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A novel development process aims at finding solutions for lightweight stiffened shell structures and their efficient production. To respect the strong interdependency of structural design and production planning, particularly observed for composite structures, it is of high interest to start considering production effects in early development phases. This integrated approach requires an integrated representation of structure and production. The purpose of this study is to investigate the scope of relevant data and to find a structure for its representation.

The development task is analyzed and a system of so-called solution dimensions is presented, which covers all important aspects of stiffened shell structures and their production. An integrated product data model is developed to cover all of the solution dimensions.

The product data model consists of five coherent partial models. It is explained how these models are defined and how they are connected to each other. An academic example of an aircraft fuselage panel is used to demonstrate the definition process. It is shown how even complex structural concepts are defined systematically.

It is explained how this integrated product data model is used in a software project for the development of aircraft fuselage structures.

The presented approach for the definition and representation of stiffened shell structures enables the developer, e.g. of aircraft fuselage, to respect the crucial criterion of manufacturability from early development phases on. Further, new design approaches, e.g. as inspired by topology optimization, can be considered.

The purpose of this paper is to present the first results of tests where cracks lie in the tension field of a shear forced aluminium panel. The paper's main focus is on the crack propagation behavior and possible 3D‐effects caused by the bending of the plate. A simplified numerical approach is presented to confirm the observed phenomena.

Experiments have been performed to investigate the influence of buckling on accidental damages. A simplified numerical approach is presented and compared to the experimental results.

It can be shown that the crack propagates due to buckling of the plate. The principal stress of the neutral axis of the plate has significant influence on the crack propagation.

Investigations of stability problems and damage tolerance behavior of metallic structures have been realized but mainly separately. This paper shows that cracks propagate due to buckling and that both phenomena influence one another considering accidental damages. The paper presents the first experimental and numerical results of cracked aluminium panels subjected to cyclic shear load.

In former work, test results of cracks in aluminium panels under cyclic shear buckling showed that cracks in the tensile stress field of a buckle propagate. The main influencing factor for the crack growth rate is the maximum principle stress. A simplified approach for crack propagation analyses based on this finding showed limitations for application on larger cracks because it disregarded the increasing out-of-plane deformation for larger cracks as well as stress redistributions. The purpose of this paper is to improve the results of the simplified approach with the help of finite element method (FEM).

An approach for crack propagation based on FEM is presented taking into account the mutual interaction of cracks and buckling. The finite element (FE) model, which is described in detail, respects the boundary conditions of the test-set-up. Different initial crack positions, loads and panel thicknesses are analyzed. Results of the stress intensity factors KI calculated by the ABAQUS® FE model provide a function which is used to run a crack propagation analysis based on Forman law.

The results of the FE-based crack propagation solution are in good agreement with test results and improve the prediction of the simplified approach. It is not restricted in terms of panel thickness, crack position or applied shear load.

Limitations of the FE-based crack propagation solution compared to the experimental results are discussed. These are, the sensitivity of crack propagation analyses to initial crack length and deviations of the experimental settings from the ideal FE model.

The interaction of cracks and buckling in aluminium shells is mainly disregarded both in research and industrial work, but can be of interest considering, accidental damages in fuselage side shells. Cracks propagate under shear load as it was shown in former work. The FE modeling of the tests presented in this paper proves the mutual interactions of crack propagation and buckling deformation.

In general two main types of criteria are essential for the sizing of aircraft structural panels, namely, stability and damage tolerance. The way these criteria act and interact is very different for metallic and composite building blocks. While interaction of both types of criteria is relatively clear for composite parts, this is normally not the case for metallic ones. What is common for both is the fact that, if an interaction occurs, the impact is essential. The paper aims to discuss these issues.

This is a survey paper.

There is a strong mutual influence of buckling and damage in many cases.

It shows the significance of both, buckling and damage as a combined phenomenon.

This paper seeks to focus on material combinations for flexible matrix composites (FMCs) and the production methods thereof. These materials enable a high flexibility in one direction while being very stiff in the other.

Tested were rubber, silicone and thermoplastic elastomer matrices with carbon fibers using different production methods. These tests focused on the impregnation of the fibers with the different matrices and the orthotropy of the produced materials.

In the paper, a production capability for large quantities of easy to use off‐the‐shelf material was developed. The produced material handles similar to prepreg material known from “classical” composite materials. Test specimens were manufactured and characterized for mechanical properties using tensile tests.

These FMC materials are envisaged for a new pneumatic actuation system for an aircraft's droop nose to replace the electro‐mechanical system designed in the SADE and SmartLED projects. Combining a tube‐like geometry and a variable fiber‐angle lay‐up enables a wide range of deformation possibilities (large design freedom of movement behaviour).

The University of Hannover has studied the fundamental factors affecting the process of vibratory feeding. From this research a step‐by‐step procedure for the design of vibratory bowl feeders has been derived.

The purpose of this paper is to describe the pre‐design and sizing of a smart leading edge section which is developed in the project SADE (Smart High Lift Devices for Next Generation Wings), which is part of the seventh framework program of the EU.

The development of morphing technologies in SADE concentrates on the leading and trailing edge high‐lift devices. At the leading edge a smart gap and step‐less droop nose device is developed. For the landing flap a smart trailing edge of the flap is in the focus of the research activities. The main path in SADE follows the development of the leading edge section and the subsequent wind tunnel testing of a five meter span full‐scale section with a chord length of three meters in the wind tunnel T‐101 at the Russian central aero‐hydrodynamic institute (TsAGI) in Moscow.

The presented paper gives an overview over the desired performance and requirements of a smart leading edge device, its aerodynamic design for the wind tunnel tests and the structural pre‐design and sizing of the full‐scale leading edge section which will be tested in the wind tunnel.

SADE aims at a major step forward in the development and evaluation of the potential of morphing airframe technologies.

The purpose of this paper is to focus on a morphing architecture, conceived to produce droop nose effect, thus preserving high lift performance and laminar flow.

A numerical approach was adopted. On the base of preliminary aerodynamic requirements, the main aspects of the actuation architecture were defined and then assessed through a genetic approach.

Two different working modalities of mentioned architecture were identified: the former implying the use of an actuator, the latter taking advantage of a tailored elastic element, able to actuate morphing under the action of aerodynamic loads, without the aid of actuators.

The research presented in this work refers to an optimisation process currently tailored on preliminary aerodynamic requirements (leading edge vertical displacement maximisation, leading edge radius increase).

The research shows the possibility of producing morphing on the leading edge zone, actuating droop nose effect on metallic (constant and pice wise thickness) and composite skins.