Thin Composite Panels Research Articles

Using artificial neural networks and strain gauges for the determination of static loads on a thin square fully-constrained composite marine panel subjected to a large central displacement

Current methods of estimating the behaviour of marine composite structures under pressure due to slamming as a result of high waves are based on trial and error or oversimplification. Normally under these conditions the non-linearities of these structures are often neglected and, in order to compensate, an overestimated safety factor is employed. These conservative approaches can result in heavier and overdesigned structures. In this paper, a new semi-empirical method is proposed that overcomes some of these problems. This work involved the use of an artificial neural network (ANN) combined with strain gauge data to enable real-time in-service load monitoring of large marine structural panels. Such a tool has other important applications, such as monitoring slamming or other transient hydrostatic loads that can ultimately affect fatigue life. To develop this system, a glass fibre-reinforced polymer (GFRP) composite panel was used due to its potential for providing a non-linear response to pressure or slamming loads. It was found that the ANN was able to predict normal loads applied at different locations on the panel accurately. This method is also capable of predicting loads on the marine structure in real time.

Cantilever Beam Static and Dynamic Response Comparison with Mid-Point Bending for Thin MDF Composite Panels

A new cantilever beam apparatus has been developed to measure static and vibrational properties of small and thin samples of wood or composite panels. The apparatus applies a known displacement to a cantilever beam, measures its static load, then releases it into its natural first mode of transverse vibration. Free vibrational tip displacements as a function of time were recorded. This paper compares the test results from the cantilever beam static bending and vibration with standard mid-point simply supported bending samples. Medium density fiberboard panels were obtained from four different commercial sources. Comparisons were made using a set of fiberboard panels with thicknesses of 8.1, 4.5, 3.7, and 2.6 mm and nominal densities of 700, 770, 780, and 830 kg/m3, respectively. Cantilever beam static modulus and dynamic modulus of elasticity linearly correlated well but were consistently higher than standard mid-point bending modulus of elasticity having linear correlations of 1.12:1 and 1.26:1, respectively. The higher strain rates of both the static and vibrating cantilever beam could be the primary reason for the slightly higher dynamic modulus values. The log decrement of the displacement was also used to calculate the damping ratio for the cantilever beam. As expected, damping ratio had a slightly decreasing slope as density increased. This paper discusses the new apparatus and initial results.

Active control of geometrically nonlinear transient vibrations of laminated composite cylindrical panels using piezoelectric fiber reinforced composite

This paper addresses the analysis of active constrained layer damping (ACLD) of geometrically nonlinear transient vibrations of laminated thin composite cylindrical panels using piezoelectric-fiber- reinforced composite (PFRC) materials. The constraining layer of the ACLD treatment is considered to be made of the PFRC materials. The Golla–Hughes–McTavish (GHM) method has been implemented to model the constrained viscoelastic layer of the ACLD treatment in time domain. The Von Karman type-nonlinear strain-displacement relations and a simple first-order shear deformation theory are used for deriving this electromechanical coupled problem. A three-dimensional finite element (FE) model of smart composite panels integrated with the patches of such ACLD treatment has been developed to demonstrate the performance of these patches on enhancing the damping characteristics of thin symmetric and antisymmetric laminated cylindrical panels in controlling the geometrically nonlinear transient vibrations. The numerical results indicate that the ACLD patches significantly improve the damping characteristics of both symmetric and antisymmetric panels for suppressing the geometrically nonlinear transient vibrations of the panels. The effect of the shallowness angle of the panels on the control authority of the patches has also been investigated.

Response of preloaded thin composite panels subjected to underwater explosive loading

The effect of preloading on thin composite plates subjected to underwater explosive loading has been studied through computational simulations. In this study the effects of three types of in plane preloading are considered: (1) directly applied compression, (2) indirectly applied compression, and (3) directly applied tension. The effects of the preloading conditions are assessed using the plate center point deflection, material damage, and delamination evolution. The results show that for thin composite plates subjected to underwater shock loading conditions there is a minimal effect of preload on the response of the plates or the amount of damage and delamination sustained.

Semi-continuous approach for the modeling of thin woven composite panels applied to oblique impacts on helicopter blades

In aeronautics, passenger safety and reliability of structures are essential aspects. In the specific case of helicopters, blades are subjected to impact loadings. Modeling these phenomena continue to be difficult and experimental tests often replace the prediction. The following work will focus on the experimental and numerical study of an oblique impact on the skin of the blade. It is equivalent in a first approach to an impact on a sandwich panel made up of a foam core and a thin woven composite skin. The objectives are to identify the mechanisms of damage in the skin for this kind of loading and to develop a representative modeling of the chronology of damage adapted to the modeling of the complete structure. Thus, a semi-continuous F.E. explicit modeling has been developed. It relies on the development of a specific damageable element at the bundles scale. Satisfactory numerical results are obtained. They allow the identification of the damage mechanism of the woven skin.

The use of a controlled multiple quasi-static indentation test to characterise through-thickness penetration of composite panels

During the quasi-static indentation of thin composite panels, well-defined flaps (sometimes called “petals”) can develop on the exit face as a consequence of through-thickness penetration of the panel; such flaps can also be seen in impact tests. The flaps develop as four triangles, with the apex of each triangle at the point of impact. In this work, thin panels of CFRP with a 0/90 configuration have been subjected to quasi-static indentation tests and the development of the flaps has been monitored. The results show that the dependence of the flap compliance is proportional to the square of the flap length, which is in agreement with theoretical predictions. The determination of the compliance/crack-length relationship enables a toughness value for fracture of the composite panel to be derived that is directly relevant to through-thickness penetration of the panel.

Dynamic fiber Bragg gratings based health monitoring system of composite aerospace structures

The main purpose of the current work is to develop a new system for structural health monitoring of composite aerospace structures based on real-time dynamic measurements, in order to identify the structural state condition. Long-gauge Fibre Bragg Grating (FBG) optical sensors were used for monitoring the dynamic response of the composite structure. The algorithm that was developed for structural damage detection utilizes the collected dynamic response data, analyzes them in various ways and through an artificial neural network identifies the damage state and its location. Damage was simulated by slightly varying locally the mass of the structure (by adding a known mass) at different zones of the structure. Lumped masses in different locations upon the structure alter the eigen-frequencies in a way similar to actual damage. The structural dynamic behaviour has been numerically simulated and experimentally verified by means of modal testing on two different composite aerospace structures. Advanced digital signal processing techniques, e.g. the wavelet transform (WT), were used for the analysis of the dynamic response for feature extraction. WT's capability of separating the different frequency components in the time domain without loosing frequency information makes it a versatile tool for demanding signal processing applications. The use of WT is also suggested by the no-stationary nature of dynamic response signals and the opportunity of evaluating the temporal evolution of their frequency contents. Feature extraction is the first step of the procedure. The extracted features are effective indices of damage size and location. The classification step comprises of a feed-forward back propagation network, whose output determines the simulated damage location. Finally, dedicated training and validation activities were carried out by means of numerical simulations and experimental procedures. Experimental validation was performed initially on a flat stiffened panel, representing a section of a typical aeronautical structure, manufactured and tested in the lab and, as a second step, on a scaled up space oriented structure, which is a composite honeycomb plate, used as a deployment base for antenna arrays. An integrated FBG sensor network, based on the advantage of multiplexing, was mounted on both structures and different excitation positions and boundary conditions were used. The analysis of operational dynamic responses was employed to identify both the damage and its position. The system that was designed and tested initially on the thin composite panel, was successfully validated on the larger honeycomb structure. Numerical simulation of both structures was used as a support tool at all the steps of the work providing among others the location of the optical sensors used. The proposed work will be the base for the whole system qualification and validation on an antenna reflector in future work.

Research Situation of the Unilateral Contact Buckling of Thin Wall Steel-Concrete Composite Panels

Thin wall steel-concrete composite wall panels can be used as bearing member and also as maintenance structural plates, which can satisfy functional requirements of building including bearing capacity, heat insulation and preservation. The problem relating unilateral contact buckling of thin wall steel composite panels has received attention from many construction engineers. The investigation progress is analyzed in general and the investigation questions are illustrated. The research approach and technical routes of delaminated critical load are also presented.

Two-stage design process of a frame-panel land vehicle structure employing topology and cross section optimization

A new structural design technology has been recently developed to build a new type of land vehicle in military use. While thick panels are only employed in conventional land vehicle structures, solid frames combined with relatively thin composite panels are employed for the new type of vehicle. The structural integrity of the new vehicle structure type is mainly guaranteed by the solid frames while composite panels are used to protect passengers and equipments of the vehicle. To design such a frame-panel structure, frame design needs to be done first. In this paper, a two-stage design process is proposed employing topology and cross section optimization methods. Overall frame arrangement of the new vehicle structure is obtained by the topology optimization in the first design stage and the detailed dimensions of the frames are obtained by the cross section optimization in the second design stage.

Active damping of laminated thin cylindrical composite panels using vertically/obliquely reinforced 1–3 piezoelectric composites

This paper deals with the analysis of active constrained layer damping (ACLD) of laminated thin composite panels using vertically and obliquely reinforced 1–3 piezoelectric composite materials as the material of the constraining layer of the ACLD treatment. A finite element model has been developed for analyzing the ACLD of laminated antisymmetric cross-ply and antisymmetric angle-ply thin composite panels integrated with the patches of such ACLD treatment. Both in-plane and out-of-plane actuations of the constraining layer of the ACLD treatment have been utilized for deriving the finite element model. The analysis revealed that the vertical actuation dominates over the in-plane actuation. Particular emphasis has been placed on investigating the performance of the patches when the orientation angle of the piezoelectric fibers of the constraining layer is varied in the two mutually orthogonal vertical planes. The analysis revealed that the vertically reinforced 1–3 piezoelectric composites which are in general being used for the distributed sensors can be potentially used for the distributed actuators of high performance light-weight smart thin composite panels.

Manufacturing and Testing of High Performance Sheet Moulding Compound

The paper describes the manufacture of thin composite panels using high performance sheet moulding compound (SMC). Topics discussed within the paper include characterisation of curing and flow behaviour of SMC material, tooling design concept and determination of suitable processing conditions for compression moulding. A Full scale “Burst test” was carried out to evaluate the mechanical performance of SMC panels. The overall performance of the SMC panels was satisfactory with all panels failed beyond the specification value. The main failure mode was a through-thickness cracking. In addition, a geometrical non-linear numerical analysis was also carried out to investigate the stress distribution and deflection behaviour of SMC panel during “Burst testing”.

Investigation for the Reduction of In-plane Compressive Strength in Preconditioned Thin Composite Panels

This paper investigates the effects of precondition characteristics on the residual compressive strength (RCS) of 16-ply carbon/epoxy panels through the establishment of their compressive and buckling response characteristics. The preconditions of varying sizes include impact damage and embedded artificial delaminations of both circular and elliptical shapes. A sequence of prebuckling, local and global buckling, and postbuckling is found in both the longitudinal and transverse directions. The possibility of delamination propagation is examined using the response characteristics on the basis of the sequences. A damage threshold is found at the 25% panel width or 455mm 2damage area. Both the global buckling loads and the RCSs suffer substantial degradation but in a different manner, depending on the threshold. For small damages, the global buckling loads are unaffected but the RCSs suffer substantial degradation, though there is little evidence for the delamination propagation. For large damages, while the global buckling loads are reduced moderately with an increase of the damages due likely to the unstable global buckling-to-postbuckling transition, the RCSs do not suffer further reduction, though evidence is found for the delamination propagation. These panels fail soon after the unstable global bucklingtopostbuckling transition. Panels with multiple delaminations suffer a greater reduction in RCS than the impact-damaged panels. A combination of the delamination size and number is shown to have the most dominant effect on the RCSs. It is demonstrated that the present method of embedding artificial delaminations proves to be very useful for studying the RCS of impact-damaged panels via the establishment of response characteristics and their links to the effects of the preconditions on them.

Modeling Hail Ice Impacts and Predicting Impact Damage Initiation in Composite Structures

Hail ice impacts are a realistic threat to exposed composite structures such as aircraft fuselage and wing skins, leading-edgeandcontrolsurfaces,enginenacelles,andfanblades.Toshedsomelightontothislittle-understood (and unavoidable ) threat, experimental, numerical, and analytical investigations have been conducted. Experiments in whichcarbon/epoxy compositepanelswereimpacted by icespheresathigh velocity (30‐200m/s)wereconducted to measure 1)the impact energy at which damage initiates, 2 ) the multiple failure modes exhibited by thin composite panels over a range of impact velocity, and 3 ) the elastic response of a composite panel resulting from impact. Subsequent numerical analyses of the last case were performed and were validated through correlation with experimental data. Insights gained from the numerical analyses were used to compose a mechanics-based formula that predicts the initiation of damage formation. This formula, derived using a global energy balance, provides a cost-effective (i.e., lower number of tests needed ) means by which the impact damage resistance of composite structures and material types can be established.

Instability of thin composite isogrid panel during autoclave cure cycle

This paper describes the analyses performed to determine why nominally flat composite isogrid panels designed for the Clementine I satellite were warping during the autoclave cure cycle. Analytical models indicated that the largest factor was the mismatch in the coefficient of thermal expansion of different parts of the tooling. The next most important factor was the coefficient of thermal expansion mismatch between the skin and rib materials of the composite isogrid itself. A new tooling design eliminated the first factor, producing a rib-only panel with no warping. An analytical model was developed, which predicted that warping caused by the rib/skin mismatch would be 3.31 mm over a length of 147.32 cm. A panel with the skin was manufactured and found to have a warp of 3.80 mm. Further improvements in producing flat panels will depend on finding appropriate materials for the rib and skin.

Genetic Algorithm To Optimal Lay-up In ThinComposite Panels

The orientation of fibre direction in layers and number of layers of composites play a major role in determining the strength and stiffness. Thus the basic design problem is to determine the optimum stacking sequence of the composite laminate. Many methods are available at present for the design optimisation of structural systems. However, these methods are based on mathematical programming techniques involving gradient search and direct search. These methods assume the variables are continuous. In this paper a different search and optimisation algorithm, known as the Genetic algorithm of Goldberg [1], has been successfully applied to obtain optimal fibre orientation of multi-layer plates, which considers the angle of fibre orientation as a discrete variable. The principle of GA is applied to obtain optimal layers and the orientation of fibres for stiffened square plates for both the symmetrical and antisymmetrical orientation of fibres subjected to uniaxial, biaxial and shear loading. Plates composing two to nine layers were analysed for their critical buckling load.

A new model for the FDTD analysis of the shielding performances of thin composite structures

A new model is proposed for the transient analysis of the electromagnetic field penetration through air-embedded conductive structures realized by thin multilayered composite panels. A magnetic field controlled formulation is developed in the frequency-domain to express the tangential components of the electric field on the external faces of the composite slab as a function of the tangential components of the magnetic field by means of the surface and transfer impedances of the thin panel coated on a perfect magnetic medium. The corresponding time-domain model is obtained by applying the inverse Fourier transform to the field quantities; an efficient piecewise linear convolution procedure is developed for the numerical calculation of the resulting convolution integrals. The model is implemented in one-dimensional (1-D) and two-dimensional (2-D) FDTD codes and applied to the analysis of different shielding configurations, both in the frequency and in the time domain.

Interlaminar dynamic crack propagation

Effective computational procedures are established and used to characterize dynamic crack propagation along material interfaces between isotropic and orthotropic materials. We first simulate a dynamic fracture experiment in which crack propagation occurs along an interface between PMMA and steel in a dynamically loaded bend specimen. In the analysis, the dynamic energy release rate and mixed-mode stress intensity factors are extracted from finite element field solutions using suitably formulated conservation integrals. These variables form the basis for defining an interface fracture criterion under dynamic conditions. In order to propagate the crack according to such a criterion, an iterative procedure is utilized to determine the correct crack tip velocity history. An important computed result, which is consistent with experimental observation, is that the energy release rate decreases as the crack propagation accelerates. The physical interpretation of this result is that less energy is absorbed by the moving crack as its velocity increases. In the second part of our study, the computational procedures are modified for the dynamic fracture analysis of a thin composite panel consisting of differently oriented orthotropic laminae. Here we investigate the interplay between delamination and buckling of the more complex structure. It is assumed that the panel contains an initial finite interlaminar crack and is subjected to a uniaxial compressive load. Without any crack extension, two buckling modes are observed under quasistatic conditions. One is characterized by an overall panel buckling and the other is dominated by a local ligament buckling near the crack. Coupling of the two modes produces not only a lower critical buckling load but also an unstable post-buckling behavior. Therefore, an embedded delamination is seen to create imperfection sensitivity, leading to limit load type behavior. Such a condition is potentially troublesome for compression loaded thin composite panels. In addition, the large compressive load may also trigger a dynamic propagation of the embedded delamination. We have simulated this unstable dynamic crack growth using the iterative method and a simplified delamination criterion. In the simulation, the energy release rate, the mixed-mode stress intensity factors and the resulting crack tip speed are obtained. These results predict that an existing embedded delamination can weaken and significantly alter the post-buckling behavior of composite panel.

A method for determination of elastic constants of a unidirectional lamina from ultrasonic bulk velocity measurements on [0/90] cross-ply composites

In this paper we introduce an inversion method to determine elastic moduli of a unidirectional lamina from ultrasonic time-delay measurements on [0/90] cross-ply composites. A lamina is a single ply (layer) of unidirectional fibers in a multiply composite plate. This method is based on a geometrical acoustic model relating the lamina properties with time-delay data for a [0/90] cross-ply composite. Interference due to multiple reflections within the laminate is assumed to be negligible. It is shown via numerical simulation that the reconstructed lamina moduli are highly accurate for composites with transversely isotropic laminae even for data containing large random scatter. For composites with orthotropic laminae, the method gives a very close approximation to the true lamina properties with reasonable accuracy. Experimental time-delay measurements in [0/90]2s cross-ply metal matrix composites (SiC/Ti–15V–3Al–3Cr–3Sn) are made in different incident planes. To avoid the effect of wave interaction within the laminate, we used the first arrival of the ultrasonic signal for time-delay measurements. From the [0/90]2s composite measurements, the elastic moduli of the lamina are determined and compared to those of a thin unidirectional composite panel measured independently by ultrasonic plate modes. The agreement between the two independent measurements is good. Utilization of measurements in the 45° plane in [0/90] composites for determination of lamina moduli is also discussed.

Flutter analysis of anisotropic panels with patched cracks

A finite element approach is developed to analyze the panel flutter problems of thin plate-like composite panels with patched cracks. The panel studied is a compound structure that comprises three layers including a thin cracked panel, an adhesive, and a thin patch. A 48-degree-of-f reedom (DOF) triangular crack patching element is derived for numerically analyzing the compound structure. The aerodynamic pressure acting on the panel is estimated using linearized piston theory with aerodynamic damping effect neglected. By a proper tailoring of the materials, very high flutter and/or divergence boundary could usually be obtained for the composite panels. The existence of a crack usually reduces the aeroelastic stability boundary; however, some exceptions were found for the composite panel within certain range of specific filament orientation. The deterioration in flutter/divergence performance due to a crack can, in general, be cured by means of patching, and anisotropic patching is more effective as compared to the isotropic patching provided that the tailoring of the structural parameters is done correctly. I. Introduction A RELIABLE and accurate flutter analysis is of primary importance in the flight vehicle design. The onset of flutter would quickly develop into either catastrophic structural failure or undesirable limit cycle type oscillation. The local panel flutter of the vehicle could be essential for a global flutter to occur. For instance, the initiation of cracks in wing structure is not unusual as they may be produced by either molding, fatigue, or impact damage.1'2 These defects could induce extremely high stresses around the cracktip regions and, thus, precipitate structural failure. When cracks are present, this local structural problem may have significant influence on the local flutter/divergence of the panel structure. The onset of the local aeroelastic instability may cause the failure of the cracked panel and cumulatively affect the global structure as well as the airload distribution and finally result in

Crystallization and processing behaviour of the polyamide J-1 polymer and composites reinforced with Kevlar fibre

The influence of thermal processing on the crystalline behaviour of J-1 polymer, a PACM-based polyamide homopolymer used as a matrix for advanced composites, was studied using differential scanning calorimetry (DSC), cross polarized microscopy, wide angle X-ray scattering (WAXS), and dynamic mechanical analysis (DMA). A double melting peak appears in DSC thermograms of neat polymer processed at cooling rates above 60°C min −1, accompanied by an additional crystallization peak above the glass transition. As the cooling rate is decreased, only a single melting peak occurs. This single peak corresponds to the lower temperature peak of two appearing in fast cooled J-1 polymer. Slower cooling rates produce a more ordered semicrystalline structure in the neat polymer, providing higher tensile strength and greater solvent resistance than the rapidly quenched material. WAXS patterns show the slow cooled J-1 polymer to have higher crystallinity than the quenched material. The equilibrium extent of crystallinity was determined to be about 20%, and the heat of fusion for 100% crystalline material is 150 J g −1. The density of the crystalline and non-crystalline phases are approximately 1.09 and 1.04 g cm −3 respectively. In order to study the behaviour of the composite, Kevlar fibre tows melt impregnated with J-1 polymer were wound into sheets and laminated into unidirectional composites in an autoclave process. Cracks due to internal stresses appeared in thin composite panels cooled at 10°C min −1. A thermal expansion differential was determined to be the primary cause of these stresses. By cooling the laminates slowly from the melt temperature, thin, crack-free composite panels were produced by permitting greater stress relaxation of the polymer and decreasing thermal gradients in the laminates.