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Article
Publication date: 25 February 2014

Aleksandr Cherniaev

The genetic algorithm (GA) technique is widely used for the optimization of stiffened composite panels. It is based on sequential execution of a number of specific…

Abstract

Purpose

The genetic algorithm (GA) technique is widely used for the optimization of stiffened composite panels. It is based on sequential execution of a number of specific operators, including the encoding of particular design variables. For instance, in the case of a stiffened composite panel, the design variables that need to be encoded are: the number of plies and their stacking sequences in the panel skin and stiffeners. This paper aims to present a novel, implicit, heuristic approach for encoding composite laminates and, through its use, demonstrates an improvement in the optimization process.

Design/methodology/approach

The stiffened panel optimization has been formulated as a constrained discrete minimum-weight design problem. GAs, which use both new encoding schemes and those previously described in the literature, have been used to find near-optimal solutions to the formulated problem. The influence of the new encoding scheme on the searching capabilities of the GA has been investigated using comparative analysis of the optimization results.

Findings

The new encoding scheme allows the definition of stacking sequences in composites using shorter symbolic representations as compared with standard encoding operators and, as a result of this, a reduction in the problem design space. According to numerical experiments performed in this work, this feature enables GA to obtain near-optimal designs using smaller population sizes than those required if standard encoding schemes are used.

Originality/value

The approach to encoding laminates presented in this paper is based on the original heuristics. In the context of GA-based optimization of stiffened composite panels, the use of the new approach rather than the standard encoding technique can lead to a significant reduction in computational time employed.

Details

Engineering Computations, vol. 31 no. 1
Type: Research Article
ISSN: 0264-4401

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Article
Publication date: 6 July 2010

Fulvio Romano, Josè Fiori and Umberto Mercurio

This paper's aim is to focus on the design, manufacture and test of a stiffened panel in composite material with integrated longitudinal foam‐filled stiffeners, spar and…

Abstract

Purpose

This paper's aim is to focus on the design, manufacture and test of a stiffened panel in composite material with integrated longitudinal foam‐filled stiffeners, spar and rib caps, using one‐shot liquid infusion (LI) process, reducing weight and number of subparts respect to metallic reference baseline P180 Avanti vertical fin.

Design/methodology/approach

Extensive activities in computational applications in order to improve the efficiency of the design process finite element analysis/structural sizing codes have led to an optimised engineering design process that resulted in a successful stiffened carbon fibre reinforced polymer panel design in terms of weight and number of parts with respect to the metallic baseline.

Findings

The composite panel has fulfilled all the design requirements (reduction of mass and number of parts with respect to the metallic reference baseline) overcoming the certification static test, and confirming the reliability of the theoretical analyses.

Research limitations/implications

The composite aircraft components, conceived as unitized structure by one‐shot process, guarantee not only a mass reduction, compared to aluminium components, but assure also the reduction of the number of subparts and of the assembly process cycle time. On the other hand, the LI technology implies the development of more specific and advanced techniques to control the manufacturing and the weight.

Practical implications

The stiffened panel is the most used component in the aircraft structures; the solution shown in this work can find applications in many parts of an aircraft.

Originality/value

The results obtained in this work can be useful to those who work in aeronautical structural departments with the aim to reduce weight and subparts of the airframe.

Details

Aircraft Engineering and Aerospace Technology, vol. 82 no. 4
Type: Research Article
ISSN: 0002-2667

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Article
Publication date: 13 April 2015

Roman Ružek, Konstantinos Tserpes and Evaggelos Karachalios

Impact and fatigue are critical loading conditions for composite aerostructures. Compression behavior after impact and fatigue is a weak point for composite fuselage panels

Abstract

Purpose

Impact and fatigue are critical loading conditions for composite aerostructures. Compression behavior after impact and fatigue is a weak point for composite fuselage panels. The purpose of this paper is to characterize experimentally the compression behavior of carbon fiber reinforced plastic (CFRP) stiffened fuselage panels after impact and fatigue.

Design/methodology/approach

In total, three panels were manufactured and tested. The first panel was tested quasi-statically to measure the reference compression behavior. The second panel was subjected to impact so as to create barely visible impact damage (BVID) at different locations, then to fatigue and finally to quasi-static compression. Finally, the third panel was subjected to impact so as to create visible impact damage (VID) at different locations and then to quasi-static compression. The panels were tested using ultrasound inspection just after manufacturing to check material quality and between different tests to detect impact and fatigue damage accumulation. During tests the mechanical behavior of the panel was monitored using an optical displacement measurement system.

Findings

Experimental results show that the presence of impact damage significantly degrades the compression behavior of the panels. Moreover, the combined effect of BVID and fatigue was proven more severe than VID.

Originality/value

The paper gives information about the compression after impact and fatigue behavior of CFRP fuselage stiffened panels, which represent the most realistic loading scenario of composite aerostructures, and describes an integrated experimental procedure for obtaining such information.

Details

International Journal of Structural Integrity, vol. 6 no. 2
Type: Research Article
ISSN: 1757-9864

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Article
Publication date: 11 November 2014

A. Sellitto, R. Borrelli, F. Caputo, A. Riccio and F. Scaramuzzino

The purpose of this paper is to investigate on the behaviour of a delaminated stiffened panel; the delamination growth is simulated via fracture elements implemented in…

Abstract

Purpose

The purpose of this paper is to investigate on the behaviour of a delaminated stiffened panel; the delamination growth is simulated via fracture elements implemented in B2000++® code based on the Modified Virtual Crack Closure Technique (MVCCT), matrix cracking and fibre failure have been also taken into account.

Design/methodology/approach

In order to correctly apply the MVCCT on the delamination front a very fine three-dimensional (3D) mesh is required very close to the delaminated area, while a 2D-shell model has been employed for the areas of minor interest. In order to couple the shell domain to the solid one, shell-to-solid coupling elements based on kinematic constraints have been used.

Findings

Results obtained with the global/local approach are in good correlation with those obtained with experimental results.

Originality/value

The global/local approach based on kinematic coupling elements in conjunction with fracture elements allows to investigate and predict the behaviour of a stiffened delaminated composite panel in an efficient and effective way.

Details

International Journal of Structural Integrity, vol. 5 no. 4
Type: Research Article
ISSN: 1757-9864

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Article
Publication date: 6 February 2017

Roman Ruzek, Martin Kadlec, Konstantinos Tserpes and Evaggelos Karachalios

Compression is critical loading condition for composite airframes. Compression behaviour of structures with or without damages is a weak point for composite fuselage panels

Abstract

Purpose

Compression is critical loading condition for composite airframes. Compression behaviour of structures with or without damages is a weak point for composite fuselage panels. This is one of the reasons for need of continuous in-service health monitoring of composite structures. The purpose of this paper is to characterize the compression panel behaviour on the base of a developed and implemented structural health monitoring (SHM) system.

Design/methodology/approach

The SHM system based on fibre optic Bragg grating (FOBG) sensors and standard resistance strain gauges (SGs) was placed onto/into (embedded or bonded) three stiffened carbon fibre reinforced polymer (CFRP) fuselage panels. The FOBG sensor system was used to monitor the structural integrity of the reference, impacted, and fatigued panels under compression loading. Both barely visible impact damage and visible impact damage were created to evaluate their influence on the panel behaviour. The functionality of the SHM system was verified through mechanical testing.

Findings

Experimental data showed the presence of impact damages significantly changes the buckling modes development and deformation behaviour of the panels. Some differences between the optical and SG sensors during buckling were observed. The buckling waves and failure development were very well indicated during loading by all sensors located on the panel surface but not by the embedded sensors. Good agreement between the data from the SGs and FOBG sensors was achieved for all sensors placed on the stringers, which did not buckle. The good reliability of FOBG sensors during the fatigue and static testing up to panel failure was verified.

Originality/value

The paper gives information about different buckling behaviour of CFRP fuselage stiffened panels in compression. The paper gives detailed information about measured signals from different sensors based on their location on/in the panel structure for realistic loading scenario of composite aerostructures. The paper gives an integrated overview of sensors placement considering possibilities to predicate structure behaviour.

Details

International Journal of Structural Integrity, vol. 8 no. 1
Type: Research Article
ISSN: 1757-9864

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Article
Publication date: 7 March 2016

Tobias Bach, Tanja Führer, Christian Willberg and Sascha Dähne

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…

Abstract

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.

Design/methodology/approach

For fast sizing and optimization, linear static finite element (FE) models are used to obtain inner loads of the structural components. The inner loads and the geometry are passed to a software, where a comprehensive set of analytical failure criteria is applied for the design of the structure. In addition to conventional design processes, the objects of stiffened panels like skin and stringer are not optimized separately and discrete layups can be considered for composites. The module is connected to a design environment, where an automated steering of the overall process and the generation of the FE models is implemented.

Findings

The exemplary application on a transport aircraft wing shows the functionality of the developed module.

Originality/value

The weight benefit of not optimizing skin and stringer separately was shown. Furthermore, with the applied approach, a fast investigation of different aircraft configurations is possible without constraining too many design variables as it often occurs in other optimization processes. The flexibility of the module allows numerous investigations on influence of design concepts and failure criteria on the mass and layout of aircraft wings.

Details

Aircraft Engineering and Aerospace Technology: An International Journal, vol. 88 no. 2
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 9 March 2010

Euripidis Mistakidis

The purpose of this paper is to provide the research and practising engineers with insight on the benefits of using low‐yield point steel with respect to ordinary steel as…

Abstract

Purpose

The purpose of this paper is to provide the research and practising engineers with insight on the benefits of using low‐yield point steel with respect to ordinary steel as a construction material for shear wall panels. The paper seeks to focus on the behaviour of such panels when installed in new or existing structures in order to improve their seismic performance.

Design/methodology/approach

Finite element models are applied in order to approximate the structural response of low‐yield steel panels, used for seismic applications. Owing to the specific characteristics of the problem at hand, geometric and material nonlinearities have to be accurately considered. For comparison reasons, low‐yield point steel and ordinary steel are considered as construction materials for the aforementioned panels. The paper examines both the case of “pure shear” steel panel and also the more realistic case that the panel is encased in the surrounding frame.

Findings

The paper reaches a number of interesting conclusions. The beneficial behaviour of low‐yield steel panels with respect to ordinary steel panels is verified. Comments are made distinguishing the differences in the behaviour of panels surrounded by strong elements (“encased” panels) compared with that of panels submitted to pure shear. Finally, the improved seismic behaviour of existing structures retrofitted by shear wall panels is verified.

Originality/value

The paper exhibits numerically the advantages of low‐yield point steel with respect to ordinary steel as a construction material for panels and, second, contributes to the comprehension of the realistic panel behaviour of encased panels. More specifically, the paper focuses on the differences in the behaviour of encased steel panels with respect to the “pure shear” steel panels.

Details

Engineering Computations, vol. 27 no. 2
Type: Research Article
ISSN: 0264-4401

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Article
Publication date: 1 April 1984

I. Vayas

During the 1970s 4 steel bridges in Australia, England, Austria and Germany, failed due to the buckling of their compressed plates. As a result of these failures much…

Abstract

During the 1970s 4 steel bridges in Australia, England, Austria and Germany, failed due to the buckling of their compressed plates. As a result of these failures much research, both theoretical and experimental, has been initiated.

Details

Engineering Computations, vol. 1 no. 4
Type: Research Article
ISSN: 0264-4401

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Article
Publication date: 1 July 1954

John H. Argyris

The methods developed in sections (5)–(7) are now applied to the determination of the torsional cum flexural failing stresses in two panels. The following points should be noted:

Abstract

The methods developed in sections (5)–(7) are now applied to the determination of the torsional cum flexural failing stresses in two panels. The following points should be noted:

Details

Aircraft Engineering and Aerospace Technology, vol. 26 no. 7
Type: Research Article
ISSN: 0002-2667

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Article
Publication date: 1 December 1949

H.L. Cox and Mrs M.J. Windle

IN the present note a comparison is made between normal aluminium alloys and alloys with increased values of the modulus of elasticity for covering the upper surfaces of…

Abstract

IN the present note a comparison is made between normal aluminium alloys and alloys with increased values of the modulus of elasticity for covering the upper surfaces of wings of moderately thick sections, particularly of the smooth wing type. This comparison is intended to form the basis for the design of test panels for experimental verification of the theoretical conclusions.

Details

Aircraft Engineering and Aerospace Technology, vol. 21 no. 12
Type: Research Article
ISSN: 0002-2667

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