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At least 127 records · Page 7

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NASA NTRSDate not supplied

Progressive Damage Failure Analysis of a Multi-Stringer Post-Buckled Panel

Validation of a progressive damage finite element analysis model using the CompDam material model was performed on a large multi-stringer panel subjected to compressive loading. The panel had a Teflon insert embedded between the skin panel and stiffener to cause an ultimate failure mode of stiffener separation. The compressive loading of the panel caused the skin between the stiffeners to locally buckle before any damage began to occur. The post-buckled behavior of the panel was the driving mechanism to the onset of damage and the damage growth that led to skin-stiffener separation. An integrated global-local modeling approach was used to validate several aspects the overall behavior of the panel from start to failure. The global region captured the pre-buckled stiffness of the test panel within 10% and the buckling load of the skin panel at the critical location within 2%. The skin’s five half-wave buckled mode shape was accurately predicted by the model. The integrated local model captured the load corresponding to the onset of damage within 5% of the average test data. Comparisons are made for key aspects of the damage morphology, such as a growth pattern that included a significant matrix split in the first ply of the skin and a migration of the delamination from the skin-stiffener interface to the skin’sply1-ply2 interface. The ultimate failure mode was shown to be an unstable growth of delamination damage under the stiffeners which led to a peak load of the analysis that was approximately 5% above the average peak load of the testing. The validated global-local modeling approach used on the multi-stringer post-buckled panel with a Teflon insert used a methodology and lessons learned using smaller specimens applied to a subcomponent that captured several aspects of a modern aircraft design.

Jason Action↗

DOE OSTI2026

Characterizing electrical panel capacity, breaker space, and loads in U.S. single-family homes

Electrical panels distribute electricity to appliances, equipment, and other building loads and maintain safety by preventing overloads of branch circuits, feeders, and utility service wires connecting homes to the grid. Load calculations in the National Electrical Code (NEC) determine the capacity of installed electrical panels in the United States, which are rated in amps (A). Circuit breakers within electrical panels affect the magnitude and number of loads that panels can serve. In home retrofits and renovations, existing electrical panels may not accommodate new electrical loads, either due to insufficient electrical capacity or a lack of available breaker spaces. Homes with these panel constraints may need to replace existing panels with new panels that have greater capacity and breaker space.

24 POWER TRANSMISSION AND DISTRIBUTION↗

NASA NTRS1972

Effects of eccentricities and lateral pressure on the design of stiffened compression panels

An analysis for determining the effects of eccentricities and lateral pressure on the design of stiffened compression panels is presented. The four types of panel stiffeners considered are integral, zee, integral zee, and integral tee. Mass-strength curves, which give the mass of the panel necessary to carry a specified load, are given along with related design equations needed to calculate the cross-sectional dimensions of the minimum-mass-stiffened panel. The results of the study indicate that the proportions of the panels are geometrically similar to the proportions of panels designed for no eccentricity or lateral pressure, but no cross-sectional dimensions are greater, resulting in significantly increased mass. The analytical minimum-mass designs of zee-stiffened panels are compared with designs from experimentally derived charts. An assumed eccentricity of 0.001 times the length of the panel is used to correlate the analytical and experimental data. Good correlation between the experimentally derived and the analytical curves is obtained for the range of loading where materials yield governs the design. At lower loads the mass given by the analytical curve using this assumed eccentricity is greater than that given by the experimental results.

Giles, G. L.↗

NASA NTRS1973

Experimental investigation of orthotropic panel flutter at arbitrary yaw angles, and comparison with theory

Flutter characteristics for yaw angles between 15 deg and 90 deg were determined experimentally for two types of corrugation-stiffened panels: those with weak twisting stiffness and those with strong twisting stiffness. By mounting the panels on a remotely controlled turntable, good definition of the flutter boundaries was obtained by rotating the panels into and out of flutter. Flutter tests were conducted at M = 2 and M = 1.6. Before testing, vibration tests and analyses were also performed. The experimental flutter data is compared with flutter theory for orthotropic panels utilizing quasi-steady aerodynamics. Five different corrugated panels were tested consisting of one single skin panel having a length-to-width ratio of 5 on clamped supports and four different square double skin panels on discrete flexible supports. The investigation indicated that flutter speed for corrugated panels is highly dependent on yaw angle. Reasonable flutter correlation between analysis and test was obtained for moderate yaw angles, but extreme sensitivity to structural parameters made the correlation at large yaw angles uncertain.

Shyprykevich, P.↗

NASA NTRS1976

Buckling of Cylindrical Panels Under Axial Compression

This paper investigates the effects of boundary conditions and panel width on the axially compressive buckling behavior of unstiffened, isotropic, circular cylindrical panels. Numerical results are presented for eight different sets of boundary conditions along the straight edges of the panels. For all sets of boundary conditions except one (SSI), the results show that the panel buckling loads monotonically approach the complete cylinder buckling load from above as the panel width is increased. Low buckling loads, sometimes less than half the complete cylinder buckling load, are found for simply supported panels with free in-plane edge displacements (SSI). The SSI buckling loads are below the complete cylinder load even for '360deg panels'. It is also observed that the prevention of circumferential edge displacement is the most important in-plane boundary condition from the point of view of increasing the buckling load, and that the prevention of edge rotation (i.e. clamping) in the circumferential direction also significantly increases the buckling load. Parametric studies are also performed to determine the effects of variations in panel length and thickness on the buckling loads.

Sobel, L. H.↗

NASA NTRS1977

Development of fire-resistant wood structural panels

Structural panels made with Xylok 210 resin as the binder had a burn-through resistance at least equal to the structural panels made with Kerimid 500. Therefore, because of its comparative ease of handling, Xylok 210 was selected as the resin binder to provide the baseline panel for the study of a means of improving the flame-spread resistance of the structural panels. The final resin-filler system consisted of Xylok 210 binder with the addition of ammonium oxalate and ammonium phosphate to the strands of the surface layers, using 24% of each salt based upon the air-dry weight of the strands. This system resulted in a panel with a flame-spread code of about 60, a Class 2 classification. A standard phenolic based structural panel had a flame-spread greater than 200 for laboratory prepared panels. The burn-through tests indicated an average burn-through time of 588 seconds for the specimens made with the final system. This compares to an average burn-through time of 287 seconds for the standard phenolic base structural specimen. One full-size panel was made with the final system.

Vaughan, T. W.↗

NASA NTRS1977

Multipurpose Panel Display Device Investigation

A multipurpose panel was developed to provide a flexible control and a LED display panel with easily changeable nomenclature for use in applications where panel space is limited, but where a number of similar subsystems must be controlled, or where basic panel nomenclature and functions must be changed rapidly, as in the case of between mission changes of space shuttle payloads. In the first application, panel area limitations are overcome by time sharing a central control panel among several subsystems. In the latter case, entire control panel changes are effected by simply replacing a memory module, thereby reducing the extent of installation and checkout procedures between missions. Several types of control technologies (other than LED's) which show potential in meeting criteria for overcoming limitations of the panel are assessed.

Sliwa, R.↗

NASA NTRS1982

Postbuckling behavior of selected flat stiffened graphite-epoxy panels loaded in compression

Results of an experimental study of the postbuckling behavior of selected flat stiffened graphite-epoxy panels loaded in compression are presented. The postbuckling response and failure characteristics of undamaged panels and panels damaged by low-speed impact are described. Each panel had four equally-spaced I-spaced stiffeners and 16- or 24-ply quasi-isotropic skins. Panels with three different stiffener spacings were tested. Some undamaged specimens supported as much as three times their initial buckling load before failing. Failure of all panels initiated in a skin-stiffener interface region. Analytical results obtained from a nonlinear general shell finite element analysis computer code correlate well with typical postbuckling test results up to failure. The analytical modeling detail necessary to predict accurately the response of a panel is described. Test results show that low-speed impact damage can reduce the postbuckling strength of a stiffened panel and that the skin-stiffener interface region is more sensitive to impact damage than the skin midway between stiffeners.

Starnes, J. H., Jr.↗

NASA NTRS1983

The response of cylindrical panels fabricated from symmetrically and unsymmetrically laminated composite materials

Equations are developed which govern the deflection response of long cylindrical panels subjected to a line load. The line load is directed toward the center of curvature of the panel, is located at an arbitrary point along the arc length of the panel, and is included at an arbitrary angle relative to the radial direction. Only the geometrically linear problem is considered and the spatial dependence in the problem is reduced to one independent variable, specifically, the arc length along the panel. The problem is thus solvable in closed form. Both symmetrically laminated and the less common unsymmetrically laminated simply supported panels are studied. The unsymmetrically laminated case was considered because the natural shape of an unsymmetric laminate is cylindrical. Results are presented which show the influence of the location and inclination of the line load on panel deflection. Shallow and deep panels are considered. Both the symmetric and unsymmetric panels exhibit similar behavior, the unsymmetric configurations being less stiff. Limited experimental results are presented.

Carper, D. M.↗

NASA NTRS1985

Effect of low-velocity or ballistic impact damage on the strength of thin composite and aluminum shear panels

Impact tests were conducted on shear panels fabricated from 6061-T6 aluminum and from woven fabric prepreg of Du Pont Kevlara fiber/epoxy resin and graphite fiber/epoxy resin. The shear panels consisted of three different composite laminates and one aluminum material configuration. Three panel aspect ratios were evaluated for each material configuration. Composite panels were impacted with a 1.27-cm (0.05-in) diameter aluminum sphere at low velocities of 46 m/sec (150 ft/sec) and 67 m/sec (220 ft/sec). Ballistic impact conditions consisted of a tumbled 0.50-caliber projectile impacting loaded composite and aluminum shear panels. The results of these tests indicate that ballistic threshold load (the lowest load which will result in immediate failure upon penetration by the projectile) varied between 0.44 and 0.61 of the average failure load of undamaged panels. The residual strengths of the panels after ballistic impact varied between 0.55 and 0.75 of the average failure strength of the undamaged panels. The low velocity impacts at 67 m/sec (220 ft/sec) caused a 15 to 20 percent reduction in strength, whereas the impacts at 46 m/sec (150 ft/sec) resulted in negligible strength loss. Good agreement was obtained between the experimental failure strengths and the predicted strength with the point stress failure criterion.

Farley, G. L.↗

NASA NTRS1984

Sizing-stiffened composite panels loaded in the postbuckling range

Stiffened panels are widely used in aircraft structures such as wing covers, fuselages, control surfaces, spar webs, bulkheads, and floors. The detailed sizing of minimum-weight stiffened panels involves many considerations. Use of composite materials introduces additional complexities. Many potential modes of failure exist. Analyses for these modes are often not trivial, especially for those involving large out-of-plane displacements. Accurate analyses of all potential failure modes are essential. Numerous practical constraints arise from manufacturing/cost considerations and from damage tolerance, durability, and stiffness requirements. The number of design variables can be large when lamina thicknesses and stacking sequence are being optimized. A significant burden is placed on the sizing code due to the complex analyses, practical constraints, and number of design variables. On the other hand, sizing weight-efficient panels without the aid of an automated procedure is almost out of the question. The sizing code postbuckled Open-Stiffener Optimum Panels (POSTOP) has been developed to aid in the design of minimum-weight panels subject to the considerations mentioned above. Developed for postbuckled composite panels, POSTOP may be used for buckling resistant panels and metallic panels as well. The COPES/CONMIN optimizer is used in POSTOP although other options such as those in the ADS system could be substituted with relative ease. The basic elements of POSTOP are shown. Some of these elements and usage of the program are described.

Biggers, S. B.↗

NASA NTRS1990

A study of the structural efficiency of fluted core graphite-epoxy panels

The structural efficiency of compression-loaded graphite-epoxy sandwich panels with fluted cores is studied to determine their weight saving potential. Graphite-epoxy equilateral triangular elements are used to construct the fluted cores for the sandwich panels. Two panel configurations are considered. One configuration has two layers of triangular elements in the fluted core and the second configuration has only one layer of triangular elements in the core. An optimization code is used to find the minimum weight design for each panel configuration. Laminate ply orientations are limited to approx. 45, 0, and 90 deg. A constraint on the axial stiffness is included in the design process so the panel will conform to typical constraints for aircraft wing structures. Minimum thickness requirements for each laminate and maximum allowable strains are also included. A comparison is made of the calculated structural efficiency of the fluted core panels to the structural efficiency of aluminum transport aircraft structures and simple blade-stiffened graphite-epoxy panels. Limited experimental results are also included for comparison with the analytical predictions and to identify the critical failure mechanisms of graphite-epoxy fluted-core sandwich panels.

Jegley, Dawn C.↗

NASA NTRS1990

Radiometric tests on wet and dry antenna reflector surface panels

The results of X-band noise temperature tests on two types of antenna surface panels are presented. The first type tested was a solid antenna panel, while the second type was a perforated panel with 3/16-in.-diameter holes. Measurements were made at 8.45 GHz using an X-band radiometric system. Included in this article are measured noise temperature contributions from: (1) thermal diffusive white paint on solid and perforated panels, and (2) water sprayed on both painted and unpainted perforated panels. Experiments on perforated panels were restricted to the 3/16-in.-diameter hole panels formerly used on Deep Space Network 64-m antennas. Rigorous calibration equations, applicable to a variety of antenna panel and dichroic plate test configurations, are presented. It was demonstrated that an accurate, stable radiometric measurement system of the type used for the results of this research makes it possible to obtain information that would be much more difficult to obtain using other techniques.

Otoshi, T. Y.↗

NASA NTRS1989

Space Station Freedom solar array panels plasma interaction test facility

The Space Station Freedom Power System will make extensive use of photovoltaic (PV) power generation. The phase 1 power system consists of two PV power modules each capable of delivering 37.5 KW of conditioned power to the user. Each PV module consists of two solar arrays. Each solar array is made up of two solar blankets. Each solar blanket contains 82 PV panels. The PV power modules provide a 160 V nominal operating voltage. Previous research has shown that there are electrical interactions between a plasma environment and a photovoltaic power source. The interactions take two forms: parasitic current loss (occurs when the currect produced by the PV panel leaves at a high potential point and travels through the plasma to a lower potential point, effectively shorting that portion of the PV panel); and arcing (occurs when the PV panel electrically discharges into the plasma). The PV solar array panel plasma interaction test was conceived to evaluate the effects of these interactions on the Space Station Freedom type PV panels as well as to conduct further research. The test article consists of two active solar array panels in series. Each panel consists of two hundred 8 cm x 8 cm silicon solar cells. The test requirements dictated specifications in the following areas: plasma environment/plasma sheath; outgassing; thermal requirements; solar simulation; and data collection requirements.

Martin, Donald F.↗

NASA NTRS1991

User's guide to a system of finite-element supersonic panel flutter programs

The utilization and operation of a set of six computer programs for the prediction of panel flutter at supersonic speeds by finite element methods are described. The programs run individually to determine the flutter behavior of a flat panel where the finite elements which model the panel each have four degrees of freedom (DOF), a curved panel where the finite elements each have four DOF, and a curved panel where the finite elements each have six DOF. The panels are assumed to be of infinite aspect ratio and are subjected to either simply-supported or clamped boundary conditions. The aerodynamics used by these programs are based on piston theory. Application of the program is illustrated by sample cases where the number of beam finite elements equals four, the in-plane tension parameter is 0.0, the maximum camber to panel length ratio for a curved panel case is 0.05, and the Mach number is 2.0. This memorandum provides a user's guide for these programs, describes the parameters that are used, and contains sample output from each of the programs.

Woolley, Christine L.↗

NASA NTRS1990

Acoustic testing of high temperature panels

Results are presented of a series of thermal-acoustic tests conducted on the NASA Langley Research Center Thermal-Acoustic Test Apparatus to (1) investigate techniques for obtaining strain measurements on metallic and carbon-carbon materials at elevated temperature; (2) document the dynamic strain response characteristics of several superalloy honeycomb thermal protection system panels at elevated temperatures of up to 1200 F; and (3) determine the strain response and sonic fatigue behavior of four carbon-carbon panels at both ambient and elevated temperatures. A second study tested four carbon-carbon panels to document panel dynamic response characteristics at ambient and elevated temperature, determine time to failure and faliure modes, and collect continuous strain data up to panel failure. Strain data are presented from both types of panels, and problems encountered in obtaining reliable strain data on the carbon-carbon panels are described. The failure modes of the carbon-carbon panels are examined.

Leatherwood, Jack D.↗

NASA NTRS1993

Nuclear thermal propulsion technology: Results of an interagency panel in FY 1991

NASA LeRC was selected to lead nuclear propulsion technology development for NASA. Also participating in the project are NASA MSFC and JPL. The U.S. Department of Energy will develop nuclear technology and will conduct nuclear component, subsystem, and system testing at appropriate DOE test facilities. NASA program management is the responsibility of NASA/RP. The project includes both nuclear electric propulsion (NEP) and nuclear thermal propulsion (NTP) technology development. This report summarizes the efforts of an interagency panel that evaluated NTP technology in 1991. Other panels were also at work in 1991 on other aspects of nuclear propulsion, and the six panels worked closely together. The charters for the other panels and some of their results are also discussed. Important collaborative efforts with other panels are highlighted. The interagency (NASA/DOE/DOD) NTP Technology Panel worked in 1991 to evaluate nuclear thermal propulsion concepts on a consistent basis. Additionally, the panel worked to continue technology development project planning for a joint project in nuclear propulsion for the Space Exploration Initiative (SEI). Five meetings of the panel were held in 1991 to continue the planning for technology development of nuclear thermal propulsion systems. The state-of-the-art of the NTP technologies was reviewed in some detail. The major technologies identified were as follows: fuels, coatings, and other reactor technologies; materials; instrumentation, controls, health monitoring and management, and associated technologies; nozzles; and feed system technology, including turbopump assemblies.

Clark, John S.↗

NASA NTRS1993

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.↗