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At least 325 records · Page 18

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NASA NTRS1989

Effects of fatigue and environment on residual strengths of center-cracked graphite/epoxy buffer strip panels

The effects of fatigue, moisture conditioning, and heating on the residual tension strengths of center-cracked graphite/epoxy buffer strip panels were evaluated using specimens made with T300/5208 graphite epoxy in a 16-ply quasi-isotropic layup, with two different buffer strip materials, Kevlar-49 or S-glass. It was found that, for panels subjected to fatigue loading, the residual strengths were not significantly affected by the fatigue loading, the number of repetitions of the loading spectrum, or the maximum strain level. The moisture conditioning reduced the residual strengths of the S-glass buffer strip panels by 10 to 15 percent below the ambient results, but increased the residual strengths of the Kevlar-49 buffer strip panels slightly. For both buffer strip materials, the heat increased the residual strengths of the buffer strip panels slightly over the ambient results.

Bigelow, Catherine A.↗

NASA NTRS1987

Superplastic forming/weld-brazing of Ti-6Al-4V panels for enhanced structural efficiency

An account is given of the results obtained by NASA-sponsored studies which exploit the processing advantages of superplastic forming (SPF) and joining by weld-brazing (WB) for the fabrication of Ti-alloy skin-stiffened structural components. The buckling strengths of panels with stiffeners having the geometry of a conventional hat shape are compared with those of panels having advanced-geometry stiffereners. Both single- and multiple-stiffened panels have been evaluated at room temperature; the SPF/WB panels with advanced-geometry stiffeners developed up to 60-percent higher buckling strength than panels with conventional-geometry stiffeners. Selective stiffener reinforcement with a boron-reinforced Al-matrix MMC was also evaluated.

Royster, Dick M.↗

NASA NTRS1988

Low temperature optical testing of CFRP telescope panels

Since 1984, low temperature optical tests were made of very lightweight mirror panels for use in balloon and space infrared and submillimeter telescopes. In order to accomplish this testing, an ambient pressure 0.5 meter test chamber operating from 20 to -80 C, developed techniques for measuring non-optical quality mirrors with phase modulation 10.6 micron interferometry, and created the interferogram reduction program. During the course of the program, nineteen mirrors from four manufactures were tested: carbon fiber reinforced plastic (CFRP) aluminum honeycomb sandwich panel mirrors, a CFRP sandwich panel with an added glass facesheet, and carbon fiber reinforced glass panels. The results of the panel development and test program are summarized.

Hoffmann, William F.↗

NASA NTRS1988

SiO overcoating and polishing of CFRP telescope panels

Development of carbon fiber reinforces plastic (CFRP) panel overcoating and polishing is structured in two parts. The first part utilized a short series of experiments to determine the feasibility of overcoating and polishing CFRP panels, and the second part employes a systematic approach to optimize techniques learned. Questions which required answers in the initial investigation are summarized. Tests were performedin the Steward Observatory's 2.2 Meter Vacuum Coating Chamber and began with 3 cm square pieces of CFRP facesheet material. Next, a 10 cm square and one-inch-thick CFPR-Aluminum core panel was tested. Tests were then conducted on a 0.5-meter-square Dornier panel (QUAD 4) with CFRP facesheets on two-inch aluminum Flexcore. To complete the initial study, a previously characterized 0.5 m Dornier panel (QUAD 23) was coated and hand polished. The mirror's optical performance was not affected by the SiO coating.

Woida, Patrick↗

NASA NTRS1990

Numerical and experimental simulations of the postbuckling response of laminated anisotropic panels

Numerical simulations are used in conjuction with experiments to study the buckling and postbuckling responses and failure initiation of flat unstiffened composite panels. The numerical simulations are conducted using two-dimensional shear-flexible finite elements. The effect of the laminate stacking sequence on the buckling and postbuckling responses is studied. Correlation between numerical and experimental results is good through buckling, but the numerical models overestimate the postbuckling stiffness of the panels. To explain the discrepancies in the postbuckling stiffnesses, analytic sensitivity derivatives are calculated and are used to study the sensitivity of the buckling and postbuckling responses to variations in different material and lamination parameters of the panel. Experimental results indicate that failure occurs along a nodal line. Numerical results show that the location of failure initiation corresponds to that of the maximum transverse shear strain energy density in the panel, which occurs at the edge of the panel at a nodal line.

Noor, Ahmed K.↗

NASA NTRS1991

Making Large Suction Panels For Laminar-Flow Control

Perforated titanium panels used to identify and resolve issues related to manufacture. Recently, relatively large suction panels with aerodynamically satisfactory surface perforations and with surface contours and smoothness characteristics necessary for Laminar-Flow Control (LFC) designed, fabricated, and tested. Requirements of production lines for commercial transport airplanes carefully considered in development of panels. Sizes of panels representative of what is used on wing of commercial transport airplane. Tests of perforated panels in transonic wind tunnel demonstrated aerodynamic stability at flight mach numbers.

Maddalon, Dal V.↗

NASA NTRS1990

Panel methods: An introduction

Panel methods are numerical schemes for solving (the Prandtl-Glauert equation) for linear, inviscid, irrotational flow about aircraft flying at subsonic or supersonic speeds. The tools at the panel-method user's disposal are (1) surface panels of source-doublet-vorticity distributions that can represent nearly arbitrary geometry, and (2) extremely versatile boundary condition capabilities that can frequently be used for creative modeling. Panel-method capabilities and limitations, basic concepts common to all panel-method codes, different choices that were made in the implementation of these concepts into working computer programs, and various modeling techniques involving boundary conditions, jump properties, and trailing wakes are discussed. An approach for extending the method to nonlinear transonic flow is also presented. Three appendices supplement the main test. In appendix 1, additional detail is provided on how the basic concepts are implemented into a specific computer program (PANAIR). In appendix 2, it is shown how to evaluate analytically the fundamental surface integral that arises in the expressions for influence-coefficients, and evaluate its jump property. In appendix 3, a simple example is used to illustrate the so-called finite part of the improper integrals.

Erickson, Larry L.↗

NASA NTRS1991

Large Deformation Behavior of Long Shallow Cylindrical Composite Panels

An exact solution is presented for the large deformation response of a simply supported orthotropic cylindrical panel subjected to a uniform line load along a cylinder generator. The cross section of the cylinder is circular and deformations up to the fully snapped through position are investigated. The orthotropic axes are parallel to the generator and circumferential directions. The governing equations are derived using laminated plate theory, nonlinear strain-displacement relations, and applying variational principles. The response is investigated for the case of a panel loaded exactly at midspan and for a panel with the load offset from midspan. The mathematical formulation is one dimensional in the circumferential coordinate. Solutions are obtained in closed-form. An experimental apparatus was designed to load the panels. Experimental results of displacement controlled tests performed on graphite-epoxy curved panels are compared with analytical predictions.

Carper, Douglas M.↗

NASA NTRS1989

Sound transmission loss of integrally damped, curved panels

Results are reported from acoustic transmission-loss measurements on 13 curved Al and composite aircraft-type panels (height 0.81 m, arc length 1.32 m, and curvature radius 2.29 m) without and with integral damping of various types. The fabrication of the panels and the test procedures are described and illustrated with photographs, and the results are presented in graphs. It is found that the loss of a curved panel exceeds the mass-law-predicted loss for a flat panel of the same material and thickness at frequencies below the ring frequency. At higher frequencies, the curved-panel loss is lower than the mass-law loss, being proportional to 20 time the log of thickness. Integral damping is found to be effective both below and above the ring frequency, but different mechanisms are responsible in each case.

Grosveld, Ferdinand W.↗

NASA NTRS1991

Detailed analysis and test correlation of a stiffened composite wing panel

Nonlinear finite element analysis techniques are evaluated by applying them to a realistic aircraft structural component. A wing panel from the V-22 tiltrotor aircraft is chosen because it is a typical modern aircraft structural component for which there is experimental data for comparison of results. From blueprints and drawings supplied by the Bell Helicopter Textron Corporation, a very detailed finite element model containing 2284 9-node Assumed Natural-Coordinate Strain (ANS) elements was generated. A novel solution strategy which accounts for geometric nonlinearity through the use of corotating element reference frames and nonlinear strain displacements relations is used to analyze this detailed model. Results from linear analyses using the same finite element model are presented in order to illustrate the advantages and costs of the nonlinear analysis as compared with the more traditional linear analysis. Strain predictions from both the linear and nonlinear stress analyses are shown to compare well with experimental data up through the Design Ultimate Load (DUL) of the panel. However, due to the extreme nonlinear response of the panel, the linear analysis was not accurate at loads above the DUL. The nonlinear analysis more accurately predicted the strain at high values of applied load, and even predicted complicated nonlinear response characteristics, such as load reversals, at the observed failure load of the test panel. In order to understand the failure mechanism of the panel, buckling and first ply failure analyses were performed. The buckling load was 17 percent above the observed failure load while first ply failure analyses indicated significant material damage at and below the observed failure load.

Davis, D. Dale, Jr.↗

NASA NTRS1992

Large-amplitude finite element flutter analysis of composite panels in hypersonic flow

Consideration is given to a finite-element approach for determining the nonlinear flutter characteristics of 3D thin laminated composite panels using the full third-order-piston transverse-loading aerodynamic theory. The unsteady hypersonic aerodynamic theory and the von Karman large-deflection-plate theory are used to formulate the aeroelasticity problem. Nonlinear flutter analyses are performed to assess the influence of the higher-order aerodynamic theory on the structure's limit-cycle amplitude and the dynamic pressure of the flow velocity. A solution procedure is presented to solve the nonlinear panel flutter and large-amplitude free-vibration finite-element equations. Nonlinear flutter analyses are performed for different boundary support-conditions and for various system parameters. Linear finite-element flutter for isotropic and composite panels and large-amplitude isotropic panel flutter results are compared with existing classical solutions. The large-amplitude panel flutter results using the full third-order piston aerodynamic theory are presented to assess the influence of the nonlinear aerodynamic theory.

Gray, Carl E., Jr.↗

NASA NTRS1991

Thermal-structural panel buckling tests

A titanium-matrix-composite (TMC) hat-stiffened panel of 61 cm sq area and 3.175 cm thick was nondestructively tested to 649 C to examine its buckling characteristics. Compressive loads were applied to the panel in a 978.6 kN uniaxial load frame system. High-temperature testing was performed using quartz lamp heating. A single-strain-age force/stiffness buckling prediction technique was developed to predict panel buckling loads. For the monolithic panel, these test predictions correlated within 10 percent with a finite-element buckling analyses performed elsewhere. Comparisons between force/stiffness predictions and analyses for the TMC panel are in progress.

Thompson, Randolph C.↗

NASA NTRS1992

Study of compression-loaded and impact-damaged structurally efficient graphite-thermoplastic trapezoidal-corrugation sandwich and semisandwich panels

The structural efficiency of compression-loaded trapezoidal-corrugation sandwich and semisandwich 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 semisandwich 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 3000 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. The results of testing impact damage panels are also discussed.

Jegley, Dawn C.↗

NASA NTRS1992

Reliability of stiffened structural panels: Two examples

The reliability of two graphite-epoxy stiffened panels that contain uncertainties is examined. For one panel, the effect of an overall bow-type initial imperfection is studied. The size of the bow is assumed to be a random variable. The failure mode is buckling. The benefits of quality control are explored by using truncated distributions. For the other panel, the effect of uncertainties in a strain-based failure criterion is studied. The allowable strains are assumed to be random variables. A geometrically nonlinear analysis is used to calculate a detailed strain distribution near an elliptical access hole in a wing panel that was tested to failure. Calculated strains are used to predict failure. Results are compared with the experimental failure load of the panel.

Stroud, W. Jefferson↗

NASA NTRS1993

SCBUCKLE user's manual: Buckling analysis program for simple supported and clamped panels

The program SCBUCKLE calculates the buckling loads and mode shapes of cylindrically curved, rectangular panels. The panel is assumed to have no imperfections. SCBUCKLE is capable of analyzing specially orthotropic symmetric panels (i.e., A(sub 16) = A(sub 26) = 0.0, D(sub 16) = D(sub 26) = 0.0, B(sub ij) = 0.0). The analysis includes first-order transverse shear theory and is capable of modeling sandwich panels. The analysis supports two types of boundary conditions: either simply supported or clamped on all four edges. The panel 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 applied loads can be divided into two parts: a preload component; and a variable (eigenvalue-dependent) component. The analysis is based on the modified Donnell's equations for shallow shells. The governing equations are solved by Galerkin's method.

Cruz, Juan R.↗

NASA NTRS1993

Shear and compression buckling analysis for anisotropic panels with centrally located elliptical cutouts

An approximate analysis for buckling of biaxial- and shear-loaded anisotropic panels with centrally located elliptical cutouts is presented in the present paper. The analysis is composed of two parts, a prebuckling analysis and a buckling analysis. The prebuckling solution is determined using Lekhnitskii's complex variable equations of plane elastostatics combined with a Laurent series approximation and a boundary collocation method. The buckling solution is obtained using the principle of minimum potential energy. A by-product of the minimum potential energy equation is an integral equation which is solved using Gaussian quadrature. Comparisons with documented experimental results and finite element analyses indicate that the approximate analysis accurately predicts the buckling loads of square biaxial- and shear-loaded panels having elliptical cutouts with major axes up to sixty percent of the panel width. Results of a parametric study are presented for shear- and compression-loaded rectangular anisotropic panels with elliptical cutouts. The effects of panel aspect ratio, cutout shape, cutout size, cutout orientation, laminate anisotropy, and combined loading on the buckling load are examined.

Britt, V. O.↗

NASA NTRS1991

Effect of temperature on nonlinear two-dimensional panel flutter using finite elements

A finite element formulation and solution procedures is presented for limit-cycle motions of 2D panels subjected simultaneously to thermal and aerodynamic loads. The thermal load is described by a steady-state temperature distribution and the quasi-steady first-order piston theory is used for aerodynamic pressure. The von Karman nonlinear strain-displacement relationship is used for the large panel deflections. Three temperature distributions are evaluated: (1) uniform; (2) symmetric sinusoidal varying temperature along panel length; and (3) linearly varying temperature through panel thickness. The influence of these three temperature distributions on limit-cycle motions and flutter boundaries of a simply supported 2D panel is investigated.

Mei, Chuh↗

NASA NTRS1946

Compressive Strength of 24S-T Aluminum-alloy Flat Panels with Longitudinal Formed Hat-section Stiffeners

Results are presented for a part of a test program on 24S-T aluminum alloy flat compression panels with longitudinal formed hat-section stiffeners. This part of the program is concerned with panels in which the thickness of the stiffener materials is 0.625 times the skin thickness. The results, presented in tabular and graphical form, show the effect of the relative dimensions of the panel on the buckling stress and the average stress at maximum load. Comparative envelope curves are presented for hat-stiffened and Z-stiffened panels having the same ratio of stiffener thickness to sheet thickness. These curves provide some indication of the relative structural efficiencies of the two types of panel.

Schuette, Evan H↗