SEARCH · Search NASA

Results for “Composite Structures”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 289 records · Page 16

All sources NASA NTRS DOE OSTI

NASA NTRS2008

Factors Influencing Progressive Failure Analysis Predictions for Laminated Composite Structure

Progressive failure material modeling methods used for structural analysis including failure initiation and material degradation are presented. Different failure initiation criteria and material degradation models are described that define progressive failure formulations. These progressive failure formulations are implemented in a user-defined material model for use with a nonlinear finite element analysis tool. The failure initiation criteria include the maximum stress criteria, maximum strain criteria, the Tsai-Wu failure polynomial, and the Hashin criteria. The material degradation model is based on the ply-discounting approach where the local material constitutive coefficients are degraded. Applications and extensions of the progressive failure analysis material model address two-dimensional plate and shell finite elements and three-dimensional solid finite elements. Implementation details are described in the present paper. Parametric studies for laminated composite structures are discussed to illustrate the features of the progressive failure modeling methods that have been implemented and to demonstrate their influence on progressive failure analysis predictions.

Knight, Norman F., Jr.↗

NASA NTRS1987

Innovative design of composite structures: Use of curvilinear fiber format to improve structural efficiency

To increase the effectiveness and efficiency of fiber-reinforced materials, the use of fibers in a curvilinear rather than the traditional straightline format is explored. The capacity of a laminated square plate with a central circular hole loaded in tension is investigated. The orientation of the fibers is chosen so that the fibers in a particular layer are aligned with the principle stress directions in that layer. Finite elements and an iteration scheme are used to find the fiber orientation. A noninteracting maximum strain criterion is used to predict load capacity. The load capacities of several plates with different curvilinear fibers format are compared with the capacities of more conventional straightline format designs. It is found that the most practical curvilinear design sandwiches a group of fibers in a curvilinear format between a pair of +/-45 degree layers. This design has a 60% greater load capacity than a conventional quasi-isotropic design with the same number of layers. The +/-45 degree layers are necessary to prevent matrix cracking in the curvilinear layers due to stresses perpendicular to the fibers in those layers. Greater efficiencies are achievable with composite structures than now realized.

Hyer, M. W.↗

NASA NTRS1975

Design of composite structural components

Aspects related to the evolution of design are examined. A schematic of the evolution of the design objective and a schematic of designing a structural component from fiber composites are presented. A description of conventional design procedures is given, taking into account design handbooks, design guides, various sources of design information, and typical design data. Optimization methods are discussed, giving attention to parametric studies, the structural index method, aspects of structural synthesis, and questions of simultaneous material and structural design optimization.

Chamis, C. C.↗

NASA NTRS2004

Allowables for Structural Composites

In order to obtain the most benefit from building a structure with composite laminates, the strength of the laminate must be known. Based on the "weakest link" theory, the lower strength numbers obtained from testing are the ones to be used for design and analysis. The strength value to be used is determined by a statistical analysis of the test data, and is known as an allowable. MIL-HDBK- 17 outlines procedures to follow for determining these allowables. There are two types of statistically determined allowables, A- Basis and B-Basis. A-Basis is defined as a strength value at which only 1 in 100 specimens will fail with a 95% confidence level. B-Basis is a strength value at which only 10 in 100 specimens will fail with a 95% confidence level. As more specimens are tested a higher value of strength can be used as a valid allowable. Composites are highly process dependent and show much strength variation with environment, so it is critical to test materials and environments that are representative of hardware. Either using data obtained from a previous test series, or extrapolation to a certain temperature is highly discouraged.

Nettles, Alan T.↗

NASA NTRS1982

Composite structural materials

Research in the basic composition, characteristics, and processng science of composite materials and their constituents is balanced against the mechanics, conceptual design, fabrication, and testing of generic structural elements typical of aerospace vehicles so as to encourage the discovery of unusual solutions to problems. Detailed descriptions of the progress achieved in the various component parts of his program are presented.

Ansell, G. S.↗

NASA NTRS2004

Validation of Design and Analysis Techniques of Tailored Composite Structures

Aeroelasticity is the relationship between the elasticity of an aircraft structure and its aerodynamics. This relationship can cause instabilities such as flutter in a wing. Engineers have long studied aeroelasticity to ensure such instabilities do not become a problem within normal operating conditions. In recent decades structural tailoring has been used to take advantage of aeroelasticity. It is possible to tailor an aircraft structure to respond favorably to multiple different flight regimes such as takeoff, landing, cruise, 2-g pull up, etc. Structures can be designed so that these responses provide an aerodynamic advantage. This research investigates the ability to design and analyze tailored structures made from filamentary composites. Specifically the accuracy of tailored composite analysis must be verified if this design technique is to become feasible. To pursue this idea, a validation experiment has been performed on a small-scale filamentary composite wing box. The box is tailored such that its cover panels induce a global bend-twist coupling under an applied load. Two types of analysis were chosen for the experiment. The first is a closed form analysis based on a theoretical model of a single cell tailored box beam and the second is a finite element analysis. The predicted results are compared with the measured data to validate the analyses. The comparison of results show that the finite element analysis is capable of predicting displacements and strains to within 10% on the small-scale structure. The closed form code is consistently able to predict the wing box bending to 25% of the measured value. This error is expected due to simplifying assumptions in the closed form analysis. Differences between the closed form code representation and the wing box specimen caused large errors in the twist prediction. The closed form analysis prediction of twist has not been validated from this test.

Jegley, Dawn C.↗

NASA NTRS1977

Evaluation of flawed composite structure under static and cyclic loading

This paper presents the results of a program investigating the effects of initial defects on the fatigue and fracture response of composite laminates. The structural laminates investigated were a typical angle-ply laminate, a polar/hoop-wound pressure vessel laminate, and a typical engine fan blade laminate. Defects investigated were full- and half-penetration circular holes, full- and half-penetration slits, and countersink holes. Results are presented showing the effects of the defect size and type on the static fracture strength, fatigue performance, and residual static strength. The results of inspection procedures are shown, describing the effect of cyclic and static loadings on damage propagation in composite laminates. The data in this study were used to define proof test levels as a qualification procedure in composite structure subjected to cyclic loading.

Porter, T. R.↗

NASA NTRS1992

Multidisciplinary tailoring of hot composite structures

A computational simulation procedure is described for multidisciplinary analysis and tailoring of multilayered multimaterial hot composite engine structural components subjected to simultaneous multiple discipline-specific thermal, structural, vibration, and acoustic loadings including the effect of aggressive environments. The simulation is based on a 3D finite element analysis technique in conjunction with structural mechanics codes, thermal/acoustic analysis methods, and tailoring procedures. The integrated multidisciplinary simulation procedure is general-purpose including the coupled effects of nonlinearities in structure geometry, material, loading, and environmental complexities. The composite material behavior is assessed at all composite scales, i.e., the laminate/ply/constituents (fiber/matrix), via a nonlinear material characterization hygro-thermomechanical model. Sample tailoring cases exhibiting nonlinear material/loading/environmental behavior of aircraft engine fan blades, are presented. The various multidisciplinary loadings lead to different tailored designs, even those opposite of each other, as in the case of minimum material cost versus minimum structure weight and in the case of minimum vibration frequency versus minimum acoustic noise.

Singhal, Surendra N.↗

NASA NTRS1992

A micromechanics-based progressive failure model for laminated composite structures

A micromechanics-based mechanistic model for progressive failure analysis of laminated composite shell structures is developed. A finite-element program for laminated shell analysis is coupled with a micromechanical elasticity solution for predicting failure and effective composite properties. The laminate model is based on a third-order shell theory which allows for both transverse normal and shearing deformations. Results are presented for the progressive failure of a (0/90(4)s graphite/epoxy laminate under uniaxial tension. The effect of fiber volume fraction in the 90-deg plies on laminate behavior was studied, and it was found that decreasing the fiber-volume fraction in the off-axis plies did not affect the laminate stiffness or the ultimate failure load, but significantly increased the load at which initial failure occurred.

Averill, R. C.↗

NASA NTRS1984

Composite structural materials

Progress is reported in studies of constituent materials composite materials, generic structural elements, processing science technology, and maintaining long-term structural integrity. Topics discussed include: mechanical properties of high performance carbon fibers; fatigue in composite materials; experimental and theoretical studies of moisture and temperature effects on the mechanical properties of graphite-epoxy laminates and neat resins; numerical investigations of the micromechanics of composite fracture; delamination failures of composite laminates; effect of notch size on composite laminates; improved beam theory for anisotropic materials; variation of resin properties through the thickness of cured samples; numerical analysis composite processing; heat treatment of metal matrix composites, and the RP-1 and RP2 gliders of the sailplane project.

Ansell, G. S.↗

NASA NTRS1991

A model for the progressive failure of laminated composite structural components

Laminated continuous fiber polymeric composites are capable of sustaining substantial load induced microstructural damage prior to component failure. Because this damage eventually leads to catastrophic failure, it is essential to capture the mechanics of progressive damage in any cogent life prediction model. For the past several years the authors have been developing one solution approach to this problem. In this approach the mechanics of matrix cracking and delamination are accounted for via locally averaged internal variables which account for the kinematics of microcracking. Damage progression is predicted by using phenomenologically based damage evolution laws which depend on the load history. The result is a nonlinear and path dependent constitutive model which has previously been implemented to a finite element computer code for analysis of structural components. Using an appropriate failure model, this algorithm can be used to predict component life. In this paper the model will be utilized to demonstrate the ability to predict the load path dependence of the damage and stresses in plates subjected to fatigue loading.

Allen, D. H.↗

NASA NTRS1984

ACEE composite structures technology

Toppics addressed include: advanced composites on Boeing commercial aircraft; composite wing durability; damage tolerance technology development; heavily loaded wing panel design; and pressure containment and damage tolerance in fuselages.

Quinlivan, John T.↗

NASA NTRS1993

Multidisciplinary tailoring of hot composite structures

A computational simulation procedure is described for multidisciplinary analysis and tailoring of layered multi-material hot composite engine structural components subjected to simultaneous multiple discipline-specific thermal, structural, vibration, and acoustic loads. The effect of aggressive environments is also simulated. The simulation is based on a three-dimensional finite element analysis technique in conjunction with structural mechanics codes, thermal/acoustic analysis methods, and tailoring procedures. The integrated multidisciplinary simulation procedure is general-purpose including the coupled effects of nonlinearities in structure geometry, material, loading, and environmental complexities. The composite material behavior is assessed at all composite scales, i.e., laminate/ply/constituents (fiber/matrix), via a nonlinear material characterization hygro-thermo-mechanical model. Sample tailoring cases exhibiting nonlinear material/loading/environmental behavior of aircraft engine fan blades, are presented. The various multidisciplinary loads lead to different tailored designs, even those competing with each other, as in the case of minimum material cost versus minimum structure weight and in the case of minimum vibration frequency versus minimum acoustic noise.

Singhal, Surendra N.↗

NASA NTRS1978

Composite structural materials

The purpose of the RPI composites program is to develop advanced technology in the areas of physical properties, structural concepts and analysis, manufacturing, reliability and life prediction. Concommitant goals are to educate engineers to design and use composite materials as normal or conventional materials. A multifaceted program was instituted to achieve these objectives.

Ansell, G. S.↗

NASA NTRS1985

Damage and Repair of Composite Structures

NASA technical memorandum reports on research in damage and repair of composite-material structures of types used in airplane. Research includes: Identifying defect areas, determining whether particular kind of damage weakens structure in manner that degrades performance in intended application, developing repair procedures, and testing to demonstrate given repair procedure enables component to serve during remainder of its design life.

Deaton, J. W.↗

NASA NTRS2021

Latent Cure Epoxy Resins for Reliable Joints in Secondary-Bonded Composite Structures

In high-performance polymer matrix composite assemblies, adhesive bonding is generally superior to mechanical fastening in structural performance and manufacturing efficiency. However, adhesive bonds are susceptible to minute levels of contamination accumulated during assembly that can lead to unpredictable, weak bonds. Current methods of measuring bond strength are all destructive mechanical tests. To overcome these challenges, redundant load paths (e.g., mechanical fasteners) are often implemented in secondary-bonded, primary-structures, which can greatly reduce structural performance. This study investigated reformulated aerospace epoxy matrix resins with stoichiometric offset to inhibit cure of the matrix resin prior to assembly. Inhibited resins can reflow and mix across the joint interface, which eliminates the material discontinuity and forms a homogenous joint with reliable fracture properties. The goal of this study was to develop and demonstrate secondary composite assemblies that are mechanically and microscopically indistinguishable from a co-cured composite joint. This article describes the development of latent epoxy resins, the fabrication of test articles, and the mechanical properties measured from experimental joints compared with conventional, co-cured laminates. Methods of in-line quality control using and infrared spectroscopy and post-assembly forensics are also described. The final mode-II fracture toughness measured from precracked AERoBOND specimens was similar to that measured from co-cured laminates indicating that later cure epoxy materials could be a suitable replacement for secondary bonding.

Polymer-matrix composites (PMCs)↗

NASA NTRS1979

Evaluation of flawed composite structural components under static and cyclic loading

The effects of initial defects on the fatigue and fracture response of graphite-epoxy composite laminates are presented. The structural laminates investigated were a typical angle ply laminate, a polar/hoop wound pressure vessel laminate, and a typical engine fan blade laminate. Defects investigated were full and half penetration circular holes, full and half penetration slits, and countersink holes. The effects of the defect size and type on the static fracture strength, fatigue performance, and residual static strength are shown as well as the results of loadings on damage propagation in composite laminates. The data obtained were used to define proof test levels as a qualification procedure in composite structure subjected to cyclic loading.

Porter, T. R.↗

NASA NTRS1993

Caul and method for bonding and curing intricate composite structures

The invention disclosed here is a method for forming and curing an intricate structure of criss-crossing composite stringers and frames that are bonded to a skin panel. A structure constructed in accordance with the invention would be well-suited for use as a portion of an aircraft fuselage, a boat hull, or the like. The method is preferably practiced by applying uncured composite stringers to an uncured composite sheet panel. This is followed by placing cured frames crosswise over the stringers. The frames have openings at the locations where they intersect with the stringers which enables the frames to come into direct contact with the skin along most of their length. During the forming and curing process, the stringers are covered with a plurality of cauls, and the entire assembly of skin panel, stringers, frames and cauls is subjected to a vacuum bagging and curing process. The cauls serve to maintain both part shape and to control the flow of resin within the stringers as they are cured. Further, they probably eliminate the need for intermediate protective materials between the vacuum bag and the stringers.

Willden, Kurtis S.↗