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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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The computer programs developed to simulate the RSECS (Representative Shuttle Environmental Control System) were described. These programs were prepared to provide pretest predictions, post-test analysis and real time problem analysis for RSECS test planning and evaluation.
The dynamic verification of spacecraft relies heavily on ground-based tests. These tests usually simulate flight environments or validate analytical models used in establishing design loads and in designing control algorithms. They also provide security against failures resulting from unanticipated or unmodeled hardware behavior. Future orbital antennas, space stations, and solar power systems are likely to be of sizes difficult to test using current ground test technology. In addition to size, other factors such as low natural frequencies, lightweight construction, and the presence of many structural joints, cause significant sensitivity of the test process to the earth-gravity environment. Yet, accuracy requirements on the verification process will be more stringent because of modern flexible-structure control approaches. This paper describes some of the problems and discusses research on potential solutions. The importance of an integrated ground test, analysis, and flight test program is emphasized. An ongoing research program of this type focusing on a 60-meter, deployable, truss-beam test article is described.
The SCD 01 (Satelite de Coleta de Dados 01) is a spin stabilized low Earth orbit satellite dedicated to the collection and distribution of environmental data. It was completely developed at the Brazilian Institute for Space Research (INPE) and is scheduled to be launched in 1992. The SCD 01 passive thermal control design configuration is presented and the thermal analysis results are compared with the temperatures obtained from a Thermal Balance Test. The correlation between the analytical and experimental results is considered very good. Numerical flight simulations show that the thermal control design can keep all the subsystem temperatures within their specified temperature range.
For vacuum-system or test-article analysis it is often desirable to know the species and partial pressures of the vacuum gases. Residual gas or Partial Pressure Analyzers (PPA's) are commonly used for this purpose. These are mass spectrometer-type instruments, most commonly employing quadrupole filters. These instruments can be extremely useful, but they should be used with caution. Depending on the instrument design, calibration procedures, and conditions of use, measurements made with these instruments can be accurate to within a few percent, or in error by two or more orders of magnitude. Significant sources of error can include relative gas sensitivities that differ from handbook values by an order of magnitude, changes in sensitivity with pressure by as much as two orders of magnitude, changes in sensitivity with time after exposure to chemically active gases, and the dependence of the sensitivity for one gas on the pressures of other gases. However, for most instruments, these errors can be greatly reduced with proper operating procedures and conditions of use. In this paper, data are presented illustrating performance characteristics for different instruments and gases, operating parameters are recommended to minimize some errors, and calibrations procedures are described that can detect and/or correct other errors.
Determining deterioration characteristics of the Space Shuttle crew escape system pyrotechnic components loaded with hexanitrostilbene would enable us to establish a hardware life-limit for these items, so we could better plan our equipment use and, possibly, extend the useful life of the hardware. We subjected components to accelerated-age environments to determine degradation characteristics and established a hardware life-limit based upon observed and calculated trends. We extracted samples using manufacturing lots currently installed in the Space Shuttle crew escape system and from other NASA programs. Hardware included in the study consisted of various forms and ages of mild detonating fuse, linear shaped charge, and flexible confined detonating cord. The hardware types were segregated into 5 groups. One was subjected to detonation velocity testing for a baseline. Two were first subjected to prolonged 155 F heat exposure, and the other two were first subjected to 255 F, before undergoing detonation velocity testing and/or chromatography analysis. Test results showed no measurable changes in performance to allow a prediction of an end of life given the storage and elevated temperature environments the hardware experiences. Given the lack of a definitive performance trend, coupled with previous tests on post-flight Space Shuttle hardware showing no significant changes in chemical purity or detonation velocity, we recommend a safe increase in the useful life of the hardware to 20 years, from the current maximum limits of 10 and 15 years, depending on the hardware.
Linear thermal expansion measurements of nine samples of Lead Magnesium Niobate (PMN) electroceramic material were recently performed in support of NASA's Terrestrial Planet Finder Coronagraph (TPF-C) mission. The TPF-C mission is a visible light coronagraph designed to look at roughly 50 stars pre- selected as good candidates for possessing earth-like planets. Upon detection of an earth-like planet, TPF-C will analyze the visible-light signature of the planet's atmosphere for specific spectroscopic indicators that life may exist there. With this focus, the project's primary interest in PMN material is for use as a solid-state actuator for deformable mirrors or compensating optics. The nine test samples were machined from three distinct boules of PMN ceramic manufactured by Xinetics Inc. Thermal expansion measurements were performed in 2005 at NASA Jet Propulsion Laboratory (JPL) in their Cryogenic Dilatometer Facility. All measurements were performed in vacuum with sample temperature actively controlled over the range of 270K to 3 10K. Expansion and contraction of the test samples with temperature was measured using a JPL developed interferometric system capable of sub-nanometer accuracy. Presented in this paper is a discussion of the sample configuration, test facilities, test method, data analysis, test results, and future plans.
Tests were run both in forward and in reverse thrust modes with a bellmouth inlet, five accelerating inlets (one hard wall and four treated) with a design throat Mach number of 0.79 at the takeoff condition, and four low Mach inlets (one hard wall and three treated) with a design throat Mach number of 0.6 at the takeoff condition. Unsuppressed and suppressed inlet radiated noise levels were measured at conditions representative of QCSEE takeoff, approach, and reverse thrust operations. Measured aerodynamic performance of the accelerating inlet is also included. The test objectives, facility, configurations, are described as well as the data analysis, results, and comparisons.
Activities in modifying a structural mathematical model to correlate with its modal test results are summarized. Results from the Viking and Galileo spacecraft, and developments in the parameter estimation of structural mathematical models of large flexible structures using multiple boundary condition tests are presented.
This paper presents an assessment of the aerothermodynamic environment around an 8.3 meter High Energy Atmospheric Reentry Test (HEART) vehicle. This study generated twelve nose shape configurations and compared their responses at the peak heating trajectory point against the baseline nose shape. The heat flux sensitivity to the angle of attack variations are also discussed. The possibility of a two-piece Thermal Protection System (TPS) design at the nose is also considered, as are the surface catalytic affects of the aeroheating environment of such configuration. Based on these analyses, an optimum nose shape is proposed to minimize the surface heating. A recommendation is also made for a two-piece TPS design, for which the surface catalytic uncertainty associated with the jump in heating at the nose-IAD juncture is reduced by a minimum of 93%. In this paper, the aeroshell is assumed to be rigid and the inflatable fluid interaction effect is left for future investigations
To provide NASA with a suite of materials strength, fracture toughness and crack growth rate test results for use in remaining life calculations for the vessels described above, Southwest Research Institute® (SwRI®) was contracted in two phases to obtain relevant material property data from a representative vessel. An initial characterization of the strength, fracture and fatigue crack growth properties was performed in Phase 1. Based on the results and recommendations of Phase 1, a more extensive material property characterization effort was developed in this Phase 2 effort. This Phase 2 characterization included additional strength, fracture and fatigue crack growth of the multilayer vessel and head materials. In addition, some more limited characterization of the welds and heat affected zones (HAZs) were performed. This report
Elastic-plastic finite element analyses of room temperature burst tests on four superalloy disks were conducted and reported in this paper. Two alloys, Rene 104 (General Electric Aircraft Engines) and Alloy 10 (Honeywell Engines & Systems), were studied. For both alloys an advanced dual microstructure disk, fine grain bore and coarse grain rim, were analyzed and compared with conventional disks with uniform microstructures, coarse grain for Rene 104 and fine grain for Alloy 10. The analysis and experimental data were in good agreement up to burst. At burst, the analysis underestimated the speed and growth of the Rene 104 disks, but overestimated the speed and growth of the Alloy 10 disks. Fractography revealed that the Alloy 10 disks displayed significant surface microcracking and coalescence in comparison to Rene 104 disks. This phenomenon may help explain the differences between the Alloy 10 disks and the Rene 104 disks, as well as the observed deviations between analytical and experimental data at burst.
It is necessary to define the launch vehicle buffet loads to ensure that structural components and vehicle subsystems possess adequate strength, stress, and fatigue margins when the vehicle structural dynamic response to buffet forcing functions are considered. In order to obtain these forcing functions, the accepted method is to perform wind-tunnel testing of a rigid model instrumented with hundreds of unsteady pressure transducers designed to measure the buffet environment across the desired frequency range. The buffet wind-tunnel test program for the Ares Crew Launch Vehicle employed 3.5 percent scale rigid models of the Ares I and Ares I-X launch vehicles instrumented with 256 unsteady pressure transducers each. These models were tested at transonic conditions at the Transonic Dynamics Tunnel at NASA Langley Research Center. The ultimate deliverable of the Ares buffet test program are buffet forcing functions (BFFs) derived from integrating the measured fluctuating pressures on the rigid wind-tunnel models. These BFFs are then used as input to a multi-mode structural analysis to determine the vehicle response to buffet and the resulting buffet loads and accelerations. This paper discusses the development of the Ares I and I-X rigid buffet model test programs from the standpoint of model design, instrumentation system design, test implementation, data analysis techniques to yield final products, and presents normalized sectional buffet forcing function root-mean-squared levels.
Analysis of heat transfer data obtained during Saturn 5/Skylab multiple protuberance heat transfer test
Combustion and emissions testing of United Technologies Research Center’s (UTRC’s) small core combustor in NASA GRC’s CE-5 flametube test facility comprising of LTO and cruise conditions has been completed. Sufficient NOx emissions data was collected to formulate correlation equations for assessing reductions in NOx for LTO against CAEP/6 and cruise against 2005 best-in-class. Back-to-back testing of Jet-A fuel and a 50/50 blend of Jet-A with an alternative fuel demonstrated a measurable effect of fuel composition on non-volatile particulate matter emissions. Post-test analysis indicates LTO NOx emissions >80% below CAEP6 in UTRC single-sector tests of small-core N+3 combustor. Documentation describing N+3 combustor concept selected, design approach, initial flametube test & analysis results; and preferred concept recommended for higher fidelity research has also been completed. This Lean Burn combustor concept shows potential to meet/exceed NASA’s Far Term LTO NOX reduction goals, but further investigation of dynamic characteristics of the combustor concept are required to ensure good operability over the entire flight regime conditions.
One of the primary goals of the Asteroid Threat Assessment Program (ATAP) is to understand the ablation phenomena of meteors (e.g., spallation, evaporation, melt flow, etc.). Therefore, a test campaign was executed at the Hypersonic Material Test System (HYMETS) facility on samples of an ordinary chondrite (e.g., Tamdakht) and samples of basalt from the Columbia plateau. High melt viscosities and low ablation rates are predicted for basalt due to higher concentrations of silica and alumina in the matrix phase. However, high-speed video and post-test analysis reveal that basalt experiences greater rates of ablation than Tamdakht. Post-test analysis of emission spectra reveals the presence of volatile species that may explain the discrepancy between model predictions and observed phenomena.
Finite difference computer techniques have been used to calculate the detailed performance of air to air two-dimensional ejectors with symmetric variable area mixing sections and coaxial coverging primary nozzles. The analysis of the primary nozzle assumed correct expansion of the flow and is suitable for subsonic and slightly supersonic velocity levels. The variation of the mixing section channel walls is assumed to be gradual so that the static pressure can be assumed uniform on planes perpendicular to the axis. A test program was run to provide two-dimensional ejector test data for verification of the computer analysis. A primary converging nozzle with a discharge geometry of 0.125 inch x 8.0 inch was supplied with 600 SCFM of air at about 35 psia and 180 F. This nozzle was combined with two mixing section geometries with throat sizes of 1.25 inch x 8.0 inch and 1.875 inch x 8.0 inch and was tested at a total of 11 operating points.
Finite difference computer techniques have been used to calculate the detailed performance of air-to-air two-dimensional ejectors with symmetric variable area mixing sections and coaxial converging primary nozzles. The analysis of the primary nozzle assumed correct expansion of the flow and is suitable for subsonic and slightly supersonic velocity levels. The variation of the mixing section channel walls is assumed to be gradual so that the static pressure can be assumed uniform on planes perpendicular to the axis. A test program was run to provide two-dimensional ejector test data for verification of the computer analysis. A primary converging nozzle with a discharge geometry of 3.18 mm x 203 mm was supplied with 0.340 kg/sec of air at about 2.43 bar and 356 K. This nozzle was combined with two mixing section geometries with throat sizes of 31.8 mm x 203 mm and 47.6 mm x 203 mm and was tested at a total of 11 operating points.
A stitched RFI graphite-epoxy panel with a fuel access door was analyzed using a finite element analysis and loaded to failure in compression. The panel was initially 56-inches long and 36.75-inches wide and the oval access door was 18-inches long and 15-inches wide. The panel was impact damaged with impact energy of 100 ft-lb prior to compressive loading; however, no impact damage was detectable visually or by A-scan. The panel carried a failure load of 695,000 Ib and global failure strain of .00494 in/in. Analysis indicated the panel would fail due to collapse at a load of 688,100 Ib. The test data indicate that the maximum strain in a region near the access door was .0096 in/in and analysis indicates a local surface strain of .010 in/in at the panel's failure load. The panel did not fail through the impact damage, but instead failed through bolt holes for attachment of the access door in a region of high strain.