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At least 145 records · Page 8

Impact-damaged graphite-thermoplastic trapezoidal-corrugation sandwich and semi-sandwich panels

The results of a study of the effects of impact damage on compression-loaded trapezoidal-corrugation sandwich and semi-sandwich graphite-thermoplastic panels are presented. 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 in this study. Panels were designed, fabricated and tested. The panels were made using the manufacturing process of thermoforming, a less-commonly used technique for fabricating composite parts. Experimental results for unimpacted control panels and panels subjected to impact damage prior to loading are presented. Little work can be found in the literature about these configurations of thermoformed panels.

Jegley, D.↗

Compressive and shear buckling analysis of metal matrix composite sandwich panels under different thermal environments

Combined inplane compressive and shear buckling analysis was conducted on flat rectangular sandwich panels using the Raleigh-Ritz minimum energy method with a consideration of transverse shear effect of the sandwich core. The sandwich panels were fabricated with titanium honeycomb core and laminated metal matrix composite face sheets. The results show that slightly slender (along unidirectional compressive loading axis) rectangular sandwich panels have the most desirable stiffness-to-weight ratios for aerospace structural applications; the degradation of buckling strength of sandwich panels with rising temperature is faster in shear than in compression; and the fiber orientation of the face sheets for optimum combined-load buckling strength of sandwich panels is a strong function of both loading condition and panel aspect ratio. Under the same specific weight and panel aspect ratio, a sandwich panel with metal matrix composite face sheets has much higher buckling strength than one having monolithic face sheets.

Ko, William L.↗

On acoustic radiation from a vibrating panel

An experimental and numerical study of the response and radiation from an acoustically loaded aircraft panel is presented. In the experiment, the panel is excited by a normally incident, harmonic wave. Measurements of the panel response and the resulting transmitted pressure are made in both the near- and acoustic far-fields. The numerical computations model the experiment and in particular account for the full coupling between the panel and the surrounding three dimensional acoustic fluid. The results demonstrate that for a sufficiently high excitation level, the panel response becomes nonlinear. The nonlinearity is characterized by the appearance of harmonics and subharmonics in the power spectral densities of the panel motion and consequently in the resulting acoustic radiation. The primary characteristic of the far-field acoustic radiation is the increase in harmonic content relative to the fundamental and subharmonics with increasing distance from the panel. This is shown to be a result of linear and weakly nonlinear wave propagation effects. The numerical results show that the radiated far-field pressure is strongly dependent on the position of the measurement point with respect to the panel center. The experimental and numerical results are in good qualitative agreement.

Frendi, Abdelkader↗

Control of Thermal Deflection, Panel Flutter and Acoustic Fatigue at Elevated Temperatures Using Shape Memory Alloys

The High Speed Civil Transport (HSCT) will have to be designed to withstand high aerodynamic load at supersonic speeds (panel flutter) and high acoustic load (acoustic or sonic fatigue) due to fluctuating boundary layer or jet engine acoustic pressure. The thermal deflection of the skin panels will also alter the vehicle's configuration, thus it may affect the aerodynamic characteristics of the vehicle and lead to poor performance. Shape memory alloys (SMA) have an unique ability to recover large strains completely when the alloy is heated above the characteristic transformation (austenite finish T(sub f)) temperature. The recovery stress and elastic modulus are both temperature dependent, and the recovery stress also depends on the initial strain. An innovative concept is to utilize the recovery stress by embedding the initially strained SMA wire in a graphite/epoxy composite laminated panel. The SMA wires are thus restrained and large inplane forces are induced in the panel at elevated temeperatures. By embedding SMA in composite panel, the panel becomes much stiffer at elevated temperatures. That is because the large tensile inplane forces induced in the panel from the SMA recovery stress. A stiffer panel would certainly yield smaller dynamic responses.

Mei, Chuh↗

Evaluation of a Composite Sandwich Fuselage Side Panel with Damage and Subjected to Internal Pressure

The results from an experimental and analytical study of a composite sandwich fuselage side panel for a transport aircraft are presented. The panel has two window cutouts and three frames, and has been evaluated with internal pressure loads that generate biaxial tension loading conditions. Design limit load and design ultimate load tests have been performed on the graphite-epoxy sandwich panel with the middle frame removed to demonstrate the suitability of this two-frame design for supporting the prescribed biaxial loading conditions with twice the initial frame spacing of 20 inches. The two-frame panel was damaged by cutting a notch that originates at the edge of a cutout and extends in the panel hoop direction through the window-belt area. This panel with a notch was tested in a combined-load condition to demonstrate the structural damage tolerance at the design limit load condition. The two panel configurations successfully satisfied all design load requirements in the experimental part of the study, and the three-frame and two-frame panel responses are fully explained by the analysis results. The results of this study suggest that there is potential for using sandwich structural concepts with greater than the usual 20-in.-wide frame spacing to further reduce aircraft fuselage structural weight.

Rouse, Marshall↗

Some Measurements of Noise Transmission and Stress Response of a 0.020-Inch Duralumin Panel in the Presence of Air Flow

Noise transmission measurements were made for a 0.020-inch panel with and without air flow on its surface. Tests were conducted with both an absorbent and reverberant chamber behind the panel. Panel stresses for some of these tests were also determined. Noise spectra obtained inside the absorbent chamber with flow attached and flow not attached to the panel appeared to contain several peaks corresponding in frequency to panel vibration modes. These peaks were notably absent when the chamber was reverberant. The noise reduction through the test panel measured with the aid of an absorbent chamber for the flow-not-attached case is in general agreement with values predicted by the theoretical weight law, which assumes negligible panel stiffness. Corresponding data for the flow-attached case do not follow the weight law but rather indicate less noise reduction at the high frequencies. The main stress responses of the panel without air flow occurred at its fundamental vibration mode. In the presence of air flow the main response occurs in a vibration mode having a node line perpendicular to the direction of air flow.

Kantarges, George T.↗

Open-Mode Debonding Analysis of Curved Sandwich Panels Subjected to Heating and Cryogenic Cooling on Opposite Faces

Increasing use of curved sandwich panels as aerospace structure components makes it vital to fully understand their thermostructural behavior and identify key factors affecting the open-mode debonding failure. Open-mode debonding analysis is performed on a family of curved honeycomb-core sandwich panels with different radii of curvature. The curved sandwich panels are either simply supported or clamped, and are subjected to uniform heating on the convex side and uniform cryogenic cooling on the concave side. The finite-element method was used to study the effects of panel curvature and boundary condition on the open-mode stress (radial tensile stress) and displacement fields in the curved sandwich panels. The critical stress point, where potential debonding failure could initiate, was found to be at the midspan (or outer span) of the inner bonding interface between the sandwich core and face sheet on the concave side, depending on the boundary condition and panel curvature. Open-mode stress increases with increasing panel curvature, reaching a maximum value at certain high curvature, and then decreases slightly as the panel curvature continues to increase and approach that of quarter circle. Changing the boundary condition from simply supported to clamped reduces the magnitudes of open-mode stresses and the associated sandwich core depth stretching.

Ko, William L.↗

Active Control of Turbulent Boundary Layer Induced Sound Radiation from Multiple Aircraft Panels

The objective of this work is to experimentally investigate active structural acoustic control of turbulent boundary layer (TBL) induced sound radiation from multiple panels on an aircraft sidewall. One possible approach for controlling sound radiation from multiple panels is a multi-input/multi-output scheme which considers dynamic coupling between the panels. Unfortunately, this is difficult for more than a few panels, and is impractical for a typical aircraft which contains several hundred such panels. An alternative is to implement a large number of independent control systems. Results from the current work demonstrate the feasibility of reducing broadband radiation from multiple panels utilizing a single-input/single-output (SISO) controller per bay, and is the first known demonstration of active control of TBL induced sound radiation on more than two bays simultaneously. The paper compares sound reduction for fully coupled control of six panels versus independent control on each panel. An online adaptive control scheme for independent control is also demonstrated. This scheme will adjust for slow time varying dynamic systems such as fuselage response changes due to aircraft pressurization, etc.

Gibbs, Gary P.↗

Size Effects in Impact Damage of Composite Sandwich Panels

Panel size has a large effect on the impact response and resultant damage level of honeycomb sandwich panels. It has been observed during impact testing that panels of the same design but different panel sizes will show large differences in damage when impacted with the same impact energy. To study this effect, a test program was conducted with instrumented impact testing of three different sizes of sandwich panels to obtain data on panel response and residual damage. In concert with the test program. a closed form analysis method was developed that incorporates the effects of damage on the impact response. This analysis method will predict both the impact response and the residual damage of a simply-supported sandwich panel impacted at any position on the panel. The damage is incorporated by the use of an experimental load-indentation curve obtained for the face-sheet/honeycomb and indentor combination under study. This curve inherently includes the damage response and can be obtained quasi-statically from a rigidly-backed specimen or a specimen with any support conditions. Good correlation has been obtained between the test data and the analysis results for the maximum force and residual indentation. The predictions can be improved by using a dynamic indentation curve. Analyses have also been done using the MSC/DYTRAN finite element code.

Dobyns, Alan↗

Improving Strength of Postbuckled Panels Through Stitching

The behavior of blade-stiffened graphite-epoxy panels with impact damage is examined to determine the effect of adding through-the-thickness stitches in the stiffener flange-to-skin interface. The influence of stitches is evaluated by examining buckling and failure for panels with failure loads up to 3.5 times greater than buckling loads. Analytical and experimental results from four configurations of panel specimens are presented. For each configuration, two panels were manufactured with skin and flanges held together with through-the-thickness stitches introduced prior to resin infusion and curing and one panel was manufactured with no stitches holding the flange to the skin. No mechanical fasteners were used for the assembly of any of these panels. Panels with and without low-speed impact damage were loaded to failure in compression. Buckling and failure modes are discussed. Stitching had little effect on buckling loads but increased the failure loads of impact-damaged panels by up to 30%.

Jegley, Dawn C.↗

Space deployable domed solar concentrator with foldable panels and hinge therefor

A space deployable solar energy concentrator is formed of a dome-shaped arrangement of compactly stowable flat panel segments mounted on a collapsible, space-deployable support structure of interconnected linear components. The support structure is comprised of a plurality of tensioned, curvilinear edge strips which extend in a radial direction from a prescribed vertex of a surrounding umbrella-like framework of radially extending rib members. Between a respective pair of radially-extending, curvilinear edge strips an individual wedge-shaped panel section is formed of a plurality of multi-segment lens panel strips each of which is supported in tension between the pair of edge strips by a pair of circumferentially extending catenary cord members connected to a pair of ribs of the surrounding umbrella-like framework. A respective lens panel strip is comprised of a plurality of flat, generally rectangular-shaped, energy-directing panels arranged side-by-side in the circumferential direction of the dome. Adjacent panels are interconnected by flexible U-shaped hinges which overlap opposing edges of adjacent panels and engage respective cylindrically-shaped, load distribution bars that slide within the flexible hinges. Because each U-shaped hinge is flexible, it is permitted to shift in the circumferential direction of the panel section to facilitate stowage and deployment of the dome.

Grayson, Fred G.↗

Buckling Testing and Analysis of Honeycomb Sandwich Panel Arc Segments of a Full-Scale Fairing Barrel: 8-ply Out-of-Autoclave Facesheets - Part 3

Four honeycomb sandwich panels, representing 1/16th arc segments of a 10 m diameter barrel section of the heavy lift launch vehicle, were manufactured under the NASA Composites for Exploration program and the NASA Constellation Ares V program. Two configurations were chosen for the panels: 6-ply facesheets with 1.125 in. honeycomb core and 8-ply facesheets with 1.000 in. honeycomb core. Additionally, two separate carbon fiber/epoxy material systems were chosen for the facesheets: inautoclave IM7/977-3 and out-of-autoclave T40-800B/5320-1. Smaller 3- by 5-ft panels were cut from the 1/16th barrel sections. These panels were tested under compressive loading at the NASA Langley Research Center. Furthermore, linear eigenvalue and geometrically nonlinear finite element analyses were performed to predict the compressive response of the 3- by 5-ft panels. This manuscript summarizes the experimental and analytical modeling efforts pertaining to the panel composed of 8-ply, T40-800B/5320-1 facesheets (referred to as Panel C). To improve the robustness of the geometrically nonlinear finite element model, measured surface imperfections were included in the geometry of the model. Both the linear and nonlinear, two-dimensional (2-D) and three-dimensional (3-D), models yield good qualitative and quantitative predictions. Additionally, it was predicted correctly that the panel would fail in buckling prior to failing in strength.

heavy lift launch vehicles↗

Experimental Evaluation of Optically Polished Aluminum Panels on the Deep Space Network's 34 Meter Antenna

The potential development of large aperture ground?based "photon bucket" optical receivers for deep space communications has received considerable attention recently. One approach currently under investigation is to polish the aluminum reflector panels of 34?meter microwave antennas to high reflectance, and accept the relatively large spotsize generated by state of?the?art polished aluminum panels. Theoretical analyses of receiving antenna pointing, temporal synchronization and data detection have been addressed in previous papers. Here we describe the experimental effort currently underway at the Deep Space Network (DSN) Goldstone Communications Complex in California, to test and verify these concepts in a realistic operational environment. Two polished aluminum panels (a standard DSN panel polished to high reflectance, and a custom designed aluminum panel with much better surface quality) have been mounted on the 34 meter research antenna at Deep?Space Station 13 (DSS?13), and a remotely controlled CCD camera with a large CCD sensor in a weather?proof container has been installed next to the subreflector, pointed directly at the custom polished panel. The point?spread function (PSF) generated by the Vertex polished panel has been determined to be smaller than the sensor of the CCD camera, hence a detailed picture of the PSF can be obtained every few seconds, and the sensor array data processed to determine the center of the intensity distribution. In addition to estimating the center coordinates, expected communications performance can also been evaluated with the recorded data. The results of preliminary pointing experiments with the Vertex polished panel receiver using the planet Jupiter to simulate the PSF generated by a deep?space optical transmitter are presented and discussed in this paper.

optical receiver↗

Spaced out: An economic framework to explore the impacts of PV panel spacing on large-scale farming in Colorado

CONTEXT Agrivoltaic systems co-locate solar technologies with agricultural operations on an integrated plot of land and potentially provide benefits to both energy and agricultural systems. To date, large-scale (>5-MW) agrivoltaic projects in the United States have been limited to grazing and ecovoltaic applications, raising questions about the impact and scalability of agrivoltaic crop systems. Many agrivoltaic designs raise the height of the solar panels to accommodate agricultural practices while keeping energy density high. However, raising the panels results in increased photovoltaic (PV) development costs, which often are higher than the economic returns of crop production underneath the panels. This leads to unfavorable project economics and the need for other agrivoltaic solutions than raising panels. OBJECTIVE To explore other solutions, we perform an initial feasibility analysis for an agrivoltaic solution that can integrate with large-scale farming practices by increasing the row spacing in between panels. Increased PV row spacing is a low-cost approach for scaling agrivoltaics to accommodate crop production and this spacing can be tailored to required crop equipment for different regions. Increasing row spacing will reduce the power density (PV installed per acre), but in areas that are not land limited, these agrivoltaic designs could be economically feasible. Our analysis establishes a framework for a feasibility analysis for where and with what crops spaced out panel agrivoltaic solutions might be economical. METHODS Using a case study for large-scale agriculture crops in Colorado, we establish a framework for wide-row agrivoltaic economic feasibility analysis. We utilized the System Advisor Model to calculate technoeconomic metrics to compare different row spacing solutions and capture tradeoffs of these system designs. RESULTS AND CONCLUSIONS We find that, in some circumstances, wider row agrivoltaic solutions that allow for continued mechanized crop production can provide economic benefits over a traditional utility-scale PV system. For most crops examined in this analysis, roughly $\$$200/acre in agricultural profit justified spacing out the panels to at least 31.7 ft. to accommodate agrivoltaic configurations versus PV only configurations. Additionally, opportunities for increased agricultural revenue with agrivoltaic systems allow PV project economics to tolerate a larger range of CAPEX variability while remaining economically viable relative to the PV only configurations. SIGNIFICANCE This framework can be adapted for a wide variety of crops and regions and allows for examination of economically favorable sites for future agrivoltaic systems that utilize different configuration and expand opportunities for agrivoltaics.

14 SOLAR ENERGY↗

The effect of hypersonic nonlinear aerodynamic loading on panel flutter.

The system considered is a two-dimensional isotropic panel, or plate-column, on hinged supports, with one end spring restrained in the plane of the panel. Panel geometric nonlinearities and piston-theory aerodynamic nonlinearities are included. Results from an earlier preliminary study indicate that only two second-order nonlinear aerodynamic terms are important. The nonlinear aerodynamic terms introduce the possibility of an amplitude-sensitive instability, where the panel is unstable to disturbances of a certain magnitude but stable for smaller ones. This type of instability is examined for various panel in-plane loads and initial conditions, with other parameters having values representative of current practice. A single new interaction parameter, representing the importance of the nonlinear aerodynamic terms in comparison with the panel geometric nonlinear terms, is introduced. A parameter survey is presented, involving this parameter and the static pressure difference across the panel. Finally, the practical impact of this type of instability is discussed.

Mcintosh, S. C., Jr.↗

Acoustic fatigue and sound transmission characteristics of a ram composite panel design

An experimental study to determine the acoustic fatigue characteristics of a flat multi-layered structural panel is described. The test panel represented a proposed design for the outer skin of a research application module to be housed within the space shuttle orbiter vehicle. The test specimen was mounted in one wall of the Wyle 100,000 cu ft reverberation room and exposed to a broadband acoustic environment having an overall level of 145 db. The test panel was exposed to nine separate applications of the acoustic environment, each application consisting of 250 seconds duration. Upon completion of the ninth test run, the specimen was exposed to a simulated micrometeoroid impact near the panel center. One additional test run of 250 seconds duration was then performed to complete the overall simulation of 50 flight missions. The experimental results show that no significant fatigue damage occurred until the test specimen was exposed to a simulated micrometeoroid impact. The intermediate foam layer forming the core of the test specimen suffered considerable damage due to this impact, causing a marked variation in the dynamic characteristics of the overall test panel. During the final application of the acoustic environment, the strain and acceleration response spectra showed considerable variation from those spectra obtained prior to impact of the panel. Fatigue damage from acoustic loading however, was limited to partial de-bonding around the edges of the composite panel.

Cockburn, J. A.↗

Prediction and verification of creep behavior in metallic materials and components for the space shuttle thermal protection system. Volume 2: Phase 2 subsize panel cyclic creep predictions

A method for predicting permanent cyclic creep deflections in stiffened panel structures was developed. The resulting computer program may be applied to either the time-hardening or strain-hardening theories of creep accumulation. Iterative techniques were used to determine structural rotations, creep strains, and stresses as a function of time. Deflections were determined by numerical integration of structural rotations along the panel length. The analytical approach was developed for analyzing thin-gage entry vehicle metallic-thermal-protection system panels subjected to cyclic bending loads at high temperatures, but may be applied to any panel subjected to bending loads. Predicted panel creep deflections were compared with results from cyclic tests of subsize corrugation and rib-stiffened panels. Empirical equations were developed for each material based on correlation with tensile cyclic creep data and both the subsize panels and tensile specimens were fabricated from the same sheet material. For Vol. 1, see N75-21431.

Cramer, B. A.↗

Minimum-mass designs of stiffened graphite/polyimide compression panels

This study presents results from an analytic investigation to determine minimum-mass designs of stiffened Gr/Pi compression panels for a wide range of uniaxial compression loads. Four panel configurations are considered: (1) hat stiffened-laminated skin, (2) hat stiffened-honeycomb core sandwich skin, (3) blade stiffened-laminated skin and (4) blade stiffened-honeycomb core sandwich skin panels. Designs are generated by an automatic optimization computer code entitled PASCO (Panel Analysis and Sizing COde). The analysis portion of PASCO contains a generalized buckling analysis which calculates all critical modes without assuming any of them a priori. Mass-strength curves are presented for the four configurations along with specific designs generated for low, moderate and high values of compression load. A mass-strength curve for a titanium hat stiffened panel is provided for comparison. A 30% to 50% mass savings over titanium panels is shown for moderately and heavily loaded Gr/Pi panels.

Weaver, G. G., II↗