Lay-up Schemes Research Articles

Optimum design of thin-walled composite beams for flexural–torsional buckling problem

The objective of this research is to present formulation and solution methodology for optimum design of thin-walled composite beams. The geometric parameters and the fiber orientation of beams are treated as design variables simultaneously. The objective function of optimization problem is to maximize the critical flexural–torsional buckling loads of axially loaded beams which are calculated by a displacement-based one-dimensional finite element model. The analysis of beam is based on the classical laminated beam theory and applied for arbitrary laminate stacking sequence configuration. A micro genetic algorithm (micro-GA) is employed as a tool for obtaining optimal solutions. It offers faster convergence to the optimal results with smaller number of populations than the conventional GA. Several types of lay-up schemes as well as different beam lengths and boundary conditions are investigated in optimization problems of I-section composite beams. Obtained numerical results show more sensitivity of geometric parameters on the critical flexural–torsional buckling loads than that of fiber angle.

Influence of Angle Ply Orientation on the Flexural Strength of Basalt and Carbon Fiber Reinforced Hybrid Composites

In this paper the influence of fiber orientation of basalt and carbon inter-ply fabrics on the flexural properties of hybrid composite laminates was experimentally investigated. Four types of basalt/carbon/epoxy inter-ply hybrid composite laminates with varying angle ply orientation of reinforced basalt fiber and fixed orientation of carbon fiber were fabricated using hand lay-up technique. Three point bending test was performed according to ASTM 7264. The fracture surface analysis was carried out by scanning electron microscope (SEM). The results obtained from the four laminates were compared. Lay-up pattern of (0B/+30B/-30B/0C)S exhibits the best properties in terms of flexural strength and flexural modulus. Scanning electron microscopy results on the fracture surface showed that the interfacial de-bonding between the fibers and epoxy resin is a dominant fracture mode for all fiber lay-up schemes.

Modelling of deformation behavior of polymeric woven fabric prepreg-based composite materials in flexure and short beam interlaminar shear testing

The mechanisms of deformation behavior of woven fabric prepreg-based carbon-fiber composites with different reinforcement lay-up schemes are studied. The FEM stress-strain analysis is carried out for specimens in flexure and short beam interlaminar shear testing environments. Both methods are based on a three-point bend model. Theoretical grounds underlying the mechanisms governing the deformation of laminates in bending tests are studied. Tests were performed on specimens having different thickness and with different ratios of the distance between supports to the specimen thickness. The effects of the monolayer properties specified in the model, such as the modulus of elasticity in the z-direction and the interlaminar shear modulus, on the effective characteristics of the material are investigated. The problems of correct experimental determination of the interlaminar shear modulus related to the large thickness of specimens, the difficulty of manufacturing and the peculiarities of measurement and identification of the obtained values of deformation in shear are considered. Flexural modules versus actual monolayer thickness and unsupported lengths versus specimen thickness and width are plotted. The necessity of registration of shear deformation in three-point bend testing for proper design of experiments and execution of calculations is shown. The accuracy and adequacy of the structural and bending characteristics are verified against the testing modes. The results confirm the relevance and validity of any given factor in modeling the stress-strain state of tested specimens.

An approach for reducing cost of carbon fibre-reinforced polymer structures using primary materials

The growing use of advanced composites in the aerospace industry has led to the development of new methods and equipment for manufacturing these new materials and structures. In this relatively new field of design, behaviour of production costs relative to different design variables is a challenging task. Carbon “pre-impregnated” (prepregs) composite fibres are costlier and are widely used as main load bearing members in the aircraft structures. As a result of this, 40% of a structural part cost is about its prepreg material. As prepreg material is expensive it is necessary to optimise its usage by avoiding wastages and implementing nesting program in automated prepreg cutting machines. The primary goal of this study is to find out an alternate solution to optimise material usage and to use primary dry fibres in between prepreg layers during lay-up and to evaluate the mechanical and ultrasonic properties. Different lay-up schemes are used to mix dry fibres in between prepreg layers to prepare test specimens as per DIN 29971 standards. This research reveals a method whereby companies can reduce their product cost by 18-20% . This concept has considerable scope for application in industry as the companies can make parts at minimum cost.

Post-buckling analysis of composite beams: A simple intuitive formulation

Post-buckling analysis of composite beams with axially immovable ends is investigated using an Intuitive formulation. Intuitive formulation uses two parameters namely critical buckling load and axial stretching force developed in the post-buckled domain of composite beam. Geometric nonlinearity of von-Karman type is taken into consideration which accounts for membrane stretching action of the beam. Axial stretching force developed in post-buckled domain of composite beam is evaluated by using an axial governing equation and is expressed either in terms of lateral displacement function as an integrated value, or as a function of both axial and lateral displacement functions at any discrete location of the beam. The available expressions of critical buckling load and derived expressions of axial stretching force developed in the beam are used for obtaining an approximate closed-form expressions for the post-buckling loads of various beam boundary conditions. Numerical accuracy of the proposed analytical closed-form expressions obtained from the intuitive formulation are compared to the available finite element solutions for symmetric and asymmetric lay-up schemes of laminated composite beam. Effect of central amplitude ratio and lay-up orientation on post-buckling load variation is briefly discussed for various beam boundary conditions considered in this study.

Post-buckling analysis of composite beams: Simple and accurate closed-form expressions

The present study proposes simple, elegant, and accurate closed-form expressions for predicting the post-buckling behavior of composite beams with axially immovable ends using the Rayleigh-Ritz (R-R) method, where geometric nonlinearity arises from mid-plane stretching of the beam. The displacement fields used for the analysis are coupled using the axial equilibrium equation of the composite beam. An exhaustive set of beam boundary conditions are studied, namely, hinged-hinged, clamped-clamped, clamped-hinged, hinged-guided, and clamped-guided beams in order to prove the efficacy of the present formulation. Post-buckling analysis results are presented for laminated composite beam with symmetric and asymmetric layup schemes. Post-buckling analysis results obtained using the R-R method are compared with the results obtained from the finite element (FE) method for all layup schemes considered and comparison shows an excellent match for all the beam boundary conditions considered and thereby proves simplicity, accuracy, and robustness of present closed-form analytical expressions for predicting the buckling and post-buckling behavior of composite beams with symmetric and asymmetric layup orientations.

Hybrid/mixed assumed stress element for anisotropic laminated elliptical and parabolic shells

A variety of elliptical/parabolic dome type structures are used in important aerospace and civil structural systems such as underwater vehicles, stadium covers, exhibition halls, auditoriums and museum halls. For the analysis of such structures, a shear deformable four-noded finite element based on a hybrid/mixed assumed stress is presented in this paper. The element called iHES (improved Hybrid and Enhanced Shell element) is developed assuming the most general arbitrary orthogonal coordinate system. The element is based on a first-order shear flexible formulation and essentially consists of a combination of drilling degrees of freedom with assumed stress and enhanced strain techniques. Using the element developed here, a detailed parametric study of anisotropic elliptical and parabolic shells of various configurations is carried out to investigate the effects of aspect and height ratios as well as layer lay-up schemes.

High-temperature tensile strength of near-stoichiometric SiC/SiC composites

In an attempt to characterize mechanical properties of near-stoichiometric SiC/SiC composites, tensile tests were conducted at room temperature in air and at elevated temperature under mild oxidizing gases atmosphere. SiC/SiC composites were fabricated by forced-flow chemical vapor infiltration method using two-dimensional fabrics of carbon coated near-stoichiometric Tyranno™SA fibers. Tensile tests were conducted on composites with two types of lay-up schemes using edge-loading small specimens. The effect of lay-up orientation on the mechanical properties and fracture behavior of composites were also examined. Tensile strength of composite was slightly decreased at 1573 K, while it retained approximately 80% of the strength at room temperature. Porosity dependence on elastic modulus was clearly exhibited.

Supersonic flutter of composite skin panels of repeated-sublaminate layup

In high-speed aerospace vehicles, supersonic flutter is a well-known phenomenon of dynamic instability to which external skin panels are prone. In theory, the instability stage is expressed by the ‘flutter critical parameter’ Q crit, which is a function of the stiffness-, and dynamic pressure parameters. For a composite skin panel, gcrit can be maximised by lay-up optimisation. Repeated-sublaminate lay-up schemes possess good potential for economical lay-up optimisation because the corresponding effort is limited to a family of sublaminates of few layers only. When Q crit, is obtained for all sublaminates of a family, and the sublaminates ranked accordingly, the resulting ranking reveals not only the optimum lay-up, but also the near-optimum lay-ups, which are useful design alternatives, and the inferior lay-ups which should be avoided. In this paper, we examine sublaminate-ranking characteristics for a composite panel prone to supersonic flutter. In particular, we consider a simple supported midplane-symmetrical rectangular panel of typical aspect ratio a and flow angle ψ, and for four-layered sublaminates, obtain the Q crit-based rankings for a wide range of the number of repeats, r. From the rankings, we find that an optimum lay-up can exist for which the outermost layer is oriented wide of, rather than along, the flow. Furthermore, for many lay-ups other than the optimum and the inferior, we see that as r increases, Q crit, undergoes significant change in the course of converging. To reconcile these findings, eigenvalue-coalescence characteristics are discussed in detail for specific cases.

Large deformation finite element analysis of laminated circular composite plates

A 48 d.o.f., four-noded quadrilateral laminated composite shell finite element is particularised to a sector finite element and is used for the large deformation analysis of circular composite laminated plates. The strain-displacement relationships for the sector element are obtained by reducing those of the quadrilateral shell finite element by substituting proper values for the geometric parameters. Subsequently, the linear and tangent stiffness matrices are formulated using conventional methods. The Newton-Raphson method is employed as the nonlinear solution technique. The computer code developed is validated by solving an isotropic case for which results are available in the literature. The method is then applied to solve problems of cylindrically orthotropic circular plates. Some of the results of cylindrically orthotropic case are compared with those available in the literature. Subsequently, application is made to the case of laminated composite circular plates having different lay-up schemes. The computer code can handle symmetric/unsymmetric lay-up schemes. The large displacement analysis is useful in estimating the damage in composite plates caused by low-velocity impact.

Shear-lag effect in CFRP I-beams under three-point bending

Shear-lag effect in CFRP (carbon-fiberreinforced plastic) I-beams is investigated by applying Reissner's method concerning box beams to CFRP I-beams. This modified beam theory (MBT) assumes that the flanges are orthotropic and includes the effect of the shearing rigidity of the web and the deformation of the flanges themselves. The anisotropy of the flanges causes the decrease of the flexural rigidity of I-beams. The decrease, however, is restrained by a preventive action on the shear-lag effect in the flanges with a decrease of the anisotropy ratio due to the selection of fiber orientation. In order to verify this MBT, I-beam structures were fabricated from carbon fiber and epoxy resin using three different lay-up schemes: unidirectional, quasi-isotropic, and a ±45-deg angle-ply. The stiffness and shear-lag effect of these I-beams were measured and compared with results from the composite beam theory and MBT. The theoretical load-deflection curves and the strain distributions agreed well with the corresponding experimental results.

Optimum Design of Composite Laminates against Delamination by Ranking

In this paper, we describe a method for the optimum design of fibre rein forced composite laminates against delamination. The software developed based on this method designs laminates so that they are not susceptible to delamination when subjected to a bi-cubic distribution of inplane and/or bending loads. The laminated plate which is made of repeated sublaminates is assumed to be thin and of symmetric construction. Various layup schemes are evaluated based on the classical plate theory and quadratic delamination criterion for the interlaminar stresses developed in the laminate under load ing conditions. The optimum layup sequence which exhibits the least propensity to delaminate and the number of sublaminates required are obtained Further, a ply-drop round-off scheme is adopted to arrive at an optimum laminate thickness. As an example, a family of 0/90/45/—45 bi-directional lamination schemes are examined for different types of loads and the gains in optimising the ply orientations in a sublaminate are demon strated.

Optimum Design of Composite Laminates for Strength by Ranking

In this paper we describe a method for the optimum design of fiber rein forced composite laminates for strength by ranking. The software developed based on this method is capable of designing laminates for strength; which are subjected to inplane and/or bending loads and optionally hygrothermal loads. Symmetric laminates only are considered which are assumed to be made of repeated sublaminate construction. Various layup schemes are evaluated based on the laminated plate theory and quadratic failure cri terion for the given mechanical and hygrothermal loads. The optimum layup sequence in the sublaminate and the number of such sublaminates required are obtained. Further, a ply-drop round-off scheme is adopted to arrive at an optimum laminate thickness. As an example, a family of 0/90/45/ -45 bi-directional lamination schemes are examined for dif ferent types of loads and the gains in optimising the ply orientations in a sublaminate are demonstrated.

Shear buckling of corrugated composite panels

Corrugated composite panels are used in aerospace vehicles primarily because of their ability to withstand higher loads before buckling compared to plain sheets. Shear buckling is one of the important modes of failure. In this paper are presented details of prediction of the shear buckling loads of composite panels having sinusoidal or hat type corrugation. The given panel is idealised as an homogeneous orthotropic plate whose equivalent properties are determined based on the given parameters of the panel. In the analysis Kirchhoff-Love assumptions are used. Two opposite edges of the panel in one direction are assumed to be simply supported and the panel is assumed to be very long in the other direction. Critical buckling loads are obtained using conventional orthotropic plate theory. A large class of 0°/90°/45°-45° lamination schemes leading to quadridirectional, tridirectional and bidirectional T300/N5208 corrugated panels is examined and ranked based on buckling loads. Results indicate that by proper choice of lamination scheme, very significant increases in buckling load compared to the quasi-isotropic case can be obtained. Having the corrugation plane normal to the shorter side is significantly better than having the corrugation plane normal to the longer side. Optimum layup schemes are found for 8, 6, 4 and 2 ply cases.

THE AUTOMATION OF THE LAY-UP AND CONSOLIDATION OF PEEK*GRAPHITE FIBER COMPOSITES*

ABSTRACT The processing of PEEK/graphite fiber composites is characterized by an inordinate amount of hand labor. This greatly increases the cost, time, and inaccuracy involved in production. Automation of the lay-up and consolidation steps are a prerequisite for the widespread application and acceptance of these, as well as all, high performance materials. Automated techniques for the lay-up and consolidation of PEEK/graphite fiber composites have been developed and are presented in this paper. Discrete part lay-up schemes utilize robotic manipulators and specially constructed tables and allow any lay-up configuration to be automatically produced at high production rates. The size of the discrete parts was limited to 4 by 4 meters by design, not by any inherent difficulties. This size can be increased by enlarging the lay-up table. The length of a continuous sheet was limited by the length of the PEEK/graphite fiber prepreg material currently available. Its width was also restricted to 4 meters by design...