Stiffened Panels Research Articles

On damage behavior and stability of composite T-shaped stiffened panels under compression after impact considering impact locations

Impact locations play a critical role in determining the residual performance of the composite stiffened panel, since the local damage behavior at different locations is not identical. The objective of this paper is to systematically investigate the influence of impact locations on damage behavior, stability and residual compression strength of T-shaped stiffened CFRP panels. Low-velocity impact tests are carried out by impacting at the skin center or the flange tip. The damage patterns and post-buckling behavior are monitored by ultrasonic C-scan and digital image correlation. An effective computational framework with explicit formulations considering complicated damage mechanisms is proposed to investigate the details of damage behavior. The skin center impact causes severe fiber breakage and matrix squeezing, while the flange tip impact induces complex damage patterns including ply splitting, fiber breakage, matrix cracks, as well as extensive skin–rib interfacial delamination. Compression after impact results indicate that impact damage at the skin center has little effect on the residual performance of the structure, and the compression failure load is only reduced by 2.02%, from 276.8 kN to 271.2 kN. However, impact damage at the flange tip weakens the supporting effect of the stiffener, leading to early buckling of the local skin. The propagation of the interfacial delamination at the damaged flange tip directly triggers the premature compression failure, resulting in an 18.79% decrease in the failure load, 224.8 kN. Numerical predictions correlate well with experimental data. The reported results may be useful in structural design and maintenance.

Damage and failure analysis of composite stiffened panels under low-velocity impact and compression after impact with damp-heat aging

Abstract The low-velocity impact and compression after the impact of the composite stiffened panels were carried out after damp-heat aging. The experimental results show that reinforcing the ribs can enhance the impact resistance of test pieces after damp-heat aging. After impacting, the specimens were tested in an axial compression. The results show that the ultimate bearing capacity of the specimen is also affected by different located positions of the impact and different aging times. Compared with the intact specimen, the ultimate load-bearing capacity was reduced to 16.83, 12.10, and 17.10% with the specimen aging for 0, 45, and 90 days, respectively, while the impact position located at the intersection of longitudinal and transverse bars has the greatest influence on the damp-heat aging of specimens.

Compressive buckling and post-buckling behaviors of J-type composite stiffened panel before and after impact load

The compressive buckling and post-buckling behaviors of J-type composite stiffened panels before and after impact load were investigated through theoretical, numerical and experimental methods. In this paper, the load-bearing characteristics, including the buckling strength, ultimate strength and failure process of the intact composite stiffened panels were theoretically predicted and tested. Next, the impact tests were carried out on both the center of the skin and the back of the rib of the J-type composite stiffened panel, and then the compression test was performed on the damaged stiffened panel. The residual strength and failure process of the damaged J-type stiffened panel were obtained. By means of numerical simulation, the effects of different degrees of impact damage at the center of the skin on the buckling and post-buckling behavior of stiffened panels were revealed. Good agreement among theoretical, numerical and experimental results has been achieved. Finally, the influence of the adhesive layer, lay-up method of both the stiffened rib and skins, and geometric dimensions of the J-type ribs on structural buckling and post-buckling behaviors was systematically studied. The outcome from this research provides reasonable suggestions for improving the buckling load, ultimate load and material utilization rate of stiffened panels.

Fast analysis of curvilinearly stiffened panels operating in the postbuckling field

A semi-analytical formulation is developed for the thermomechanical postbuckling analysis of innovative aerospace panels obtained by the combination of variable stiffness skins and curvilinear stringers. A formulation is proposed in conjunction with first-order shear deformation theory to model the skin and stiffeners. The nonlinear problem is formulated in the framework of the Ritz method and an iterative-incremental procedure is employed for the solution process. The results demonstrate the excellent level of accuracy guaranteed by the proposed method, making it an ideal candidate for gathering insight into the potential of innovative stiffened panels designed to operate in the postbuckling field.

Skin-stringer separation in post-buckling of butt-joint stiffened thermoplastic composite panels

Two aeronautical thermoplastic composite stiffened panels are analysed and tested to investigate the buckling behaviour, the skin-stringer separation and the final failure mode. The panels are made of fast crystallising polyetherketoneketone carbon composite, have three stringers with an angled cap on one side, and are joined to the skin by a short-fibre reinforced butt-joint. The panels contain an initial damage in the middle skin-stringer interface representing barely visible impact damage. Finite element analysis using the virtual crack closure technique are conducted before the test to predict the structural behaviour. During the tests, the deformation of the panels is measured by digital image correlation, the damage propagation is recorded by GoPro cameras and the final failure is captured by high speed cameras. The panels show an initial three half-wave buckling shape in each bay, with damage propagation starting shortly after buckling. A combination of relatively stable and unstable damage propagation is observed until final failure, when the middle stringer separates completely and the panels fail in an unstable manner. The test results are compared to the numerical prediction, which shows great agreement for both the buckling and failure behaviour.

Artificial neural network prediction of residual compressive strength of composite stiffened panels with open crack

Openings can be introduced into composite stiffened panels during assembly or use, which can cause degradation and cracking. Therefore, the effect of the initial damage, such as open cracks, on the residual bearing capacity of stiffened panels must be investigated. In this study a back-propagation neural network model was established to predict the ultimate compressive load of single-stiffened panels with open cracks. The crack position, length, width, and angle were used as the input parameters and the dataset was established in batches via the secondary development of ABAQUS in Python. Subsequently the learning effects of artificial neural networks (ANN) with different network structures were evaluated. The results show that an ANN model with a 5-15-30-15-10-5-1structure, activated by a tanh function, not only possesses excellent learning ability but also demonstrates excellent generalization; furthermore, it can accurately and efficiently predict the ultimate compressive load of single-stiffened panels with open cracks.

Damage evolution and failure mechanism of asymmetric composite laminates under low-velocity impact and compression after impact

The damage mechanisms of the balanced symmetrical laminates under low-velocity impact (LVI) and compression after impact (CAI) loading conditions are not fully applicable to composite stiffened panels. However, studies on LVI and CAI damage evolution and failure mechanism for asymmetric composite laminates are very rare, which poses challenges to the theoretical study, strength prediction, and application of composite structures, and increases the risk of sudden failure under load. In this study, the damage evolution and failure mechanisms of asymmetric composite laminates under LVI and CAI loading conditions were studied experimentally and numerically. Ultrasonic phased array C-scan technique was used to analyze damage state, damage size, and delamination damage induced by LVI under different impact energies in impact tests. Moreover, a three-dimensional digital image correlation (3D-DIC) technique was employed to monitor the evolution of full-field displacement and strain in the asymmetric composite laminates during the CAI tests. A three-dimensional damage model considering the interaction among the interlaminar delamination damage, intralaminar matrix damage, and intralaminar fiber damage was proposed, and an interface-based cohesive behavior embedded framework in ABAQUS/Explicit was used to define and capture the interlaminar damage. The damage initiation, evolution, and propagation behaviors of different damage modes were simulated to study the complex damage and failure mechanisms of the asymmetric laminates in the LVI and CAI processes. The relationships among the different impact energies and the impact damage modes, delamination morphologies, and compression damage propagation and failure modes were discussed.

Empirical formulations for estimation of the ultimate strength of the perforated stiffened panels under the combined lateral pressure and in-plane compression

The holes artificially cut for a certain purpose can greatly reduce the ultimate strength of the perforated stiffened panels. In this paper, the ABAQUS finite element analysis program is used to investigate the effect of the position and size of the circular holes, the slenderness ratio, the aspect ratio, and the loading ratio on the ultimate strength of the perforated stiffened panels. Periodic and symmetric boundary conditions are applied to the (1/2 + 1+1/2) span model to consider the continuity of the stiffened panel in modeling. The results show that the ultimate strength of the perforated stiffened panel is highly dependent on the location and size of the hole. An empirical formula to primarily predict the ultimate strength reduction rate of perforated stiffened panel is proposed, and its accuracy is evaluated.

Experimental investigation of shear fatigue behavior of composite stiffened panels with impact damage

Shear fatigue experiments were performed to investigate the fatigue behavior and failure mechanism of composite stiffened panels. Low-velocity impact was conducted on the stiffener flange from the smooth side. Barely visible impact damage was introduced and it decreased the buckling and failure loads of the stiffened panels by 13.5% and 14.4%, respectively. The delamination and residual strength did not exhibit noticeable changes when the fatigue load level was set at 40% of shear strength after impact (SSAI). However, when it was increased to 67% of SSAI, delamination grew steadily and stiffener damage caused the delamination growth rate to temporarily increase by approximately 5-fold. A clear correlation was observed between the delamination growth and sublaminate buckling. Delamination grew over the entire web of the stiffener, resulting in the ultimate failure of the stiffened panels.

Optimization of the curing cycle for the large-scale composite stiffened panel based on the response surface methodology and Fgoalattain algorithm

The autoclave curing process of the large-scale composite stiffened panel greatly affects the manufacturing quality of the component, especially the curing quality of stiffeners and the bonding quality between stiffeners and the panel. In this paper, a multi-objective optimization method is developed to optimize the curing cycle for the composite T-stiffened fuselage panel, which aims to minimize the temperature difference within the component to achieve better forming quality. The autoclave curing model considering thermal properties of composites is established and verified by the experiment first. The response surface models are then built up with the Finite Element Method (FEM) and the Box-Behnken Design (BBD). The fgoalattain algorithm is applied to solve the multi-objective optimization problem with nonlinear constraints. Results show that the maximum temperature difference of the two groups of monitoring points decreases by 22.91% and 21.42%, respectively. Besides, the total curing time is also reduced by 11 mins.

Fatigue driven delamination in composite structures: Definition and assessment of a novel fracture mechanics based computational tool

The development of numerical methodologies to investigate material damages is a crucial aspect of aerospace components design. The advent of composite materials has made this aspect more critical due to their anisotropic properties, which result in damage mechanisms difficult to control and predict. Despite the progress made in the development of computational approaches for composites failure assessment, fatigue damage remains an open issue. This work deals with the development of a robust and efficient computational tool to simulate the fatigue driven delamination in complex composite structures. The main objective is to exploit the peculiarities, in terms of mesh and load step size independency, of the well-established SMart–time XB procedure, based on the Virtual Crack Closure Technique approach, to develop a Paris Law-based module to mimic the delamination growth rate due to cyclic loads. The implemented fatigue block interacts with the already existing modules, thus simulating the fatigue driven delamination without mesh dependency and potentially accounting for the R-curve behaviour of the material. The procedure takes advantage of the solver of the commercial Finite Element code Ansys Mechanical but is fully implemented in the Ansys Parametric Design Language, resulting in an absolutely versatile and parametric procedure. The promoted numerical tool has been firstly validated at coupon level considering the literature available Single Leg Bending-based benchmark. Then, the post–buckling behaviour of a typical aeronautical stiffened panel, based on a literature experimental test, has been studied under non-constant cyclic load conditions. Comparison against literature numerical and experimental data have proved the efficiency of the proposed computational tool to simulate fatigue-driven delamination both at coupon level and in complex structures.This work deals with the development of a robust and efficient computational tool to simulate the fatigue driven delamination in complex composite structures. The main objective is to exploit mesh and load independency peculiarities of the well-established SMart–time XB procedure, to develop a Paris Law-based module to mimic the delamination growth due to cyclic loads. The procedure is fully implemented in the Ansys Parametric Design Language, resulting in a versatile and parametric procedure. The promoted tool has been firstly validated at coupon level. Then, the post–buckling behaviour of a stiffened panel, based on a literature experimental test, has been investigated.

Empirical formulae assessment of ultimate strength for perforated stiffened panels under longitudinal compression

The present paper studies the influence of cut-outs of local plates on the ultimate strength of stiffened panels based on the tests and numerical simulations. Four intact and perforated stiffened panels subjected to uniaxial compression loads were tested firstly. The initial imperfections of the specimens have been measured before testing, which were used in the FE (finite element) analysis. The tensile tests are used to determine the properties of material. The experimental results are used to verify the modelling technology of FE models. Based on the numerical simulations, the parametric studies with various configurations including the initial twisting distortion, boundary condition and the dimensions of opening are performed to investigate their influences on the collapse behaviours of the stiffened panels. It is found that the residual ultimate strength of perforated stiffened panel mainly depends on the transverse length of opening. Using numerical database, the empirical formulae of ultimate strength of perforated stiffened are developed.

Dynamic instability analysis of perforated stiffened laminated composite panels subjected to non-uniform in-plane edge load under hygrothermal condition

A robust and computationally efficient finite element (FE) formulation has been developed in an effort to understand the parametric instability behaviour of a perforated stiffened composite panel subjected to a pulsating non-uniform edge load under a hygrothermal environment. The panel is modelled using a 9-noded heterosis element to avoid shear locking, whereas the stringer uses a 3-noded isoparametric beam element with a displacement compatibility condition employed at the skin–stringer​ interface. Due to the non-uniform distribution of stress in the perforated panel as a result of exposure to environmental and mechanical loading conditions, a unique dynamic technique has been implemented to account for the instability performance by employing two types of boundary conditions. In the current investigation, an unstiffened panel is initially obtained based on the highest value of the dynamic instability opening (DIO) and the unstable excitation frequency of the panel for different lamina schemes, cutout shapes, and loading conditions. The obtained unstiffened panel is further investigated for various stiffener configurations to determine a highly unstable perforated stiffened panel based on the DIO value and instability zone frequency range. Detailed parametric studies are then carried out considering the cutout sizes, cutout orientation, panel’s aspect ratio, and stiffener depth to understand the parametric instability behaviour of the panel in a hygrothermal environment. It is found that moisture significantly influences the dynamic instability behaviour of the perforated stiffened panel compared to the temperature exposure by shifting the instability zone towards the lower frequency. The width of the instability region of the stiffened panel broadens abruptly as the magnitude of the applied static in-plane edge load enhances. Also, for certain combinations of the dynamic and static load factors, the lower bound of the instability region reduces to zero, resulting in fluttering. In general, it is observed that stiffer the panel, wider is the instability zone with a higher unstable excitation frequency range.

Probabilistic multi-scale design of 2D plain woven composites considering meso-scale uncertainties

This paper presents a novel multi-scale design strategy for 2D plain woven composites considering the manufacturing process. The design methodology integrates the meso-structural geometry of woven fabric composites to maximise the buckling load of composite stiffened panels at the macro-level. A numerical model to simulate the autoclave manufacturing process of woven composites is developed, in which a new geometric modelling method is proposed to construct a more realistic RVE model for 2D plain woven composites accounting for the compaction effects. Deterministic and robust design optimisations are conducted to find solutions with both high performances and high robustness to uncertainties. In particular, this study considers the morphological uncertainties of the meso-scale fabric structure, i.e. yarns’ waviness and porosity incurred during manufacturing and its influence on the structural integrity of the parts. The optimisation results obtained from the deterministic and robust design problems are presented. A parametric analysis is conducted to characterise the sensitivity of the fabric properties to the geometry of yarns and related uncertainties. It is shown that the space between yarns has a significant negative impact on the structural buckling load and negative impact on the cure-induced residual stresses.

Ultimate strength assessment of stiffened panels using Equivalent Single Layer approach under combined in-plane compression and shear

Equivalent Single Layer (ESL) approach is extended to model ultimate strength of stiffened panels under a combination of in-plane compression and shear. Ultimate strength under this combined loading depends on the loading path and could be lower than for only uni-axial compression since shear load can produce axial forces. Thus, to account for this effect, the recent ESL model is extended to include A13 stiffness component as non-zero value. Procedure to accurately obtain A13 is presented. In ESL approach, a stiffened panel is transformed into a two-dimensional (2D) single layer with the same stiffness obtained from unit cell simulations. To obtain non-linear stiffness matrix of ESL, elastic–plastic material properties and initial imperfection were applied to the unit cell. ESL responses were validated by comparing numerical and experimental results from the literature. Several stiffened panel configurations were analyzed to obtain different collapse modes. Combined loads were applied for shear to compression ratio of 0, 1, and 2. Lateral pressure loading was also considered in the simulations. Analyses were carried out based on the load sequences consisting of: 1) compression and shear loaded simultaneously and 2) shear applied first, followed by compression. The results show that the modified ESL can well capture the effect of shear load on ultimate strength in comparison to a detailed 3D FEM model of stiffened panels. The accuracy of the ESL varies depending on the collapse mode of stiffened panels.

Study on compressive bearing capacity and efficiency of composite stiffened panels with different cross-sections

The post-buckling behavior and load bearing efficiency of composite panels with three kinds of stiffeners subject to the axial compression load were investigated. The progressive damage analysis models considering the failure of fiber/matrix, debonding of adhesive interface and stiffness degradation were established. The buckling load, failure load and failure process for three types of stiffened panels were derived. The numerical results coincide with tests, and the relative errors are all less than 6%, which indicates the FEM model given in this paper is reliable and practical. Further, the load bearing efficiency of these three kinds of stiffened panels with the same weight was discussed. Results reveal that the bearing capacity and efficiency of I-shape stiffened panel is the highest, and its bearing efficiency reaches 154.5N/g. The results of the paper provide a reference for engineering application of composite stiffened structure design.

A novel methodology to research the residual compressive mechanical properties of composite stiffened panel after low velocity impact based on quantification damage simulation

AbstractComposite stiffened panels have become one of the most usual components in the design process of the aircraft's structure due to their splendid mechanical properties in the last decades. The main purpose of this work is investigating and presenting a more precise method to forecast the residual compressive mechanical behavior of the composite stiffened panel. First of all, an axial compression after impact (CAI) test is performed to understand the failure mechanism of different type of L‐shaped stiffened carbon fiber reinforced polymer panels with barely visible impact damage, which is introduced by the low velocity impact test. Subsequently, a finite element model, which contains a quantification damage model, is established in ABAQUS software based on the CAI test results to predict the mechanical behavior of these composite stiffened panels during the CAI test process. Both of the experimental and predicted results demonstrate that the propagation of the delamination caused by LVI is going to exert a destructive influence on the total failure of the composite stiffened panel, and the prediction results of the residual compressive strength of these composite stiffened panels are well consistent to that measured in the CAI test.

Linear, nonlinear and post buckling analysis of a stiffened panel with cutouts

Stiffeners are typically used to strengthen structures that rely on plates for load bearing. Such structures are subjected to various kinds of in-plane and out-of-plane loading over its lifespan. Furthermore, cutouts in aerospace structures are considered for practical and design reasons. In the present paper, finite element method is used for the analysis of a flat rectangular stiffened plate. Static analysis is performed on three panels with different stiffener cross sections (I section, hat section and rectangular section), and the panel with the least deformation and stress is chosen for further buckling analysis. Linear, non-linear and post buckling analysis of the selected stiffened panel under in-plane compression is carried out and its load and end – shortening response is evaluated till collapse. Furthermore, the effect of rectangular and circular central cutouts on the buckling behaviour of the stiffened panel is investigated. By varying the dimensions of the cutouts, a parametric study on buckling and ultimate collapse strength is performed. The study takes into account an aluminium alloy with isotropic material properties and a carbon fibre reinforced polymer (CFRP) with orthotropic material properties.

An enhanced elastoplastic damage coupled model for compression characteristics analysis of continuous fiber-reinforced thermoplastic composite stiffened panel

In this work, the elastoplastic behaviors, damage characteristics, and debonding forms of continuous fiber-reinforced thermoplastic (CFRTP) composite stiffened panel under uniaxial compression load have been studied. The enhanced elastoplastic damage coupled model (EPDM) was established to model the mechanical properties of CFRTP composites which were based on generalized Hill yield criterion, the associated flow rules, the power hardening law, Chang-Chang failure criterion, and continuous damage mechanics (CDM). The forward Euler algorithm was implemented in Abaqus via user-defined subroutine VUMAT. A series of experiments including tensile, compression, and interlaminar shear tests of laminates were performed to validate the material model. Subsequently, the developed EPDM combined with cohesive zone model (CZM) was applied to simulate the nonlinear mechanical responses of CFRTP composite stiffened panel. Moreover, an axial compression test of CFRTP composite stiffened panel was also carried out, which confirmed that the simulation results were in agreement with the corresponding experimental results. Furthermore, based on the established finite element model, the buckling knockdown factor (KDF) and plastic region development of the CFRTP composite stiffened panel were also analyzed. Overall, the obtained results demonstrate the efficiency of the proposed EPDM for the compression characteristics analysis of the CFRTP composite structures.

Post-buckling failure analysis of composite stiffened panels considering the mode III fracture

Fully utilizing the post-buckling capacity of stiffened composite panels is important to reduce the weight of aeronautical composite structures. However, owing to the complex post-buckling failure modes and mechanisms, there are no effective analysis methods. This paper explores the impacts of the mode III fracture energy on the damage behavior of the stiffener–skin interface of composite stiffened panels under post-buckling conditions. To characterize the bonding interface damage, an interface damage initiation criterion considering the effects of through-thickness compression on the shear failure was used and combined with the Reeder damage extension criterion, which considers the mode III fracture. The mode III fracture toughness characteristics of the damage evolution of the stiffener–skin interface were obtained using the edge crack torsion (ECT) test method. Moreover, the compressive deformation response, buckling load, buckling modes, failure load, and failure modes of the panel were obtained by a compression test of I-shaped stiffened panels. Based on the aforementioned experimental data and numerical model, the post-buckling behavior of the I-shaped stiffened composite panels with an open cross-section stiffener was studied. It was found that the mode III fracture energy caused by the longitudinal shear stress τ31 played a major role in the stiffener–skin interface damage extension of the stiffened panels. Therefore, the mode III fracture behavior should not be ignored in the post-buckling analysis. The numerical analysis method developed can accurately predict the damage initiation and evolution processes of the composite panel interface. The method can be effectively used for post-buckling analysis of aeronautical composite panels.