Buckling load reduction for stiffened panels due to cutouts in ribs

In aerospace structures it is common to find stiffened panels with transverse supporting structures, e.g. wing ribs or fuselage frames. Incorporating cutouts into these supporting structures to allow the stringers to pass through freely considerably reduces the buckling load of the panels. It is shown that a minor modification in the fabrication of the stiffened panel gives most of the advantages of cutouts while still giving a buckling load close to that of a panel with no cutouts

Design of structures by a splitting method

A simplified method for the design of multi-spar wing boxes is presented. In typical multi-spar wing boxes the spars divide the boxes into cells. In the method presented these are analyzed individually, with adjacent cells taking their share of the stiffnesses of the common spar wall. This splitting method yields a design method that is computationally much quicker than designing a complete wing box, because each cell is considered separately from the others, except for linking between their design variables. The critical buckling load factor of the assembled structure when designed in this way will usually exceed the design load factor and otherwise will be equal to it, i.e. the design is guaranteed to be conservative

Buckling and vibration of stiffened panels or single plates with clamped ends

An efficient method for the buckling and vibration analysis of plates or stiffened panels with clamped ends is presented. The method uses Lagrangian multipliers to couple sinusoidal modes with appropriate half-wavelengths of response, thereby enforcing the end conditions at discrete point supports. Clamped ends can usually be modelled accurately using only a few point supports, while arguments from symmetry often enable some of the required end conditions to be satisfied without explicitly applying constraints. In such cases few half-wavelengths are needed to obtain excellent accuracy. Solutions obtained for the simple limiting case of single plates are exact or within 1% of the classical or other reported solutions. Solutions obtained for stiffened panels are in close agreement with those obtained using finite element analysis

Optimum design and testing of a postbuckled stiffened panel

The efficient, industrially used, linear elastic preliminary design software VICONOPT is employed to design a stiffened panel with a post-buckled reserve of strength. The initial buckling mode is a local skin mode in longitudinal compression with allowance being made for the effects of an initial overall imperfection. The resulting panel has been analyzed using the non-linear FE package ABAQUS and four laboratory specimens have been tested to failure. The similarity of the experimental failure with the VICONOPT and ABAQUS predictions suggests that VICONOPT can give a satisfactory preliminary design. While neither model matches completely the boundary conditions found in a real aircraft compression panel, it is suggested that the VICONOPT model may be a better representation than either the ABAQUS model or the experimental tests

Analysis and testing of a postbuckled stiffened panel

The suitability of using the ef cient, linear elastic design softwareVICONOPT for the analysisof a stiffened panel with a postbuckling reserve of strength is investigated. A longitudinallycompressed panel, which initially buckled in a local skin mode, was analyzed with allowance being made for the effects of an initial overall imperfection. The panel was also analyzed using the nonlinear nite element package ABAQUS, and four laboratory specimens that represent the panel were tested to failure. The similarity of the experimental failure with the VICONOPT and ABAQUS predictions indicates that VICONOPT can give satisfactory analysis results for use in preliminary design

Optimum design and testing of a post-buckled stiffened panel

The efficient, industrially used, linear elastic preliminary design software VICONOPT is employed to design a stiffened panel with a post-buckled reserve of strength. The initial buckling mode is a local skin mode in longitudinal compression with allowance being made for the effects of an initial overall imperfection. The resulting panel has been analyzed using the non-linear FE package ABAQUS and four laboratory specimens have been tested to failure. The similarity of the experimental failure with the VICONOPT and ABAQUS predictions suggests that VICONOPT can give a satisfactory preliminary design. While neither model matches completely the boundary conditions found in a real aircraft compression panel, it is suggested that the VICONOPT model may be a better representation than either the ABAQUS model or the experimental tests