Design Of Aircraft Structures Research Articles

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.

A non-conservative [formula omitted] solution for early and robust fault diagnosis in aircraft control surface servo-loops

The presented work is undertaken within the FP7-ADDSAFE (Advanced Fault Diagnosis for Sustainable Flight Guidance and Control) project, a European collaborative project that aims to propose new fault diagnosis techniques for AIRBUS aircraft that could significantly advance the aircraft performance, e.g. by optimizing the aircraft structural design (weight saving) or decreasing its environmental footprint (e.g. less fuel consumption and noise). The paper discusses the design of a model-based fault detection scheme for robust and early detection of faults in aircraft control surfaces servo-loop. The proposed strategy consists of two fault detectors: The first fault detector is based on a H−/H∞ residual generator that maximizes sensitivity to any kind of control surface servo-loop faults whilst simultaneously minimizes the influence of unknown inputs. The second fault detector consists of a pure H∞ residual generator that is sensitive to a restricted set of faults and robust to unknown inputs. By such a structured strategy, it is shown that it is possible to discriminate between different fault types occurring in the control surfaces servo-loop. Monte-Carlo campaigns from a highly representative simulator provided by AIRBUS as well as experimental results obtained on AIRBUS test facilities demonstrate the fault detection performance, robustness and viability of the proposed technique.

비틀림 비선형성을 갖는 2차원 익형의 모델링 및 Bifurcation 해석

Recent developments for high altitude, long endurance conventional UAVs(HALE UAVs) have revealed new issues regarding aircraft structure design and analysis. First of all, due to intensive mission requirements, the structures of HALE UAVs have lightweight and very flexible main wing with high aspect ratio, and slender fuselage. For this kind of structures, aeroelastic characteristics are different from conventional aircrafts. Hence, currently developed analysis methods are not suitable to fully understand strucutral dynamics of the very flexible aircraft, and to guarantee structural reliability. Therefore, various structural studies considering nonlinear behaviors which are generally ignored for the conventional aircraft strucutral analyis have been attracting researchers interests. Nonlinear flutter of the very flexible wing is one of the subject to be studied in combination with strong coupling between aeroelastic characteristics and flight dynamics. Herein, as preliminary study, modeling and nonlinear system analysis of the 2D airfoild with torsional nonlinearity have been discussed.

The Static Strength Reliability Analysis of Composite Aircraft Structures

The static strength design of aircraft structure is the most basic design principles. The basic applicable conditions of composite structures static strength reliability analysis using the current methods were mentioned in this paper, and compliance analysis of these basic conditions was carried out. In addition, further studies were also conducted to research the influence of the number of full-scale structural specimens and environmental factors on the composite structures static strength reliability.

Finite element modeling of mechanically fastened composite-aluminum joints in aircraft structures

A three-dimensional, solid finite element model of a composite-aluminum single-lap bolted joint with a countersunk titanium fastener is developed. The model includes progressive damage behavior of the composite and a plasticity model for the metals. The response to static loading is compared to experimental results from the literature. It is shown that the model predicts the initiation and the development of the damage well, up to failure load. The model is used to evaluate the local force–displacement responses of a number of single-lap joints installed in a hybrid composite-aluminum wing-like structure. A structural model is made where the fasteners are represented by two-node connector elements which are assigned the force–displacement characteristics determined by local models. The behavior of the wing box is simulated for bending and twisting loads applied together with an increased temperature and the distribution of fastener forces and the progressive fastener failure is studied. It is shown that the fastener forces caused by the temperature difference are of significant magnitude and should be taken into account in the design of hybrid aircraft structures. It is concluded that, the account of the non-linear response of the joints results in a less conservative load distribution at ultimate failure load.

Modelling the impact behaviour of sandwich structures with folded composite cores

The European aircraft industry has strong interest in novel structural concepts for future aircraft fuselage with lower fabrication costs and high performance. A critical safety issue for the design of aircraft structure is vulnerability and damage tolerance due to foreign object impact. This paper focuses on the improvement of modelling techniques for foldcore sandwich structures with composite skins under low velocity impact. A new modelling approach for the foldcore cell wall material – aramid paper – is presented, which considers the aramid papers inhomogeneous nature in the thickness direction. In combination with in-ply continuum damage mechanics and a cohesive delamination interface for the modelling of the composite surface skin the proposed modelling approach is used to reproduce drop tower impact tests performed at various impact velocities. Results and interpretation of the numerical and experimental work are presented. The strengths and weaknesses of the presented modelling approach are discussed and evaluated.

Dynamically controlled variable-fidelity modelling for aircraft structural design optimisation

Structural mass optimisation of an aircraft design is important in increasing the likelihood that a high quality airframe is designed of minimal weight whilst providing necessary resistance to load. Analysis of such structures is often performed at a single level of model fidelity, the selection of which can lead to either excessive computational time or reduced accuracy of results. Alternatively, variable-fidelity modelling may be employed to reduce such computational expense whilst maintaining accuracy, traditionally performed using predetermined levels of fidelity for specific periods of the optimisation process. This paper investigates dynamically controlled variable-fidelity modelling during aircraft conceptual design optimisation wherein fidelity is controlled as a dynamic parameter of the optimisation process. Consequently, model fidelity is adapted during optimisation to promote early discovery of promising design characteristics prior to detailed analysis of the best designs available. Models are constructed through the grouping of similar structural members within elements, thus reducing the number of degrees of freedom and subsequent computational effort required for analysis of each design. A case study is performed to verify the results of analysis and obtain benchmark results for optimisation with static model fidelity prior to the investigation of various set-ups of dynamically controlled variable-fidelity modelling. The results of this study indicate improved design quality using dynamically controlled variable-fidelity modelling compared to using static model fidelity whilst reducing the necessary computation time.

Modelización no lineal de secciones aeronaúticas con EXCEL-SOLVER©

The program MICROSOFT-EXCEL © has a great popularity in the professional world of engineering-consulting. The authors are aware of that from own experience. This is the reason that has led us to implement a nonlienar model to predict the ultimate strenght of fuselage and wings structures. The didactic methology has been developed for the subject Aircraft Structural Design in the 4th course of Aeronautical Engineering studies at the Polythecnic University of Valencia. The students show a high level of satisfaction since they learn nonlinear computational methods, valid for real structures; they are capable of the implementation with help of the very visual and handle tool MICROSOFT-EXCEL © (without programming code) and finally they solve nonlinear optimization problems with the numerical analysis complement SOLVER © .

Validation of damage modelling in composite fuselage structures under high velocity impact

The paper discusses finite element (FE) simulation methods in use for predicting impact damage in advanced composite structures. Before these simulation tools are accepted by the industry for design and certification of composite aircraft structures, composites damage models have to be developed, implemented in commercial FE codes and validation studies at specimen and substructure level have to be performed. For vulnerability analysis, the DLR uses meso-scale composites ply damage models with an energy-based delamination failure criteria in explicit FE codes to study damage progression in composite structures under impact. The methods developed are summarised in the paper and applied to predict impact damage in stiffened aircraft panels and novel foldcore sandwich structures from hard debris. Issues concerned with use of such computational methods for certification by analysis are discussed.

Testing of Composites: A Review

A B S T R A C T Manufacturers from various industries always strive to produce structures that are lighter, stronger, offer greater corrosion resistance and exhibit other required properties; than those of the traditionally available materials. Composite materials offer a solution to all the needs of these manufacturer(s). Composites are complex structures made of a variety of different polymers, metals, and ceramic materials. Composite materials are being used increasingly in a wide variety of products and applications. Ever-increasing performance demands in the industries viz. the aerospace, automotive, and environmentally sustainable-energy systems requires a wider range of physical and mechanical testing. Testing of composites for physical and mechanical properties is required to characterize and understand the performance of these materials. Testing of composite materials includes high force tension and compression, impact, flexure, shear, rheology, and fatigue tests under a range of environmental and/or experimental conditions. Mechanical testing is an important step in the building block approach to design of composite aircraft structures. Mechanical testing instruments are configured using a range of fixtures that have been developed to provide various ways to test composite materials depending on the type of material and its intended end use. In this paper various methods of testing composite materials has been discussed in details. The paper also provides a source of information about the current testing methods through which the advanced composites can be evaluated to yield satisfactory results.

AIRCRAFT-LIGHTNING ELECTRODYNAMICS USING THE TRANSMISSION LINE MODEL PART I: REVIEW OF THE TRANSMISSION LINE MODEL

This work presents a self-consistent and self-contained model to study and analyze aircraft-lightning electrodynamics. In this paper, we review the well developed and reported transmission line model of the cloud-to-ground (CG) lightning return stroke. Subsequently, the incorporation of a circuit model of the aircraft into the return stroke model is considered. The direct hit characteristics of aircraft body lightning currents for both CG and GC (ground- to-cloud) are important when designing protection, shielding and flltering systems for airborne electronic and electrical systems within the aircraft system. Moreover, the model will allow design of aircraft structure and geometry to minimize energy dissipation into the aircraft structure and systems. Basic electromagnetic theory is used to show the validity of considering the return stroke as a transverse magnetic wave along a transmission line. A distributed transmission line model for the aircraft and the return stroke channel of the lightning is used to simulate the return strokes of CG and GC ∞ashes. The efiects of the aircraft geometry with sharp edges are included in the computation of aircraft capacitance values, both distributed as well as lumped values.

Fracture mechanics in aircraft failure analysis: Uses and limitations

Fracture mechanics is generally accepted as essential to the design of metallic aircraft structures, notably for predicting and validating the fatigue crack growth and residual strength properties. Applications of Linear Elastic Fracture Mechanics (LEFM) began in the early 1970s, and would be expected by now to have reached a mature status. Hence one might also expect that LEFM methods can often be used for analysing service fatigue failures. However, the DSTO’s and NLR’s experience is that only non-LEFM analyses are directly useful for determining service failure fatigue crack growth (FCG) behaviour. LEFM analyses can assist in analysing FCG in full-scale, component and specimen tests, and these analyses can be used – within limits – to predict some aspects of the FCG behaviour of in-service structures and components. The DSTO has developed a hybrid method, the Effective Block Approach (EBA) using both LEFM and non-LEFM analyses, from full-scale testing. The EBA enables a wide range of FCG predictions for high performance aircraft.

Simulation of Pylon Emergency Break-Away of Large Commercial Aircraft

Most large commercial aircraft engines are hanged below wings. When the aircraft makes an emergent landing process, at the same time, the landing gear cannot work or the centrifugal force becomes unbalanced as the fan blades fly away. In order to ensure safe landing, overall tank crack and fuel leak should be avoid at the crack area. Access to a large number of emergent landing accidents, emergency break-away is essential. In this paper, the break-away position, which locates between the pylon and the wing, is mainly considered. We choose true size in the models of wing, pylon and engine. High strength steel is employed for the bolts which connect pylon and wing. Earth and lake are employed concrete and pure water respectively. SPH method is applied in the case that the aircraft lands on the lake. Whats more, different landing cases have been analyzed. By constantly adjusting the size of pins, a set of conclusions of the emergent landing problem are obtained in the simulation process. The locus of centroid of engine and pylon is obtained, and then the condition which may achieve safe flight and avoid the secondary damage to wings can be chosen, from which we can provide reasonable designing strength of the wing box and accordingly provide reference for the design of aircraft structure.

An agent-based collaborative design framework for feature-based design of aircraft structural parts

Multiple and concurrent iterations during the design phase of aircraft structural parts significantly increase the lead-time and cost of new aircraft products' development. Effective communication, collaboration and coordination of designers and software tools used by them are key mechanisms to address this challenge. This article presents an agent-based collaborative design framework to facilitate the collaboration of feature-based aircraft structural parts design and analysis tools, including detailed design, machining feature recognition, feature-based manufacturability evaluation and cost estimation. A new kind of attribute adjacency graph called dynamic holistic attribute adjacency graph (DHAAG) is established for local feature recognition. A dynamic machining feature relation matrix (DMFRM) is created for incremental manufacturability analysis. Process plans are generated based on DMFRM for cost estimation, and the result will feedback to designer for improving design. A prototype system has been implemented and demonstrated through a case study in collaboration with a major aircraft manufacturer in China.

항공기 복합재 구조에 적용된 두꺼운 적층판의 손상 허용 기준 평가

최근 국내에서 미국과 상호항공안전협정 체결을 위한 소형 항공기가 연구 개발 중이다. 연구 대상 항공기는 경량화 하여 연료 절감을 위해 전기체 복합재료가 적용되었다. 그러나 복합재 구조는 외부의 충격 손상에 취약한 구조이다. 따라서 항공기 구조물은 충격 손상에 대한 압축 파손 강도를 고려하여 손상 허용 설계가 반드시 수행되어야 한다. 이는 복합재 구조 항공기 인증에 매우 중요한 요소이다. 본 연구에서는 항공기 복합재 구조에 적용된 두꺼운 적층판에 대한 손상 허용 연구를 수행하였다. 두꺼운 적층판의 세 가지 형태인 손상이 없는 시편, 구멍 손상 및 충격 손상이 적용된 시편의 압축 하중 하에서 손상 허용 기준이 평가되었다. Recently, development of a small aircraft has been carried out for the BASA(Bilateral Aviation Safety Agreement) program in Korea. This aircraft adopted all composite structures for environmental friendly by low fuel consumption due to its lightness behavior. However the composite structure has disadvantage which is very weak against impact damages. Therefore, damage allowable design of aircraft structure must be performed considering compressive fracture strength. This point is very important for certification of composite structure aircraft. In this paper, it is performed the research on damage tolerance of thick laminate adopting aircraft structure. The damage tolerance of three different types of thick laminates such as no damage, open hole and impact damage is evaluated under compression loading.

Numerical Study on Impact Response of Aircraft Sandwich Wing Made of Fiber-Metal Laminate Face-Sheets Subjected to Bird Strike

Aircraft structures are frequently subjected to impacts from objects such as runway debris and birds. In new aircraft structural design, Fiber Metal Laminates (FMLs) play a significant role due to their excellent mechanical properties, particularly the impact properties. In this study, the aircraft sandwich wing with FML face-sheets are analyzed by finite element model for simulating the bird strike. The numerical simulations of bird strike impact are performed adopting a lagrangian approach to design the wing by MSC/PATRAN FE code. The numerical obtained results are compared with the results in the literature for validation of the model. The effect of fiber orientations, fiber types, metal types in FML face sheets in sandwich wing on impact responses are investigated. The impact responses are illustrated by displacement history, contact force history and energy absorption. According to these results, the sandwich panel with FML skin is suitable structure for energy absorption (that is the most important factor in impact phenomena). The lay-ups with titanium metal layer with aramid fibers are the best.

Structural Optimization Research of Composite Aircraft Based on Different Wing Stiffness Constraints

Wing stiffness, mainly including bending stiffness and torsion stiffness, is very important in aircraft structure design. In order to research the relation between wing stiffness and optimization results, this paper makes optimization analysis to a composite T-tail configuration and a composite flying wing configuration based on different wing stiffness constraints, and some valuable conclusions are achieved.

Evolving fuselage designs by incorporating SHM technologies

AbstractResearch efforts at Bombardier Aerospace have investigated design constraints that need to be addressed when considering the design of a future unconventional fuselage concept. Using a composite design platform integrated with structural health monitoring (SHM) technologies, the design space is broadened and there are many more unknowns to investigate from our current knowledge of the subject to date. The principal idea is to identify structural zones and quantify stress levels in areas of concern, referred to as ‘hot-spots’, by integrating an SHM system at a conceptual design level. This new concept would require an assessment of benefits and detriments in order to evaluate the certification processes, impact on maintenance, operation, and ownership costs.The integration of an SHM system using onboard sensors introduces many challenges, such as the requirement for multiple sensors in the structure and additional systems weight. The investigation shows how the aircraft structural design would be impacted and how design, stress, supply-chain, manufacturing, and the systems departments need to be harmonised in order to design a feasible and integrated SHM-structural fuselage concept.

The 2nd International conference on aircraft structural design special issue

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Educational Software for Stress Analysis of an Aircraft Fuselage

A stress analysis software based on MATLAB graphical user interface (GUI) has been developed. The developed software can be used to estimate internal loads on a fuselage and to compute the stresses at any point along the length of the fuselage of a given aircraft. The generalized formulation allows user to perform stress analysis even on tapered fuselages. The software is expected to be a useful tool for effective teaching learning process of courses on aircraft structures and aircraft structural design.