Stiffener Interface Research Articles

A unified modeling strategy of the stability and progressive damage behavior of CFRP double-blade composite stiffened structures (DCSS) under uniaxial compression

As a typical thin-walled structure, CFRP double-blade composite stiffened structure (DCSS) is widely used in the aerospace field and its stability and failure behavior demand a comprehensive understanding. In this work, a novel, unified modeling strategy is proposed for predicting the stability behavior and progressive damage processes of the DCSS under compression. A user-defined materials subroutine of the based-strain 3D-Hashin criterion and continuum damage mechanics model is developed using the FORTRAN language to simulate the intra-laminar damage initiation and propagation processes in the DCSS skin, stiffeners, and triangle zones. Moreover, the damage initiation and the extension of the inter-laminar in the skin and the skin–stiffener interface are revealed using the cohesive zone model. Furthermore, various experimental methods are employed to obtain the buckling, post-buckling and collapse behavior of the DCSS. The digital image correlation technique is used to describe the buckling and post-buckling displacement fields and the failure behavior of stiffened structures. Strain gauges are adopted to accurately acquire the buckling behavior of the DCSS. Finally, the buckling load, ultimate load, displacement data and failure modes of the stiffened structure obtained from the developed 3D finite element model are verified with the experimental results. The high correlation between the numerical and experimental results confirms the reliability of the proposed modeling strategy.

Laser-based pretreatment of composite T-joints for improved pull-off strength and toughness

This study demonstrates the efficiency of two CO2 surface laser treatment strategies for improving the strength and toughness of the adhesive bonding of T-joints. The first strategy involves uniform laser treatment to the surface, and the second strategy involves laser patterning by alternatively applying high- and low-power laser irradiation to the T-joint stiffener and skin surfaces. All the laser-treated T-joints showed progressive failure modes during pull-off test, unlike the PP surface-treated joints that exhibited brittle failure. The uniform laser-treated joints exhibited improved strength and toughness because contaminants were removed and the fibers were exposed to be directly bonded with adhesive. The joints subjected to the laser-patterning strategy demonstrated a substantial improvement in terms of the strength and toughness, reaching values of 2.2 and 12 times those of the PP surface-treated T-joints, respectively. This is because of the crack arresting at the transition between the two treatments at the stiffener interface, inducing crack migration from the stiffener–adhesive interface to the skin–adhesive interface while forming adhesive ligaments. The formation, deformation, and breakage of these ligaments dissipated more energy, hence increasing the effective joint toughness and strength.

Local high edge observation for the bonding line monitoring of a complex composite beam

The most part of defects in composite structures carrying attached subelements is the disbond at the interface, as the skin/stringer sections. This is sometimes due to a nonoptimal manufacturing process or sometimes due to accidental object impacts during service. It has been verified that structural discontinuities within an elastic medium under mechanical loads can cause analogous discontinuities within the strain field. Starting from this analysis, the present work investigates the effect of artificially induced kissing bond areas just at the in the skin–stiffener interface of an aeronautical complex composite beam. This research uses longitudinal strain values, acquired at the locations where distributed fiber optic sensors are installed. The applied methodology uses different strain-based features providing local high edge observation both in time and spatial domains. Their autocorrelations are, in the end, computed to improve signal-to-noise ratio. The local high edge observation algorithm is proposed that proves its capability to monitor disbond being at the same time load and baseline independent.

Constraint method for laminated composite flat stiffened panel analysis using variational asymptotic method (VAM)

Abstract In this work, a new methodology is developed for constructing the stiffened panel structure. Here we introduce the constraint method to integrate the stiffener and plate. The geometric nonlinear analysis is performed to analyze the stiffened panel behavior in terms of load-deformation response curve. Variational Asymptotic Method (VAM) is applied to analyze composite flat stiffened panels. The VAM development begins with the 3-D nonlinear strain energy functional which splits into a 1-D through-the-thickness analysis and a 2-D plate analysis. The 2-D constitutive law that is an output of 1-D analysis is used as an input to perform the 2-D nonlinear plate analysis. The integration of skin and the stiffener technique is implemented by introducing a constraint matrix. At the skin and stiffener interface, the developed constraint matrix is used in realization of 3-D displacements of stiffener and skin are same. The proposed approach is simple and reliable to address the complexity in analyzing stiffened structures. This methodology is developed using a computational symbolic tool Mathematica and the implemented computer program named as NASSVAM (Nonlinear Analysis of Stiffened Structures using Variational Asymptotic Method). The obtained results from NASSVAM are compared and showed good agreement with the 3-D FEA results.

Analysis of Mode-I crack growth arresting mechanism in curved laminated joint

The critical damage for carbon fiber structures, which are being employed in aircraft structures, is known as delamination/crack. Fasteners are commonly installed to arrest the delamination by clamping the laminate together. Fasteners are installed in each corner of the delaminated zone to provide significant arrest capability, shifting the failure mode away from delamination under most conditions. The Finite Element Analysis (FEA) model is constructed to study the effectiveness of the fastener as crack arrest mechanism. The FEA results show that the fastener provides significant crack retardation capability in Mode-I condition. An analytical model is developed for the delamination embedded between the skin and stiffener interface of the joint. The fasteners are modeled with spring elements. The analysis is solved with Virtual Crack Closure Technique (VCCT) approach. The primary objective of the current research work is to enhance the safety of bonded joint by providing arrest mechanisms.

Analysis of Mode I and Mode II Crack Growth Arrest Mechanism with Z-Fibre Pins in Composite Laminated Joint

This paper presents the numerical study of the mode I and mode II interlaminar crack growth arrest in hybrid laminated curved composite stiffened joint with Z-fibre reinforcement. A FE model of hybrid laminated skin-stiffener joint reinforced with Z-pins is developed to investigate the effect of Z- fibre pins on mode I and mode II crack growth where the delamination is embedded inbetween the skin and stiffener interface. A finite element model was developed using S4R element of a 4-node doubly curved thick shell elements to model the composite laminates and non linear interface elements to simulate the reinforcements. The numerical analyses revealed that Z-fibre pinning were effective in suppressing the delamination growth when propagated due to applied loads. Therefore, the Z-fibre technique effectively improves the crack growth resistance and hence arrests or delays crack growth extension.

Investigation Analysis of Crack Growth Arresting with Fasteners in Hybrid Laminated Skin-Stiffener Joint

In recent years carbon fibre-reinforced polymers (CFRP) emerged its increasing demand in aerospace engineering. Due to their high specific strength to weight ratio, these composites offer more characteristics and considerable advantages compared to metals. Metals, unlike composites, offer plasticity effects to evade high stress concentrations during postbuckling. Under compressive load, composite structures show a wide range of damage mechanisms where a set of damage modes combined together might lead to the eventual structural collapse. Crack is one of the most critical damages in fiber composites, which are being employed in primary aircraft structures. A parametric study is conducted to investigate the arrest mechanism of the delamination or crack growth with installation of multiple fasteners when the delamination is embedded in between the skin and stiffener interface.

Vibration Analysis of Curvilinearly Stiffened Composite Panel Subjected to In-Plane Loads

This paper presents an efficient finite element approach to study the prestressed vibration mode results of a curvilinearly stiffened composite panel subjected to various in-plane loads. The present method models the plate and the stiffener separately, which allows the stiffener element nodes to not coincide with the plate shell-element nodes. The stiffness and mass matrices of a stiffener are transformed to those of the plate through the displacement compatibility conditions at the plate–stiffener interface via finite element interpolation. Convergence and validation studies have been conducted to verify the present method in the finite element vibration analysis by using the examples from the existing literature. Prestressed vibration mode results are examined for a stiffened composite panel with arbitrarily shaped composite stiffeners in the presence of the in-plane normal and shear loads. Numerical results show the possible benefits of using curvilinear stiffeners to improve the vibration response by increasing both the fundamental frequency and the buckling load through changing the vibration and buckling mode shapes with a negligible or even no weight penalty. The stiffener depth ratio is found to increase both the fundamental frequency and the buckling load, but only up to a certain value. Any further increase in the stiffener depth ratio causes stiffener buckling before plate buckling, and it leads the free and prestressed vibration mode results to behave as the vibration mode results for a plate with simply supported boundary conditions along the stiffeners.

Impact Analysis of Embedded Delamination Location in Hybrid Curved Laminated Composite Stiffened Panel

Modern, aero structures are predominantly of curved construction characterized by a skin and stiffeners. The latest generation of large passenger aircraft also uses mostly composite material in their primary structure and there is trend towards the utilization of bonding of subcomponents. The presence of delamination is a major problem in composite laminated panels and so, it is of great concern to both the academic and aeronautical industrial worlds Indeed delamination can strongly affect the material strength and, sometimes, can cause their breaking up in service. A Pre-damaged configuration is loaded to study the delamination location and mode for delamination initiation and propagation. A parametric study is conducted to investigate the effect of the location of the delamination propagation when delamination is embedded inbetween plies of the skin-stiffener interface, with the cases i) delamination located at front and inbetween plies of the skin-stiffener interface ii) delamination located in middle and inbetween plies of the skin-stiffener interface iii) delamination located at the end and inbetween plies of the skin- stiffener interface. Further the influence of the location of the delamination on load carrying capacity of the panel is investigated. The effect of location of debonds on crack growth and collapse behavior is analyzed using analysis tool. An analysis tool is applied that includes an approach for predicting interlaminar damage initiation and interlaminar damage growth as well as in-plane damage mechanisms to predict the design of defect free panel.

Analytical prediction of Mode I stress intensity factors for cracked panels containing bonded stiffeners

An analytical model for stress intensity factors in cracked skin panels containing bonded stiffening elements, based on the principles of superposition and displacement compatibility, is presented. The methodology permits assessment of the influence of stiffening elements ahead, behind, and over top of the skin panel crack tip, and allows the influence of stiffener interface disbond size to be studied. Validation of the model is made using experimental data from various literature sources. The influence of stiffener geometry and interface disbond size on damage tolerance is discussed using predictions from the presented model.

Design of composite stiffener run-outs for damage tolerance

A major concern in stiffener run-out regions, where the stiffener is terminated due to a cut-out, intersecting rib, or some other structural feature which interrupts the load path, is the relatively weak skin–stiffener interface in the absence of mechanical fasteners. More damage tolerant stiffener run-outs are clearly required and these are investigated in this paper. Using a parametric finite element analysis, the run-out region was optimised for stable debonding crack growth. The modified run-out, as well as a baseline configuration, were manufactured and tested. Damage initiation and propagation was investigated in detail using state-of-the-art monitoring equipment including Acoustic Emission and Digital Image Correlation. As expected, the baseline configuration failed catastrophically. The modified run-out showed improved crack-growth stability, but subsequent delamination failure in the stiffener promptly led to catastrophic failure.

Compression and post-buckling damage growth and collapse analysis of flat composite stiffened panels

Experimental and numerical investigations were conducted into the damage growth and collapse behaviour of composite blade-stiffened structures. Four panel types were tested, consisting of two secondary-bonded skin–stiffener designs in both undamaged and pre-damaged configurations. The pre-damaged configurations were manufactured by replacing the skin–stiffener adhesive with a centrally located, full-width Teflon strip. All panels were loaded in compression to collapse, which was characterised by complex post-buckling deformation patterns and ply damage, particularly in the stiffener. For the pre-damaged panels, significant crack growth was seen in the skin–stiffener interface prior to collapse, which caused a reduction in load-carrying capacity. In the numerical analysis of the undamaged panels, collapse was predicted using a ply failure degradation model, and a global–local approach that monitored a strength-based criterion in the skin–stiffener interface. The pre-damaged models were analysed with ply degradation and a method for capturing interlaminar crack growth based on multi-point constraints controlled using the Virtual Crack Closure Technique. The numerical approach gave close correlation with experimental results, and allowed for an in-depth analysis of the damage growth and failure mechanisms contributing to panel collapse. The successful prediction of collapse under the combination of deep post-buckling deformations and several composite damage mechanisms has application for the next generation of composite aircraft designs.

Numerical Analysis of Stiffener Runout Sections

The recent trend of incorporating more composite material in primary aircraft structures has highlighted the vulnerability of stiffened aerostructures to through-thickness stresses, which may lead to delamination and debonding at the skin–stiffener interface, leading to collapse. Stiffener runout regions are particularly susceptible to this problem and cannot be avoided due to the necessity to terminate stiffeners at rib intersections or at cutouts, interrupting the stiffener load path. In this paper, experimental tests relating to two different stiffener runout specimens are presented and the failure modes of both specimens are discussed in detail. A thinner-skinned specimen showed sudden and unstable crack propagation, while a thicker-skinned specimen showed initially unstable but subsequent stable crack growth. Detailed finite element models of the two specimens are developed, and it is shown how such models can explain and predict the behaviour and failure mode of stiffener runouts. The models contain continuum shell elements to model the skin and stiffener, while cohesive elements using a traction-separation law are placed at the skin–stiffener interface to effectively model the debonding which promotes structural failure.

A new approach for the large deflection finite element analysis of isotropic and composite plates with arbitrary orientated stiffeners

A new approach for the large deflection analysis of isotropic and composite arbitrary orientated stiffened plates is presented. Non-linear equilibrium equations are derived using the principle of virtual work applied to a continuum with a total Lagrangian description of motion. Eight node isoparametric plate elements are combined with three node beam elements, using the concept of equal displacements at the plate–stiffener interface, to represent the stiffened plate. The stiffness of the beam element is computed first irrespective of its position within the plate element and then transferred to the plate nodes depending on its orientation and position within the plate element. The Newton–Raphson incremental-iterative solution technique is used to obtain the non-linear response path. Results obtained by the present approach are compared with those available in the open literature to demonstrate the validity and efficiency of the proposed approach and good agreement is found in all the investigated cases.

The Behavior of Compressively Loaded Stiffener Runout Specimens – Part I: Experiments

The development of the next generation of civil and military transport aircraft will inevitably see an increased use of advanced carbon fibre composite material in the primary structure if performance targets are to be met. One concern in this development is the vulnerability of co-cured and co-bonded stiffened structures to through-thickness stresses at the skin–stiffener interfaces, particularly in stiffener runout regions. These regions are a consequence of the requirement to terminate stiffeners at cutouts, rib intersections, or other structural features which interrupt the stiffener load path. This work presents the results of an experimental programme investigating the failure of thick-sectioned stiffener runout specimens loaded in uniaxial compression. For all tests, failure initiated at the edge of the runout and propagated across the skin–stiffener interface. It was found that the failure load of each specimen was greatly influenced by intentional changes in the geometric features of these specimens. High frictional forces at the edge of the runout were also deduced from a fractographic analysis, indicating a predominantly Mode II initial failure mode.

The Behavior of Compressively Loaded Stiffener Runout Specimens – Part II: Finite Element Analysis

The termination of stiffeners in composite aircraft structures give rise to regions of high interlaminar shear and peel stresses as the load in the stiffener is diffused into the skin. This is of particular concern in co-cured composite stiffened structures where there is a relatively low resistance to through-thickness stress components at the skin–stiffener interface. In Part I, experimental results of tested specimens highlighted the influence of local design parameters on their structural response. Indeed some of the observed behavior was unexpected. There is a need to be able to analyse a range of changes in geometry rapidly to allow the analysis to form an integral part of the structural design process. This work presents the development of a finite element methodology for modelling the failure process of these critical regions. An efficient thick shell element formulation is presented and this element is used in conjuction with the Virtual Crack Closure Technique (VCCT) to predict the crack growth characteristics of the modelled specimens. Three specimens were modelled and the qualitative aspects of crack growth were captured successfully. The shortcomings in the quantitative correlation between the predicted and observed failure loads are discussed. There was evidence to suggest that high through-thickness compressive stresses enhanced the fracture toughness in these critical regions.

Linear static analysis of composite hat-stiffened laminated shells using finite elements

Analysis of composite stiffened panels used in aerospace, ship and other engineering structures by the method of finite elements has been presented. The formulation presented is based on the concept of equal displacements at the shell–stiffener interface. An eight-noded isoparametric shell element is used in association with three-noded curved beam element for the formulation of the stiffened panel element. First-order shear deformation theory is used in the present study. The stiffness matrix of the stiffener is computed first irrespective of its position within the plate/shell element. The stiffness matrix so computed for the stiffener is then transferred to that of the plate/shell element nodes depending on its position and orientation within the element before assembling the element stiffness matrix. A generalised formulation for the hat-shaped stiffener is presented. The investigation is restricted to linear static analysis of composite stiffened panels and parametric studies have been made to study the various aspects of laminated composite shell with open and closed shaped stiffeners.

FINITE ELEMENT TRANSIENT DYNAMIC ANALYSIS OF LAMINATED STIFFENED SHELLS

The present work describes the transient dynamic response of unstiffened/stiffened composite plates/shells using finite element method. Composite panels find wide applications in aerospace, marine and other engineering because of its high strength to weight ratios. These structures are often subjected to air-blast loading, underwater shock etc., which requires a thorough dynamic response analysis under such loading. A modified approach of shell and stiffener modelling is adopted here using an eight-noded isoparametric quadratic element for the shell and a three-noded curved stiffener element for the stiffeners on the concept of equal displacements at the shell–stiffener interface. The present formulation obviates the need for imposing the mesh line along the stiffeners; rather it accommodates the stiffeners elegantly anywhere placed arbitrarily inside the element with computational efficiency. Newmarks method for direct time integration has been adopted for the solution of the governing equation for undamped motion. The transient dynamic response of stiffened and unstiffened structures subjected to various kinds of time variant loading has been studied and the results are compared with the published ones.

Failure of thick-skinned stiffener runout sections loaded in uniaxial compression

Recent efforts towards the development of the next generation of large civil and military transport aircraft within the European community have provided new impetus for investigating the potential use of composite material in the primary structure. One concern in this development is the vulnerability of co-cured stiffened structures to through-thickness stresses at the skin–stiffener interfaces particularly in stiffener runout regions. These regions are an inevitable consequence of the requirement to terminate stiffeners at cutouts, rib intersections or other structural features which interrupt the stiffener load path. In this respect, thicker-skinned components are more vulnerable than thin-skinned ones. This work presents an experimental and numerical study of the failure of thick-sectioned stiffener runout specimens loaded in uniaxial compression. The experiments revealed that failure was initiated at the edge of the runout and propagated across the skin–stiffener interface. High frictional forces at the edge of the runout were also deduced from a fractographic analysis and it is postulated that these forces may enhance the fracture toughness of the specimens. Finite element analysis using an efficient thick-shell element and the Virtual Crack Closure Technique was able to qualitatively predict the crack growth characteristics for each specimen.

Efficient stress solutions at skin stiffener interfaces of compositestiffened panels

Solutions for the stresses at the skin stiffener interface of composite stiffened panels are developed. Initially, the governing partial differential equations for the stresses are set up and the solutions are found. These solutions are then used in conjunction with an energy minimization approach to determine unknown constants in the analyticaly determined stress expressions. Results of this method are compared to boundary element solutions and are shown to be in good agreement. The present method has the advantage of being in closed form and very efficient. It can therefore be used to screen design candidates and give an accurate idea of the stress field near dropped plies without resorting to time-consumin g finite element or other solutions. Nomenclature Ai,Cn,Dn = constants in stress expressions EQ, E = constants describing far-field stress dependence f(x),Fi(x),F2(x) = functions describing in-plane stress dependence g(z) = function describing out-of-plane stress dependence