Richard Degenhardt

A Finite Element Methodology for Analysing Degradation and Collapse in Postbuckling Composite Aerospace Structures

A methodology for analysing the degradation and collapse in postbuckling composite structures is proposed. One aspect of the methodology predicts the initiation of interlaminar damage using a strength criterion applied with a global-local analysis technique. A separate approach represents the growth of a pre-existing interlaminar damage region with user-defined multi-point constraints that are controlled based on the Virtual Crack Closure Technique. Another aspect of the approach is a degradation model for in-plane ply damage mechanisms of fiber fracture, matrix cracking, and fiber-matrix shear. The complete analysis methodology was compared to experimental results for two fuselage-representative composite panels tested to collapse. For both panels, the behavior and structural collapse were accurately captured, and the analysis methodology provided detailed information on the development and interaction of the various damage mechanisms.

Investigation of Buckling Behavior of Composite Shell Structures with Cutouts

Thin-walled cylindrical composite shell structures can be applied in space applications, looking for lighter and cheaper launcher transport system. These structures are prone to buckling under axial compression and may exhibit sensitivity to geometrical imperfections. Today the design of such structures is based on NASA guidelines from the 1960’s using a conservative lower bound curve generated from a database of experimental results. In this guideline the structural behavior of composite materials may not be appropriately considered since the imperfection sensitivity and the buckling load of shells made of such materials depend on the lay-up design. It is clear that with the evolution of the composite materials and fabrication processes this guideline must be updated and / or new design guidelines investigated. This need becomes even more relevant when cutouts are introduced to the structure, which are commonly necessary to account for access points and to provide clearance and attachment points for hydraulic and electric systems. Therefore, it is necessary to understand how a cutout with different dimensions affects the buckling load of a thin-walled cylindrical shell structure in combination with other initial geometric imperfections. In this context, this paper present some observations regarding the buckling load behavior vs. cutout size and radius over thickness ratio, of laminated composite curved panels and cylindrical shells, that could be applied in further recommendations, to allow identifying when the buckling of the structure is dominated by the presence of the cutout or by other initial imperfections.

AN ANALYSIS TOOL FOR DESIGN AND CERTIFICATION OF POSTBUCKLING COMPOSITE AEROSPACE STRUCTURES

In aerospace, carbon-fiber-reinforced polymer (CFRP) composites and postbuckling skin-stiffened structures are key technologies that have been used to improve structural efficiency. However, the application of composite postbuckling structures in aircraft has been limited due to concerns related to both the durability of composite structures and the accuracy of design tools. In this work, a finite element analysis tool for design and certification of aerospace structures is presented, which predicts collapse by taking the critical damage mechanisms into account. The tool incorporates a global–local analysis technique for predicting interlaminar damage initiation, and degradation models to capture the growth of a pre-existing interlaminar damage region, such as a delamination or skin–stiffener debond, and in-plane ply damage mechanisms such as fiber fracture and matrix cracking. The analysis tool has been applied to single- and multistiffener fuselage-representative composite panels, in both intact and predamaged configurations. This has been performed in a design context, in which panel configurations are selected based on their suitability for experimental testing, and in an analysis context for comparison with experimental results as being representative of aircraft certification studies. For all cases, the tool was capable of accurately capturing the key damage mechanisms contributing to final structural collapse, and suitable for the design of next-generation composite aerospace structures.

Dealing with Imperfection Sensitivity of Composite Structures Prone to Buckling

The Space industry demands for lighter and cheaper launcher transport systems. Structural weight reduction by exploitation of structural reserves in composite launcher structures contributes to this aim, however, it requires accurate, fast and experimentally validated stability analysis of real structures under realistic loading conditions. Structures in space applications can be imperfection sensitive because their maximum load is often equal or close to the first buckling load. The current design guidelines were developed only for metallic structures and are from 1968. For composites structures no appropriate guidelines exist. To fill this gap DLR developed a promising “Single Perturbation Load Approach” which exploits the worst imperfections idea efficiently. In the running EU project DESICOS (New Robust DESIgn Guideline for Imperfection Sensitive COmposite Launcher Structures) this approach will be further investigated and combined with a stochastic approach resulting in a future design approach. This chapter deals with the state-of-the-art in buckling of imperfection sensitive composite structures, recent investigations on the new design approach, and the DESICOS project. It describes the line of actions of the new design approach, and specifies the theoretical and experimental work to be carried out.

CYCLIC BUCKLING TESTS OF PRE-DAMAGED CFRP STRINGER-STIFFENED PANELS

Experimental results obtained from cyclic buckling and postbuckling tests of pre-damaged stiffened CFRP panels are presented in this paper. This work was conducted within the COCOMAT project funded by the EU with the objective of contributing to the reduction of structural weight at safe design. COCOMAT was targeted at establishing a new design scenario for composite stiffened panels which are understood as part of an aircraft fuselage. This design scenario aimed at exploiting considerable reserves in the load carrying capacity in fiber composite fuselage structures by accurate simulation of collapse. The project results cover an experimental database, improved slow and fast computational tools, as well as design guidelines. A reliable simulation of the collapse load requires taking degradation into account. For the validation of the tools, a sound database of experiments is needed which gives information about the progress of damage during the loading process. In this context, the present paper focuses on the investigation of pre-damaged stringer-stiffened panels under cyclic axial compression. A set of four panels of the same design was split into two variants which differ only in the position of an artificial Teflon disbond beneath a stringer. One panel of each variant was tested statically until collapse in one step as reference, while the other panel was tested cyclically with different amplitudes. Before the test, all test structures were assessed by ultrasonic inspection and the geometric imperfections were measured. During the test, advanced measurement systems such as the ARAMIS system for the measurement of the buckling pattern and thermography for monitoring the skin-stringer separation were utilized in addition to strain gauge measurements and the record of the load shortening data. The test structures, their preparations for testing, the buckling test facility, and the measurement systems used are described. The test results as to the influence of the cyclic loading on the damage progression in the skin-stringer connection are presented and discussed.

Experiments to Detect Damage Progression in Axially Compressed CFRP Panels under Cyclic Loading

The aircraft industry strives for significantly reduced development and operating costs. Reduction of structural weight at safe design is one possibility to reach this objective which is aimed by the running EU project COCOMAT. The main objective of COCOMAT is a future design scenario for composite curved stiffened panels which are understood as parts of real aircraft structures. This design scenario exploits considerable reserve carrying capacities in fibre composite fuselage structures by accurate simulation of collapse. The project results will comprise an experimental data base, improved slow and fast computational tools as well as design guidelines. A reliable simulation of the collapse load requires also taking degradation into account. For the validation of the tools a sound database of experiments are needed which give information about the progress of damage during the loading process. This paper focuses on experimental results of four nominally identical CFRP panels tested within the COCOMAT project at the buckling test facility of the Institute of Composite Structures and Adaptive Systems (DLR). In a first step, three of the four panels were loaded several thousand times. Each time the panel was loaded beyond global buckling and was unloaded to zero. Finally, all panels were tested until collapse. During the tests, advanced measurement systems such as High-Speed-ARAMIS, thermography and Lamb-waves were applied. The test results given in this paper may be used as benchmarks.

METAMODELING METHODOLOGY FOR POSTBUCKLING SIMULATION OF DAMAGED COMPOSITE STIFFENED STRUCTURES WITH PHYSICAL VALIDATION

A metamodeling methodology has been proposed for postbuckling simulation of stiffened composite structures with integrated degradation scenarios. The presence of artificial damage in between the outer skin and stiffeners has been simulated as softening of the material properties in predetermined regions in the structure. The proposed methodology for the fast design procedure of axially or torsionally loaded stiffened composite structures is based on response surface methodology (RSM) and design and analysis of computer experiments (DACE). Numerical analyses have been parametrically sampled by means of ANSYS/LS-DYNA probabilistic design toolbox extracting the load-shortening response curves in the preselected domain of interest. These response curves have been simplified using piece-wise linear approximation identifying the buckling and postbuckling stiffness ratios along with the values of the skin and the stiffener buckling loads. Three stiffened panel designs and a closed box structure with preselected damage scenarios have been elaborated and validated with the tests performed within the COCOMAT project. The resulting design procedure provides a time effective design tool for preliminary study and for elaboration of the optimum design guidelines of composite stiffened structures with material degradation restraints.

Design and Analysis of Composite Panels

European aircraft industry demands for reduced development and operating costs, by 20% and 50% in the short and long term, respectively. Contributions to this aim are provided by the completed project POSICOSS (5thFP) and the running follow-up project COCOMAT (6thFP), both supported by the European Commission. As an important contribution to cost reduction a decrease in structural weight can be reached by exploiting considerable reserves in primary fibre composite fuselage structures through an accurate and reliable simulation of postbuckling up to collapse. The POSICOSS team developed fast procedures for postbuckling analysis of stiffened fibre composite panels, created comprehensive experimental data bases and derived design guidelines. COCOMAT builds up on the POSICOSS results and considers in addition the simulation of collapse by taking degradation into account. The results comprise an extended experimental data base, degradation models, improved certification and design tools as well as design guidelines. The projects POSICOSS and COCOMAT develop improved tools which are validated by experimental results obtained during the projects. Because the new tools must consider a wide range of different aspects a lot of different structures had to be tested. These structures were designed under different design objectives. For the design process the consortium applied already available simulation tools and brought in their own design experience. This paper deals with the design process within both projects and the analysis procedure applied within this task. It focuses on the experience of DLR on the design and analysis of stringer stiffened CFRP panels gained in the frame of these projects.

Experimental Nondestructive Test for Estimation of Buckling Load on Unstiffened Cylindrical Shells Using Vibration Correlation Technique

Nondestructive methods, to calculate the buckling load of imperfection sensitive thin-walled structures, such as large-scale aerospace structures, are one of the most important techniques for the evaluation of new structures and validation of numerical models. The vibration correlation technique (VCT) allows determining the buckling load for several types of structures without reaching the instability point, but this technique is still under development for thin-walled plates and shells. This paper presents and discusses an experimental verification of a novel approach using vibration correlation technique for the prediction of realistic buckling loads of unstiffened cylindrical shells loaded under axial compression. Four different test structures were manufactured and loaded up to buckling: two composite laminated cylindrical shells and two stainless steel cylinders. In order to characterize a relationship with the applied load, the first natural frequency of vibration and mode shape is measured during testing using a 3D laser scanner. The proposed vibration correlation technique allows one to predict the experimental buckling load with a very good approximation without actually reaching the instability point. Additional experimental tests and numerical models are currently under development to further validate the proposed approach for composite and metallic conical structures.

The Influence of Geometrical Parameters on the Buckling Behavior of Conical Shell by the Single Perturbation Load Approach

Since the development of the first theories to predict the buckling induced by axial compression in shells sensitive to imperfections, a significant discrepancy between theoretical and experimental results has been observed. Donnell and Koiter are among the first authors demonstrating, for these structures, the relevant influence of the geometrical imperfections on the reduction of the buckling load. Currently, the preliminary design of imperfections sensitive shell structures used in space applications is carried out according to the NASA SP-8007guideline. However, several studies have proven that this guideline leads to over-conservative design configurations when considering the geometrical and material imperfections existing in real cones. Since the pioneer work of Arbocz, alternative methods have been investigated to overcome this issue. Among the different approaches, in this paper, the Single Perturbation Load Approach (SPLA), originally developed byHühne as a deterministic way to calculate the knock-down factor of imperfection sensitive shells, is further studied. Indeed, a numerical investigation about the application of the SPLA to the simulation of the mechanical behavior of imperfection sensitive composite conical structures under axial compression is presented. This study is related to part of the work performed in the frame of the European Union (EU) project DESICOS.