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Priorities for research in advanced laminated fibrous composite materials are presented. Supporting evidence is presented in two bodies, including a general literature survey and a survey of aerospace composite hardware and service experience. Both surveys were undertaken during 1977-1979. Specific results and conclusions indicate that a significant portion of contemporary published research diverges from recommended priorites.
Priorities for standard compression testing research in advanced laminated fibrous composite materials are presented along with a state of the art survey (completed in 1979) including history and commentary on industrial test methods. Historically apparent research priorities and consequent (lack of) progress are supporting evidence for newly derived priorities.
This paper presents buckling and postbuckling results for plates loaded by inplane shear. The buckling results have been plotted to show the effects of thickness on the stress coefficient for aluminum plates. Results are given for various length-to-width ratios. Postbuckling results for thin plates with transverse shearing flexibility are compared to results from classical theory. The problems considered are the postbuckling response of plates in shear made of aluminum and of a + or - 45 deg graphite-epoxy laminate. Thus the materials are isotropic and orthotropic, respectively. The plates are considered to be long with side edges simply supported, with various inplane edge conditions, and the plates are subject to a constant shearing displacement along the side edges. Characteristic curves presenting the average shear stress resultant as a function of the applied displacement are given. These curves indicate that change in inplane edge conditions influence plate postbuckling stiffness and that transverse shearing is important in some cases.
The micromechanical model of Aboudi (1982, 1984, and 1986) is used to predict the time-dependent creep behavior of viscoelastic-matrix composites with unidirectional elastic-fiber reinforcement. The derivation of the model and its numerical implementation are outlined, and predictions for T300/934 CFRP at 72 and 250 F are compared with experimental measurements in extensive tables and graphs. Good agreement is demonstrated, and the applicability of the present approach to composite laminates and hybrid materials such as ARALL is indicated.
Diamines polymerized with dianhydrides or diacyl halides to produce fire-resistant polymers. Properties of polymers useful as matrices in laminated composite-material aircraft structural components that resist fire and exhibit desirable mechanical properties as high tensile strength. Diamines used as encapsulating or potting compounds when reacted with epoxy resins. Free phosphonic acids, byproducts of synthesis of diamines used as complexing agents to extract metals from aqueous solutions or as corrosion inhibitors for metals.
A recent variational mechanics analysis of microcracking damage in cross-ply laminates of the form /(S)/90n/s, where (S) is any orthotropic sublaminate much stiffer than /90n/, has been extended to account for the presence of delaminations emanating from the tips of microcracks in the /90 2n/T sublaminate. The new two-dimensional stress analysis is used to calculate the total strain energy, effective modulus, and longitudinal thermal expansion coefficient for a laminate having microcracks and delaminations. These results are used to calculate the energy release rate for the initiation and growth of a delamination induced by a matrix microcrack. At low crack densities, /(S)/90n/s laminates are expected to fail by microcracking and to show little or no delamination. At some critical crack density, which is a function of laminate structure and material properties, the energy release rate for delamination exceeds that for microcracking and delamination is predicted to dominate over microcracking. A quasi-three-dimensional model is used to predict the propagation of arbitrarily shaped delamination fronts. All predictions agree with experimental observations.
The application of composite materials to aircraft construction has provided the designer with increased flexibility. The orientation of plies can be tailored to provide additional aeroelastic performance unobtainable with an isotropic material. A tailored laminate is made up of plies of several orientations, usually 0 deg, 45 deg, -45 deg, and 90 deg. The direction of the 0 deg plies, does not need to be oriented with the leading edge, but can be varied to obtain a wide variety of structural properties. Also, the number of plies of each orientation varies from one zone to another on the planform. Thus, a thick laminate with mainly 0 deg plies may form the root zone, and a thinner laminate with mainly +45 deg plies may form the leading edge zone. Tailored laminates were designed using complicated optimization programs. Unfortunately, many tailored designs must be modified before they are manufactured. The modification adds weight and decreases performance. One type of modification is ply interleaving, an overlap of plies between zones on the laminate. These interleaves are added to ensure that zones with varying ply percentages can be connected without loss of strength. In this paper, the constraints needed to eliminate interleaves in the laminate optimization process will be described and implemented in a structural optimization problem. The method used has the potential to prevent changes to composite laminates late in the design cycle.
The high speed civil transport is a commercial aircraft that is expected to carry 300 passengers at Mach 2.4 over a range of more than 6000 nautical miles. With the existing commercial structural material technology (i.e., aluminum) the performance characteristics of the high speed civil transport would not be realized. Therefore there has been a concerted effort in the development of light weight materials capable of withstanding elevated temperatures for long duration. Thermoplastic composite materials are such candidate materials and the understanding of how these materials perform over the long term under harsh environments is essential to safe and effective design. The matrix dominated properties of thermoplastic composites are most affected by both time and temperature. There is currently an effort to perform short term testing to predict long term behavior of in-plane mechanical properties E22 (transverse modulus of elasticity) and G12 (shear modulus). Out-of-plane properties such as E33, G13, and G23 are inherently more difficult to characterize. This is especially true for the out-of-plane shear modulus G23 and hence there is no existing acceptable standard test method. Since G23 is the most matrix dominated property, it is essential that a test method be developed. A shear test methodology is developed to do just that. The test method, called the double notched specimen, along with the previously developed shear gage was tested at room temperature. Mechanical testing confirmed the attributes of the methodology. A finite element parametric study was conducted for specimen optimization. Moire interferometry, a high sensitivity laser optical method, was used for full-field analysis of the specimen. From this work, material parameters will be determined and thus enable the prediction of long term material behavior of laminates subjected to general loading states.
This Technical Memorandum examines the effects of heat and absorbed moisture on the open hole compression strength of carbon/epoxy laminates with the material and layup intended for the Ares I composite interstage. The knockdown due to temperature, amount of moisture absorbed, and the interaction between these two are examined. Results show that temperature is much more critical than the amount of moisture absorbed. The environmental knockdown factor was found to be low for this material and layup and thus obtaining a statistically significant number for this value needs to be weighed against a program s cost and schedule since basis values, damage tolerance, and safety factors all contribute much more to the overall knockdown factor.
MAT162, a laminated composite failure material model developed by The Material Sciences Corporation for the commercial finite element software LS-Dyna, is widely used within the aerospace industry to predict damage events under a range of dynamic conditions. The material model involves numerous inputs consisting of both physical material properties and numerical calibration parameters. Due to the large number material card inputs, often there is a lack of uniqueness to MAT162 material cards that limits the predictive capability to only the directly calibrated space. To expand this space, MAT162 requires a prudent and robust calibration process in which significant parameters are calibrated to high confidence damage events observed in experiment. Critical to this success is fully defining the material properties correctly, namely the fiber crush (SFC) and fiber shear (SFS) values, prior to calibrating the numerical parameters. In this paper, the effect of the determination of SFS and SFC on subsequent calibration steps is examined using two different experimental techniques.
New analytical techniques and a processability laminate have been used in the development of a material specification for 350 F curing epoxy-graphite materials. The development of the specification required the evaluation of prepregs which had variations in their processing behavior that were detected only in production. Chemical techniques were used to analyze the prepregs and establish limits on the concentrations of the various chemical constituents. A batch acceptance test laminate was designed to determine suitability of prepregs for production useage. The new specification requires that the processability laminate be ultrasonically and micrographically inspected for voids and porosity. Incorporation of these and other requirements will help to insure material consistency and processability.
The effect of low velocity impact on the strength of laminates fabricated from graphite/epoxy and Kevlar 49/epoxy composite materials was studied. The test laminates were loaded statically either in uniaxial tension or compression when impact occurred to evaluate the effect of loading on the initiation of damage and/or failure. Typical aircraft service conditions such as runway debris encountered during landing were simulated by impacting 1.27-cm-diameter projectiles normal to the plane of the test laminates at velocities between 5.2 and 48.8 m/s.
Generalized plane strain finite element analysis is used to predict reduction of interlaminar normal stresses when a U-shaped cap is bonded to the edge of a laminate. Three-dimensional composite material failure criteria are used in a progressive laminate failure analysis to predict failure loads of laminates with different edge cap designs. In an experimental program, symmetric 11-layer graphite-epoxy laminates with a one-layer cap of Kevlar-epoxy cloth are shown to be 130 to 140 percent stronger than uncapped laminates under static tensile and tension-tension fatigue loading. In addition, the coefficient of variation of the static tensile failure load decreases from 24 to 8 percent when edge caps are added. The predicted failure load calculated with the finite element results is 10 percent lower than the actual failure load. For both capped and uncapped laminates, actual failure loads are much lower than those predicted using classical lamination theory stresses and a two-dimensional failure criterion. Possible applications of the free edge reinforcement concept are described, and future research is suggested.
Progressive failure material modeling methods used for structural analysis including failure initiation and material degradation are presented. Different failure initiation criteria and material degradation models are described that define progressive failure formulations. These progressive failure formulations are implemented in a user-defined material model for use with a nonlinear finite element analysis tool. The failure initiation criteria include the maximum stress criteria, maximum strain criteria, the Tsai-Wu failure polynomial, and the Hashin criteria. The material degradation model is based on the ply-discounting approach where the local material constitutive coefficients are degraded. Applications and extensions of the progressive failure analysis material model address two-dimensional plate and shell finite elements and three-dimensional solid finite elements. Implementation details are described in the present paper. Parametric studies for laminated composite structures are discussed to illustrate the features of the progressive failure modeling methods that have been implemented and to demonstrate their influence on progressive failure analysis predictions.
Standard testing methods for composite laminates reinforced with continuous networks of braided, woven, or stitched fibers have been evaluated. The microstructure of these textile' composite materials differs significantly from that of tape laminates. Consequently, specimen dimensions and loading methods developed for tape type composites may not be applicable to textile composites. To this end, a series of evaluations were made comparing testing practices currently used in the composite industry. Information was gathered from a variety of sources and analyzed to establish a series of recommended test methods for textile composites. The current practices established for laminated composite materials by ASTM and the MIL-HDBK-17 Committee were considered. This document provides recommended test methods for determining both in-plane and out-of-plane properties. Specifically, test methods are suggested for: unnotched tension and compression; open and filled hole tension; open hole compression; bolt bearing; and interlaminar tension. A detailed description of the material architectures evaluated is also provided, as is a recommended instrumentation practice.
A simple, inexpensive method for bonding solid objects exploits hydroxide-catalyzed hydration and dehydration to form silicate-like networks in thin surface and interfacial layers between the objects. The method can be practiced at room temperature or over a wide range of temperatures. The method was developed especially to enable the formation of precise, reliable bonds between precise optical components. The bonds thus formed exhibit the precision and transparency of bonds formed by the conventional optical-contact method and the strength and reliability of high-temperature frit bonds. The method also lends itself to numerous non-optical applications in which there are requirements for precise bonds and/or requirements for bonds, whether precise or imprecise, that can reliably withstand severe environmental conditions. Categories of such non-optical applications include forming composite materials, coating substrates, forming laminate structures, and preparing objects of defined geometry and composition. The method is applicable to materials that either (1) can form silicate-like networks in the sense that they have silicate-like molecular structures that are extensible into silicate-like networks or (2) can be chemically linked to silicate-like networks by means of hydroxide-catalyzed hydration and dehydration. When hydrated, a material of either type features surface hydroxyl (-OH) groups. In this method, a silicate-like network that bonds two substrates can be formed either by a bonding material alone or by the bonding material together with material from either or both of the substrates. Typically, an aqueous hydroxide bonding solution is dispensed and allowed to flow between the mating surfaces by capillary action. If the surface figures of the substrates do not match precisely, bonding could be improved by including a filling material in the bonding solution. Preferably, the filling material should include at least one ingredient that can be hydrated to have exposed hydroxyl groups and that can be chemically linked, by hydroxide catalysis, to a silicate-like network. The silicate-like network could be generated in situ from the filling material and/or substrate material, or could be originally present in the bonding material.
Design approaches and materials are described from which are fabricated pyrostatic graphite/epoxy (Gr/Ep) laminates that show improved retention of graphite particulates when subjected to burning. Sixteen hybridized plus two standard Gr/Ep laminates were designed, fabricated, and tested in an effort to eliminate the release of carbon (graphite) fiber particles from burned/burning, mechanically disturbed samples. The term pyrostatic is defined as meaning mechanically intact in the presence of fire. Graphite particulate retentive laminates were constructed whose constituent materials, cost of fabrication, and physical and mechanical properties were not significantly different from existing Gr/Ep composites. All but one laminate (a Celion graphite/bis-maleimide polyimide) were based on an off-the-shelf Gr/Ep, the AS-1/3501-5A system. Of the 16 candidates studied, four thin (10-ply) and four thick (50-ply) hybridized composites are recommended.
A method of bonding substrates by hydroxide-catalyzed hydration/dehydration involves applying a bonding material to at least one surface to be bonded, and placing the at least one surface sufficiently close to another surface such that a bonding interface is formed between them. A bonding material of the invention comprises a source of hydroxide ions, and may optionally include a silicate component, a particulate filling material, and a property-modifying component. Bonding methods of the invention reliably and reproducibly provide bonds which are strong and precise, and which may be tailored according to a wide range of possible applications. Possible applications for bonding materials of the invention include: forming composite materials, coating substrates, forming laminate structures, assembly of precision optical components, and preparing objects of defined geometry and composition. Bonding materials and methods of preparing the same are also disclosed.