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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.

A method is presented for calculating influence coefficients in multispar wings of any plan form. Shear deflexion, chordwise bending and taper are taken into account. All the bending material is concentrated at the rib and spar booms and the skin is assumed to carry only shear. The method is particularly useful in the early stages of design as it is rapid and gives a good internal load distribution. The choice of the method, the assumptions made, and the idealization of the structure are discussed. The structure is reduced to a system of beams and torque boxes, the latter connected to the beams by vertical shear at their corner nodes. The internal forces are first expressed as functions of the displacements at the nodes by means of stiffness coefficients. Next the equations of equilibrium, at each node, are established. Finally the conditions of compatibility are brought in. The solution of this system of equations gives the matrix of influence coefficients. Wherever vertical loads only are applied it is possible to solve for the moment equations of straight beams separately from the remainder of the structure, hence from the stiffness matrix. For convenience, these straight beams are organized into sub‐matrices whose moment equations are ‘reduced’. The specific end conditions also allow further reductions. The internal loads can be obtained by substituting the influence coefficients back into the sub‐matrices and by considering the equilibrium of the beams. The angular displacements of the nodes may also be calculated in a similar manner by using the ‘reduced’ rows from the sub‐matrices.

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.

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.

A conceptual design method for composite material stiffened panels used in aircraft tail structures and unmanned aircraft has been developed to bear compression and shear loads.

The method is based on classical laminated theory to fulfil the requirement of building a fast design tool, necessary for this preliminary stage. The design criterion is local and global buckling happen at the same time. In addition, it is considered that the panel does not fail due to crippling, stiffeners column buckling or other manufacturing restrictions. The final geometry is determined by minimising the area and, consequently, the weight of the panel.

The results obtained are compared with a classical method for sizing stiffened panels in aluminium. The weight prediction is validated by weight reductions in aircraft structures when comparing composite and aluminium alloys.

The work is framed in conceptual design field, so hypotheses like material or stiffeners geometry shall be taken a priori. These hypotheses can be modified if it is necessary, but even so, the methodology continues being applicable.

The procedure presented in this paper allows designers to know composite structure weight of aircraft tails in commercial aviation or any lifting surface in unmanned aircraft field, even for unconventional configurations, in early stages of the design, which is an aid for them.

The contribution of this paper is the development of a new rapid methodology for conceptual design of composite panels and the feasible application to aircraft tails and also to unmanned aircraft.

THE paper suggests a new approach to the calculation of the bending strength of semi‐monocoque (and possibly fully mono‐coque) aircraft fuselage shells. The method described uses directly load‐deflexion graphs obtained from laboratory compression tests on panels representative of the shell at the section under consideration, and is analogous to the calculation of the form factor for a solid beam section, which factor is dependent on the shape of the section and on the stress‐strain relationship of the material of which the beam is made.

Ceramic materials and glasses have become important in modern industry as well as in the consumer environment. Heat resistant ceramics are used in the metal forming processes or as welding and brazing fixtures, etc. Ceramic materials are frequently used in industries where a wear and chemical resistance are required criteria (seals, liners, grinding wheels, machining tools, etc.). Electrical, magnetic and optical properties of ceramic materials are important in electrical and electronic industries where these materials are used as sensors and actuators, integrated circuits, piezoelectric transducers, ultrasonic devices, microwave devices, magnetic tapes, and in other applications. A significant amount of literature is available on the finite element modelling (FEM) of ceramics and glass. This paper gives a listing of these published papers and is a continuation of the author's bibliography entitled “Finite element modelling of ceramics and glass” and published in Engineering Computations, Vol. 16, 1999, pp. 510‐71 for the period 1977‐1998.

The form of the paper is a bibliography. Listed references have been retrieved from the author's database, MAKEBASE. Also Compendex has been checked. The period is 1998‐2004.

Provides a listing of 1,432 references. The following topics are included: ceramics – material and mechanical properties in general, ceramic coatings and joining problems, ceramic composites, piezoceramics, ceramic tools and machining, material processing simulations, fracture mechanics and damage, applications of ceramic/composites in engineering; glass – material and mechanical properties in general, glass fiber composites, material processing simulations, fracture mechanics and damage, and applications of glasses in engineering.

This paper makes it easy for professionals working with the numerical methods with applications to ceramics and glasses to be up‐to‐date in an effective way.

The deformation of a large fuselage panel is unavoidable due to its weak-stiffness and low-rigidity. Sometimes, the assembly accuracy of the panel is out of tolerance. The purpose of this paper is to propose a method to predict and correct the assembly deformation of a large fuselage panel during digital assembly by using a finite element (FE) analysis and partial least squares regression (PLSR) method.

A FE model is proposed to optimize the layout of load-transmitting devices to reduce panel deformation after the process of hoisting and supporting. Furthermore, another FE model is established to investigate the deformation behavior of the panel. By orthogonal simulations, the position error data of measurement points representing the precision of the panel are obtained. Then, a mathematical model of the relationship between the position errors of measurement points on the panel and the displacements of numerical control positioners is developed based on the PLSR method.

The case study shows that the model has a high level of computing accuracy and that the proposed method is an efficient way to correct the panel deformation in digital assembly.

The results of this study will enhance the understanding of the deformation behavior of a panel in aircraft digital assembly and help to improve the assembly precision systematically and efficiently.

CONSIDER a fuselage or wing structure in the form of a reinforced cylindrical tube. We shall base our analysis of the equilibrium conditions of this structure upon the assumptions outlined in 2.6. In particular referring in the first place to a skin panel lying between adjacent stringers and rings, we remark that this panel carries only shear stresses and is free from external forces. It follows, as we have observed before, that this panel must therefore be in a state of uniform shearing and so must apply uniform shear flows at its lines of juncture with the adjacent panels and the reinforcing stringers and rings. The equilibrium conditions to be satisfied at a stringer‐skin joint are now clear. The panels adjacent to the stringer apply different, but uniform, shear flows, to the line of attachment. The reaction from the stringer is determined by the rate of variation of its end load, for this clearly gives the rate of load input into the stringer. Adopting a consistent sign convention for the shear flows in the several skin panels we can thus enunciate the following theorem: