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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 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1997

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↗
NASA NTRS1960

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.↗
NASA NTRS1999

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.↗
NASA NTRS2002

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.↗
NASA NTRS2003

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↗
NASA NTRS2007

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.↗
NASA NTRS1989

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.↗
NASA NTRS2014

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↗
NASA NTRS2011

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↗
NASA NTRS1972

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.↗
NASA NTRS1972

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.↗
NASA NTRS1975

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.↗
NASA NTRS1979

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↗
NASA NTRS1980

Solar panel thermal cycling testing by solar simulation and infrared radiation methods

For the solar panels of the European Space Agency (ESA) satellites OTS/MAROTS and ECS/MARECS the thermal cycling tests were performed by using solar simulation methods. The performance data of two different solar simulators used and the thermal test results are described. The solar simulation thermal cycling tests for the ECS/MARECS solar panels were carried out with the aid of a rotatable multipanel test rig by which simultaneous testing of three solar panels was possible. As an alternative thermal test method, the capability of an infrared radiation method was studied and infrared simulation tests for the ultralight panel and the INTELSAT 5 solar panels were performed. The setup and the characteristics of the infrared radiation unit using a quartz lamp array of approx. 15 sq and LN2-cooled shutter and the thermal test results are presented. The irradiation uniformity, the solar panel temperature distribution, temperature changing rates for both test methods are compared. Results indicate the infrared simulation is an effective solar panel thermal testing method.

Nuss, H. E.↗
NASA NTRS1979

A higher order panel method for linearized supersonic flow

The basic integral equations of linearized supersonic theory for an advanced supersonic panel method are derived. Methods using only linear varying source strength over each panel or only quadratic doublet strength over each panel gave good agreement with analytic solutions over cones and zero thickness cambered wings. For three dimensional bodies and wings of general shape, combined source and doublet panels with interior boundary conditions to eliminate the internal perturbations lead to a stable method providing good agreement experiment. A panel system with all edges contiguous resulted from dividing the basic four point non-planar panel into eight triangular subpanels, and the doublet strength was made continuous at all edges by a quadratic distribution over each subpanel. Superinclined panels were developed and tested on s simple nacelle and on an airplane model having engine inlets, with excellent results.

Ehlers, F. E.↗
NASA NTRS1982

Flightweight radiantly and actively cooled panel: Thermal and structural performance

A 2- by 4-ft flightweight panel was subjected to thermal/structural tests representative of design flight conditions for a Mach 6.7 transport and to off-design conditions simulating flight maneuvers and cooling system failures. The panel utilized Rene 41 heat shields backed by a thin layer of insulation to radiate away most of the 12 Btu/ft2-sec incident heating. A solution of ethylene glycol in water circulating through tubes in an aluminum-honeycomb-sandwich panel absorbed the remainder of the incident heating (0.8 Btu/sq ft-sec). The panel successfully withstood (1) 46.7 hr of radiant heating which included 53 thermal cycles and 5000 cycles of uniaxial inplane loading of + or - 1200 lfb/in; (2) simulated 2g-maneuver heating conditions and simulated cooling system failures without excessive temperatures on the structural panel; and (3) the extensive thermal/structural tests and the aerothermal tests reported in NASA TP-1595 without significant damage to the structural panel, coolant leaks, or hot-gas ingress to the structural panel.

Shore, C. P.↗
NASA NTRS1980

Compression panel studies for supersonic cruise vehicles

Results of analytical and experimental studies are summarized for titanium, boron fiber reinforced aluminum matrix composite, Borsic fiber reinforced aluminum matrix composite, and titanium sheathed Borsic fiber reinforced aluminum matrix composite stiffened panels. The results indicate that stiffened panels with continuous joints (i.e., brazed, diffusion bonded or adhesive bonded joints) are more structurally efficient than geometrically similar panels with discrete joints (i.e., spotwelded or bolted joints). In addition, results for various types of fiber reinforced aluminum matrix stiffened panels indicate that titanium sheathed Borsic fiber reinforced aluminum matrix composite panels are the most structurally efficient. Analytical results are also presented for graphite fiber reinforced polyimide matrix composite stiffened panels and superplastically formed and diffusion bonded titanium sandwich panels.

Mcwithey, R. R.↗
NASA NTRS1983

Fabrication and evaluation of cold/formed/weldbrazed beta-titanium skin-stiffened compression panels

The room temperature and elevated temperature buckling behavior of cold formed beta titanium hat shaped stiffeners joined by weld brazing to alpha-beta titanium skins was determined. A preliminary set of single stiffener compression panels were used to develop a data base for material and panel properties. These panels were tested at room temperature and 316 C (600 F). A final set of multistiffener compression panels were fabricated for room temperature tests by the process developed in making the single stiffener panels. The overall geometrical dimensions for the multistiffener panels were determined by the structural sizing computer code PASCO. The data presented from the panel tests include load shortening curves, local buckling strengths, and failure loads. Experimental buckling loads are compared with the buckling loads predicted by the PASCO code. Material property data obtained from tests of ASTM standard dogbone specimens are also presented.

Royster, D. M.↗