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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 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Fabrication and evaluation of superplastically formed/weld-brazed corrugated compression panels with beaded webs

A study was made to investigate the feasibility of superplastically forming corrugated panels with beaded webs and to demonstrate the structural integrity of these panels by testing. The test panels in the study consist of superplastically formed titanium alloy Ti-6Al-4V half-hat elements that are joined by weld-brazing to titanium alloy Ti-6Al-4V caps to form either single-corrugation compression panels or multiple-corrugation compression panels. Stretching and subsequent thinning of the titanium sheet during superplastic forming is reduced by approximately 35 percent with a shallow half-hat die concept instead of a deep die concept and results in a more uniform thickness across the beaded webs. The complete panels are tested in end compression at room temperature and the results compared with analysis. The heavily loaded panels failed at loads approaching the yield strength of the titanium material. At maximum load, the caps wrinkled locally accompanied with separation of the weld-braze joint in the wrinkle. None of the panels tested, however, failed catastrophically in the weld-braze joint. Experimental test results are in good agreement with structural analysis of the panels.

Royster, D. M.↗
NASA NTRS1989

A finite element method for nonlinear panel flutter

Panel flutter is the self excited oscillations of a plate in supersonic flow. Linear theory gives no information about the panel's deflections and stresses. Hence, the service life of the panel cannot be predicted by linear methods. Nonlinear structural theory determines the limit cycle oscillating frequency, and also panel deflections and stresses. Panel fatigue life, therefore, can be predicted. For a more thorough understanding of panel flutter behavior, the geometric nonlinearity effects due to large deflections must be considered in the formulation. The Galerkin's method was used in the investigation of limit cycle oscillations of panels in supersonic flow in the spatial domain, and the panel deflections. All nonlinear flutter methods and results presently available were limited to a linear or linearized aerodynamic theory. Extension of the finite element flutter formulation to include nonlinear hypersonic aerodynamic loading and finite element nonlinear panel flutter results are presented.

Mei, Chuh↗
NASA NTRS1995

Sensitivity analysis in multipole-accelerated panel methods for potential flow

In the design of an airframe, the effect of changing the geometry on resulting computations is necessary for design optimization. The geometry is defined in terms of a series of design variables, including design variables to define the wing planform, tail, canard, pylon, and nacelle. Design optimization in this research is based on how these design variable affect the potential flow. The potential flow is computed as a function of the geometry and location of a series of panels describing the airframe, which are in turn a function of the design variables. Multipole accelerated panel methods improve the computational complexity of the problem and thus are an attractive approach. To utilize the methods in design optimization, it was necessary to define the appropriate sensitivity derivatives. The overhead incurred from finding the sensitivity derivatives in conjunction with the original computation should be small. This research developed the background for multipole-accelerated panel methods and the framework for finding sensitivity derivatives in the methods. Potential flow panel codes are commonly used for powered-lift aerodynamic predictions for three dimensional geometries. Given an airframe which has been discretized into a series of panels to define the airframe geometry, potential is computed as a function of the influence of all panels on all other panels. This is a computationally intensive problem for which efficient solutions are desired to improve the computational time and to allow greater resolution by use of more panels. One such solution is the use of hierarchical multipole methods which entail approximations of the effects of far-field terms. Hierarchical multipole methods have become prevalent in molecular dynamics and gravitational physics, and have been introduced into the fields of capacitance calculations, computational fluid dynamics, and electromagnetics. The methods utilize multipole expansions to describe the effect of bodies (i.e. particles, astrophysical bodies, panels, etc.) within a sphere on points distant from the sphere, where the influence diminishes as a function of distance. The expansions are exact with infinite series, however, for practical computations, the series are truncated and accuracy is selected based on the number of terms retained in the expansions. A hierarchical tree structure groups bodies together based on proximity to allow definition of multipole expansions for each group. The multipole expansions are then used to compute the effect of the bodies in a group on distant bodies.

Leathrum, James F., Jr.↗
NASA NTRS1997

Compression Response of a Sandwich Fuselage Keel Panel With and Without Damage

Results are presented from an experimental and analytical study of a sandwich fuselage keel panel with and without damage. The fuselage keel panel is constructed of graphite-epoxy skins bonded to a honeycomb core, and is representative of a highly loaded fuselage keel structure. The face sheets of the panel contain several terminated or dropped plies along the length of the panel. The results presented provide a better understanding of the load distribution in damaged and undamaged thick-face-sheet composite sandwich structure with dropped plies and of the failure mechanisms of such structure in the presence of low-speed impact damage and discrete-source damage. The impact-damage condition studied corresponds to barely visible impact damage (BVID), and the discrete-source damage condition studied is a notch machined through both face sheets. Results are presented from an impact-damage screening study conducted on another panel of the same design to determine the impact energy necessary to inflict BVID on the panel. Results are presented from compression tests of the panel in three conditions: undamaged; BVID in two locations; and BVID in two locations and a notch through both face sheets. Surface strains in the face sheets of the undamaged panel and the notched panel obtained experimentally are compared with finite element analysis results. The experimental and analytical results suggest that for the damage conditions studied, discrete-source damage influences the structural performance more than BVID.

McGowan, David M.↗
NASA NTRS1959

Elevated-Temperature Tests Under Static and Aerodynamic Conditions on Honeycomb-Core Sandwich Panels

Stainless-steel honeycomb-core sandwich panels which differed primarily in skin thicknesses were tested at elevated temperatures under static and aerodynamic conditions. The results of these tests were evaluated to determine the insulating effectiveness and structural integrity of the panels. The static radiant-heating tests were performed in front of a quartz-tube radiant heater at panel skin temperatures up to 1,5000 F. The aerodynamic tests were made in a Mach 1.4 heated blowdown wind tunnel. The tunnel temperature was augmented by additional heat supplied by a radiant heater which raised the panel surface temperature above 8000 F during air flow. Static radiant-heating tests of 2 minutes duration showed that all the panels protected the load-carrying structure about equally well. Thin-skin panels showed an advantage for this short-time test over thick-skin panels from a standpoint of weight against insulation. Permanent inelastic strains in the form of local buckles over each cell of the honeycomb core caused an increase in surface roughness. During the aero- dynamic tests all of the panels survived with little or no damage, and panel flutter did not occur.

Groen, Joseph M.↗
NASA NTRS1961

Experimental Investigation at Mach Number 3.0 of the Effects of Thermal Stress and Buckling on the Flutter of Four-Bay Aluminum Alloy Panels with Length-Width Ratios of 10

Skin-stiffener aluminum alloy panels consisting of four bays, each bay having a length-width ratio of 10, were tested at a Mach number of 3.0 at dynamic pressures ranging from 1,500 psf to 5,000 psf and at stagnation temperatures from 300 F to 655 F. The panels were restrained by the supporting structure in such a manner that partial thermal expansion of the skins could occur in both the longitudinal and lateral directions. A boundary faired through the experimental flutter points consisted of a flat-panel portion, a buckled-panel portion, and a transition point at the intersection of the two boundaries. In the region where a panel must be flat when flutter occurs, an increase in panel skin temperature (or midplane compressive stress) makes the panel more susceptible to flutter. In the region where a panel must be buckled when flutter occurs, the flutter trend is reversed. This reversal in trend is attributed to the panel postbuckling behavior.

Dixon, Sidney C.↗
NASA NTRS2000

Development of Electrostatically Clean Solar Array Panels

Certain missions require Electrostatically Clean Solar Array (ECSA) panels to establish a favorable environment for the operation of sensitive scientific instruments. The objective of this program was to demonstrate the feasibility of an ECSA panel that minimizes panel surface potential below 100mV in LEO and GEO charged particle environments, prevents exposure of solar cell voltage and panel insulating surfaces to the ambient environment, and provides an equipotential, grounded structure surrounding the entire panel. An ECSA panel design was developed that uses a Front Side Aperture-Shield (FSA) that covers all inter-cell areas with a single graphite composite laminate, composite edge clips for connecting the FSA to the panel substrate, and built-in tabs that interconnect the FSA to conductive coated coverglasses using a conductive adhesive. Analysis indicated the ability of the design to meet the ECSA requirements. Qualification coupons and a 0.5m x 0.5m prototype panel were fabricated and tested for photovoltaic performance and electrical grounding before and after exposure to acoustic and thermal cycling environments. The results show the feasibility of achieving electrostatic cleanliness with a small penalty in mass, photovoltaic performance and cost, with a design is structurally robust and compatible with a wide range of current solar panel technologies.

Stern, Theodore G.↗
NASA NTRS2006

A Numerical and Experimental Study of Compression-Loaded Composite Panels With Cutouts

Results from a numerical and experimental study on the effects of laminate orthotropy and circular cutout size on the response of compression-loaded composite curved panels are presented. Several 60-in-radius composite panels with four different laminate configurations were tested with cutout diameters that range from 10% to 60% of the panel width. Finite-element analyses were performed for each panel in order to identify the effects boundary conditions, measured initial geometric imperfections and thickness variations had on the nonlinear and buckling behavior of the panels. The compression-loaded panels considered herein exhibited two separate types of behavior depending on the laminate stacking sequence and cutout size. More specifically, some of the panels exhibited the classical snap-through type buckling response; however, some of the panels exhibited a monotonically increasing stable response and achieved compressive loads in excess of twice the predicted linear bifurcation buckling load. In general, the finite-element analyses were able to predict accurately the nonlinear response and buckling loads of the panels and the prebuckling and postbuckling out-of-plane deformations and strains.

Thornburgh, Robert P.↗
NASA NTRS2006

A Mesh Refinement Study on the Impact Response of a Shuttle Leading-Edge Panel Finite Element Simulation

A study was performed to examine the influence of varying mesh density on an LS-DYNA simulation of a rectangular-shaped foam projectile impacting the space shuttle leading edge Panel 6. The shuttle leading-edge panels are fabricated of reinforced carbon-carbon (RCC) material. During the study, nine cases were executed with all possible combinations of coarse, baseline, and fine meshes of the foam and panel. For each simulation, the same material properties and impact conditions were specified and only the mesh density was varied. In the baseline model, the shell elements representing the RCC panel are approximately 0.2-in. on edge, whereas the foam elements are about 0.5-in. on edge. The element nominal edge-length for the baseline panel was halved to create a fine panel (0.1-in. edge length) mesh and doubled to create a coarse panel (0.4-in. edge length) mesh. In addition, the element nominal edge-length of the baseline foam projectile was halved (0.25-in. edge length) to create a fine foam mesh and doubled (1.0-in. edge length) to create a coarse foam mesh. The initial impact velocity of the foam was 775 ft/s. The simulations were executed in LS-DYNA for 6 ms of simulation time. Contour plots of resultant panel displacement and effective stress in the foam were compared at four discrete time intervals. Also, time-history responses of internal and kinetic energy of the panel, kinetic and hourglass energy of the foam, and resultant contact force were plotted to determine the influence of mesh density.

Fasanella, Edwin L.↗
NASA NTRS2007

Evaluation of a Highly Anticlastic Panel with Tow Overlaps

A rectangular, variable-stiffness panel with tow overlaps was manufactured using an advanced tow placement machine. The cured panel had large anticlastic imperfections, with measured amplitudes of over two times the average panel thickness. These imperfections were not due to the overall steered-fiber layup or the tow overlaps, but instead resulted from local asymmetries in the laminate that were caused by a manufacturing oversight. In the nominal panel layup, fiber angles vary linearly from 60 degrees on the panel axial centerline to 30 degrees on the parallel edges. A geometrically nonlinear analysis was performed with a -280 degree Fahrenheit thermal load to simulate the postcure cooldown to room temperature. The predicted geometric imperfections correlated well with the measured panel shape. Unique structural test fixtures were then developed which greatly reduced these imperfections, but they also caused prestresses in the panel. Surface imperfections measured after the panel was installed in the test fixtures were used with nonlinear finite element analyses to predict these fixturing-induced prestresses. These prestresses were also included in structural analyses of panel end compression to failure, and the analytical results compared well with test data when both geometric and material nonlinearities were included.

Wu, K. Chauncey↗
NASA NTRS2005

Heat exchanger panel

The present invention relates to a heat exchanger panel which has broad utility in high temperature environments. The heat exchanger panel has a first panel, a second panel, and at least one fluid containment device positioned intermediate the first and second panels. At least one of the first panel and the second panel have at least one feature on an interior surface to accommodate the at least one fluid containment device. In a preferred embodiment, each of the first and second panels is formed from a high conductivity, high temperature composite material. Also, in a preferred embodiment, the first and second panels are joined together by one or more composite fasteners.

Warburton, Robert E.↗
NASA NTRS2011

Pressure Testing of a Minimum Gauge PRSEUS Panel

Advanced aircraft configurations that have been developed to increase fuel efficiency require advanced, novel structural concepts capable of handling the unique load conditions that arise. One such concept is the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) developed by the Boeing Company. The PRSEUS concept is being investigated by NASA s Environmentally Responsible Aviation (ERA) Program for use in a hybrid-wing body (HWB) aircraft. This paper summarizes the analysis and test of a PRSEUS panel subjected to internal pressure, the first such pressure test for this structural concept. The pressure panel used minimum gauge skin, with stringer and frame configurations consistent with previous PRSEUS tests. Analysis indicated that for the minimum gauge skin panel, the stringer locations exhibit fairly linear response, but the skin bays between the stringers exhibit nonlinear response. Excellent agreement was seen between nonlinear analysis and test results in the critical portion at the center of the panel. The pristine panel was capable of withstanding the required 18.4 psi pressure load condition without exhibiting any damage. The impacted panel was capable of withstanding a pressure load in excess of 28 psi before initial failure occurred at the center stringer, and the panel was capable of sustaining increased pressure load after the initial failure. This successful PRSEUS panel pressure panel test was a critical step in the building block approach for enabling the use of this advanced structural concept on future aircraft, such as the HWB.

Lovejoy, Andrew J.↗
NASA NTRS2013

Nonlinear Analysis and Post-Test Correlation for a Curved PRSEUS Panel

The Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept, developed by The Boeing Company, has been extensively studied as part of the National Aeronautics and Space Administration's (NASA s) Environmentally Responsible Aviation (ERA) Program. The PRSEUS concept provides a light-weight alternative to aluminum or traditional composite design concepts and is applicable to traditional-shaped fuselage barrels and wings, as well as advanced configurations such as a hybrid wing body or truss braced wings. Therefore, NASA, the Federal Aviation Administration (FAA) and The Boeing Company partnered in an effort to assess the performance and damage arrestments capabilities of a PRSEUS concept panel using a full-scale curved panel in the FAA Full-Scale Aircraft Structural Test Evaluation and Research (FASTER) facility. Testing was conducted in the FASTER facility by subjecting the panel to axial tension loads applied to the ends of the panel, internal pressure, and combined axial tension and internal pressure loadings. Additionally, reactive hoop loads were applied to the skin and frames of the panel along its edges. The panel successfully supported the required design loads in the pristine condition and with a severed stiffener. The panel also demonstrated that the PRSEUS concept could arrest the progression of damage including crack arrestment and crack turning. This paper presents the nonlinear post-test analysis and correlation with test results for the curved PRSEUS panel. It is shown that nonlinear analysis can accurately calculate the behavior of a PRSEUS panel under tension, pressure and combined loading conditions.

Gould, Kevin↗
NASA NTRS2013

Buckling Testing and Analysis of Honeycomb Sandwich Panel Arc Segments of a Full-Scale Fairing Barrel Part 1: 8-Ply In-Autoclave Facesheets: 8-Ply In-Autoclave Facesheets - Part 1

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 Space Launch Systems 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 3.00- by 5.00-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 analysis was performed to predict the compressive response of the 3.00- by 5.00-ft panels. This manuscript summarizes the experimental and analytical modeling efforts pertaining to the panel composed of 8-ply, IM7/977-3 facesheets (referred to Panel A). 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 models yield good qualitative and quantitative predictions. Additionally, it was predicted correctly that the panel would fail in buckling prior to failing in strength. Furthermore, several imperfection studies were performed to investigate the influence of geometric imperfections, fiber misalignments, and three-dimensional (3 D) effects on the compressive response of the panel.

Myers, David E.↗
NASA NTRS2013

Buckling Testing and Analysis of Honeycomb Sandwich Panel Arc Segments of a Full-Scale Fairing Barrel: 6-Ply In-Autoclave Facesheets - Part 2

Four honeycomb sandwich panel types, representing 1/16th arc segments of a 10-m diameter barrel section of the Heavy Lift Launch Vehicle (HLLV), were manufactured and tested 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: in-autoclave 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 (LaRC). Furthermore, linear eigenvalue and geometrically nonlinear finite element analyses were performed to predict the compressive response of each 3- by 5-ft panel. This manuscript summarizes the experimental and analytical modeling efforts pertaining to the panels composed of 6-ply, IM7/977-3 facesheets (referred to as Panels B-1 and B-2). 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 models yield good qualitative and quantitative predictions. Additionally, it was correctly predicted that the panel would fail in buckling prior to failing in strength. Furthermore, several imperfection studies were performed to investigate the influence of geometric imperfections, fiber angle misalignments, and three-dimensional (3-D) effects on the compressive response of the panel.

Pineda, Evan J.↗
NASA NTRS2015

Development of an Aerodynamic Analysis Method and Database for the SLS Service Module Panel Jettison Event Utilizing Inviscid CFD and MATLAB

This paper describes the development, testing, and utilization of an aerodynamic force and moment database for the Space Launch System (SLS) Service Module (SM) panel jettison event. The database is a combination of inviscid Computational Fluid Dynamic (CFD) data and MATLAB code written to query the data at input values of vehicle/SM panel parameters and return the aerodynamic force and moment coefficients of the panels as they are jettisoned from the vehicle. The database encompasses over 5000 CFD simulations with the panels either in the initial stages of separation where they are hinged to the vehicle, in close proximity to the vehicle, or far enough from the vehicle that body interference effects are neglected. A series of viscous CFD check cases were performed to assess the accuracy of the Euler solutions for this class of problem and good agreement was obtained. The ultimate goal of the panel jettison database was to create a tool that could be coupled with any 6-Degree-Of-Freedom (DOF) dynamics model to rapidly predict SM panel separation from the SLS vehicle in a quasi-unsteady manner. Results are presented for panel jettison simulations that utilize the database at various SLS flight conditions. These results compare favorably to an approach that directly couples a 6-DOF model with the Cart3D Euler flow solver and obtains solutions for the panels at exact locations. This paper demonstrates a method of using inviscid CFD simulations coupled with a 6-DOF model that provides adequate fidelity to capture the physics of this complex multiple moving-body panel separation event.

Applebaum, Michael P.↗
NASA NTRS2019

Modeling and Simulation Techniques for the NASA SLS Service Module Panel Separation Event; from Loosely-Coupled Euler to Fully-Coupled 6-DOF, Time-Accurate, Navier-Stokes Methodologies

An aerodynamic database has been generated for use by the Orion Multi-Purpose Crew Vehicle (MPCV) Program to analyze Service Module (SM) panel jettison from the NASA SLS vehicle. The database is a combination of CFD data for the panel aerodynamic coefficients, and MATLAB code written to query the CFD data. The Cart3D inviscid CFD flow solver was used to generate the panel aerodynamic coefficients for static panel orientations and free stream conditions that can occur during the jettison event. The MATLAB code performs the multivariate interpolation to obtain aerodynamic coefficients. The MATLAB code uses input for SM panel parameters and returns the SM panel aerodynamic force and moment coefficients for use with a Six-Degree-of-Freedom (6-DOF) motion solver to model the jettison event. This paper examines the accuracy of the sequential-static database approach by modeling the panel jettison event with a fully-coupled, time-dependent, viscous, moving-body CFD simulation. The fully-coupled simulation is obtained using the Loci/Chem unstructured Navier-Stokes CFD solver. The results show that the fully-coupled approach agrees well with the loosely-coupled database/6-DOF approach, indicating that unsteady effects are minimal for the panel jettison event. These results suggest that the database/6-DOF approach is sufficient. In addition, this paper presents the development of an uncertainty model for use in Monte Carlo analysis of the panel jettison event. Here viscous CFD simulations are obtained with Loci/Chem and compared to the inviscid CFD forces and moments. An uncertainty model based on model-form error and numerical error is presented.

Hall, Leslie H.↗
NASA NTRS2019

Buckling Testing and Analysis of Honeycomb Sandwich Panel Arc Segments of a Full-Scale Fairing Barrel Part 4: Six-ply Out-of-Autoclave Facesheets

Four honeycomb sandwich panel types, representing 1/16th arc segments of a 10-m diameter barrel section of the Heavy Lift Launch Vehicle (HLLV), were manufactured and tested 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: in-autoclave IM7/977-3 and out-of-autoclave T40-800b/5320-1. Smaller 3 ft. 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 (LaRC). Furthermore, linear eigenvalue and geometrically nonlinear finite element analyses were performed to predict the compressive response of each 3 ft. by 5 ft. panel. This manuscript summarizes the experimental and analytical modeling efforts pertaining to the panels composed of 6-ply, T40-800b/5320-1 facesheets (referred to as Panels D). 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 models yield good qualitative and quantitative predictions. Additionally, it was correctly predicted that the panel would fail in buckling prior to failing in strength. Furthermore, three-dimensional (3D) effects on the compressive response of the panel were studied.

David E Myers↗