ALBERT

Description: Conventional sandwich structure fabrication methods are labor intensive and high in cost. A low cost method is needed to produce lightweight sandwich structures. Sundstrand has developed a series of in situ composite fabrication methods in which the raw materials (skin and core materials) are placed in a closed mold, and the component is produced in one heating cycle. Internal pressure is generated by chemical agents during the thermal cycles, which consolidates the skins and produces the foam core. The finished part is a net-shape composite sandwich structure with skins and a foamed core. The in situ process reduces cost by eliminating several secondary operations that are used in conventional fabrication methods. Further, a strong molecular bond is produced between the core and skin, which eliminates adhesive bonding and prevents a weak bond section in the sandwich structure. In this investigation, we evaluated the feasibility of the in situ process using thermoset materials currently under consideration for commercial airplane fuselage applications, such as keel sections. The materials used were Hercules 855340 toughened epoxy resin in both liquid and powder forms, and 3M Scotchply PR500 resin, manufactured by 3M Corporation, in powder form. We successfully foamed these resins and produced experimental panels with AS-4/855340 Hercules prepreg skins. Chopped fibers were added to the core to increase performance of the foam. Mechanical property testing on these panels showed properties competitive with other foams. Additional experiments are required to optimize the in situ foam core sandwiches for specific properties and applications.

Description: Three-dimensional element analyses of (0/theta/-theta)s graphite epoxy laminates, where theta = 15, 20, 25, 30, and 45 deg, subjected to axial tensile load, were performed. The interlaminar stresses in the theta/-theta interface were calculated with and without a matrix crack in the central -theta plies. The interlaminar normal stress changes from a small compressive stress when no matrix crack is present to a high tensile stress at the intersection of the matrix crack and the free edge. The analysis of local delamination from the -theta matrix crack indicates a high strain energy release rate and a localized Mode I component near the free edge, within one-ply distance from the matrix crack. To examine the stress state causing the matrix cracking, the maximum principal normal stress in a plane perpendicular to the fiber direction in the -theta ply was calculated in an uncracked laminate. The corresponding shear stress parallel to the fiber was also calculated. The principal normal stress at the laminate edge increased through the ply thickness and reached a very high tensile value at the theta/-theta interface indicating that the crack in the -theta ply may initiate at the theta/-theta interface.

Description: The tensile creep and creep-recovery behavior of a hot-pressed unidirectional SiC-fiber/Si3N4-matrix composite was investigated at 1200 C in air, in order to determine how various sustained and cyclic creep loading histories would influence the creep rate, accumulated creep strain, and the amount of strain recovered upon specimen unloading. The data accumulated indicate that the fundamental damage mode for sustained tensile creep at stresses of 200 and 250 MPa was periodic fiber fracture and that the creep life and the failure mode at 250 MPa were strongly influenced by the rate at which the initial creep stress was applied. Cyclic loading significantly lowered the duration of primary creep and the overall creep-strain accumulation. The implications of the results for microstructural and component design are discussed.

Description: Microstructural changes occurring during sliding wear of self-mated Al2O3 SiC whisker-reinforced composites were studied using optical, scanning electron microscopy and transmission electron microscopy. Pin-on-disc specimens were slid in air at 2.7 m/s sliding velocity under a 26.5 N load for 1 h. Wear tests were conducted at 23, 600, 800 and 1200 C. Mild wear with a wear factor of 2.4 x 10 exp -7 - 1.5 x 10 exp -6 cu mm /N per m was experienced at all test temperatures. The composite showed evidence of wear by fatigue mechanisms at 800 C and below. Tribochemical reaction (SiC oxidation and reaction of SiO2 and Al2O3) leads to intergranular failure at 1200 C. Distinct microstructural differences existing at each test temperature are reported.

Description: This paper examines the use of a thin layer of Ultra High Molecular Weight Polyethylene (UHMWPE) on the outer surface of carbon/epoxy composite materials as a method of improving impact resistance and damage tolerance through hybridization. Flat 16-ply laminates as well as honeycomb sandwich structures with eight-ply facesheets were tested in this study. Instrumented drop-weight impact testing was used to inflict damage upon the specimens. Evaluation of damage resistance included instrumented impact data, visual examination, C-scanning and compression after impact (CAI) testing. The results show that only one lamina of UHMWPE did not improve the damage tolerance (strength retention) of the 16-ply flat laminate specimens or the honeycomb sandwich beams, however, a modest gain in impact resistance (detectable damage) was found for the honeycomb sandwich specimens that contained an outer layer of UHMWPE.

Description: Chemical and hydrodynamic aspects of wetting and interfacial phenomena during the solidification processing of metal-matrix composites are reviewed. Significant experimental results on fiber-matrix interactions and wetting under equilibrium and non-equilibrium conditions in composites of engineering interest have been compiled, based on a survey of the recent literature. Finally, certain aspects of wetting relevant to stir-casting and infiltration processing of composites are discussed.

Description: The inelastic deformation mechanisms were evaluated for a model titanium-based, fiber-reinforced composite: a beta titanium alloy (Ti-15V-3Al-3Cr-3Sn) reinforced with SiC (SCS-6) fibers. The primary emphasis of this article is to illustrate the sequence in which damage and plasticity evolved for this system. The mechanical responses and the results of detailed microstructural evaluations for the 0(8), 90(8), and +/- 45(2s) line oriented laminates are provided. It is shown that the characteristics of the reaction zone around the fiber play a very important role in the way damage and plasticity evolve, particularly in the microyield regime of deformation, and must be included in any realistic constitutive model. Fiber-matrix debonding was a major damage mode for the off-axis systems. The tension test results are also compared with the predictions of a few constitutive models.

Description: CTE (coefficient of thermal expansion) mismatch-induced stresses as they affect the fiber-matrix bond integrity of Al2O3 fiber-reinforced superalloy composites are examined. Of the three individual stress components, only the radial stress directly affects the integrity of the fiber-matrix interface. It is noted that a compressive radial stress leads to a clamping action on the fiber and is therefore beneficial to the integrity of the fiber-matrix bond. A radial tensile stress, on the other hand, can cause debonding of the fiber from the matrix for a weak fiber-matrix bond.

Description: Microstructural aspects of alloy solidification within the interstices of porous compacts of platelet-shaped single crystals of alpha-SiC, when the latter are infiltrated with a hot metal under pressure, have been described. Microstructural evidence is presented of selective reorientation of platelets and nonhomogeneous solute distribution under shear of pressurized melt, of constrained growth of primary solid within finite width zones, and of the modulation of coring due to microsegregation as a result of variations in the pore size of compacts.

Description: This paper investigates the unnotched tensile properties of two-dimensional (2D) triaxial braid-reinforced composites both experimentally and analytically. The materials are graphite fibers in an epoxy matrix. Three different reinforcing fiber architectures were considered. There were considerable differences in the observed elastic constants from different size strain gage and extensometer readings. Larger strain gages gave more consistent results and correlated better with the extensometer readings. Experimental strains correlated reasonably well with analytical predictions in the longitudinal, 0 deg, fiber direction but not in the transverse direction. Tensile strength results were not always predictable even in reinforcing directions. Minor changes in braid geometry led to disproportionate strength variations. The unit cell structure of the triaxial braid was discussed with the assistance of computer analysis of the microgeometry. Photomicrographs of braid geometry were used to improve upon the computer graphics representations of unit cells. These unit cells were used to predict the elastic moduli with various degrees of sophistication. The simple and the complex analyses were generally in agreement, but none adequately matched the experimental results for all the braids.

Description: Concerns related to the proper preparation of composite specimens for microstructural analysis are examined. Proper preparation will minimize the amount of surface and subsurface damage at each stage of the procedure so that the microstructural features of the final-polished specimen can be accurately determined as a function of the composite's response to processing, testing, or service conditions. This requires that an optimum combination of abrasive type, size, and bond be applied during each grinding, lapping, and polishing step. Machine settings, such as polishing speed, force, and relative polishing direction, are also important. Guidelines are given for each step of the six-stage specimen preparation process: sectioning, planar grinding, sample integrity, polishing, and etching.

Description: A modified sedimentation process was used in the production of a functionally gradient material (FGM), NiAl/Al2O3. A simple finite element model was used to guide our design and fabrication efforts by estimating residual stress states as a function of composite structure. This approach could lead to tailored designs that enhance or avoid specific residual stress states. Thermal cycling tests were factored into the model to predict time dependent or steady-state internal temperature and stress profiles. Four-point bend tests were conducted to establish the mechanical load-displacement behavior of a single interlayer FGM at room temperature, 800 and 1000 K. Room temperature bend strength of the FGM was 3-4 times that of the base NiAl. At elevated temperatures, composite fracture occurred in a gradual, noncatastrophic mode involving NiAl retardation of a succession of cracks originating in the alumina face.

Description: The next generation of hypersonic vehicles (NASP, SSTO) that require reusable thermal protection systems will experience acreage surface temperatures in excess of 1100 C. More important, they will experience a more severe physical environment than the Space Shuttle due to non-pristine launching and landing conditions. As a result, maintenance, inspection, and replacement factors must be more thoroughly incorporated into the design of the TPS. To meet these requirements, an advanced thermal protection system was conceived, designated 'TOPHAT'. This system consists of a toughened outer ceramic matrix composite (CMC) attached to a rigid reusable surface insulator (RSI) which is directly bonded to the surface. The objective of this effort was to evaluate this concept in an aeroconvective environment, to determine the effect of impacts to the CMC material, and to compare the results with existing thermal protection systems.

Description: A model is described for predicting the wear behavior of whisker reinforced ceramics. The model was successfully applied to a silicon carbide whisker reinforced alumina ceramic composite subjected to sliding contact. The model compares the friction forces on the whiskers due to sliding, which act to pull or push them out of the matrix, to the clamping or compressive forces on the whiskers due to the matrix, which act to hold the whiskers in the composite. At low temperatures, the whiskers are held strongly in the matrix and are fractured into pieces during the wear process along with the matrix. At elevated temperatures differential thermal expansion between the whiskers and matrix can cause loosening of the whiskers and lead to pullout during the wear process and to higher wear. The model, which represents the combination of elastic stress analysis and a friction heating analysis, predicts a transition temperature at which the strength of the whiskers equals the clamping force holding them in the matrix. Above the transition the whiskers are pulled out of the matrix during sliding, and below the transition the whiskers are simply fractured. The existence of the transition gives rise to a dual wear mode or mechanism behavior for this material which was observed in laboratory experiments. The results from this model correlate well with experimentally observed behavior indicating that the model may be useful in obtaining a better understanding of material behavior and in making material improvements.

Description: The objective of this study is to determine the energy dissipation processes in polymer-matrix composites during impact of ballistic projectiles. These processes include heat, fiber deformation and breakage, matrix deformation and fracture, and interfacial delamination. In this study, experimental measurements were made, using specialized specimen designs and test methods, to isolate the energy consumed by each of these processes during impact in the ballistic range. Using these experiments, relationships between material parameters and energy dissipation were examined. Composites with the same matrix but reinforced with Kevlar, PE, and graphite fabric were included in this study. These fibers were selected based on the differences in their intrinsic properties. Matrix cracking was found to be one of the most important energy absorption mechanisms during impact, especially in ductile samples such as Spectra-900 PE and Kevlar-49 reinforced polymer. On the contrary, delamination dominated the energy dissipation in brittle composites such as graphite reinforced materials. The contribution from frictional forces was also investigated and the energy partitioning among the different processes evaluated.

Description: The uniaxial response of a continuous fiber elastic-perfectly plastic composite is modeled herein as a two-element composite cylinder. An axisymmetric analytical micromechanics solution is obtained for the rate-independent elastic-plastic response of the two-element composite cylinder subjected to tensile loading in the fiber direction for the case wherein the core fiber is assumed to be a transversely isotropic elastic-plastic material obeying the Tsai-Hill yield criterion, with yielding simulating fiber failure. The matrix is assumed to be an isotropic elastic-plastic material obeying the Tresca yield criterion. It is found that there are three different circumstances that depend on the fiber and matrix properties: fiber yield, followed by matrix yielding; complete matrix yield, followed by fiber yielding; and partial matrix yield, followed by fiber yielding, followed by complete matrix yield. The order in which these phenomena occur is shown to have a pronounced effect on the predicted uniaxial effective composite response.

Description: Load-controlled isothermal and nonisothermal fatigue lives of a (0-deg)s SiC/Ti-15-3 were evaluated at temperatures between 150 and 550 C and a target strain range of about 0.45 percent. In nonisothermal fatigue tests, load was first cycled at minimum temperature and then temperature was cycled at zero load. For fatigue tests with peak temperatures at or above 300 C, fatigue life was dramatically reduced compared to that at 150 C. The shortest life was produced by the nonisothermal test with the greatest temperature range (Delta T = 400 C) and highest peak temperature (T(max) = 550 C). Vacuum testing showed that much of the life reduction under isothermal and nonisothermal conditions was related to environmental effects, although the nature of the fatigue-environment interaction was decidedly different for the isothermal and nonisothermal test cycles which were studied.

Description: Fiber push-out tests have been performed on a ceramic matrix composite consisting of carborundum-sintered SiC fibers, with a BN coating, embedded in a reaction-bonded SiC matrix. Analysis of the push-out data, utilizing the most complete theory presently available, shows that one of the fiber/coating/matrix interfaces has a low fracture energy (one-tenth that of the fiber) and a moderate sliding resistance of about 8 MPa. The debonded sliding interface shows some continuous but minor abrasion, which appears to increase the sliding resistance, but overall the system exhibits very clean smooth sliding. The tensile response of a full-scale composite is then modeled using data obtained here and known fiber strengths to demonstrate the good composite behavior predicted for this material.

Description: Silicon carbide SCS-6 fibers in a Ti24A1 + 11Nb matrix were subjected to off axis loading in a 'thin-slice' pushout test, resulting in various combinations of shear, radial compression, and tension along the fibers as a function of orientation angle. The load necessary for debonding decreased as the orientation angle increased, whereas the average frictional sliding stress after 60 s of sliding remained relatively constant for orientation angles less than 30 deg. Analyses of the specimen bending stresses and of the contact stresses by finite element modeling and thin plate theory are presented.

Description: This report describes a thermal-vacuum outgassing model and test protocol for predicting outgassing times and dimensional changes for polymer matrix composites. Experimental results derived from 'control' samples are used to provide the basis for analytical predictions to compare with the outgassing response of Long Duration Exposure Facility (LDEF) flight samples. Coefficient of thermal expansion (CTE) data are also presented. In addition, an example is given illustrating the dimensional change of a 'zero' CTE laminate due to moisture outgassing.

Description: Bolts and rivets provide a means of load transfer in the construction of aircraft. However, they give rise to stress concentrations and are often the source and location of static and fatigue failures. Furthermore, fastener holes are prone to cracks during take-off and landing. These cracks present the most common origin of structural failures in aircraft. Therefore, accurate determination of the contact stresses associated with such loaded holes in mechanically fastened joints is essential to reliable strength evaluation and failure prediction. As the laminate is subjected to loading, the contact region, whose extent is not known, develops between the fastener and the hole boundary through this contact region, which consists of slip and no-slip zones due to friction. The presence of the unknown contact stress distribution over the contact region between the pin and the composite laminate, material anisotropy, friction between the pin and the laminate, pin-hole clearance, combined bearing-bypass and shear loading, and finite geometry of the laminate result in a complex non-linear problem. In the case of bearing-bypass loading in compression, this non-linear problem is further complicated by the presence of dual contact regions. Previous research concerning the analysis of mechanical joints subjected to combined bearing-bypass and shear loading is non-existent. In the case of bearing-bypass loading only, except for the study conducted by Naik and Crews (1991), others employed the concept of superposition which is not valid for this non-linear problem. Naik and Crews applied a linear finite element analysis with conditions along the pin-hole contact region specified as displacement constraint equations. The major shortcoming of this method is that the variation of the contract region as a function of the applied load should be known a priori. Also, their analysis is limited to symmetric geometry and material systems, and frictionless boundary conditions. Since the contact stress distribution and the contact region are not known a priori, they did not directly impose the boundary conditions appropriate for modelling the contact and on-contact regions between the fastener and the hole. Furthermore, finite element analysis is not suitable for iterative design calculations for optimizing laminate construction in the presence of fasteners under complex loading conditions. In this study, the solution method developed by Madenci and Ileri (1992a,b) has been extended to determine the contact stresses in mechanical joints under combined bearing-bypass and shear loading, and bearing-bypass loading in compression resulting in dual contact regions.

Description: A thermal-vacuum outgassing model and test protocol for predicting outgassing times and dimensional changes for polymer matrix composites is described. Experimental results derived from a 'control' sample are used to provide the basis for analytical predictions to compare with the outgassing response of Long Duration Exposure Facility (LDEF) flight samples.

Description: The electronics module cover for the leading edge (Row D 9) experiment M0003-8 was fabricated from T300 graphite/934 epoxy unidirectional prepreg tape in a (O(sub 2), +/- 45, O(sub 2), +/- 45, 90, 0)(sub s) layup. This 11.75 in x 16.75 in panel was covered with thermal control coatings in three of the four quadrants with the fourth quadrant uncoated. The composite panel experienced different thermal cycling extremes in each quadrant due to the different optical properties of the coatings and bare composite. The panel also experienced ultraviolet (UV) and atomic oxygen (AO) attack as well as micrometeoroid and space debris impacts. An AO reactivity of 0.99 x 10(exp -24) cm(sup 3)/atom was calculated for the bare composite based on thickness loss. The white urethane thermal control coatings (A276 and BMS 1060) prevented AO attack of the composite substrate. However, the black urethane thermal control coating (Z306) was severely eroded by AO, allowing some AO attack of the composite substrate. An interesting banding pattern on the AO eroded bare composite surface was investigated and found to match the dimensions of the graphite fiber tow widths as prepregged. Also, erosion depths were greater in the darker bands. Five micrometeoroid/space debris impacts were cross sectioned to investigate possible structural damage as well as impact/AO interactions. Local crushing and delaminations were found to some extent in all of the impacts. No signs of coating undercutting were observed despite the extensive AO erosion patterns seen in the exposed composite material at the impact sites. An extensive microcrack study was performed on the panel along with modeling of the thermal environment to estimate temperature extremes and thermal shock. The white coated composite substrate displayed almost no microcracking while the black coated and bare composite showed extensive microcracking. Significant AO erosion was seen in many of the cracks in the bare composite.

Description: This report presented published LDEF micrometeoroid/debris impact data in a nomogram format useful for estimating the total number of hits that could be expected on a space structure as a function of time in orbit, angular location relative to ram, and exposed surface area. Correction factors accounting for different altitudes are given. These are normalized to the average LDEF altitude. Examples on how to use the nomograph are also included. In addition, impact data and damage areas observed on composite laminates (experiment AO 180) are discussed.

Description: Braided composite materials have potential for application in aircraft structures. Fuselage frames, floor beams, wing spars, and stiffeners are examples where braided composites could find application if cost effective processing and damage tolerance requirements are met. Another important consideration for braided composites relates to their mechanical properties and how they compare to the properties of composites produced by other textile composite processes being proposed for these applications. Unfortunately, mechanical property data for braided composites do not appear extensively in the literature. Data are presented in this paper on the mechanical characterization of 2D triaxial braid, 2D triaxial braid plus stitching, and 3D (through-the-thickness) braid composite materials. The braided preforms all had the same graphite tow size and the same nominal braid architectures, (+/- 30 deg/0 deg), and were resin transfer molded (RTM) using the same mold for each of two different resin systems. Static data are presented for notched and unnotched tension, notched and unnotched compression, and compression after impact strengths at room temperature. In addition, some static results, after environmental conditioning, are included. Baseline tension and compression fatigue results are also presented, but only for the 3D braided composite material with one of the resin systems.

Description: Results are presented from an experimental evaluation of the combined effects of temperature and humidity cycling on AS4/3501-6 composites (unstitched, Kevlar 29 stitched, and S-2 glass stitched uniweave fabric) and AS4/E905L composites (2-D, S-2 glass stitched 2-D, and 3-D braided fabric). The AS4/3501-6 uniweave material had a quasi-isotropic layup, whereas the AS4/E905L materials were braided in a (+/-30 deg/0 deg)(sub s) orientation. Data presented include compression strengths and compression-compression fatigue results for uncycled composites and cycled composites (160, 480, 720, and 1280 cycles from 140 deg F at 95 percent relative humidity to -67 deg F). To observe the presence of microcracking within the laminates, photomicrographs were taken of each material type at the end of each cycling period. Microcracks were found to be more prevalent within stitched laminates, predominantly around individual stitches. The glass stitched laminates showed significant microcracking even before cycling. Less microcracking was evident in the Kevlar stitched materials, whereas the unstitched uniweave material developed microcracks only after cycling. The 3-D braid did not develop microcracks. The static compression strengths of the unstitched and Kevlar stitched uniweave materials were degraded by about 10 percent after 1280 temperature/humidity cycles, whereas the reduction in compression strength for the glass stitched uniweave was less than 3 percent. The reduction in compression strength for the glass stitched 2-D braid was less than 8 percent. The unstitched 2-D and 3-D braids did not lose strength from temperature/humidity cycling. The compression-compression fatigue properties of all six material types were not affected by temperature/humidity cycling.

Description: The thermal expansion behavior of Long Duration Exposure Facility (LDEF) metal matrix composite materials was studied by (1) analyzing the flight data that was recorded on orbit to determine the effects of orbital time and heating/cooling rates on the performance of the composite materials, and (2) characterizing and comparing the thermal expansion behavior of post-flight LDEF and lab-control samples. The flight data revealed that structures in space are subjected to nonuniform temperature distributions, and thermal conductivity of a material is an important factor in establishing a uniform temperature distribution and avoiding thermal distortion. The flight and laboratory data showed that both Gr/Al and Gr/Mg composites were stabilized after prolonged thermal cycling on orbit. However, Gr/Al composites showed more stable thermal expansion behavior than Gr/Mg composites and offer advantages for space structures particularly where very tight thermal stability requirements in addition to high material performance must be met.

Description: Long Duration Exposure Facility (LDEF) Experiment A0175 involved the non-instrumented exposure of seven carbon-fiber reinforced resin-matrix advanced composite panels contained in two trays - A7 and A1. These two trays were located, respectively, on the leading and trailing faces of LDEF, obliquely oriented to the RAM (Row 9) and WAKE (Row 3) directions. The identity and location of the seven panels, which consisted of six flat laminates of the following material systems are shown: carbon/epoxy (T300/934), carbon/bismaleimide (T300/F178), and carbon/polyimide (C6000/LARC-160 and C6000/PMR-15), plus one bonded honeycomb sandwich panel (T300/934 face sheets and Nomex core) patterned after the Space Shuttle payload bay door construction. These material systems were selected to represent a range of then-available matrix resins which, by virtue of their differing polymer chemistry, could conceivably exhibit differing susceptibility to the low-earth orbit (LEO) environment. The principal exposure conditions of the LDEF environment at these tray locations are shown. Noteworthy to some of the observations discussed is the four-orders-of magnitude difference in the atomic oxygen (AO) fluence, which made a shallow incidence angle (approximately 22 deg) to Tray A7, while Tray A1 on the trailing face was essentially shielded from AO exposure. This evaluation focused on determining the individual and relative suitability of a variety of resin-matrix composite systems for long-term space structural applications. This was accomplished primarily by measuring and comparing a range of engineering mechanical properties on over 300 test coupons sectioned from the flight panels and from identical control panels, and tested at ambient and elevated temperatures. This testing was supported by limited physical characterization, involving visual examination of flight panel surface features, measurements of weight loss and warpage, and examination for changes in internal integrity (micro cracking, delamination) by ultrasonic c-scan and polished cross-sections.

Description: Concluded remarks are: (1) advanced carbon-carbon (ACC) substrate fabrication technology in good shape; (2) ACC coating improvements satisfactory but additional work needed; (3) non-destructive test techniques to monitor hardware during operational life needed; and (4) cost reduction approaches a high priority.

Description: There were two components, experimental and analytical, to this investigation of triaxially braided textile composite materials. The experimental portion of the study centered on measuring the materials' longitudinal and transverse tensile moduli, Poisson's ratio, and strengths. The identification of the damage mechanisms exhibited by these materials was also a prime objective of the experimental investigation. The analytical portion of the investigation utilized the Textile Composites Analysis (TECA) model to predict modulus and strength. The analytical and experimental results were compared to assess the effectiveness of the analysis. The figures contained in this paper reflect the presentation made at the conference. They may be divided into four sections: a definition of the material system tested; followed by a series of figures summarizing the experimental results (these figures contain results of a Moire interferometry study of the strain distribution in the material, examples and descriptions of the types of damage encountered in these materials, and a summary of the measured properties); a description of the TECA model follows the experimental results (this includes a series of predicted results and a comparison with measured values); and finally, a brief summary completes the paper.

Description: In the current environment, new technology must be cost-effective in addition to improving operability. Various approaches have been used to determine the 'hurdle' or 'breakthrough' return that must be achieved to gain customer commitment for a new product or aircraft, or in this case, a new application of the technology. These approaches include return-on-investment, payback period, and addition to net worth. An easily understood figure-of-merit and one used by our airline customers is improvement in direct operating cost per seat-mile. Any new technology must buy its way onto the aircraft through reduction in direct operating cost (DOC).

Description: Interactive Stiffened Panel Analysis (ISPAN) modules, written in FORTRAN, were developed to provide an easy to use tool for creating finite element models of composite material stiffened panels. The modules allow the user to interactively construct, solve and post-process finite element models of four general types of structural panel configurations using only the panel dimensions and properties as input data. Linear, buckling and post-buckling solution capability is provided. This interactive input allows rapid model generation and solution by non finite element users. The results of a parametric study of a blade stiffened panel are presented to demonstrate the usefulness of the ISPAN modules. Also, a non-linear analysis of a test panel was conducted and the results compared to measured data and previous correlation analysis.

Description: Composite structures are used for a wide variety of aerospace applications. Practical structures contain cutouts and these structures are subjected to in-plane and out-of-plane loading conditions. Structurally efficient designs for composite structures require a thorough understanding of the effects of cutouts on the response of composite plates subjected to inplane or out-of-plane loadings. Most investigations of the behavior of composite plates with cutouts have considered in-plane loadings only. Out-of-plane loadings suchas bending or twisting have received very limited attention. The response of homogeneous plates (e.g., isotropic or orthotropic plates) subjected to bending or twisting moments has been studied analytically. These analyses are for infinite plates and neglect finite-plate effects. Recently, analytical and experimental studies were conducted to determine the effects of cutouts on the response of laminated composite plates subjected to bending moments. No analytical or experimental results are currently available for the effects of cutouts on the response of composite laminates subjected to twisting moments.

Description: A methodology and attendant computer code have been developed and are described to computationally simulate the uncertain behavior of composite structures. The uncertain behavior includes buckling loads, stress concentration factors, displacements, stress/strain etc., which are the consequences of the inherent uncertainties (scatter) in the primitive (independent random) variables (constituent, ply, laminate and structural) that describe the composite structures. The computer code, IPACS (Integrated Probabilistic Assessment of Composite Structures), can handle both composite mechanics and composite structures. Application to probabilistic composite mechanics is illustrated by its uses to evaluate the uncertainties in the major Poisson's ratio and in laminate stiffness and strength. IPACS application to probabilistic structural analysis is illustrated by its use to evaluate the uncertainties in the buckling of a composite plate, in the stress concentration factor in a composite panel and in the vertical displacement and ply stress in a composite aircraft wing segment.

Description: In this study, the response of laminates with stress concentrations is explored. Automated Tow Placed (ATP, also known as Fiber Placement) laminates are compared to conventional tape layup manufacturing. Previous tensile fracture tests on fiber placed laminates show an improvement in tensile fracture of large notches over 20 percent compared to tape layup laminates. A hierarchial modeling scheme is presented. In this scheme, a global model is developed for laminates with notches. A local model is developed to study the influence of inhomogeneities at the notch tip, which are a consequence of the fiber placement manufacturing technique. In addition, a stacked membrane model was developed to study delaminations and splitting on a ply-by-ply basis. The results indicate that some benefit with respect to tensile fracture (up to 11 percent) can be gained from inhomogeneity alone, but that the most improvement may be obtained with splitting and delaminations which are more severe in the case of fiber placement compared to tape layup. Improvements up to 36 percent were found from the model for fiber placed laminates with damage at the notch tip compared to conventional tape layup.

Description: Many computer simulations in engineering and science -- and especially in computational fluid dynamics (CFD) -- produce huge quantities of numerical data. These data are often so large as to make even relatively simple post-processing of this data unwieldy. The data, once computed and quality-assured, is most likely analyzed by only a few people. As a result, much useful numerical data is under-utilized. Since future state-of-the-art simulations will produce even larger datasets, will use more complex flow geometries, and will be performed on more complex supercomputers, data management issues will become increasingly cumbersome. My goal is to provide software which will automate the present and future task of managing and post-processing large turbulence datasets. My research has focused on the development of these software tools -- specifically, through the development of a very high-level language called 'Tensoral'. The ultimate goal of Tensoral is to convert high-level mathematical expressions (tensor algebra, calculus, and statistics) into efficient low-level programs which numerically calculate these expressions given simulation datasets. This approach to the database and post-processing problem has several advantages. Using Tensoral the numerical and data management details of a simulation are shielded from the concerns of the end user. This shielding is carried out without sacrificing post-processor efficiency and robustness. Another advantage of Tensoral is that its very high-level nature lends itself to portability across a wide variety of computing (and supercomputing) platforms. This is especially important considering the rapidity of changes in supercomputing hardware.

Description: The Douglas Aircraft/NASA Act contract has been focused over the past three years at developing a materials, manufacturing, and cost base for stitched/Resin Transfer Molded (RTM) composites. The goal of the program is to develop RTM and stitching technology to provide enabling technology for application of these materials in primary aircraft structure with a high degree of confidence. Presented in this paper will be the progress to date in the area of manufacturing and associated cost values of stitched/RTM composites.

Description: Under NASA's Novel Composites for Wing and Fuselage Applications (NCWFA) program, Contract NAS1-18784, Grumman is evaluating the structural efficiency of graphite/epoxy cross-stiffened panel elements fabricated using innovative textile preforms and cost effective Resin Transfer Molding (RTM) and Resin Film Infusion (RFI) processes. Two three-dimensional woven preform assembly concepts have been defined for application to a representative window belt design typically found in a commercial transport airframe. The 3D woven architecture for each of these concepts is different; one is vertically woven in the plane of the window belt geometry and the other is loom woven in a compressed state similar to an unfolded eggcrate. The feasibility of both designs has been demonstrated in the fabrication of small test element assemblies. These elements and the final window belt assemblies will be structurally tested, and results compared.

Description: Composite materials that are subjected to complex loads have traditionally been fabricated with multidirectionally oriented prepreg tape materials. Some of the problems associated with this type of construction include low delamination resistance, poor out-of-plane strength, and labor intensive fabrication processes. Textile reinforced composites with through-the-thickness reinforcement have the potential to solve some of these problems. Recently, a relatively new class of noncrimp fabrics designated as multiaxial warp knits have been developed to minimize some of the high cost and damage tolerance concerns. Multiple stacks of warp knit fabrics can be knitted or stitched together to reduce layup labor cost. The through-the-thickness reinforcement can provide significant improvements in damage tolerance and out-of-plane strength. Multilayer knitted/stitched preforms, in conjunction with resin transfer molding (RTM), offer potential for significant cost savings in fabrication of primary aircraft structures. The objectives of this investigation were to conduct RTM processing studies and to characterize the mechanical behavior of composites reinforced with three multiaxial warp knit fabrics. The three fabrics investigated were produced by Hexcel and Milliken in the United States, and Saerbeck in Germany. Two resin systems, British Petroleum E9O5L and 3M PR 500, were characterized for RTM processing. The performance of Hexcel and Milliken quasi-isotropic knitted fabrics are compared to conventional prepreg tape laminates. The performance of the Saerbeck fabric is compared to uniweave wing skin layups being investigated by Douglas Aircraft Company in the NASA Advanced Composites Technology (ACT) program. Tests conducted include tension, open hole tension, compression, open hole compression, and compression after impact. The effects of fabric defects, such as misaligned fibers and gaps between tows, on material performance are also discussed. Estimated material and labor cost savings are projected for the Saerbeck fabric as compared to uniweave fabric currently being used by Douglas in the NASA ACT wing development program.

Description: This paper presents the first results in an assessment of the strength, stiffness, and damage tolerance of stiffened wing and fuselage subcomponents. Under this NASA funded program, 10 large wing and fuselage panels, variously fabricated by automated tow placement and dry-stitched preform/resin transfer molding, are to be tested. The first test of an automated tow placement six-longeron fuselage panel under shear load was completed successfully. Using NASTRAN finite-element analysis the stiffness of the panel in the linear range prior to buckling was predicted within 3.5 percent. A nonlinear analysis predicted the buckling load within 10 percent and final failure load within 6 percent. The first test of a resin transfer molding six-stringer wing panel under compression was also completed. The panel failed unexpectedly in buckling because of inadequate supporting structure. The average strain was 0.43 percent with a line load of 20.3 kips per inch of width. This strain still exceeds the design allowable strains. Also, the stringers did not debond before failure, which is in contrast to the general behavior of unstitched panels.

Description: Advanced resin systems and 3D textile preforms are being evaluated at Lockheed Aeronautical Systems Company (LASC) under NASA's Advanced Composites Technology (ACT) Program. This work is aimed towards the development of low-cost, damage-tolerant composite fuselage structures. Resin systems for resin transfer molding and powder epoxy towpreg materials are being evaluated for processability, performance and cost. Three developmental epoxy resin systems for resin transfer molding (RTM) and three resin systems for powder towpregging are being investigated. Various 3D textile preform architectures using advanced weaving and braiding processes are also being evaluated. Trials are being conducted with powdered towpreg, in 2D weaving and 3D braiding processes for their textile processability and their potential for fabrication in 'net shape' fuselage structures. The progress in advanced resin screening and textile preform development is reviewed here.

Description: This paper provides a brief overview of the NASA Advanced Composites Technology (ACT) Program. Critical technology issues that must be addressed and solved to develop composite primary structures for transport aircraft are delineated. The program schedule and milestones are included. Work completed in the first 3 years of the program indicates the potential for achieving composite structures that weigh less and are cost effective relative to conventional aluminum structure. Selected technical accomplishments are noted. Readers who are seeking more in-depth technical information should study the other papers included in these proceedings.

Description: The Boeing 777 is the first of a new family of wide body airplanes. The new large twin is sized to accommodate 360 to 390 passengers in typical two-class configurations and planned growth beyond that. The 777 offers airlines three engine options, extremely attractive operating costs, and compatibility with existing airport gates and taxiways. The 777 has a wingspan of nearly 197 feet and is offered with a wing-tip folding mechanism that will reduce the span to 156 feet. Extensive use of advance composite is included in the 777. The application range from fiberglass fairing to primary structures. The 777 empennage includes vertical fin and a horizontal stabilizer. The material used for the empennage is a new, toughened epoxy materials. The material provides outstanding resistance to impact damage.

Description: The automated fiber placement process has been in development at Hercules since 1980. Fiber placement is being developed specifically for aircraft and other high performance structural applications. Several major milestones have been achieved during process development. These milestones are discussed in this paper. The automated fiber placement process is currently being demonstrated on the NASA ACT program. All demonstration projects to date have focused on fiber placement of transport aircraft fuselage structures. Hercules has worked closely with Boeing and Douglas on these demonstration projects. This paper gives a description of demonstration projects and results achieved.

Description: Textile laminates, developed a number of years ago, have recently been shown to be applicable to primary aircraft structures for both small and large components. Such structures have the potential to reduce acquisition costs but require advanced automated processing to keep costs controlled while verifying product reliability and assuring structural integrity, durability and affordable life-cycle costs. Recently, resin systems and graphite-reinforced woven shapes have been developed that have the potential for improved RTM processes for aircraft structures. Ciba-Geigy, Brochier Division has registered an RTM prepreg reinforcement called 'Injectex' that has shown effectivity for aircraft components. Other novel approaches discussed are thermotropic resins producing components by injection molding and ceramic polymers for long-duration hot structures. The potential of such materials and processing will be reviewed along with initial information/data available to date.

Description: Flexible manufacturing methods are needed to reduce the cost of using advanced composites in structural applications. One method that allows for this is the stretch forming of long discontinuous fiber materials with thermoplastic matrices. In order to exploit this flexibility in an economical way, a thorough understanding of the relationship between manufacturing and component performance must be developed. This paper reviews some of the recent work geared toward establishing this understanding. Micromechanics models have been developed to predict the formability of the material during processing. The latest improvement of these models includes the viscoelastic nature of the matrix and comparison with experimental data. A finite element scheme is described which can be used to model the forming process. This model uses equivalent anisotropic viscosities from the micromechanics models and predicts the microstructure in the formed part. In addition, structural models have been built to account for the material property gradients that can result from the manufacturing procedures. Recent developments in this area include the analysis of stress concentrations and a failure model each accounting for the heterogeneous material fields.

Description: The objective of this study was to examine the effect of the unit cell architecture on the mechanical response of textile reinforced composite materials. Specifically, the study investigated the effect of unit cell size on the tensile properties of 2D triaxially braided graphite epoxy laminates. The figures contained in this paper reflect the presentation given at the conference. They may be divided into four sections: (1) a short definition of the material system tested; (2) a statement of the problem and a review of the experimental results; (3) experimental results consist of a Moire interferometry study of the strain distribution in the material plus modulus and strength measurements; and (4) a short summary and a description of future work will close the paper.

Description: Work ongoing under the NASA Langley - Advanced Composite Technology (ACT) program is discussed. The primary emphasis of the work centers around the development and characterization of graphite fiber that has been impregnated with an epoxy powder. Four epoxies have been characterized in towpreg form as to their weaveability and braidability. Initial mechanical properties have been generated on each resin system. These include unidirectional as well as 8-harness satin cloth. Initial 2D and 3D weaving and braiding trials will be reported on as well as initial efforts to develop towpreg suitable for advanced tow placement.

Description: Composite sandwich structures are being considered for primary structure in aircraft such as subsonic and high speed civil transports. The response of sandwich structures must be understood and predictable to use such structures effectively. Buckling is one of the most important response mechanisms of sandwich structures. A simple buckling analysis is derived for sandwich structures. This analysis is limited to flat, rectangular sandwich panels loaded by uniaxial compression (N(sub x)) and having simply supported edges. In most aerospace applications, however, the structure's geometry, boundary conditions, and loading are usually very complex. Thus, a general capability for analyzing the buckling behavior of sandwich structures is needed. The present paper describes and evaluates an improved buckling analysis for cylindrically curved composite sandwich panels. This analysis includes orthotropic facesheets and first-order transverse shearing effects. Both simple support and clamped boundary conditions are also included in the analysis. The panels can be subjected to linearly varying normal loads N(sub x) and N(sub y) in addition to a constant shear load N(sub xy). The analysis is based on the modified Donnell's equations for shallow shells. The governing equations are solved by direct application of Galerkin's method. The accuracy of the present analysis is verified by comparing results with those obtained from finite element analysis for a variety of geometries, loads, and boundary conditions. The limitations of the present analysis are investigated, in particular those related to the shallow shell assumptions in the governing equations. Finally, the computational efficiency of the present analysis is considered.

Description: The structural efficiency of compression-loaded trapezoidal-corrugation sandwich and semi-sandwich composite panels is studied to determine their weight savings potential. Sandwich panels with two identical face sheets and a trapezoidal corrugated core between them, and semi-sandwich panels with a corrugation attached to a single skin are considered. An optimization code is used to find the minimum weight designs for critical compressive load levels ranging from 3,000 to 24,000 lb/in. Graphite-thermoplastic panels based on the optimal minimum weight designs were fabricated and tested. A finite-element analysis of several test specimens was also conducted. The results of the optimization study, the finite-element analysis, and the experiments are presented.

Description: AS4/PEEK towpreg and IM7/Radel 8320 slit tape were used to make flat panels by automated tow placement. The panels were tested in notched and un-notched tension, notched and un-notched compression and compression after impact (CAI) at room temperature and under hot/wet conditions (notched and un-notched compression and CAI only). The properties were compared with AS4/PEEK tape laminate properties found in the literature. The tow placed AS4/PEEK material was stronger in tension but weaker in compression than the AS4/PEEK tape laminates. The tow placed AS4/PEEK was stronger but less stiff than the tow placed IM7/Radel 8320 in all compression tests. The IM7/Radel performed better in all other mechanical tests. The IM7/Radel outperformed the AS4/PEEK in all CAI tests.

Description: The HSCT plane, envisioned to have a lifetime of over 60,000 flight hours and to travel at speeds in excess of Mach 2, is the source of intensive study at NASA. In particular, polymer matrix composites are being strongly considered for use in primary and secondary structures due to their high strength to weight ratio and the options of property tailoring. However, an added difficulty in the use of polymer based materials is that their properties change significantly over time, especially at the elevated temperatures that will be experienced during flight, and prediction of properties based on irregular thermal and mechanical loading is extremely difficult. This study focused on one aspect of long-term polymer composite behavior: physical aging. When a polymer is cooled to below its glass transition temperature, the material is not in thermodynamic equilibrium and the free volume and enthalpy evolve over time to approach their equilibrium values. During this time, the mechanical properties change significantly and this change is termed physical aging. This work begins with a review of the concepts of physical aging on a pure polymer system. The effective time theory, which can be used to predict long term behavior based on short term data, is mathematically formalized. The effects of aging to equilibrium are proven and discussed. The theory developed for polymers is then applied first to a unidirectional composite, then to a general laminate. Comparison to experimental data is excellent. It is shown that the effects of aging on the long-term properties of composites can be counter-intuitive, stressing the importance of the development and use of a predictive theory to analyze structures.

Description: Increasingly, composite materials are being used in advanced structural applications because of the significant weight savings they offer when compared to more traditional engineering materials. The higher cost of composites must be offset by the increased performance that results from reduced structural weight if these new materials are to be used effectively. At present, there is considerable interest in fabricating solid rocket motor cases out of composite materials, and capitalizing on the reduced structural weight to increase rocket performance. However, one of the difficulties that arises when composite materials are used is that composites can develop significant amounts of internal damage during low velocity impacts. Such low velocity impacts may be encountered in routine handling of a structural component like a rocket motor case. The ability to assess the reduction in structural integrity of composite motor cases that experience accidental impacts is essential if composite rocket motor cases are to be certified for manned flight. While experimental studies of the post-impact performance of filament wound composite motor cases haven been proven performed (2,3), scaling impact data from small specimens to full scale structures has proven difficult. If such a scaling methodology is to be achieved, an increased understanding of the damage processes which influence residual strength is required. The study described herein was part of an ongoing investigation of damage development and reduction of tensile strength in filament wound composites subjected to low velocity impacts. The present study, which focused on documenting the damage that develops in filament wound composites as a result of such impacts, included two distinct tasks. The first task was to experimentally assess impact damage in small, filament wound pressure bottles using x-ray radiography. The second task was to study the feasibility of using digital image processing techniques to assist in determining the 3-D distribution of damage from stereo x-ray pairs.

Description: Durability and damage tolerance may have different connotations to people from different industries and with different backgrounds. Damage tolerance always refers to a safety of flight issue where the structure must be able to sustain design limit loads in the presence of damage and return to base safely. Durability, on the other hand, is an economic issue where the structure must be able to survive a certain life under load before the initiation of observable damage. Delamination is typically the observable damage mechanism that is of concern for durability, and the growth and accumulation of delaminations through the laminate thickness is often the sequence of events that leads to failure and the loss of structural integrity.

Description: A composite material incorporates high strength, high modulus fibers in a matrix (polymer, metal, or ceramic). The fibers may be oriented in a manner to give varying in-plane properties (longitudinal, transverse-stress, strain, and modulus of elasticity). The lay-up of the composite laminates is such that a center line of symmetry and no bending moment exist through the thickness. The laminates are tabbed, with either aluminum or fiberglass, and are ready for tensile testing. The determination of the tensile properties of resin matrix composites, reinforced by continuous fibers, is outlined in ASTM standard D 3039, Tensile Properties of Oriented Fiber Composites. The tabbed flat tensile coupons are placed into the grips of a tensile machine and load-deformation curves plotted. The load-deformation data are translated into stress-strain curves for determination of mechanical properties (ultimate tensile strength and modulus of elasticity).

Description: The following questions are addressed: What are high performance composites?; How are they used today?; What are their properties?; How do you make them?; and What are the future technology needs for composites? In addition, samples of composite reinforcements such as glass, carbon and Kevlar fibers, matrix materials, and fabricated composite parts will be demonstrated and made available.

Description: The Naval Surface Warfare Center has developed processes for the preparation of mullite (3(Al2O3)(dot)2(SiO2)) whiskers and mullite-whisker felt. Three patents on the technology were issued in 1990. The processes are based on chemical reactions between AlF3, Al2O3, and SiO2. The felt is formed in-situ during the processing of shaped powdered precursors. It consists of randomly oriented whiskers which are mutually intergrown forming a rigid structure. The microstructure and properties of the felt and size of the whiskers can be modified by varying the amount of Al2O3 in the starting mixture. Loose mullite whiskers can be used as a reinforcement for polymer-, metal-, and ceramic-matrix composites. The felt can be used as preforms for fabricating composite materials as well as for thermal insulation and high temperature, chemically stable filters for liquids (melts) and gases.

Description: Carbon/phenolic composites are used extensively as ablative insulating materials in the nozzle region of solid rocket motors. The current solid rocket motor (RSRM) on the space shuttle is fabricated from woven rayon cloth which is carbonized and then impregnated with the phenolic resin. These plies are layed up in the desired configuration and cured to form the finished part. During firing, the surface of the carbon/phenolic insulation is exposed to 5000 F gases from the rocket exhaust. The resin pyrolizes and the material chars to a depth which progresses with time. The rate of charring and erosion are generally predictable, and the insulation depth is designed to allow adequate safety margins over the firing time of the motor. However, anomalies in the properties and response of the carbon/phenolic materials can lead to severe material damage which may decrease safety margins to unacceptable levels. Three macro damage modes which were observed in fired nozzles are: ply lift, 'wedge out', and pocketing erosion. Ply lift occurs in materials with plies oriented nearly parallel to the surface. The damage occurs in a region below the charred material where material temperatures are relatively low - about 500 F. Wedge out occurs at the intersection of nozzle components whose plies are oriented at about 45 deg. The corner of the block of material breaks off along a ply interface. Pocketing erosion occurs in material with plies oriented normal to the surface. Thermal expansion is restrained in two directions resulting in large tensile strains and material failure normal to the surface. When a large section of material is removed as a result of damage, the insulation thickness is reduced which may lead to failure of the nozzle due to excessive heating of critical components. If these damage events cannot be prevented with certainty, the designer must increase the thickness of the insulator thus adding to both weight and cost. One of the difficulties in developing a full understanding of these macro damage mechanisms is that the loading environment and the material response to that environment are extremely complex. These types of damage are usually only observed in actual motor firings. Therefore, it is difficult and expensive to evaluate the reliability of new materials. Standard material tests which measure mechanical and thermal properties of test specimens can only provide a partial picture of how the material will respond in the service environment. The development of the ANALOG test procedure which can combine high heating rates and mechanical loads on a specimen will improve the understanding of the interactive effects of the various loads on the system. But a mechanistic model of material response which can account for the heterogeneity of the material, the progression of various micromechanical damage mechanisms, and the interaction of mechanical and thermal stresses on the material is required to accurately correlate material tests with response to service environments. A model based on fundamental damage mechanisms which is calibrated and verified under a variety of loading conditions will provide a general tool for predicting the response of rocket nozzles. The development of a micromechanical simulation technique was initiated and demonstrated to be effective for studying across-ply tensile failure of carbon/phenolic composites.

Description: The material-adaptive three-dimensional analysis of inhomogeneous structures based on the meso-volume concept and application of deficient spline functions for displacement approximations is proposed. The general methodology is demonstrated on the example of a brick-type mosaic parallelepiped arbitrarily composed of anisotropic meso-volumes. A partition of each meso-volume into sub-elements, application of deficient spline functions for a local approximation of displacements and, finally, the use of the variational principle allows one to obtain displacements, strains, and stresses at anypoint within the structural part. All of the necessary external and internal boundary conditions (including the conditions of continuity of transverse stresses at interfaces between adjacent meso-volumes) can be satisfied with requisite accuracy by increasing the density of the sub-element mesh. The application of the methodology to textile composite materials is described. Several numerical examples for woven and braided rectangular composite plates and stiffened panels under transverse bending are considered. Some typical effects of stress concentrations due to the material inhomogeneities are demonstrated.

Description: NASA's Advanced Composite Technology (ACT) Program was initiated in 1990 with the purpose of developing less costly composite aircraft structures. A number of innovative materials and processes were evaluated as a part of this effort. Chief among them are composite materials reinforced with textile preforms. These new forms of composite materials bring with them potential testing problems. Methods currently in practice were developed over the years for composite materials made from prepreg tape or simple 2-D woven fabrics. A wide variety of 2-D and 3-D braided, woven, stitched, and knit preforms were suggested for application in the ACT program. The applicability of existing test methods to the wide range of emerging materials bears investigation. The overriding concern is that the values measured are accurate representations of the true material response. The ultimate objective of this work is to establish a set of test methods to evaluate the textile composites developed for the ACT Program.

Description: A probabilistic model utilizing random material characteristics to predict damage evolution in textile laminates is presented. Model is based on a division of each ply into two sublaminas consisting of cells. The probability of cell failure is calculated using stochastic function theory and maximal strain failure criterion. Three modes of failure, i.e. fiber breakage, matrix failure in transverse direction, as well as matrix or interface shear cracking, are taken into account. Computed failure probabilities are utilized in reducing cell stiffness based on the mesovolume concept. A numerical algorithm is developed predicting the damage evolution and deformation history of textile laminates. Effect of scatter of fiber orientation on cell properties is discussed. Weave influence on damage accumulation is illustrated with the help of an example of a Kevlar/epoxy laminate.

Description: One of the greatest difficulties in developing detailed models of the mechanical response of textile reinforced composites is an accurate model of the reinforcing elements. In the case of elastic property prediction, the variation of fiber position may not have a critical role in performance. However, when considering highly localized stress events, such as those associated with cracks and holes, the exact position of the reinforcement probably dominates the failure mode. Models were developed for idealized reinforcements which provide an insight into the local behavior. However, even casual observations of micrographical images reveals that the actual material deviates strongly from the idealized models. Some of the deviations and causes are presented for triaxially braided and three dimensionally woven textile composites. The necessary modeling steps to accommodate these variations are presented with some examples. Some of the ramifications of not accounting for these discrepencies are also addressed.

Description: A new type of braided fiber seal was developed for high temperature engine applications. Development work performed includes seal design, fabrication, leakage flow testing, and flow resistance modeling. This new type of seal utilizes the high flow resistance of tightly packed fibers and the conformability of textile structures. The seal contains a core part with aligned fibers, and a sheath with braided fiber layers. Seal samples are made by using the conventional braiding process. Leakage flow measurements are then performed. Mass flow rate versus the simulated engine pressure and preload pressure is recorded. The flow resistance of the seal is analyzed using the Ergun equation for flow through porous media, including both laminar and turbulent effects. The two constants in the Ergun equation are evaluated for the seal structures. Leakage flow of the seal under the test condition is found to be in the transition flow region. The analysis is used to predict the leakage flow performance of the seal with the determined design parameters.

Description: The study on the permeability of an aligned fiber bundle is the key building block in modeling the permeability of advanced woven and braided preforms. Available results on the permeability of fiber bundles in the literature show that a substantial difference exists between numerical and analytical calculations on idealized fiber packing structures, such as square and hexagonal packing, and experimental measurements on practical fiber bundles. The present study focuses on the variation of the permeability of a fiber bundle under practical process conditions. Fiber bundles are considered as containing openings and fiber clusters within the bundle. Numerical simulations on the influence of various openings on the permeability were conducted. Idealized packing structures are used, but with introduced openings distributed in different patterns. Both longitudinal and transverse flow are considered. The results show that openings within the fiber bundle have substantial effect on the permeability. In the longitudinal flow case, the openings become the dominant flow path. In the transverse flow case, the fiber clusters reduce the gap sizes among fibers. Therefore the permeability is greatly influenced by these openings and clusters, respectively. In addition to the porosity or fiber volume fraction, which is commonly used in the permeability expression, another fiber bundle status parameter, the ultimate fiber volume fraction, is introduced to capture the disturbance within a fiber bundle.

Description: A technique based on matching the refractive index of an invading liquid to that of a fiber mat was used to study entrapment of air ('voids') that occurs during forced in-plane radial flow into nonwoven multifilament glass networks. The usefulness of this technique is demonstrated in quantifying and mapping the air pockets. Experiments with a series of fluids with surface tensions varying from 28 x 10(exp -3) to 36 x 10(exp -3) N/m, viscosities from 45 x 10(exp -3) to 290 x 10(exp -3) Pa.s, and inlet flow rates from 0.15 x 10(exp -6) to 0.75 x 10(exp -6) m(exp 3)/s, showed that void content is a function of the capillary number characterizing the flow process. A critical value of capillary number, Ca = 2.5 x 10(exp -3), identifies a zone below which void content increases exponentially with decreasing capillary number. Above this critical value, negligible entrapment of voids is observed. Similar experiments carried out on surface treated nonwoven mats spanning a range of equilibrium contact angles from 20 deg to 78 deg showed that there is a critical contact angle above which negligible entrapment is observed. Below this value, there is no apparent effect of contact angle on the void fraction - capillary number relationship described earlier. Studies on the effect of filament wettability, and fluid velocity and viscosity on the size of the entrapment (voids) were also carried out. These indicate that larger sized entrapments which envelop more than one pore are favored by a low capillary number in comparison to smaller, pore level bubbles. Experiments were carried out on deformed mats - imposing high permeability spots at regular intervals on a background of low permeability. The effect of these spatial fluctuations in heterogeneity of the mat on entrapment is currently being studied.

Description: New textile composite materials development, production, and application are discussed. Topics covered include: super-high-strength, super-high-modulus fibers, filaments, and materials manufactured on their basis; heat-resistant and nonflammable fibers, filaments, and textile fabrics; fibers and textile fabrics based on fluorocarbon poylmers; antifriction textile fabrics based on polyfen filaments; development of new types of textile combines and composite materials; and carbon filament-based fabrics.

Description: An engineering approach for the application of textile composites to a structural component is addressed. The main objective is to improve impact resistance of composite blades by using some form of 3-D reinforcement. Project goals, results, and conclusions are discussed.

Description: Efferent sorption methods of organism detoxication (by medical trend) are presented. Recently, specialists have shown their keen interest in the problem of treating exogenous and endogenous intoxications. This was stipulated by the growing production and accumulation of chemical products for industrial, agricultural, and domestic needs. To solve this problem the industrial production of carbon fibrous adsorbents was developed and implemented at NII Chimvolokno in St. Petersburg. A description of the carbon fibers is given. Also, application of modern composite materials for manufacturing compression-distraction apparatus used for setting fractured bones is described.

Description: The development of modern technologies demands the creation of new nonmetallic, fibrous materials with specific properties. The fibers and materials developed by NII 'Chimvolokno', St. Petersburg, can be divided into two groups. The first group includes heat-resistant fibers, fire-resistant fibers, thermotropic fibers, fibers for medical application, and textile structures. The second group contains refractory fibers, chemoresistant and antifriction fibers, fibers on the basis of polyvinyl alcohol, microfiltering films, and paperlike and nonwoven materials. In cooperation with NPO 'Chimvolokno' MYTITSHI, we developed and started producing heat-resistant high-strength fibers on the base of polyhetarearilin and aromatic polyimides (SVM and terlon); heat-resistant fibers on the base of polyemede (aramid); fire-retardant fibers (togilen); chemoresistant and antifriction fibers on the basis of homo and copolymers of polytetrafluoroethylene (polyfen and ftorin); and water soluble, acetylated, and high-modulus fibers from polyvinyl alcohol (vylen). Separate reports will deal with textile structures and thermotropic fibers, as well as with medical fibers. One of the groups of refractory fibers carbon fibers (CF) and the corresponding paperlike nonwoven materials are discussed in detail. Also, composite materials (CM) and their base, which is the subject of the author's research since 1968, is discussed.

Description: Three-dimensional (3D) woven preforms are currently being considered for use as primary structural components. Lack of technology to properly manufacture, characterize and predict mechanical properties, and predict damage mechanisms leading to failure are problems facing designers of textile composite materials. Two material systems with identical specifications but different manufacturing approaches are investigated. One manufacturing approach resulted in an irregular (nonuniform) preform geometry. The other approach yielded the expected preform geometry (uniform). The objectives are to compare the mechanical properties of the uniform and nonuniform angle interlock 3D weave constructions. The effect of adding layers of laminated tape to the outer surfaces of the textile preform is also examined. Damage mechanisms are investigated and test methods are evaluated.

Description: Fundamentally, the braiding process is a highly efficient, low cost method for combining single yarns into circumferential shapes, as evidenced by the number of applications for continuous sleeving. However, this braiding approach cannot fully demonstrate that it can drastically reduce the cost of complex shape structural preforms. Factors such as part geometry, machine design and configuration, materials used, and operating parameters are described as key cost drivers and what is needed to minimize their effect on elevating the cost of structural braided preforms.

Description: An overview is presented of the research on textile composites at Katholieke Universiteit Leuven. Three dimensionally woven sandwich fabric preforms are investigated for delamination resistant sandwich structures, velvet woven 2.5 dimensional fabrics for delamination resistant laminates, and knitted fabrics with good drapability for laminates of complex shape.

Description: New technology generally serves two main goals of the automotive industry: one is to enable vehicles to comply with various governmental regulations and the other is to provide a competitive edge in the market. The latter goal can either be served through improved manufacturing and design capabilities, such as computer aided design and computer aided manufacturing, or through improved product performance, such as anti-lock braking (ABS). Although safety features are sometimes customer driven, such as the increasing use of airbags and ABS, most are determined by regulations as outlined by the Federal Motor Vehicle Safety Standards (FMVSS). Other standards, set by the Environmental Protection Agency, determine acceptable levels of emissions and fuel consumption. State governments, such as in California, are also setting precedent standards, such as requiring manufacturers to offer zero-emission vehicles as a certain fraction of their sales in the state. The drive to apply new materials in the automobile stems from the need to reduce weight and improve fuel efficiency. Topics discussed include: new lightweight materials; types of automotive materials; automotive composite applications; the role for composite materials in automotive applications; advantages and disadvantages of composite materials; material substitution economics; economic perspective; production economics; and composite materials production economics.

Description: The performance of interlinked, multi-layer fabrics and near net shape preforms for engineering applications, woven on a 48 shaft dobby loom using glass, aramid, and carbon continuous filament yarns is assessed. The interlinking was formed using the warp yarns. Two basic types of structure were used. The first used a single warp beam and hence each of the warp yarns followed a similar path to form four layer interlinked reinforcements and preforms. In the second two warp beams were used, one for the interlinking yarns which pass from the top to the bottom layer through-the-thickness of the fabric and vice versa, and the other to provide 'straight' yarns in the body of the structure to carry the axial loading. Fabrics up to 15mm in thickness were constructed with varying amounts of through-the-thickness reinforcement. Tapered T and I sections were also woven, with the shaping produced by progressive removal of ends during construction. These fabrics and preforms were impregnated with resin and cured to form composite samples for testing. Using these two basic types of construction, the influence of reinforcement construction and the proportion and type of interlinking yarn on the performance of the composite was assessed.

Description: The effects of low and high velocity impact on thick hybrid composites (THC's) were experimentally compared. Test Beams consisted of CFRP skins which were bonded onto an interleaved syntactic foam core and cured at 177 C (350 F). The impactor tip for both cases was a 16 mm (0.625 inch) steel hemisphere. In spite of the order of magnitude difference in velocity ranges and impactor weights, similar relationships between impact energy, damage size, and residual strength were found. The dependence of the skin compressive strength on damage size agree well with analytical open hole models for composite laminates and may enable the prediction of ultimate performance for the damaged composite, based on visual inspection.

Description: The NASA Langley Research Center is conducting and sponsoring research to explore the benefits of textile reinforced composites for civil transport aircraft primary structures. The objective of this program is to develop and demonstrate the potential of affordable textile reinforced composite materials to meet design properties and damage tolerance requirements of advanced aircraft structures. In addition to in-house research, the program includes major participation by the aircraft industry and aerospace textile companies. The major program elements include development of textile preforms, processing science, mechanics of materials, experimental characterization of materials, and development and evaluation of textile reinforced composite structural elements and subcomponents. The NASA Langley in-house research is focused on science-based understanding of resin transfer molding (RTM), development of powder-coated towpreg processes, analysis methodology, and development of a performance database on textile reinforced composites. The focus of the textile industry participation is on development of multidirectional, damage-tolerant preforms, and the aircraft industry participation is in the areas of innovative design concepts, cost-effective fabrication, and testing of textile reinforced composite structural elements and subcomponents. Textile processes such as 3-D weaving, 2-D and 3-D braiding, and knitting/stitching are being compared with conventional laminated tape processes for improved damage tolerance. Through-the-thickness reinforcements offer significant damage tolerance improvements. However, these gains must be weighed against potential loss in in-plane properties such as strength and stiffness. Analytical trade studies are underway to establish design guidelines for the application of textile material forms to meet specific loading requirements. Fabrication and testing of large structural components are required to establish the full potential of textile reinforced composite materials. The goals of the NASA Langley-sponsored research program are to demonstrate technology readiness with subscale composite components by 1995 and to verify the performance of full-scale composite primary aircraft structural components by 1997. The status of textile reinforced composite structural elements under development by Boeing, Douglas, Lockheed, and Grumman are presented. Included are braided frames and woven/stitched wing and fuselage panels.

Description: The long-term performance of polymer-based composites in the space environment is discussed. Both thermoset and thermoplastic matrix composites are included in this discussion. Previous efforts on the space environmental effects on composites are briefly reviewed. Focus of this review is placed on the effects of hygrothermal stresses, atomic oxygen, ultraviolet (UV), and space debris/micrometeoroid impacts along with the potential synergism. Potential approaches to estimating the residual strength of polymer composites after exposure to atomic oxygen erosion or space debris/micrometeoroid impact are evaluated. New ground-based data are then utilized to illustrate the effects of atomic oxygen and thermal cycling on the failure behavior of polymer composites. Finally, research needs, challenges, and opportunities in the field of space environmental effects on composite materials are highlighted.

Description: Ply-lift and pocketing are two critical anomalies of carbon cloth phenolic composites (CCPC) in rocket nozzle applications. Ply lift occurs at low temperatures when the A/P and in-plane permeabilities of the composite materials are still very low and in-plane porous paths are blocked. Pocketing occurs at elevated temperatures when in-plane permeability is reduced by the A/P compressive stress. The thermostructural response of CCPC in a rapid heating environment involves simultaneous heat, mass, and momentum transfer along with the degradation of phenolic resin in a multiphase system with temperature- and time-dependent material properties as well as dynamic processing conditions. Three temperature regions represent the consequent chemical reactions, material transformations, and property transitions, and provide a quick qualitative method for characterizing the thermostructural behavior of a CCPC. In order to optimize the FM5939 LDCCP (low density carbon cloth phenolic) for the nozzle performance required in the Advanced Solid Rocket Motor (ASRM) program, a fundamental study on LDCCP materials was conducted. The cured composite has a density of 1.0 +/- 0.5 gm/cc which includes 10 to 25 percent void volume. The weight percent of carbon microballoon is low (7-15 percent). However, they account for approximately one third of the volume and historically their percentages have not been controlled very tightly. In addition, the composite properties show no correlation with microballoon weight percent or fiber properties (e.g. fiber density or fiber moisture adsorption capacity). Test results concerning the ply-lift anomaly in the MNASA motor firings were: (1) Steeper ply angle (shorter path lenght) designs minimized/eliminated by lifting, (2) material with higher void volume ply lifted less frequently, (3) materials with high (greater than 9 percent) microballoon content had a higher rate of ply lifting, and (4) LDCCP materials failed at microballoon-resin interfaces. The objectives of this project are: (1) to investigate the effects of carbon microballoon and cabosil fillers as well as fiber heat treatment on plylift-related mechanical properties, (2) to develop a science-based thermostructural process model for the carbon phenolics. The model can be used in the future for the selection of the improved ASRM materials, (3) to develop the micro-failure mechanisms for the ply-lift initiation and propagation processes during the thermoelastic region of phenolic degradation, i.e. postcuring and devolatilization.

Description: Future proposed NASA missions with the need for large deployable or erectable precision structures will require solutions to many technical problems. The Jet Propulsion Laboratory (JPL) is developing new technologies in Adaptive Structures to meet these challenges. The technology requirements, approaches to meet the requirements using Adaptive Structures, and the recent JPL research results in Adaptive Structures are described.

Description: Structural components produced from laminated CMC (ceramic matrix composite) materials are being considered for a broad range of aerospace applications that include various structural components for the national aerospace plane, the space shuttle main engine, and advanced gas turbines. Specifically, these applications include segmented engine liners, small missile engine turbine rotors, and exhaust nozzles. Use of these materials allows for improvements in fuel efficiency due to increased engine temperatures and pressures, which in turn generate more power and thrust. Furthermore, this class of materials offers significant potential for raising the thrust-to-weight ratio of gas turbine engines by tailoring directions of high specific reliability. The emerging composite systems, particularly those with silicon nitride or silicon carbide matrix, can compete with metals in many demanding applications. Laminated CMC prototypes have already demonstrated functional capabilities at temperatures approaching 1400 C, which is well beyond the operational limits of most metallic materials. Laminated CMC material systems have several mechanical characteristics which must be carefully considered in the design process. Test bed software programs are needed that incorporate stochastic design concepts that are user friendly, computationally efficient, and have flexible architectures that readily incorporate changes in design philosophy. The CCARES (Composite Ceramics Analysis and Reliability Evaluation of Structures) program is representative of an effort to fill this need. CCARES is a public domain computer algorithm, coupled to a general purpose finite element program, which predicts the fast fracture reliability of a structural component under multiaxial loading conditions.

Description: Chemical compatibility between Fe-19.8Cr-4.8Al (weight percent), which is the base composition for the commercial superalloy MA956, and several carbides, borides, nitrides, oxides, and silicides was analyzed from thermodynamic considerations. The effect of addition of minor alloying elements, such as Ti, Y, and Y2O3, to the Fe-Cr-Al alloy on chemical compatibility between the alloy and various compounds was also analyzed. Several chemically compatible compounds that can be potential reinforcement materials and/or interface coating materials for Fe-Cr-Al based composites were identified.

Description: Over 35 different types of organic matrix composites were flown as part of 11 different experiments onboard the NASA Long Duration Exposure Facility (LDEF) satellite. This materials and systems experiment satellite flew in low-earth orbit (LEO) for 69 months. For that period, the experiments were subjected to the LEO environment including atomic oxygen (AO), ultraviolet (UV) radiation, thermal cycling, microvacuum, meteoroid and space debris (M&D), and particle radiation. Since retrieval of the satellite in January of 1990, the principal experiment investigators have been deintegrating, examining, and testing the materials specimens flown. The most detrimental environmental effect on all organic matrix composites was material loss due to AO erosion. AO erosion of uncoated organic matrix composites (OMC) facing the satellite ram direction was responsible for significant mechanical property degradations. Also, thermal cycling-induced microcracking was observed in some nonunidirectional reinforced OMC's. Thermal cycling and outgassing caused significant but predictable dimensional changes as measured in situ on one experiment. Some metal and metal oxide-based coatings were found to be very effective at preventing AO erosion of OMC's. However, M&D impacts and coating fractures which compromised these coatings allowed AO erosion of the underlying OMC substrates. The findings for organic matrix composites flown on the LDEF are summarized and the LEO environmental factors, their effects, and the influence on space hardware design factors for LEO applications are identified.

Description: Over 200 graphite/aluminum and graphite/magnesium composites were flown on the leading and trailing edges of LDEF on the Advanced Composites Experiment. The performance of these composites was evaluated by performing scanning electron microscopy and x-ray photoelectron spectroscopy of exposed surfaces, optical microscopy of cross sections, and on-orbit and postflight thermal expansion measurements. Graphite/aluminum and graphite/magnesium were found to be superior to graphite/polymer matrix composites in that they are inherently resistant to atomic oxygen and are less susceptible to thermal cycling induced microcracking. The surface foils on graphite/aluminum and graphite/magnesium protect the graphite fibers from atomic oxygen and from impact damage from small micrometeoroid or space debris particles. However, the surface foils were found to be susceptible to thermal fatigue cracking arising from contamination embrittlement, surface oxidation, or stress risers. Thus, the experiment reinforced requirements for carefully protecting these composites from prelaunch oxidation or corrosion, avoiding spacecraft contamination, and designing composite structures to minimize stress concentrations. On-orbit strain measurements demonstrated the importance of through-thickness thermal conductivity in composites to minimize thermal distortions arising from thermal gradients. Because of the high thermal conductivity of aluminum, thermal distortions were greatly reduced in the LDEF thermal environment for graphite/aluminum as compared to graphite/magnesium and graphite/polymer composites. The thermal expansion behavior of graphite/aluminum and graphite/magnesium was stabilized by on-orbit thermal cycling in the same manner as observed in laboratory tests.

Description: Damped composite laminates were fabricated by co-curing viscoelastic damping film with graphite/epoxy prepreg plies. The dynamic response of the damped plates was measured using an impulse response technique and compared with the response of similar undamped laminates. Modal damping was computed from the frequency response data. Micrographs of the damped laminates showed that the damping layers retained their integrity during the fabrication process. The layers significantly increased the damping in the composite laminates. The use of the constrained viscoelastic film as an integral part of composite structures appears to be a feasible approach to passive vibration control. Composite plates manufactured with co-cured damping layers may have commercial applications in cases where light weight, strength, and vibration and noise reduction are important considerations.

Description: A finite element micromechanics approach was utilized to investigate the thermally induced stress fields in continuous fiber reinforced polymer matrix composites at temperatures typical of spacecraft operating environments. The influence of laminate orientation was investigated with a simple global/local formulation. Thermal stress calculations were used to predict probable damage initiation locations, and the results were compared to experimentally observed damage in several epoxy matrix composites. The influence of an interphase region on the interfacial stress states was investigated.

Description: A pressure box test fixture was designed and fabricated to evaluate the effects of internal pressure, biaxial tension loads, curvature, and damage on the fracture response of composite fuselage structure. Previous work in composite fuselage tension damage tolerance, performed during NASA contract NAS1-17740, evaluated the above effects on unstiffened panels only. This work extends the tension damage tolerance testing to curved stiffened fuselage crown structure that contains longitudinal stringers and circumferential frame elements. The pressure box fixture was designed to apply internal pressure up to 20 psi, and axial tension loads up to 5000 lb/in, either separately or simultaneously. A NASTRAN finite element model of the pressure box fixture and composite stiffened panel was used to help design the test fixture, and was compared to a finite element model of a full composite stiffened fuselage shell. This was done to ensure that the test panel was loaded in a similar way to a panel in the full fuselage shell, and that the fixture and its attachment plates did not adversely affect the panel.

Description: Closed form and finite element analyses are presented for axial direction and transverse direction dimensional stability of skin/stringer panels. Several sensitivity studies are presented to illustrate the influence of various design parameters on the dimensional stability of these panels. Panel geometry, material properties (stiffness and coefficient of thermal expansion), restraint conditions and local details, such as resin fillets, all combine to influence dimensional stability, residual and assembly forces.

Description: The strength of laminated composite materials may be significantly reduced by foreign object impact induced damage. An understanding of the damage state is required in order to predict the behavior of structure under operational loads or to optimize the structural configuration. Types of damage typically induced in laminated materials during an impact event include transverse matrix cracking, delamination, and/or fiber breakage. The details of the damage state and its influence on structural behavior depend on the location of the impact. Damage in the skin may act as a soft inclusion or affect panel stability, while damage occurring over a stiffener may include debonding of the stiffener flange from the skin. An experiment to characterize impact damage resistance of fuselage structure as a function of structural configuration and impact threat was performed. A wide range of variables associated with aircraft fuselage structure such as material type and stiffener geometry (termed, intrinsic variables) and variables related to the operating environment such as impactor mass and diameter (termed, extrinsic variables) were studied using a statistically based design-of-experiments technique. The experimental design resulted in thirty-two different 3-stiffener panels. These configured panels were impacted in various locations with a number of impactor configurations, weights, and energies. The results obtained from an examination of impacts in the skin midbay and hail simulation impacts are documented. The current discussion is a continuation of that work with a focus on nondiscrete characterization of the midbay hail simulation impacts and discrete characterization of impact damage for impacts over the stiffener.

Description: The goal of the Boeing effort under the NASA ACT program is to reduce manufacturing costs of composite fuselage structure. Materials, fabrication of complex subcomponents and assembly issues are expected to drive the costs of composite fuselage structure. Several manufacturing concepts for the crown section of the fuselage were evaluated through the efforts of a Design Build Team (DBT). A skin-stringer-frame intricate bond design that required no fasteners for the panel assembly was selected for further manufacturing demonstrations. The manufacturing processes selected for the intricate bond design include Advanced Tow Placement (ATP) for multiple skin fabrication, resin transfer molding (RTM) of fuselage frames, innovative cure tooling, and utilization of low-cost material forms. Optimization of these processes for final design/manufacturing configuration was evaluated through the fabrication of several intricate bond panels. Panels up to 7 ft. by 10 ft. in size were fabricated to simulate half scale production parts. The qualitative and quantitative results of these manufacturing demonstrations were used to assess manufacturing risks and technology readiness for production.

Description: Tests were conducted on over 200 center-crack specimens to evaluate: (a) the tension-fracture performance of candidate materials and laminates for commercial fuselage applications; and (b) the accuracy of several failure criteria in predicting response. Crack lengths of up to 12 inches were considered. Other variables included fiber/matrix combination, layup, lamination manufacturing process, and intraply hybridization. Laminates fabricated using the automated tow-placement process provided significantly higher tension-fracture strengths than nominally identical tape laminates. This confirmed earlier findings for other layups, and possibly relates to a reduced stress concentration resulting from a larger scale of repeatable material inhomogeneity in the tow-placed laminates. Changes in material and layup result in a trade-off between small-notch and large-notch strengths. Toughened resins and 0 deg-dominate layups result in higher small-notch strengths but lower large-notch strengths than brittle resins, 90 deg and 45 deg dominated layups, and intraply S2-glass hybrid material forms. Test results indicate that strength-prediction methods that allow for a reduced order singularity of the crack-tip stress field are more successful at predicting failure over a range of notch sizes than those relying on the classical square-root singularity. The order of singularity required to accurately predict large-notch strength from small-notch data was affected by both material and layup. Measured crack-tip strain distributions were generally higher than those predicted using classical methods. Traditional methods of correcting for finite specimen width were found to be lacking, confirming earlier findings with other specimen geometries. Fracture tests of two stiffened panels, identical except for differing materials, with severed central stiffeners resulted in nearly identical damage progression and failure sequences. Strain-softening laws implemented within finite element models appear attractive to account for load redistribution in configured structure due to damage-induced crack tip softening

Description: We used a variational stress analysis and an energy release rate failure criterion to construct a master plot analysis of matrix microcracking. In the master plot, the results for all laminates of a single material are predicted to fall on a single line whose slope gives the microcracking toughness of the material. Experimental results from 18 different layups of AS4/3501-6 laminates show that the master plot analysis can explain all observations. In particular, it can explain the differences between microcracking of central 90 deg plies and of free-surface 90 deg plies. Experimental results from two different AS4/PEEK laminates tested at different temperatures can be explained by a modified master plot that accounts for changes in the residual thermal stresses. Finally, we constructed similar master plot analyses for previous literature microcracking models. All microcracking theories that ignore the thickness dependence of the stresses gave poor results.

Description: We subjected double cantilever beam specimens from four different composite materials to mixed-mode precracking. Three different precracking mode 1 to mode 2 ratios were used--1 to 4, 1 to 1, and 4 to 1. Following precracking the specimens were tested for mode I fracture toughness. The mixed-mode precracking often influenced the mode 1 toughness and its influence persisted for as much as 60 mm of mode 1 crack growth. We tested composites with untoughened matrices, composites with rubber-toughened matrices, and composites with interlayer toughening. Depending on material type and precracking mode ratio, the precracking could cause either a significant increase or a significant decrease in the mode 1 fracture toughness.

Description: The rheology of SiC particulate/Al-6.5 pct Si composite slurries was explored. The rheological behavior of the composite slurries shows both thixotropic and pseudoplastic behaviors. Isostructural experiments on the composite slurries revealed a Newtonian behavior beyond a high shear rate limit. The rheology of fully molten composite slurries over the low to high shear rate range indicates the existence of a low shear rate Newtonian region, an intermediate pseudoplastic region and a high shear rate Newtonian region. The isostructural studies indicate that the viscosity of a composite slurry depends upon the shearing history of a given volume of material. An unexpected shear thinning was noted for SiC particulate + alpha slurries as compared to semi-solid metallic slurries at the same fraction solid. The implications of these findings for the processing of slurries into cast components is discussed.

Description: Gas derived graphite fibers generated by the decomposition of an organic gas are joined with a suitable binder. This produces a high thermal conductivity composite material which passively conducts heat from a source, such as a semiconductor, to a heat sink. The fibers may be intercalated. The intercalate can be halogen or halide salt, alkaline metal, or any other species which contributes to the electrical conductivity improvement of the graphite fiber. The fibers are bundled and joined with a suitable binder to form a high thermal conductivity composite material device. The heat transfer device may also be made of intercalated highly oriented pyrolytic graphite and machined, rather than made of fibers.

Description: Degradation of a number of fiber/polymer composites located on the leading and trailing surfaces of LDEF where the atomic oxygen (AO) fluences ranged from 10(exp 22) to 10(exp 4) atoms/cm(sup 2), respectively, was observed and compared. While matrices of the composites on the leading edge generally exhibited considerable degradation and erosion-induced fragmentation, this 'asking' process was confined to the near surface regions because these degraded structures acted as a 'protective blanket' for deeper-lying regions. This finding leads to the conclusion that simple surface coatings can significantly retard AO and other combinations of degrading phenomena in low-Earth orbit. Micrometeoroid and debris particle impacts were not a prominent feature on the fiber composites studied and apparently do not contribute in a significant way to their degradation or alteration in low-Earth orbit.

Description: A methodology and attendant computer code were developed and are used to computationally simulate the uncertain behavior of composite structures. The uncertain behavior includes buckling loads, stress concentration factors, displacements, stress/strain, etc., which are the consequences of the inherent uncertainties (scatter) in the primitive (independent random) variables (constituent, ply, laminate, and structural) that describe the composite structures. The computer code is IPACS (Integrated Probabilistic Assessment of Composite Structures). IPACS can simulate both composite mechanics and composite structural behavior. Application to probabilistic composite mechanics is illustrated by its use to evaluate the uncertainties in the major Poisson's ratio and in laminate stiffness and strength. IPACS' application to probabilistic structural analysis is illustrated by its used to evaluate the uncertainties in the buckling of a composite plate, the stress concentration factor in a composite panel, and the vertical displacement and ply stress in a composite aircraft wing segment. IPACS' application to probabilistic design is illustrated by its use to assess the thin composite shell (pipe).

Description: In this work, we utilize fuzzy sets theory to evaluate and make predictions of flexural strength and density of NASA 6Y silicon nitride ceramic. Processing variables of milling time, sintering time, and sintering nitrogen pressure are used as an input to the fuzzy system. Flexural strength and density are the output parameters of the system. Data from 273 Si3N4 modulus of rupture bars tested at room temperature and 135 bars tested at 1370 C are used in this study. Generalized mean operator and Hamming distance are utilized to build the fuzzy predictive model. The maximum test error for density does not exceed 3.3 percent, and for flexural strength 7.1 percent, as compared with the errors of 1.72 percent and 11.34 percent obtained by using neural networks, respectively. These results demonstrate that fuzzy sets theory can be incorporated into the process of designing materials, such as ceramics, especially for assessing more complex relationships between the processing variables and parameters, like strength, which are governed by randomness of manufacturing processes.

Description: An analytical procedure is presented for determining the transient response of simply supported, rectangular laminated composite plates subjected to impact loads from airgun-propelled or dropped-weight impactors. A first-order shear-deformation theory is included in the analysis to represent properly any local short-wave-length transient bending response. The impact force is modeled as a locally distributed load with a cosine-cosine distribution. A double Fourier series expansion and the Timoshenko small-increment method are used to determine the contact force, out-of-plane deflections, and in-plane strains and stresses at any plate location due to an impact force at any plate location. The results of experimental and analytical studies are compared for quasi-isotropic laminates. The results indicate that using the appropriate local force distribution for the locally loaded area and including transverse-shear-deformation effects in the laminated plate response analysis are important. The applicability of the present analytical procedure based on small deformation theory is investigated by comparing analytical and experimental results for combinations of quasi-isotropic laminate thicknesses and impact energy levels. The results of this study indicate that large-deformation effects influence the response of both 24- and 32-ply laminated plates, and that a geometrically nonlinear analysis is required for predicting the response accurately.

Description: The utility of a recently developed analytical micromechanics model for the response of metal matrix composites under thermal loading is illustrated by comparison with the results generated using the finite-element approach. The model is based on the concentric cylinder assemblage consisting of an arbitrary number of elastic or elastoplastic sublayers with isotropic or orthotropic, temperature-dependent properties. The elastoplastic boundary-value problem of an arbitrarily layered concentric cylinder is solved using the local/global stiffness matrix formulation (originally developed for elastic layered media) and Mendelson's iterative technique of successive elastic solutions. These features of the model facilitate efficient investigation of the effects of various microstructural details, such as functionally graded architectures of interfacial layers, on the evolution of residual stresses during cool down. The available closed-form expressions for the field variables can readily be incorporated into an optimization algorithm in order to efficiently identify optimal configurations of graded interfaces for given applications. Comparison of residual stress distributions after cool down generated using finite-element analysis and the present micromechanics model for four composite systems with substantially different temperature-dependent elastic, plastic, and thermal properties illustrates the efficacy of the developed analytical scheme.