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At least 667 records · Page 37
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NASA NTRS1992

Analysis for delamination initiation in postbuckled dropped-ply laminates

The compression strength of dropped-ply, graphite-epoxy laminated plates for the delamination mode of failure is studied by analysis and corroborated with experiments. The nonlinear response of the test specimens is modeled by a geometrically nonlinear finite element analysis. The methodology for predicting delamination is based on a quadratic interlaminar stress criterion evaluated at a characteristic distance from the ply drop-off. The compression strength of specimens exhibiting a linear response is greater than the compression strength of specimens with the same layup exhibiting a geometrically nonlinear response. The analyses for both linear and nonlinear response show that severe interlaminar stress gradients occur in the interfaces at the drop-off because of the thickness/stiffness discontinuity. However, these interlaminar stress distributions are altered in the geometrically nonlinear response such that, with increasing load, their growth at the center of the laminate is retarded while their growth near the unloaded supported edge is increased.

Davila, Carlos G.↗
NASA NTRS1992

Free-edge effects in laminates under extension, bending and twisting. II - Sublaminate/layer modeling and analysis

The stress-function-based variational method of Yin (1991) is extended and modified into a combined layer/sublaminate approach applicable to a laminated strip composed of a large number of differently orientated, anisotropic elastic plies. Lekhnitskii's (1963) stress functions are introduced into two interior layers adjacent to a particular interface. The remaining layers are grouped into an upper sublaminate and a lower sublaminate. The stress functions are expanded in truncated power series of the thickness coordinate, and the differential equations governing the coefficient functions are derived by using the complementary virtual work principle. The layer/sublaminate approach limits the dimension of the eigenvalue problem to a fixed number irrespective of the number of layers in the sublaminate, so that reasonably accurate solutions of the interlaminar stresses can be computed with extreme ease. For symmetric, four-layer, angle-ply and cross-ply laminates, a comparison of the previous analysis results based on the pure layer model and new results based on two different layer/sublaminate models indicates reasonable over-all agreement in the interlaminar stresses and superior agreement in the total peeling and shearing force.

Yin, Wan-Lee↗
NASA NTRS1992

Initial postbuckling response of anisotropic laminated rectangular plates

It has been shown that anisotropic plates can have unstable postbuckling behavior resulting in potential imperfection sensitivity. The present paper quantifies the degree of instability for rectangular, simply supported asymmetrically laminated plates. The analysis is based on asymptotic Koiter-type expansion of postbuckling response. The degree of postbuckling instability is quantified in terms of the reduction in load carrying capacity in the immediate postbuckling range. For graphite-epoxy plates it is found that this measure of instability is very small unless the lamination asymmetry is very pronounced.

Johnson, Eric R.↗
NASA NTRS1992

Reliability analysis of laminated CMC components through shell subelement techniques

An updated version of the integrated design program C/CARES (composite ceramic analysis and reliability evaluation of structures) was developed for the reliability evaluation of CMC laminated shell components. The algorithm is now split in two modules: a finite-element data interface program and a reliability evaluation algorithm. More flexibility is achieved, allowing for easy implementation with various finite-element programs. The new interface program from the finite-element code MARC also includes the option of using hybrid laminates and allows for variations in temperature fields throughout the component.

Starlinger, A.↗
NASA NTRS1992

Vibration of thick composite laminates - Analytic theory and finite element approximations

Two versions of a (1,2) higher-order laminate plate theory are extended to elastodynamics, and their performance on free vibration problems for homogeneous and laminated plates is examined. Both versions are based on continuously varying through-the-thickness linear and parabolic distributions for the inplane and transverse displacements, respectively. They also use continuous parabolic distributions for the transverse shear strains. The theory is capable of providing an improved accuracy as compared to other approximate theories and accurately predicting the lowest thickness-stretch frequencies which cannot be obtained with shear-deformable theories. The theory is considered to be particularly well-suited for generating simple and efficient variationally-based displacement finite elements leading to C sup 0 elements with only five degrees of freedom per node.

Tessler, A.↗
NASA NTRS1992

Thermally-induced, geometrically nonlinear response of symmetrically laminated composite plates

This paper discusses the thermally-induced geometrically nonlinear response of symmetrically laminated composite plates. The plate response is due to a temperature increase that is uniform in the plane of the plate but has a slight gradient through the thickness. The case of a completely uniform temperature increase but with an initial out-of-plane imperfection in the plate is also considered. Because they are closely allied problems, thermal buckling and postbuckling are discussed. Using variational methods in conjunction with a Rayleigh-Ritz formulation, these responses are investigated for two laminates, a (+/- 45/0/90)s and a (+/- 45/02)s, under two different simple support conditions.

Meyers, C. A.↗
NASA NTRS1992

Analytical shape sensitivities and approximations of modal response of generally laminated tapered skew plates

In the present determination of the derivatives of natural frequencies and mode shapes of a generally laminated tapered skew plate, with respect to various shape parameters, springs are used to simulate the essential boundary conditions. The independent shape parameters are plate surface area, aspect ratio, taper ratio, and sweep angle. Eigenvalues and eigenvectors are approximated over the range of the variable using linear, exponential, and pseudoexponential approximation schemes, and are compared with reanalysis-obtained values. Numerical results are presented for symmetrically and unsymmetrically laminated plates.

Singhvi, Sarvesh↗
NASA NTRS1992

Stacking sequence optimization of simply supported laminates with stability and strain constraints

An integer programming formulation for the design of symmetric and balanced rectangular composite laminates with simply supported boundary conditions subject to buckling and strain constraints is presented. The design variables that define the stacking sequence of the laminate are ply-identity zero-one integers. The buckling constraint is linear in terms of the ply-identity design variables, but strains are nonlinear functions of these variables. A linear approximation is developed for the strain constraints so that the problem can be solved by sequential linearization using the branch and bound algorithm. Examples of graphite-epoxy plates under biaxial compression are presented. Optimum stacking sequences obtained using the linear approximation are compared with global optimum designs obtained using a genetic search procedure.

Nagendra, S.↗
NASA NTRS1992

Near-tip dual-length scale mechanics of mode-I cracking in laminate brittle matrix composites

This paper presents the preliminary results of an on-going study of the near-tip mechanics of mode-I cracking in brittle matrix composite laminates. A finite element model is developed within the context of two competing characteristic lengths present in the composite: the microstructural length (the thickness of the layers) and a macro-length (crack-length, uncracked ligament size, etc.). For various values of the parameters which describe the ratio of these lengths and the constituent properties, the stresses ahead of a crack perpendicular to the laminates are compared with those predicted by assuming the composite is homogeneous orthotropic. The results can be used to determine the conditions for which homogenization can provide a sufficiently accurate description of the stresses in the vicinity of the crack-tip.

Ballarini, R.↗
NASA NTRS1992

Exact solutions for laminated composite cylindrical shells in cylindrical bending

Analytic elasticity solutions for laminated composite cylindrical shells under cylindrical bending are presented. The material of the shell is assumed to be general cylindrically anisotropic. Based on the theory of cylindrical anisotropic elasticity, coupled governing partial differential equations are developed. The general expressions for the stresses and displacements in the laminated composite cylinders are discussed. The closed form solutions based on Classical Shell Theory (CST) and Donnell's (1933) theory are also derived for comparison purposes. Three examples illustrate the effect of radius-to-thickness ratio, coupling and stacking sequence. The results show that, in general, CST yields poor stress and displacement distributions for thick-section composite shells, but converges to the exact elasticity solution as the radius-to-thickness ratio increases. It is also shown that Donnell's theory significantly underestimates the stress and displacement response.

Yuan, F. G.↗
NASA NTRS1992

Stiffness and thermoelastic coefficients for composite laminates

Simple analytic expressions are presented for the stiffness and thermoelastic coefficients of composite laminates in terms of the material properties of the individual layers. Expressions for the derivatives of the various coefficients with respect to each of the material properties and fiber orientation angles are also included. For typical high-performance composites, numerical values are given showing the effects of the stacking sequence and the fiber orientation angle of quasi-isotropic and angle-ply laminates on the values of the various coefficients and their sensitivity derivatives. The expressions for the thermoelastic coefficients and the sensitivity derivatives are given herein for the first time.

Noor, Ahmed K.↗
NASA NTRS1991

Fracture analysis of transverse crack-tip and free-edge delamination in laminated composites

A shear deformation model including hygrothermal effects is developed for the analysis of local delaminations originating from transverse cracks in 90-deg plies located in and around the laminate midplane. A sublaminate approach is used and the model is applied to (+/- 25/90n)s T300/934 graphite/epoxy laminates for n values between 0.5 and 8, along with previously developed edge-delamination shear-deformation models. Critical loads and delamination modes are identified and compared with experimental results. Hygrothermal effects are included in all the models to make the comparisons realistic.

Armanios, Erian A.↗
NASA NTRS1991

Delamination analysis of tapered laminated composites under tensile loading

A study was conducted to analyze tapered composite laminates under tensile loading. A tapered construction made of S2/SP250 glass/epoxy laminate was used to achieve a thickness reduction using three consecutive dropped plies over a distance of 60 ply thicknesses. The principle of minimum complementary potential energy was used to determine interlaminar stresses. The interlaminar peel stress distribution shows a higher tensile intensity at the taper/thin portion juncture. The total strain energy release rate is determined using a simplified membrane model. Results are compared with a finite element simulation.

Armanios, Erian A.↗
NASA NTRS1992

Verification of a 2-D to 3-D global/local finite element method for symmetric laminates

A two-dimensional to three-dimensional global/local finite element analysis technique, as applied to both cross-ply and general symmetric composite laminates, is presented. In particular, the local areas of interest for the laminate are the straight free edge in proximity to a hole and curved free edge of a hole. Verification of this technique was accomplished by comparison to results from complete three-dimensional finite element analyses. Examination of the interlaminar stress fields and the amount of time for their determination indicated that the global/local method developed yielded accurate results and was more efficient.

Thompson, Danniella M.↗
NASA NTRS1992

Bending response of thin-walled laminated composite cylinders

Closed-form linear response solutions are presented for the bending response of thin-walled laminated composite cylinders. Displacements and intralaminar stresses are computed using the Donnell theory, and interlaminar stresses are computed using 3D equilibrium equations of elasticity. Laminates with stacking sequencies and an aluminum cylinder are used to illustrate the displacement response, intralaminar, and interlaminar stresses for two length-to-radius ratios of 1 and 5. Data obtained indicate that if only fiber breakage causes ultimate failure, material failure will occur on the compression side of the cylinders, at the ends, due to fiber compression. The interlaminar normal stress is an order of magnitude smaller than the interlaminar shear stresses and it is not likely to initiate failure.

Fuchs, H. P.↗
NASA NTRS1993

Effects of thermal and mechanical fatigue on the flexural strength of G40-600/PMR-15 cross-ply laminates

The effects of thermal and mechanical fatigue on the flexural strength of G40-600/PMR-15 cross-ply laminates with ply orientation of (0(2),90(2))2S and (90(2),0(2))2S are examined. The relative importance of shear and tensile stresses is examined by varying the span-to-depth ratios of flexural test specimens from 8 to 45. Acoustic emission signals are measured during the flexural tests in order to monitor the initiation and growth of damage. Optical microscopy is used to examine specimens for resin cracking, delamination, and fiber breaks after testing. Transverse matrix cracks and delaminations occur in all specimens, regardless of ply orientation, span-to-depth ratio, or previous exposure of specimens to thermal and mechanical fatigue. A small amount of fiber tensile fracture occurs in the outer 0 deg ply of specimens with high span-to-depth ratios. Because of the complex failure modes, the flexural test results represent the 'apparent' strengths rather than the true flexural or shear strengths for these cross-ply laminates. Thermal cycling of specimens prior to flexural testing does not reduce the apparent flexural strength or change the mode of failure. However, fewer acoustic events are recorded at all strains during flexural testing of specimens exposed to prior thermal cycling. High temperature thermal cycling (32 to 260 C, 100 cycles) causes a greater reduction in acoustic events than low temperature thermal cycling (-85 to +85 C, 500 cycles). Mechanical cycling (0 to 50 percent of the flexural strength, 100 cycles) has a similar effect, except that acoustic events are reduced only at strains less than the maximum strain applied during flexural fatigue.

Roberts, Gary D.↗
NASA NTRS1993

Metal Matrix Laminate Tailoring (MMLT) code: User's manual

The User's Manual for the Metal Matrix Laminate Tailoring (MMLT) program is presented. The code is capable of tailoring the fabrication process, constituent characteristics, and laminate parameters (individually or concurrently) for a wide variety of metal matrix composite (MMC) materials, to improve the performance and identify trends or behavior of MMC's under different thermo-mechanical loading conditions. This document is meant to serve as a guide in the use of the MMLT code. Detailed explanations of the composite mechanics and tailoring analysis are beyond the scope of this document, and may be found in the references. MMLT was developed by the Structural Mechanics Branch at NASA Lewis Research Center (LeRC).

Murthy, P. L. N.↗
NASA NTRS1991

Developments in impact damage modeling for laminated composite structures

Damage tolerance is the most critical technical issue for composite fuselage structures studied in the Advanced Technology Composite Aircraft Structures (ATCAS) program. The objective here is to understand both the impact damage resistance and residual strength of the laminated composite fuselage structure. An understanding of the different damage mechanisms which occur during an impact event will support the selection of materials and structural configurations used in different fuselage quadrants and guide the development of analysis tools for predicting the residual strength of impacted laminates. Prediction of the damage state along with the knowledge of post-impact response to applied loads will allow for engineered stacking sequencies and structural configurations; intelligent decisions on repair requirements will also result.

Dost, Ernest F.↗