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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 307 records · Page 17
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Detailed analysis and test correlation of a stiffened composite wing panel

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

Davis, D. Dale, Jr.↗
NASA NTRS1991

Analysis and testing of a soft actuation system for segmented reflector articulation and isolation

Segmented reflectors have been proposed for space-based applications such as optical communication and large-diameter telescopes. An actuation system for mirrors in a space-based segmented mirror array has been developed as part of the National Aeronautics and Space Administration-sponsored Precision Segmented Reflector program. The actuation system, called the Articulated Panel Module (APM), articulates a mirror panel in 3 degrees of freedom in the submicron regime, isolates the panel from structural motion, and simplifies space assembly of the mirrors to the reflector backup truss. A breadboard of the APM has been built and is described. Three-axis modeling, analysis, and testing of the breadboard is discussed.

Jandura, Louise↗
NASA NTRS2014

Advanced CMOS Radiation Effects Testing and Analysis

Presentation at the annual NASA Electronic Parts and Packaging (NEPP) Program Electronic Technology Workshop (ETW). The material includes an update of progress in this NEPP task area over the past year, which includes testing, evaluation, and analysis of radiation effects data on the IBM 32 nm silicon-on-insulator (SOI) complementary metal oxide semiconductor (CMOS) process. The testing was conducted using test vehicles supplied by directly by IBM.

cmos↗
NASA NTRS1999

1995 NASA High-Speed Research Program Sonic Boom Workshop: Configuration Design, Analysis, and Testing - Volume 2

The High-Speed Research Program and NASA Langley Research Center sponsored the NASA High-Speed Research Program Sonic Boom Workshop on September 12-13, 1995. The workshop was designed to bring together NASAs scientists and engineers and their counterparts in industry, other Government agencies, and academia working together in the sonic boom element of NASAs High-Speed Research Program. Specific objectives of this workshop were to: (1) report the progress and status of research in sonic boom propagation, acceptability, and design; (2) promote and disseminate this technology within the appropriate technical communities; (3) help promote synergy among the scientists working in the Program; and (4) identify technology pacing, the development C, of viable reduced-boom High-Speed Civil Transport concepts. The Workshop was organized in four sessions: Sessions 1 Sonic Boom Propagation (Theoretical); Session 2 Sonic Boom Propagation (Experimental); Session 3 Acceptability Studies-Human and Animal; and Session 4 - Configuration Design, Analysis, and Testing.

High-speed research↗
NASA NTRS2006

Structural Analysis and Testing of the Inflatable Re-entry Vehicle Experiment (IRVE)

The Inflatable Re-entry Vehicle Experiment (IRVE) is a 3.0 meter, 60 degree half-angle sphere cone, inflatable aeroshell experiment designed to demonstrate various aspects of inflatable technology during Earth re-entry. IRVE will be launched on a Terrier-Improved Orion sounding rocket from NASA s Wallops Flight Facility in the fall of 2006 to an altitude of approximately 164 kilometers and re-enter the Earth s atmosphere. The experiment will demonstrate exo-atmospheric inflation, inflatable structure leak performance throughout the flight regime, structural integrity under aerodynamic pressure and associated deceleration loads, thermal protection system performance, and aerodynamic stability. Structural integrity and dynamic response of the inflatable will be monitored with photogrammetric measurements of the leeward side of the aeroshell during flight. Aerodynamic stability and drag performance will be verified with on-board inertial measurements and radar tracking from multiple ground radar stations. In addition to demonstrating inflatable technology, IRVE will help validate structural, aerothermal, and trajectory modeling and analysis techniques for the inflatable aeroshell system. This paper discusses the structural analysis and testing of the IRVE inflatable structure. Equations are presented for calculating fabric loads in sphere cone aeroshells, and finite element results are presented which validate the equations. Fabric material properties and testing are discussed along with aeroshell fabrication techniques. Stiffness and dynamics tests conducted on a small-scale development unit and a full-scale prototype unit are presented along with correlated finite element models to predict the in-flight fundamental mod

Lindell, Michael C.↗
NASA NTRS1991

Detailed analysis and test correlation of a stiffened composite wing panel

New finite element analysis methods are examined by application to a complicated composite wing panel from the V-22 rotorcraft. A detailed FEM model with a relatively coarse mesh of 9-node elements was generated, and linear and nonlinear stress analyses, first-ply failure analyses, and buckling analyses were conducted. At low values of applied load, i.e., up to the design ultimate load of the panel, the linear stress analysis accurately predicted the strains and structural response characteristics of the panel.

Davis, D. D., Jr.↗
NASA NTRS2017

Arc Jet Test and Analysis of Asbestos Free Solid Rocket Motor Nozzle Dome Ablative Materials

Asbestos free solid motor internal insulation samples were recently tested at the MSFC Hyperthermal Arc Jet Facility. Objectives of the test were to gather data for solid rocket motor analog characterization of ablative and in-depth thermal performance of rubber materials subject to high enthalpy/pressure flow conditions. Tests were conducted over a range of convective heat fluxes for both inert and chemically reactive sub-sonic free stream gas flow. Active instrumentation included use of total calorimeters, in-depth thermocouples, and a surface pyrometer for in-situ surface temperature measurement. Post-test sample forensics involved determination of eroded depth, charred depth, total sample weight loss, and documentation of the general condition of the eroded profile. A complete Charring Material Ablator (CMA) style aero thermal analysis was conducted for the test matrix and results compared to the measured data. In general, comparisons were possible for a number of the cases and the results show a limited predictive ability to model accurately both the ablative response and the in-depth temperature profiles. Lessons learned and modeling recommendations are made regarding future testing and modeling improvements that will increase understanding of the basic chemistry/physics associated with the complicated material ablation process of rubber materials.

Clayton, J. Louie↗
NASA NTRS2016

Analysis and Testing of a Composite Fuselage Shield for Open Rotor Engine Blade-Out Protection

The Federal Aviation Administration is working with the European Aviation Safety Agency to determine the certification base for proposed new engines that would not have a containment structure on large commercial aircraft. Equivalent safety to the current fleet is desired by the regulators, which means that loss of a single fan blade will not cause hazard to the Aircraft. The NASA Glenn Research Center and The Naval Air Warfare Center (NAWC), China Lake, collaborated with the FAA Aircraft Catastrophic Failure Prevention Program to design and test lightweight composite shields for protection of the aircraft passengers and critical systems from a released blade that could impact the fuselage. LS-DYNA® was used to predict the thickness of the composite shield required to prevent blade penetration. In the test, two composite blades were pyrotechnically released from a running engine, each impacting a composite shield with a different thickness. The thinner shield was penetrated by the blade and the thicker shield prevented penetration. This was consistent with pre-test LS-DYNA predictions. This paper documents the analysis conducted to predict the required thickness of a composite shield, the live fire test from the full scale rig at NAWC China Lake and describes the damage to the shields as well as instrumentation results.

Composite↗
NASA NTRS1994

Testing and analysis of flat and curved panels with multiple cracks

An experimental and analytical investigation of multiple cracking in various types of test specimens is described in this paper. The testing phase is comprised of a flat unstiffened panel series and curved stiffened and unstiffened panel series. The test specimens contained various configurations for initial damage. Static loading was applied to these specimens until ultimate failure, while loads and crack propagation were recorded. This data provides the basis for developing and validating methodologies for predicting linkup of multiple cracks, progression to failure, and overall residual strength. The results from twelve flat coupon and ten full scale curved panel tests are presented. In addition, an engineering analysis procedure was developed to predict multiple crack linkup. Reasonable agreement was found between predictions and actual test results for linkup and residual strength for both flat and curved panels. The results indicate that an engineering analysis approach has the potential to quantitatively assess the effect of multiple cracks in the arrest capability of an aircraft fuselage structure.

Broek, David↗
NASA NTRSDate not supplied

Design, Instrumentation, and Data Analysis for the SLS Core Stage Green Run Test Series

The Space Launch System (SLS) Core Stage (CS) Thrust Vector Control (TVC) system is comprised of eight mechanical feedback Shuttle heritage Type III TVC actuators and four RS-25 engines, each attached to a Shuttle heritage gimbal block/bearing. The actuators are powered by a Shuttle-derived hydraulic Core Auxiliary Power Unit (CAPU), and integrated with an all-new Core Stage thrust structure. The actuators are interfaced to the SLS Vehicle Management (VM) software via an all-new TVC Actuator Control (TAC) avionics subsystem. Despite the significant test and flight experience of the Shuttle hardware, the SLS Green Run ambient and hot fire test activities revealed a number of new findings associated with the dynamic response of the TVC integrated system. Test responses suggested that the TVC system did not meet its performance specifications and its step and frequency responses exhibited unexpected departures from prior lab tests and modeled behavior. This paper is the fifth installment in a seven-paper series surveying the design, engineering, test validation, and flight performance of the Core Stage Thrust Vector Control system. In this paper, the design of the TVC analyses conducted during the Core Stage Green Run test series are discussed in detail. Throughout the course of the test activities, the SLS flight control team worked diligently with the Core Stage contractor to revise test command profiles and ensure sufficient instrumentation was available to collect data. Post-test analysis combined the Green Run modal, ambient, and hot fire test data, MSFC 2- axis Core Stage TVC Inertial Load Simulator (ILS) data, Hardware-In-the-Loop (HWIL) Systems Integration Lab (SIL) results, and actuator Acceptance Testing Procedure (ATP) responses. These data were used to characterize the response, validate critical math models of the TVC subsystem, and isolate the probable cause of the unexpected responses. Through comprehensive analysis of the available test data sources, the integrated team identified the dominant contributors to the observed response and developed test-correlated rationale for vehicle flight control system performance, ultimately leading to a confident posture for the Artemis I mission.

John H. Wall↗
NASA NTRS2005

Test and Analysis of a Hyper-X Carbon-Carbon Leading Edge Chine

During parts production for the X43A Mach 10 hypersonic vehicle nondestructive evaluation (NDE) of a leading edge chine detected on imbedded delamination near the lower surface of the part. An ultimate proof test was conducted to verify the ultimate strength of this leading edge chine part. The ultimate proof test setup used a pressure bladder design to impose a uniform distributed pressure field over the bi-planar surface of the chine test article. A detailed description of the chine test article and experimental test setup is presented. Analysis results from a linear status model of the test article are also presented and discussed. Post-test inspection of the specimen revealed no visible failures or areas of delamination.

Smith, Russell W.↗
NASA NTRS1977

Substructure coupling for dynamic analysis and testing

Fixed interface and free interface methods of substructure coupling for dynamic analysis are discussed. Three methods for reducing the number of coordinates required by fixed interface methods are introduced. Matrix ordinary differential equations are employed to improve accuracy in free interface substructure coupling methods.

Craig, R. R., Jr.↗
NASA NTRS2012

TPSAS-NF1676L-13678-DND

An area that shows promise in enhancing structural integrity of aircraft and aerospace structures is integral stitched composite technology. The most recent generation of this technology is the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept. The goal of the PRSEUS concept is to provide damage-containment capability for composite structures while reducing overall structural weight. The National Aeronautics and Space Administration, the Boeing Company, and the Federal Aviation Administration (FAA) have partnered in an effort to assess the damage-containment features of a full-scale curved PRSEUS panel using the FAA Full-Scale Aircraft Structural Test Evaluation and Research facility. The background, test plan, fixture modifications, pre-test analysis, and planned experimental procedure were presented at the 2011 Aircraft Airworthiness and Sustainment Conference. This follow-on paper and presentation will provide details of the experimental procedure, test results, nondestructive inspection results, and preliminary test analysis comparison. The test program included three phases of loading and inspections: Phase I, as-built; Phase II, with barely visible impact damage (BVID); and Phase III, with a two-bay notch severing the central stiffener. Axial tension, internal pressure, and combined axial tension and internal pressure load conditions were applied during each phase using load levels that demonstrate compliance with the strength, deformation, and damage-tolerance requirements of Title 14 Code of Federal Regulations Part 25. Pressure loads were based on an operating pressure of 9.2 psi, designated as 1P, and the axial loads were based on a design limit load (DLL) of 227 kip. The Phase I test results provided a baseline. First, 50 percent limit-load levels were applied for each of the three load cases to verify proper load introduction and repeatability by examining strain and displacement results. Next, three limit-load cases were applied: limit pressure of 12.2 psi (1.33 P), axial DLL of 227 kip, and combined 1P pressure and axial DLL. Linear strain and displacement results were observed in all load cases. For Phase-II testing, the panel was impacted to introduce BVID with an energy of 40 ft-lb, with the impact located between the central stringer stitch line and flange edge. The impact created non-penetrating visual damage with fiber breaks and delamination between the skin and stringer flange. Limit-load conditions were applied followed by pressure overload of 18.4 psi (2P) and ultimate load conditions, defined as 1.5 times DLL. Inspections were performed after each load case; no damage growth was observed. For Phase-III testing, a two-bay notch was machined severing the central stiffener. The panel was then subjected to limit-load conditions, followed by combined 1P pressure while increasing axial tension load to catastrophic failure. As limit combined loads were approached, damage initiation was observed in the form of 45? splitting at 1P + 79% DLL. Axial loading was increased to 1P + 150 percent DLL causing damage to propagate to the adjacent stringer flanges. Damage was contained within the two-bay region by the stitching rows up to 1P + 176% DLL. Axial load was further increased until catastrophic failure, when a stringer rod failed at 1P + 192% DLL. These test results further verify the damage containment features of the PRSEUS concept and suggest its appropriateness for future flight vehicles.

Andrew Bergan↗
NASA NTRS2012

Stress Analysis and Testing at the Marshall Space Flight Center to Study Cause and Corrective Action of Space Shuttle External Tank Stringer Failures

After the launch scrub of Space Shuttle mission STS-133 on November 5, 2010, large cracks were discovered in two of the External Tank intertank stringers. The NASA Marshall Space Flight Center, as managing center for the External Tank Project, coordinated the ensuing failure investigation and repair activities with several organizations, including the manufacturer, Lockheed Martin. To support the investigation, the Marshall Space Flight Center formed an ad-hoc stress analysis team to complement the efforts of Lockheed Martin. The team undertook six major efforts to analyze or test the structural behavior of the stringers. Extensive finite element modeling was performed to characterize the local stresses in the stringers near the region of failure. Data from a full-scale tanking test and from several subcomponent static load tests were used to confirm the analytical conclusions. The analysis and test activities of the team are summarized. The root cause of the stringer failures and the flight readiness rationale for the repairs that were implemented are discussed.

Wingate, Robert J.↗