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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 Orbital Refueling System (ORS); force torque sensor; Plasma Motor/Generator (PMG) proof of function; voice controlled system; infrared intercommunications; superfluid helium on orbit transfer; laser docking sensor; and the Small Expendable Deployment System (SEDS) are summarized.
When the HSR program began there was widespread belief that a simple and familiar turbojet-like engine coupled to an advanced technology mixer-ejector nozzle was the propulsion system of choice for achieving FAR 36-Stage 3 noise requirements. Our ability to quickly demonstrate a practical 20(+) dB suppression nozzle was confidently presumed by many. Our rate of progress towards that objective, however, has been somewhat humbling. At the moment we are reasonably confident of achieving about 15 dB suppression with a mixer-ejector nozzle designed for a high specific thrust turbojet-like cycle. Therefore, if we make no further suppression progress and conservatively assume no new operational procedures such as programmed lapse rate (PLR), then meeting the Stage 3 goal requires a large amount of engine and/or wing oversizing which is economically prohibitive. The scenario is further aggravated by the possibility of eventually needing to comply with even more stringent regulations (Stage 4). While this status may be somewhat disappointing to some, it must be remembered that the HSR program plan involves two generations of mixer-ejector nozzles beyond the current generation I nozzle designs. It is premature to conclude that we cannot design a practical 20(+) dB mixer-ejector nozzle. On the other hand, it is prudent to consider alternative solutions to the noise problem. Thus, we are investigating four other propulsion system concepts.
The primary objective of this research was to determine the processes and feasibility of using commercial off-the-shelf PC104 hardware for flight applications. This would lead to a faster, better, and cheaper approach to low-budget programs as opposed to the design, procurement. and fabrication of space flight hardware. This effort will provide experimental evaluation with results of flight environmental testing. Also, a method and/or suggestion used to bring test hardware up to flight standards will be given. Several microgravity programs, such as the Equiaxed Dendritic Solidification Experiment, Self-Diffusion in Liquid Elements, and various other programs, are interested in PC104 environmental testing to establish the limits of this technology.
The Crew Launch Vehicle (CLV), a key component of NASA's blueprint for the next generation of spacecraft to take humans back to the moon, is being designed and built by engineers at NASA s Marshall Space Flight Center (MSFC). The vehicle s design is based on the results of NASA's 2005 Exploration Systems Architecture Study (ESAS), which called for development of a crew-launch system to reduce the gap between Shuttle retirement and Crew Exploration Vehicle (CEV) Initial Operating Capability, identification of key technologies required to enable and significantly enhance these reference exploration systems, and a reprioritization of near- and far-term technology investments. The Upper Stage Element (USE) of the CLV is a clean-sheet approach that is being designed and developed in-house, with element management at MSFC. The USE concept is a self-supporting cylindrical structure, approximately 115' long and 216" in diameter, consisting of the following subsystems: Primary Structures (LOX Tank, LH2 Tank, Intertank, Thrust Structure, Spacecraft Payload Adaptor, Interstage, Forward and Aft Skirts), Secondary Structures (Systems Tunnel), Avionics and Software, Main Propulsion System, Reaction Control System, Thrust Vector Control, Auxiliary Power Unit, and Hydraulic Systems. The ESAS originally recommended a CEV to be launched atop a four-segment Space Shuttle Main Engine (SSME) CLV, utilizing an RS-25 engine-powered upper stage. However, Agency decisions to utilize fewer CLV development steps to lunar missions, reduce the overall risk for the lunar program, and provide a more balanced engine production rate requirement prompted engineers to switch to a five-segment design with a single Saturn-derived J-2X engine. This approach provides for single upper stage engine development for the CLV and an Earth Departure Stage, single Reusable Solid Rocket Booster (RSRB) development for the CLV and a Cargo Launch Vehicle, and single core SSME development. While the RSRB design has changed since the CLV Project's inception, the USE design has remained essentially a clean-sheet approach. Although a clean-sheet upper stage design inherently carries more risk than a modified design, it does offer many advantages: a design for increased reliability; built-in extensibility to allow for commonality/growth without major redesign; and incorporation of state-of-the-art materials, hardware, and design, fabrication, and test techniques and processes to facilitate a potentially better, more reliable system. Because consideration was given in the ESAS to both clean-sheet and modified USE designs, this paper will highlight the advantages and disadvantages of both approaches and provide a detailed discussion of trades/selections made that led to the final upper stage configuration.
The Upper Stage Element of NASA s Ares I Crew Launch Vehicle (CLV) is a "clean-sheet" approach that is being designed and developed in-house, with Element management at MSFC. The USE concept is a self-supporting cylindrical structure, approximately 115 long and 216" in diameter. While the Reusable Solid Rocket Booster (RSRB) design has changed since the CLV inception, the Upper Stage Element design has remained essentially a clean-sheet approach. Although a clean-sheet upper stage design inherently carries more risk than a modified design, it does offer many advantages: a design for increased reliability; built-in extensibility to allow for commonality/growth without major redesign; and incorporation of state-of-the-art materials, hardware, and design, fabrication, and test techniques and processes to facilitate a potentially better, more reliable system.
An overview is presented of the Mars-Global Reference Atmospheric Model (Mars-GRAM 2005) and its new features. One important new feature is the "auxiliary profile" option, whereby a simple input file is used to replace mean atmospheric values from Mars-GRAM's conventional (General Circulation Model) climatology. An auxiliary profile can be generated from any source of data or alternate model output. Results are presented using auxiliary profiles produced from mesoscale model output (Southwest Research Institute's Mars Regional Atmospheric Modeling System (MRAMS) model and Oregon State University's Mars mesoscale model (MMM5) model) for three candidate Mars Science Laboratory (MSL) landing sites (Terby Crater, Melas Chasma, and Gale Crater). A global Thermal Emission Spectrometer (TES) database has also been generated for purposes of making 'Mars-GRAM auxiliary profiles. This data base contains averages and standard deviations of temperature, density, and thermal wind components, averaged over 5-by-5 degree latitude bins and 15 degree L(sub S) bins, for each of three Mars years of TES nadir data. Comparisons show reasonably good consistency between Mars-GRAM with low dust optical depth and both TES observed and mesoscale model simulated density at the three study sites. Mean winds differ by a more significant degree. Comparisons of mesoscale and TES standard deviations' with conventional Mars-GRAM values, show that Mars-GRAM density perturbations are somewhat conservative (larger than observed variability), while mesoscale-modeled wind variations are larger than Mars-GRAM model estimates. Input parameters rpscale (for density perturbations) and rwscale (for wind perturbations) can be used to "recalibrate" Mars-GRAM perturbation magnitudes to better replicate observed or mesoscale model variability.
The new Mars-GRAM auxiliary profile capability, using data from TES observations, mesoscale model output, or other sources, allows a potentially higher fidelity representation of the atmosphere, and a more accurate way of estimating inherent uncertainty in atmospheric density and winds. Figure 3 indicates that, with nominal value rpscale=1, Mars-GRAM perturbations would tend to overestimate observed or mesoscale-modeled variability. To better represent TES and mesoscale model density perturbations, rpscale values as low as about 0.4 could be used. Some trajectory model implementations of Mars-GRAM allow the user to dynamically change rpscale and rwscale values with altitude. Figure 4 shows that an mscale value of about 1.2 would better replicate wind standard deviations from MRAMS or MMM5 simulations at the Gale, Terby, or Melas sites. By adjusting the rpscale and rwscale values in Mars-GRAM based on figures such as Figure 3 and 4, we can provide more accurate end-to-end simulations for EDL at the candidate MSL landing sites.
Long-duration space flight poses many hazards to the health of the crew. Among those hazards is the physiological deconditioning of the musculoskeletal and cardiovascular systems due to prolonged exposure to microgravity. To combat the physical toll that exploration space flight may take on the crew, NASAs Human Research Program is charged with developing exercise protocols and hardware to maintain astronaut health and fitness during long-term missions. The goal of this effort is to preserve the physical capability of the crew to perform mission critical tasks in transit and during planetary surface operations. As NASA aims toward space travel outside of low-earth orbit (LEO), the constraints placed upon exercise equipment onboard the vehicle increase. Proposed vehicle architectures for transit to and from locations outside of LEO call for limits to equipment volume, mass, and power consumption. While NASA has made great strides in providing for the physical welfare of the crew, the equipment currently used onboard ISS is too large, too massive, and too power hungry to consider for long-duration flight. The goal of the Advanced Exercise Concepts (AEC) project is to maintain the resistive and aerobic capabilities of the current, ISS suite of exercise equipment, while making reductions in size, mass, and power consumption in order to make the equipment suitable for long-duration missions.
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Flash drying selective reduction process for producing dispersion-strengthened cobalt base alloys
There are many similarities between the Mars Surveyor '01 (MS '01) landing site selection process and that of Mars Pathfinder. The selection process includes two parallel activities in which engineers define and refine the capabilities of the spacecraft through design, testing and modeling and scientists define a set of landing site constraints based on the spacecraft design and landing scenario. As for Pathfinder, the safety of the site is without question the single most important factor, for the simple reason that failure to land safely yields no science and exposes the mission and program to considerable risk. The selection process must be thorough, defensible and capable of surviving multiple withering reviews similar to the Pathfinder decision. On Pathfinder, this was accomplished by attempting to understand the surface properties of sites using available remote sensing data sets and models based on them. Science objectives are factored into the selection process only after the safety of the site is validated. Finally, as for Pathfinder, the selection process is being done in an open environment with multiple opportunities for community involvement including open workshops, with education and outreach opportunities.
There are many similarities between the Mars Surveyor '01 (MS '01) landing site selection process and that of Mars Pathfinder. The selection process includes two parallel activities in which engineers define and refine the capabilities of the spacecraft through design, testing and modeling and scientists define a set of landing site constraints based on the spacecraft design and landing scenario. As for Pathfinder, the safety of the site is without question the single most important factor, for the simple reason that failure to land safely yields no science and exposes the mission and program to considerable risk. The selection process must be thorough and defensible and capable of surviving multiple withering reviews similar to the Pathfinder decision. On Pathfinder, this was accomplished by attempting to understand the surface properties of sites using available remote sensing data sets and models based on them. Science objectives are factored into the selection process only after the safety of the site is validated. Finally, as for Pathfinder, the selection process is being done in an open environment with multiple opportunities for community involvement including open workshops, with education and outreach opportunities. Additional information is contained in the original extended abstract.
There are many similarities between the Mars Surveyor '01 (MS '01) landing site selection process and that of Mars Pathfinder. The selection process includes two parallel activities in which engineers define and refine the capabilities of the spacecraft through design, testing and modeling and scientists define a set of landing site constraints based on the spacecraft design and landing scenario. As for Pathfinder, the safety of the site is without question the single most important factor, for the simple reason that failure to land safely yields no science and exposes the mission and program to considerable risk. The selection process must be thorough and defensible and capable of surviving multiple withering reviews similar to the Pathfinder decision. On Pathfinder, this was accomplished by attempting to understand the surface properties of sites using available remote sensing data sets and models based on them. Science objectives are factored into the selection process only after the safety of the site is validated. Finally, as for Pathfinder, the selection process is being done in an open environment with multiple opportunities for community involvement including open workshops, with education and outreach opportunities.
Walsh functions in image processing, rotational feature selection and pattern recognition, defining set of orthogonal transformations
The development of large-scale windpowered systems is outlined. Topics discussed include: prototype specifications development, site selection process, and selection of prototype contractor.
This volume is the first of a three volume set that discusses the structural arrangement trade study plan that will identify the most suitable configuration for an SSTO winged vehicle capable of delivering 25,000 lbs to a 220 nm circular orbit at 51.6 deg inclination. The Reusable Hydrogen Composite Tank System (RHCTS), and Graphite Composite Primary Structures most suitable for intertank, wing and thrust structures are identified. This executive summary presents the trade study process, the selection process, requirements used, analysis performed and data generated. Conclusions and recommendations are also presented.
The integration of many life support systems necessary to construct a stable habitat is difficult. The correct identification of the appropriate technologies and corresponding interfaces is an exhaustive process. Once technologies are selected secondary issues such as mechanical and electrical interfaces must be addressed. The required analytical and testing work must be approached in a piecewise fashion to achieve timely results. A repeatable process has been developed to identify and prioritize system level assessments and testing needs. This Assessment Selection Process has been defined to assess cross cutting integration issues on topics at the system or component levels. Assessments are used to identify risks, encourage future actions to mitigate risks, or spur further studies.
Area selective atomic layer deposition (AS-ALD) is a bottom-up synthesis approach with potential for deposition with molecular level precision. Here, the site-specific hydration of metal oxide substrates, combined with surface H 2 O-selective ALD processes, provides a potentially powerful path to targeted synthesis. Density functional theory (DFT) calculations are used to predict the thermodynamics of ALD precursor reactivity and hydration for (001), (101), (110), and (100) rutile SnO 2 facets as a function of temperature. Trimethylaluminum (TMA) and dimethyl aluminum isopropoxide (DMAI) dimers are predicted to react with both dehydrated and hydrated SnO 2 (001), (101), and (110) facets at ALD-relevant temperatures, while the SnO 2 (100) facet is predicted to be uniquely unreactive with TMA and DMAI monomers as well as dehydrate near 177 °C making this facet more amenable to targeted ALD. In situ ellipsometric studies of Al 2 O 3 ALD on polycrystalline SnO 2 at 150 °C are consistent with the computational predictions of rapid and unselective nucleation, in stark contrast to inhibited and selective ALD on isostructural rutile TiO 2 .