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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.

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At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

A numerical solution of Duffing's equations including the prediction of jump phenomena

Numerical methodology for the solution of Duffing's differential equation is presented. Algorithms for the prediction of multiple equilibrium solutions and jump phenomena are developed. In addition, a filtering algorithm for producing steady state solutions is presented. The problem of a rigidly clamped circular plate subjected to cosinusoidal pressure loading is solved using the developed algorithms (the plate is assumed to be in the geometrically nonlinear range). The results accurately predict regions of solution multiplicity and jump phenomena.

Moyer, E. T., Jr.↗
NASA NTRS1990

Buckling and postbuckling behavior of compression-loaded isotropic plates with cutouts

An experimental study of the buckling and postbuckling behavior of square and rectangular compression loaded aluminum plates with centrally located circular, square, and elliptical cutouts is presented. Experimental results indicate that the plates exhibit overall trends of increasing buckling strain and decreasing initial postbuckling stiffness with increasing cutout width. Corresponding plates with circular and square cutouts of the same width buckle at approximately the same strain level, and exhibit approximately the same initial postbuckling stiffness. Results show that the reduction in initial postbuckling stiffness due to a cutout generally decreases as the plate aspect ratio increases. Other results presented indicate that square plates with elliptical cutouts having a large cutout-width-to-plate-width ratio generally lose prebuckling and initial postbuckling stiffness as the cutout height increases. However, the plates buckle at essentially the same strain level. Results also indicate that postbuckling stiffness is more sensitive to changes in elliptical cutout height than are prebuckling stiffness and buckling strain.

Nemeth, Michael P.↗
NASA NTRS1990

Buckling and postbuckling behavior of compression-loaded isotropic plates with cutouts

An experimental study of buckling and postbuckling behavior of square and rectangular compression-loaded aluminum plates with centrally located circular, square, and elliptical cutouts is presented. Experimental results indicate that the plates exhibit overall trends of increasing buckling strain and decreasing initial postbuckling stiffness with increasing cutout width. Corresponding plates with circular and square cutouts of the same width buckle at approximately the same strain level, and exhibit approximately the same initial postbuckling stiffness. Results show that the reduction in initial postbuckling stiffness due to a cutout generally decreases as the plate aspect ratio increases. Other results presented in this paper indicate that square plates with elliptical cutouts having a large cutout-width-to-plate-width ratio generally lose prebuckling and initial postbuckling stiffness as the cutout height increases. However, the plates buckle at essentially the same strain level. Results also indicate that postbuckling stiffness is more sensitive to changes in elliptical cutout height than are prebuckling stiffness and buckling strain.

Nemeth, Michael P.↗
NASA NTRS1930

Bending Tests of Metal Monocoque Fuselage Construction

This thesis presents a study of the bending stress in smooth skin, aluminum alloy, true monocoque fuselage sections of varying ratio of diameters to thickness. The test specimens - circular, thin-walled cylinder were loaded to give a large bending stress in proportion to the shearing stress, in order to represent the critical section of a fuselage.

Mossman, Ralph W↗
NASA NTRS2015

Damage Arresting Composites

Under NASA's Environmentally Responsible Aviation Project (ERA) the most promising vehicle concepts and technologies that can simultaneously reduce aircraft fuel use, community noise, and emissions are being evaluated. Two key factors to accomplishing these goals are reducing structural weight and moving away from the traditional tube and wing aircraft configuration to a shape that has improved lift and less drag. The hybrid wing body (HWB) configuration produces more lift and less drag by smoothly joining the wings to the center fuselage section so it provides aerodynamic advantages. This shape, however, presents structural challenges with its pressurized, non-circular cabin subjected to aerodynamic flight loads. In the HWB, the structure of the center section where the passenger cabin would be located must support large in-plane loads as well as internal pressure on nearly-flat panels and right-angle joints. This structural arrangement does not lend itself to simple, efficient designs. Traditional aluminum and even state-of-the-art composites do not provide a solution to this challenge.

Jegley, Dawn C.↗
NASA NTRS1992

Analysis of stresses in finite anisotropic panels with centrally located cutouts

A method for analyzing biaxial- and shear-loaded anisotropic rectangular panels with centrally located circular and elliptical cutouts is presented in the present paper. The method is based on Lekhnitskii's complex variable equations of plane elastostatics combined with a boundary collocation method and a Laurent series approximation. Results are presented for anisotropic panels with elliptical cutouts and subjected to combined shear and compression loading. The effects on the stress field of panel aspect ratio, anisotropy, cutout size, and cutout orientation are addressed. Angle-ply laminates, unidirectional off-axis laminates, and ((+ or - 45/0/90)(sub 3))s, ((+ or - 45/0(sub 2))(sub 3))s, and ((+ or - 45/90(sub 2))(sub 3))s laminates are examined.

Britt, Vicki O.↗
NASA NTRS1981

Postbuckling and failure characteristics of selected flat rectangular graphite-epoxy plates loaded in compression

Results of an experimental study of the postbuckling response and failure characteristics of 16- and 24-ply quasi-isotropic and 24-ply orthotropic flat rectangular graphite-epoxy plates loaded in compression are examined. Some of the specimens had circular drilled holes or were subjected to low-speed impact damage. The ratio of failure load to buckling load was found to be higher for specimens with lower initial buckling strains than for specimens with higher initial buckling strains. Some specimens supported more than five times their initial buckling load before failing. The specimens failed along a nodal line of the buckling mode in an induced shear failure mode. Holes and impact damage had a more adverse effect on the postbuckling strength of specimens with high initial buckling strains than on the strength of specimens with lower initial buckling strains.

Starnes, J. H., Jr.↗
NASA NTRS2000

Polar Plate Theory for Orthogonal Anisotropy

Laminated fiber-reinforced (or filamentary) composites are used today for their high strength-to-weight and stiffness-to-weight ratios. However, because of the anisotropic behavior of composites, determining the response on a macroscopic scale is challenging. This is particularly evident in the evaluation of the governing differential equations of a circular disk with the fibers of the lamina oriented with rectilinear orthogonality. This includes any situation involving a composite plate of circular geometry in which out-of-plane displacements due to load are desired, such as fastener pull through loading of a composite plate. Current analysis techniques use numerical methods with rectilinear coordinate systems to solve problems with circular geometry. These analyses over predict plate stiffness by 20% and underpredict failure by 70%. Consequently, there is a need to transform classical composite plate theory to a polar coordinate system. In order to better analyze structures with circular geometries the classical composite plate equations are transformed into the plate equations for a rectilinearly anisotropic composite in polar coordinates. A composite plate is typically a laminate of fibers in rectilinear directions. Subsequent to the lay-tip the necessary geometry is cut out of a rectangular plate. In a similar manner, the derivation of the plate equation starts with the fundamental definitions of strain, displacement and curvature and incorporates the material property angular dependence into the equilibrium equations for a differential polar element. In the transformed state, the stiffness coefficients are no longer constant, adding to the complexity of the governing differential equations. This paper discusses the new derivation and evaluation of the plate equations for a circular composite disk with orthogonal rectilinear anisotropy. The resultant new three partial differential equations, which describe the circular anisotropic plate, can be used to evaluate out-of-plane displacements for given load conditions in design of composite Structures. Without the formation of these unique plate equations, design of such structures is not a precise engineering accomplishment due to the lack of precise design tools. Such structures are over designed to compensate and costly tests need to be performed after production to validate safety. Though impossible to present the mathematical solution within the limits of this paper, the application of the newly derived plate equation in its orthotropic form for a balanced symmetric laminate compared favorably with test results and surpassed the results of standard numerical methods. Within the limitations of the problem foundation the predictive model provides a mechanism for establishing out-of-plane deflection levels for a circular composite plate. Future work in this area should include incorporating the circular composite plate equations in Finite Element Models. The new anisotropic polar plate equations call be utilized as a design tool so that the design more accurately meets the requirements, thus taking full advantage of the weight savings of composites. These accomplishments have not previously been presented to the engineering community nor ever proposed to the IAF. Should the proposed paper be accepted for presentation, attendance is assured.

Bailey, Michelle D.↗
NASA NTRS1985

High discharge rate characteristics of nickel-cadmium batteries for pulse load filtering

Several tests of specially fabricated nickel-cadmium batteries having circular disk type electrodes were considered. These batteries were evaluated as filter elements between a constant current power supply and a five hertz pulsed load demanding approximately twice the power supply current during the load on portion of the cycle. Short tests lasting 10,000 cycles were conducted at up to a 21 C rate and an equivalent energy density of over 40 Joules per pound. In addition, two batteries were subjected to 10 to the 7 charge/discharge cycles, one at a 6.5 C rate and the other at a 13 C rate. Assuming an electrode to battery weight ratio of 0.5, these tests represent an energy density of about 7 and 14 Joules per pound respectively. Energy density, efficiency, capacitance, average voltage, and available capacity were tracked during these tests. After 10 to the 7 cycles, capacity degradation was negligible for one battery and about 20% for the other. Cadmium electrode failure may be the factor limiting lifetime at extremely low depth of discharge cycling. The output was examined and a simple equivalent circuit was proposed.

Gearing, G. M.↗
NASA NTRS1991

A novel concept for a combined-load test apparatus

An original concept for a combined-load test machine is presented in the paper. The concept employs a circular test specimen which may be rotated with respect to the biaxially applied loads. The ability to rotate the specimen and to vary the applied loads allows for limitless combinations of in-plane normal and shear loads. The loads generated in the test section reference frame are calculated from simple Mohr's circle relationships. In this paper, the concept is evaluated for isotropic, orthotropic, and discretely stiffened panels using elasticity and finite element methods. Strengths and weaknesses of the test machine concept are also discussed and evaluated.

Martin, Carl J.↗
NASA NTRS2015

Face Sheet/Core Disbond Growth in Honeycomb Sandwich Panels Subjected to Ground-Air-Ground Pressurization and In-Plane Loading

Typical damage modes in light honeycomb sandwich structures include face sheet/core disbonding and core fracture, both of which can pose a threat to the structural integrity of a component. These damage modes are of particular interest to aviation certification authorities since several in-service occurrences, such as rudder structural failure and other control surface malfunctions, have been attributed to face sheet/core disbonding. Extensive studies have shown that face sheet/core disbonding and core fracture can lead to damage propagation caused by internal pressure changes in the core. The increasing use of composite sandwich construction in aircraft applications makes it vitally important to understand the effect of ground-air-ground (GAG) cycles and conditions such as maneuver and gust loads on face sheet/core disbonding. The objective of the present study was to use a fracture mechanics based approach developed earlier to evaluate the loading at the disbond front caused by ground-air-ground pressurization and in-plane loading. A honeycomb sandwich panel containing a circular disbond at one face sheet/core interface was modeled with three-dimensional (3D) solid finite elements. The disbond was modeled as a discrete discontinuity and the strain energy release rate along the disbond front was computed using the Virtual Crack Closure Technique (VCCT). Special attention was paid to the pressure-deformation coupling which can decrease the pressure load within the disbonded sandwich section significantly when the structure is highly deformed. The commercial finite element analysis software, Abaqus/Standard, was used for the analyses. The recursive pressure-deformation coupling problem was solved by representing the entrapped air in the honeycomb cells as filled cavities in Abaqus/Standard. The results show that disbond size, face sheet thickness and core thickness are important parameters that determine crack tip loading at the disbond front. Further, the pressure-deformation coupling was found to have an important load decreasing effect [6]. In this paper, a detailed problem description is provided first. Second, the analysis methodology is presented. The fracture mechanics approach used is described and the specifics of the finite element model, including the fluid-filled cavities, are introduced. Third, the initial model verification and validation are discussed. Fourth, the findings from a closely related earlier study [6] are summarized. These findings provided the basis for the current investigation. Fifth, an aircraft ascent scenario from 0 to 12192 m (0 to 40000 ft) is considered and the resulting crack tip loading at the disbond front is determined. In-plane loading to simulate maneuvers and gust conditions are also considered. Sixth, the results are shown for a curved panel, which was used to simulate potential fuselage applications. Finally, a brief summary of observations is presented and recommendations for improvement are provided.

Chen, Zhi M.↗
NASA NTRS1987

An analytical study of the effects of transverse shear deformation and anisotropy on buckling loads of laminated cylinders

Buckling loads of thick-walled orthotropic and anisotropic simply supported circular cylinders are predicted using a higher-order transverse-shear deformation theory. A comparison of buckling loads predicted by the conventional first-order transverse-shear deformation theory and the higher-order theory show that the additional allowance for transverse shear deformation has a negligible effect on the predicted buckling loads of medium-thick metallic isotropic cylinders. However, the higher-order theory predicts buckling loads which are significantly lower than those predicted by the first-order transverse-shear deformation theory for certain short, thick-walled cylinders which have low through-the-thickness shear moduli. A parametric study of the effects of ply orientation on the buckling load of axially compressed cylinders indicates that laminates containing 45 degree plies are most sensitive to transverse-shear deformation effects. Interaction curves for buckling loads of cylinders subjected to axial compressive and external pressure loadings indicate that buckling loads due to external pressure loadings are as sensitive to transverse-shear deformation effects as buckling loads due to axial compressive loadings. The effects of anisotropy are important over a much wider range of cylinder geometries than the effects of transverse shear deformation.

Jegley, Dawn C.↗
NASA NTRS1992

Micromechanical combined stress analysis: MICSTRAN, a user manual

Composite materials are currently being used in aerospace and other applications. The ability to tailor the composite properties by the appropriate selection of its constituents, the fiber and matrix, is a major advantage of composite materials. The Micromechanical Combined Stress Analysis (MICSTRAN) code provides the materials engineer with a user-friendly personal computer (PC) based tool to calculate overall composite properties given the constituent fiber and matrix properties. To assess the ability of the composite to carry structural loads, the materials engineer also needs to calculate the internal stresses in the composite material. MICSTRAN is a simple tool to calculate such internal stresses with a composite ply under combined thermomechanical loading. It assumes that the fibers have a circular cross-section and are arranged either in a repeating square or diamond array pattern within a ply. It uses a classical elasticity solution technique that has been demonstrated to calculate accurate stress results. Input to the program consists of transversely isotropic fiber properties and isotropic matrix properties such as moduli, Poisson's ratios, coefficients of thermal expansion, and volume fraction. Output consists of overall thermoelastic constants and stresses. Stresses can be computed under the combined action of thermal, transverse, longitudinal, transverse shear, and longitudinal shear loadings. Stress output can be requested along the fiber-matrix interface, the model boundaries, circular arcs, or at user-specified points located anywhere in the model. The MICSTRAN program is Windows compatible and takes advantage of the Microsoft Windows graphical user interface which facilitates multitasking and extends memory access far beyond the limits imposed by the DOS operating system.

Naik, R. A.↗
NASA NTRS2005

Tow-Steered Panels With Holes Subjected to Compression or Shear Loads

Tailoring composite laminates to vary the fiber orientations within a fiber layer of a laminate to address non-uniform stress states and provide structural advantages such as the alteration of principal load paths has potential application to future low-cost, light-weight structures for commercial transport aircraft. Evaluation of this approach requires the determination of the effectiveness of stiffness tailoring through the use of curvilinear fiber paths in flat panels including the reduction of stress concentrations around the holes and the increase in load carrying capability. Panels were designed through the use of an optimization code using a genetic algorithm and fabricated using a tow-steering approach. Manufacturing limitations, such as the radius of curvature of tows the machine could support, avoidance of wrinkling of fibers and minimization of gaps between fibers were considered in the design process. Variable stiffness tow-steered panels constructed with curvilinear fiber paths were fabricated so that the design methodology could be verified through experimentation. Finite element analysis where each element s stacking sequence was accurately defined is used to verify the behavior predicted based on the design code. Experiments on variable stiffness flat panels with central circular holes were conducted with the panels loaded in axial compression or shear. Tape and tow-steered panels are used to demonstrate the buckling, post-buckling and failure behavior of elastically tailored panels. The experimental results presented establish the buckling performance improvements attainable by elastic tailoring of composite laminates.

Jegley, Dawn C.↗
NASA NTRS2005

Torsion of Noncircular Composite Cylinders

The paper presents a brief overview of the predicted deformation and failure characteristics of noncircular composite cylinders subjected to torsion. Using a numerical analysis, elliptical cylinders with a minor-to-major diameter ratio of 0.7 are considered. Counterpart circular cylinders with the same circumference as the elliptical cylinders are included for comparison. The cylinders are constructed of a medium-modulus graphite-epoxy material in a quasi-isotropic lay-up. Imperfections generated from the buckling mode shapes are included in the initial cross-sectional geometry of the cylinders. Deformations until first fiber failure, as predicted using the maximum stress failure criterion and a material degradation scheme, are presented. For increasing levels of torsion, the deformations of the elliptical cylinders, in the form of wrinkling of the cylinder wall, occur primarily in the flatter regions of the cross section. By comparison the wrinkling deformations of the circular cylinders are more uniformly distributed around the circumference. Differences in the initial failure and damage progression and the overall torque vs. twist relationship between the elliptical and circular cylinders are presented. Despite differences in the response as the cylinders are being loaded, at first fiber failure the torque and twist for the elliptical and circular cylinders nearly coincide.

Rouse, Marshall↗
NASA NTRS1988

Elasticity effects on cavitation in a squeeze film damper undergoing noncentered circular whirl

Elasticity of the liner and its effects on cavitation were numerically determined for a squeeze film damper subjected to dynamic loading. The loading was manifested as a prescribed motion of the rotor undergoing noncentered circular whirl. The boundary conditions were implemented using Elrod's algorithm which conserves lineal mass flux through the moving cavitation bubble as well as the oil film region of the damper. Computational movies were used to analyze the rapidly changing pressures and vapor bubble dynamics throughout the dynamic cycle for various flexibilities in the damper liner. The effects of liner elasticity on cavitation were only noticeable for the intermediate and high values of viscosity used in this study.

Brewe, David E.↗
NASA NTRS1978

The effect of impact damage and circular holes on the compressive strength of a graphite-epoxy laminate

Specimens were impacted by 1.27-cm-diameter aluminum spheres with speeds ranging from 52 to 101 m/s. Some specimens were impacted without any applied compressive load and then loaded to failure to determine their residual strength. Other specimens were loaded to a prescribed axial compressive strain and impacted while at that applied load. Loaded specimens that did not fail catastrophically on impact were subsequently loaded to failure to determine their residual strength. Low-velocity impact damage was found to degrade seriously the laminate static compressive strength. Low-strain compression-compression cyclic loading was found to degrade further the compressive strength of impact-damaged specimens. Specimens with circular holes having diameters up to a third of the specimen width were loaded to failure in compression. It was found that circular holes can also degrade the static compressive strength of the laminate. The effects of circular holes and impact damage on the compressive strength of the laminate are compared.

Starnes, J. H., Jr.↗