laser to optimise the mechanical performance of such joints with respect to tensile strength, fatigue, fatigue crack propagation and mechanical fracture behaviour. The relationships between the LBW process and the microstructural and mechanical properties of welds were established. Compared to state-of-the-art aerospace alloys, AZ31HP demonstrates that magnesium alloys have potential for use in structural applications, with AZ31HP being comparable to AA2024T351 and AA6061T6. Welded AZ31HP exhibits better crack resistance than the base material, so fully welded integral structures made from magnesium alloys can be used in lightweight construction.

Laser Shock Peening to Repair, Design and Manufacture Current and Future Aircraft Structures by Residual Stress Engineering

This paper will provide an overview on potential applications for the aerospace industry for repairing aircraft as well as to ensure salvage for identified hot spots in terms of fatigue and crack growth performance. Residual stress engineering is a field of engineering aiming to improve the economic and ecological impact of future aircraft structures by controlling the residual stresses induced by Laser Shock Peening (LSP). Managing the residual stresses for designing structures represents an innovative approach for next generation aircraft. Predicting crack turning induced via a LSP treatment and the optimization of the LSP treatment itself for reaching the crack growth design stress for the targeted weight benefit will be discussed. Advanced forming processes in aircraft manufacturing represent another potential area of interest and the benefits and challenges of applying laser peen forming in this context will be presented. The aeronautical industry requirements for future developments of the laser shock process will also be included for applications ranging from the repair environment to design and manufacturing of aircraft structures.

Modelling Fatigue Crack Growth Rates in Aluminium Plates Reinforced by Bonded Fibre-Metal Laminates for Fatigue Life Extension

Aircraft structural design and manufacture is moving towards lighter structures that have extended lives and improved damage tolerance. Hybrid structures are a possible solution to improve damage tolerance. They are a combination of metallic structure locally reinforced with adhesively bonded damage tolerant straps. In the present study a 3D finite element model has been developed with a bond line delamination growing under a fatigue law. A series of fatigue delamination tests on bonded aluminium were performed to provide input data. An iterative model for crack and debonding growth was developed to describe how debonding influence crack stress intensity and crack profile, which in turn influence debonding. The model predicts decrease in stress intensity on the bonded face and an overall retardation of fatigue crack growth rates. The stress intensity factor was predicted to vary through the thickness of the substrate due to the phenomenon of secondary bending and also the bridging effect caused by the presence of the reinforcing strap.

Topology optimisation via the moving iso-surface threshold method: implementation and application

AbstractTopology optimisation is a useful tool for the design of aircraft structures. This work details how the new moving iso-surface threshold (MIST) topology optimisation method works and how it can be applied to aircraft structural design. This method has been coupled with commercial finite element analysis software in a simple manner without requiring the modification of the commercial software source code. In this way the user is able to take advantage of the finite element analysis tools such as automatic mesh generation and efficient solving. The extension of the method to 3D designs is also presented. The topology results of a 2D leading-edge rib for stiffness, stress and morphing objectives and a 3D wing skin stringer for stress design demonstrate the functionality of this method.