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  • Spacecraft Propulsion and Power  (208)
  • Deutschland
  • Fisheries
  • Inorganic Chemistry
  • Organic Chemistry
  • 2000-2004  (210)
  • 2002  (210)
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  • 2000-2004  (210)
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  • 1
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    Support of marine sustainability science (2002)
    American Association for the Advancement of Science (AAAS)
    Details
    Publication Date: 2002-08-07
    Description: 〈br /〉〈span class="detail_caption"〉Notes: 〈/span〉Ajayi, Thomas -- Sherman, Kenneth -- Tang, Qisheng -- New York, N.Y. -- Science. 2002 Aug 2;297(5582):772.〈br /〉〈span class="detail_caption"〉Record origin:〈/span〉 〈a href="http://www.ncbi.nlm.nih.gov/pubmed/12162321" target="_blank"〉PubMed〈/a〉
    Keywords: Biomass ; Conservation of Natural Resources/*economics/*methods/trends ; *Ecosystem ; Europe ; Fisheries ; International Cooperation ; *Marine Biology/economics/trends ; North America ; Water Pollution/prevention & control
    Print ISSN: 0036-8075
    Electronic ISSN: 1095-9203
    Topics: Biology , Chemistry and Pharmacology , Computer Science , Medicine , Natural Sciences in General , Physics
    Published by American Association for the Advancement of Science (AAAS)
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    PAPER CURRENT
  • 2
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    A general model for designing networks of marine reserves (2002)
    American Association for the Advancement of Science (AAAS)
    Details
    Publication Date: 2002-12-10
    Description: There is debate concerning the most effective conservation of marine biodiversity, especially regarding the appropriate location, size, and connectivity of marine reserves. We describe a means of establishing marine reserve networks by using optimization algorithms and multiple levels of information on biodiversity, ecological processes (spawning, recruitment, and larval connectivity), and socioeconomic factors in the Gulf of California. A network covering 40% of rocky reef habitat can fulfill many conservation goals while reducing social conflict. This quantitative approach provides a powerful tool for decision-makers tasked with siting marine reserves.〈br /〉〈span class="detail_caption"〉Notes: 〈/span〉Sala, Enric -- Aburto-Oropeza, Octavio -- Paredes, Gustavo -- Parra, Ivan -- Barrera, Juan C -- Dayton, Paul K -- New York, N.Y. -- Science. 2002 Dec 6;298(5600):1991-3.〈br /〉〈span class="detail_caption"〉Author address: 〈/span〉Center for Marine Biodiversity and Conservation, Scripps Institution of Oceanography, La Jolla, CA 92093, USA. esala@ucsd.edu〈br /〉〈span class="detail_caption"〉Record origin:〈/span〉 〈a href="http://www.ncbi.nlm.nih.gov/pubmed/12471258" target="_blank"〉PubMed〈/a〉
    Keywords: Animals ; California ; Computer Simulation ; *Conservation of Natural Resources ; *Ecosystem ; Environment ; Fisheries ; Fishes ; Invertebrates ; *Models, Biological ; *Seawater
    Print ISSN: 0036-8075
    Electronic ISSN: 1095-9203
    Topics: Biology , Chemistry and Pharmacology , Computer Science , Medicine , Natural Sciences in General , Physics
    Published by American Association for the Advancement of Science (AAAS)
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  • 3
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    Overview of Photovoltaic Calibration and Measurement Standards at GRC (2002)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: Photovoltaic (PV) systems (cells and arrays) for spacecraft power have become an international market. This market demands accurate prediction of the solar array power output in space throughout the mission life of the spacecraft. Since the beginning of space flight, space-faring nations have independently developed methods to calibrate solar cells for power output in low Earth orbit (LEO). These methods rely on terrestrial, laboratory, or extraterrestrial light sources to simulate or approximate the air mass zero (AM0) solar intensity and spectrum.
    Keywords: Spacecraft Propulsion and Power
    Type: 17th Space Photovoltaic Research and Technology Conference; 101-104; NASA/CP-2002-211831
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  • 4
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    Measuring Rocket Engine Temperatures with Hydrogen Raman Spectroscopy (2002)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: Laser-based combustion diagnostics, such as single-pulse UV Raman spectroscopy and visible Raman spectroscopy, have been successfully applied to optically-accessible rocket-like test articles. If an independent pressure measurement is available, Raman major species concentration measurements can also provide a temperature measurement. However it is desirable to obtain a Raman-derived temperature measurement without the need for simultaneous pressure measurement, especially when chamber pressure may vary spatially. This report describes Raman temperature measurements obtained by exploiting the variation in shape of the H2 Raman spectrum. Hydrogen is advantageous since it is ubiquitous in H2-O2 systems and its Raman spectrum is simpler than for other diatomics. However the influence of high pressure on the H2 Raman spectrum must be investigated. At moderate pressures, well below those of rocket engines, the Raman spectra of O2 and N2 are known to become featureless due to collisional broadening.
    Keywords: Spacecraft Propulsion and Power
    Type: Research Reports: 2001 NASA/ASEE Summer Faculty Fellowship Program; LII-1 - LII-5; NASA/CR-2002-211840
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  • 5
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    Multi-Zone Liquid Thrust Chamber Performance Code with Domain Decomposition for Parallel Processing (2002)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: Computational Fluid Dynamics (CFD) has considerably evolved in the last decade. There are many computer programs that can perform computations on viscous internal or external flows with chemical reactions. CFD has become a commonly used tool in the design and analysis of gas turbines, ramjet combustors, turbo-machinery, inlet ducts, rocket engines, jet interaction, missile, and ramjet nozzles. One of the problems of interest to NASA has always been the performance prediction for rocket and air-breathing engines. Due to the complexity of flow in these engines it is necessary to resolve the flowfield into a fine mesh to capture quantities like turbulence and heat transfer. However, calculation on a high-resolution grid is associated with a prohibitively increasing computational time that can downgrade the value of the CFD for practical engineering calculations. The Liquid Thrust Chamber Performance (LTCP) code was developed for NASA/MSFC (Marshall Space Flight Center) to perform liquid rocket engine performance calculations. This code is a 2D/axisymmetric full Navier-Stokes (NS) solver with fully coupled finite rate chemistry and Eulerian treatment of liquid fuel and/or oxidizer droplets. One of the advantages of this code has been the resemblance of its input file to the JANNAF (Joint Army Navy NASA Air Force Interagency Propulsion Committee) standard TDK code, and its automatic grid generation for JANNAF defined combustion chamber wall geometry. These options minimize the learning effort for TDK users, and make the code a good candidate for performing engineering calculations. Although the LTCP code was developed for liquid rocket engines, it is a general-purpose code and has been used for solving many engineering problems. However, the single zone formulation of the LTCP has limited the code to be applicable to problems with complex geometry. Furthermore, the computational time becomes prohibitively large for high-resolution problems with chemistry, two-equation turbulence model, and two-phase flow. To overcome these limitations, the LTCP code is rewritten to include the multi-zone capability with domain decomposition that makes it suitable for parallel processing, i.e., enabling the code to run every zone or sub-domain on a separate processor. This can reduce the run time by a factor of 6 to 8, depending on the problem.
    Keywords: Spacecraft Propulsion and Power
    Type: Research Reports: 2001 NASA/ASEE Summer Faculty Fellowship Program; XXXV-1 - XXXV-5; NASA/CR-2002-211840
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  • 6
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    3-D Visualization in Support of Advanced Propulsion (2002)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: The tools and techniques of three-dimensional computer imaging and animation are more than just a bag of new tricks. They have the power to communicate, inspire, and move the minds of people. Through these animations, it is the intent of the author to help the Propulsion Research Center educate and inspire the public about the vast possibilities of space exploration using Fission Electric Propulsion systems.
    Keywords: Spacecraft Propulsion and Power
    Type: Research Reports: 2001 NASA/ASEE Summer Faculty Fellowship Program; XXVIII-1 - XXVIII-5; NASA/CR-2002-211840
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  • 7
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    Advanced Computing Technologies for Rocket Engine Propulsion Systems: Object-Oriented Design with C++ (2002)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: This document explores the use of advanced computer technologies with an emphasis on object-oriented design to be applied in the development of software for a rocket engine to improve vehicle safety and reliability. The primary focus is on phase one of this project, the smart start sequence module. The objectives are: 1) To use current sound software engineering practices, object-orientation; 2) To improve on software development time, maintenance, execution and management; 3) To provide an alternate design choice for control, implementation, and performance.
    Keywords: Spacecraft Propulsion and Power
    Type: Research Reports: 2001 NASA/ASEE Summer Faculty Fellowship Program; V-1 - V-22; NASA/CR-2002-211840
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  • 8
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    Magnetized Target Fusion (2002)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: Magnetized target fusion (MTF) is under consideration as a means of building a low mass, high specific impulse, and high thrust propulsion system for interplanetary travel. This unique combination is the result of the generation of a high temperature plasma by the nuclear fusion process. This plasma can then be deflected by magnetic fields to provide thrust. Fusion is initiated by a small traction of the energy generated in the magnetic coils due to the plasma's compression of the magnetic field. The power gain from a fusion reaction is such that inefficiencies due to thermal neutrons and coil losses can be overcome. Since the fusion reaction products are directly used for propulsion and the power to initiate the reaction is directly obtained from the thrust generation, no massive power supply for energy conversion is required. The result should be a low engine mass, high specific impulse and high thrust system. The key is to successfully initiate fusion as a proof-of-principle for this application. Currently MSFC is implementing MTF proof-of-principle experiments. This involves many technical details and ancillary investigations. Of these, selected pertinent issues include the properties, orientation and timing of the plasma guns and the convergence and interface development of the "pusher" plasma. Computer simulations of the target plasma's behavior under compression and the convergence and mixing of the gun plasma are under investigation. This work is to focus on the gun characterization and development as it relates to plasma initiation and repeatability.
    Keywords: Spacecraft Propulsion and Power
    Type: Research Reports: 2001 NASA/ASEE Summer Faculty Fellowship Program; XVIII-1 - XVIII-6; NASA/CR-2002-211840
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  • 9
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    Integrated Nozzle Design for the GTX RBCC Flowpath (2002)
    In:  Other Sources
    Details
    Publication Date: 2011-08-23
    Description: The development of rocket based combined cycle (RBCC) engines are highly dependent upon integrating several different modes of operation into a single system. Due to the integrated nature of the propulsion system, each operating mode relies on the same expansion system to provide thrust. A fixed geometry, altitude-compensating aft-expansion configuration is used for the GTX flowpath configuration. Initial studies on the GTX expansion designs have demonstrated the importance of a smooth, highly integrated design for propulsion system performance. Based upon the results from the initial studies, further design improvements were made to the expansion system. Nozzles designed based on both conical and streamline traced flowfields; are discussed. Results from 3-D CFD calculations on an optimized geometry are also presented. A series of cold-flow experiments are proposed to validate the CFD analysis and quantify performance of the flowpath expansions surface design. A discussion is provided of the research hardware designs and experimental test plans.
    Keywords: Spacecraft Propulsion and Power
    Type: 26th JANNAF Airbreathing Propulsion Subcommittee Meeting; Volume 1; 271-279; CPIA-Publ-713-Vol-1
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  • 10
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    Thin-Film Solar Array Earth Orbit Mission Applicability Assessment (2002)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: This is a preliminary assessment of the applicability and spacecraft-level impact of using very lightweight thin-film solar arrays with relatively large deployed areas for representative Earth orbiting missions. The most and least attractive features of thin-film solar arrays are briefly discussed. A simple calculation is then presented illustrating that from a solar array alone mass perspective, larger arrays with less efficient but lighter thin-film solar cells can weigh less than smaller arrays with more efficient but heavier crystalline cells. However, a proper spacecraft-level systems assessment must take into account the additional mass associated with solar array deployed area: the propellant needed to desaturate the momentum accumulated from area-related disturbance torques and to perform aerodynamic drag makeup reboost. The results for such an assessment are presented for a representative low Earth orbit (LEO) mission, as a function of altitude and mission life, and a geostationary Earth orbit (GEO) mission. Discussion of the results includes a list of specific mission types most likely to benefit from using thin-film arrays. NASA Glenn's low-temperature approach to depositing thin-film cells on lightweight, flexible plastic substrates is also briefly discussed to provide a perspective on one approach to achieving this enabling technology. The paper concludes with a list of issues to be addressed prior to use of thin-film solar arrays in space and the observation that with their unique characteristics, very lightweight arrays using efficient, thin-film cells on flexible substrates may become the best array option for a subset of Earth orbiting missions.
    Keywords: Spacecraft Propulsion and Power
    Type: 17th Space Photovoltaic Research and Technology Conference; 74-83; NASA/CP-2002-211831
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  • 11
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    34% Efficient InGaP/GaAs/GaSb Cell-Interconnected-Circuits for Line-Focus Concentrator Arrays (2002)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: While monolithic multi-junction cells are preferred for flat plate arrays, mechanically stacked multi-junction cells are superior for solar concentrator applications. Reasons for this are that the mechanical stacked configuration with high efficiency Gallium Antimonide cells allows utilization of a much wider range of the solar energy spectrum, and the ability to use voltage matched interconnects results in full use of low bandgap cell currents. Herein, data are presented for simple two terminal voltage-matched circuits using InGaP/GaAs/GaSb stacked cells showing 34% average circuit efficiency for a lot of 12 circuits given prismatic covers. These circuits have been designed to fit into the ultralight Stretched Lens Array being developed by NASA. With these new cell-interconnected-circuits, we project that the power density at GEO operating temperature can be increased from 296 W/m2 to 350 W/m2 while maintaining the specific power at 190 W/kg at the full wing level.
    Keywords: Spacecraft Propulsion and Power
    Type: 17th Space Photovoltaic Research and Technology Conference; 24-31; NASA/CP-2002-211831
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  • 12
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    GTX Reference Vehicle Structural Verification Methods and Weight Summary (2002)
    In:  CASI
    Details
    Publication Date: 2013-08-24
    Description: The design of a single-stage-to-orbit air breathing propulsion system requires the simultaneous development of a reference launch vehicle in order to achieve the optimal mission performance. Accordingly, for the GTX study a 300-lb payload reference vehicle was preliminary sized to a gross liftoff weight (GLOW) of 238,000 lb. A finite element model of the integrated vehicle/propulsion system was subjected to the trajectory environment and subsequently optimized for structural efficiency. This study involved the development of aerodynamic loads mapped to finite element models of the integrated system in order to assess vehicle margins of safety. Commercially available analysis codes were used in the process along with some internally developed spread-sheets and FORTRAN codes specific to the GTX geometry for mapping of thermal and pressure loads. A mass fraction of 0.20 for the integrated system dry weight has been the driver for a vehicle design consisting of state-of-the-art composite materials in order to meet the rigid weight requirements. This paper summarizes the methodology used for preliminary analyses and presents the current status of the weight optimization for the structural components of the integrated system.
    Keywords: Spacecraft Propulsion and Power
    Type: 26th JANNAF Airbreathing Propulsion Subcommittee Meeting; Volume 1; 291-303; CPIA-Publ-713-Vol-1
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  • 13
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    Dynamic Characterization of an Inflatable Concentrator for Solar Thermal Propulsion (2002)
    In:  CASI
    Details
    Publication Date: 2013-08-29
    Description: Solar-thermal propulsion is a concept for producing thrust sufficient for orbital transfers and requires innovative, lightweight structures. This note presents a description of an inflatable concentrator that consists of a torus, lens simulator, and three tapered struts. Modal testing was discussed for characterization and verification of the solar concentrator assembly. Finite element shell models of the concentrator were developed using a two-step nonlinear approach, and results were compared to test data. Reasonable model-to-test agreement was achieved for the torus, and results for the concentrator assembly were comparable to the test for several modes.
    Keywords: Spacecraft Propulsion and Power
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  • 14
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    Space and Terrestrial Photovoltaics: Synergy and Diversity (2002)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: A historical view of the research and development in photovoltaics from the perspective of both the terrestrial and the space communities is presented from the early days through the '70s and '80s and the '90s and beyond. The synergy of both communities in the beginning and once again in the present and hopefully future are highlighted, with examples of the important features in each program. The space community which was impressed by the light-weight and reliability of photovoltaics drove much of the early development. Even up to today, nearly every satellites and other scientific space probe that has been launched has included some solar power. However, since the cost of these power systems were only a small fraction of the satellite and launch cost, the use of much of this technology for the terrestrial marketplace was not feasible. It was clear that the focus of the terrestrial community would be best served by reducing costs. This would include addressing a variety of manufacturing issues and raising the rate of production. Success in these programs and a resulting globalization of effort resulted in major strides in the reduction of PV module costs and increased production. Although, the space community derived benefit from some of these advancements, its focus was on pushing the envelope with regard to cell efficiency. The gap between theoretical efficiencies and experimental efficiencies for silicon, gallium arsenide and indium phosphide became almost non-existent. Recent work by both communities have focused on the development thin film cells of amorphous silicon, CuInSe2 and CdTe. These cells hold the promise of lower costs for the terrestrial community as well as possible flexible substrates, better radiation resistance, and higher specific power for the space community. It is predicted that future trends in both communities will be directed toward advances through the application of nanotechnology. A picture is emerging in which the space and terrestrial solar cell communities shall once again share many common goals and, in fact, companies may manufacture both space and terrestrial solar cells in III-V materials and thin film materials. Basic photovoltaics research including these current trends in nanotechnology provides a valuable service for both worlds in that fundamental understanding of cell processes is still vitally important, particularly with new materials or new cell structures. It is entirely possible that one day we might have one solar array design that will meet the criteria for success in both space and on the Earth or perhaps the Moon or Mars.
    Keywords: Spacecraft Propulsion and Power
    Type: 17th Space Photovoltaic Research and Technology Conference; 202-210; NASA/CP-2002-211831
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  • 15
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    Atmospheric Pressure Spray Chemical Vapor Deposited CuInS2 Thin Films for Photovoltaic Applications (2002)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: Solar cells have been prepared using atmospheric pressure spray chemical vapor deposited CuInS2 absorbers. The CuInS2 films were deposited at 390 C using the single source precursor (PPh3)2CuIn(SEt)4 in an argon atmosphere. The absorber ranges in thickness from 0.75 - 1.0 micrometers, and exhibits a crystallographic gradient, with the leading edge having a (220) preferred orientation and the trailing edge having a (112) orientation. Schottky diodes prepared by thermal evaporation of aluminum contacts on to the CuInS2 yielded diodes for films that were annealed at 600 C. Solar cells were prepared using annealed films and had the (top down) composition of Al/ZnO/CdS/CuInS2/Mo/Glass. The Jsc, Voc, FF and (eta) were 6.46 mA per square centimeter, 307 mV, 24% and 0.35%, respectively for the best small area cells under simulated AM0 illumination.
    Keywords: Spacecraft Propulsion and Power
    Type: 17th Space Photovoltaic Research and Technology Conference; 84-90; NASA/CP-2002-211831
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  • 16
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    Boeing's High Voltage Solar Tile Test Results (2002)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: Real concerns of spacecraft charging and experience with solar array augmented electrostatic discharge arcs on spacecraft have minimized the use of high voltages on large solar arrays despite numerous vehicle system mass and efficiency advantages. Boeing's solar tile (patent pending) allows high voltage to be generated at the array without the mass and efficiency losses of electronic conversion. Direct drive electric propulsion and higher power payloads (lower spacecraft weight) will benefit from this design. As future power demand grows, spacecraft designers must use higher voltage to minimize transmission loss and power cable mass for very large area arrays. This paper will describe the design and discuss the successful test of Boeing's 500-Volt Solar Tile in NASA Glenn's Tenney chamber in the Space Plasma Interaction Facility. The work was sponsored by NASA's Space Solar Power Exploratory Research and Technology (SERT) Program and will result in updated high voltage solar array design guidelines being published.
    Keywords: Spacecraft Propulsion and Power
    Type: 17th Space Photovoltaic Research and Technology Conference; 49-54; NASA/CP-2002-211831
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  • 17
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    The Stretched Lens Array (SLA): An Ultra-Light Photovoltaic Concentrator (2002)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: A high-performance, ultralight, photovoltaic concentrator array is being developed for space power. The stretched lens array (SLA) uses stretched-membrane, silicone Fresnel lenses to concentrate sunlight onto triple-junction photovoltaic cells. The cells are mounted to a composite radiator structure. The entire solar array wing, including lenses, photovoltaic cell flex circuits, composite panels, hinges, yoke, wiring harness, and deployment mechanisms, has a mass density of 1.6 kg/sq.m. NASA Glenn has measured 27.4% net SLA panel efficiency, or 375 W/sq.m. power density, at room temperature. At GEO operating cell temperature (80 C), this power density will be 300 W/sq.m., resulting in more than 180 W/kg specific power at the full wing level. SLA is a direct ultralight descendent of the successful SCARLET array on NASA's Deep Space 1 spacecraft. This paper describes the evolution from SCARLET to SLA, summarizes the SLA's key features, and provides performance and mass data for this new concentrator array.
    Keywords: Spacecraft Propulsion and Power
    Type: 17th Space Photovoltaic Research and Technology Conference; 14-23; NASA/CP-2002-211831
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  • 18
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    CIGS2 Thin-Film Solar Cells on Flexible Foils for Space Power (2002)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: CuIn(1-x)Ga(x)S2 (CIGS2) thin-film solar cells are of interest for space power applications because of the near optimum bandgap for AM0 solar radiation in space. CIGS2 thin film solar cells on flexible stainless steel (SS) may be able to increase the specific power by an order of magnitude from the current level of 65 Wkg(sup -1). CIGS solar cells are superior to the conventional silicon and gallium arsenide solar cells in the space radiation environment. This paper presents research efforts for the development of CIGS2 thin-film solar cells on 127 micrometers and 20 micrometers thick, bright-annealed flexible SS foil for space power. A large-area, dual-chamber, inline thin film deposition system has been fabricated. The system is expected to provide thickness uniformity of plus or minus 2% over the central 5" width and plus or minus 3% over the central 6" width. During the next phase, facilities for processing larger cells will be acquired for selenization and sulfurization of metallic precursors and for heterojunction CdS layer deposition both on large area. Small area CIGS2 thin film solar cells are being prepared routinely. Cu-rich Cu-Ga/In layers were sputter-deposited on unheated Mo-coated SS foils from CuGa (22%) and In targets. Well-adherent, large-grain Cu-rich CIGS2 films were obtained by sulfurization in a Ar: H2S 1:0.04 mixture and argon flow rate of 650 sccm, at the maximum temperature of 475 C for 60 minutes with intermediate 30 minutes annealing step at 120 C. Samples were annealed at 500 C for 10 minutes without H2S gas flow. The intermediate 30 minutes annealing step at 120 C was changed to 135 C. p-type CIGS2 thin films were obtained by etching the Cu-rich layer segregated at the surface using dilute KCN solution. Solar cells were completed by deposition of CdS heterojunction partner layer by chemical bath deposition, transparent-conducting ZnO/ZnO: Al window bilayer by RF sputtering, and vacuum deposition of Ni/Al contact fingers through metal mask. PV parameters of a CIGS2 solar cell on 127 micrometers thick SS flexible foil measured under AM 0 conditions at NASA GRC were: V(sub oc) = 802.9 mV, J(sub sc) = 25.07 mA per square centimeters, FF = 60.06%, and efficiency 0 = 8.84%. For this cell, AM 1.5 PV parameters measured at NREL were: V(sub oc) = 788 mV, J(sub sc) = 19.78 mA per square centimeter, FF = 59.44%, efficiency 0 = 9.26%. Quantum efficiency curve showed a sharp QE cutoff equivalent to CIGS2 bandgap of approximately 1.50 eV, fairly close to the optimum value for efficient AM0 PV conversion in the space.
    Keywords: Spacecraft Propulsion and Power
    Type: 17th Space Photovoltaic Research and Technology Conference; 91-100; NASA/CP-2002-211831
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  • 19
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    SRM-Assisted Trajectory for the GTX Reference Vehicle (2002)
    In:  Other Sources
    Details
    Publication Date: 2011-08-23
    Description: A goal of the GTX effort has been to demonstrate the feasibility of a single stage- to- orbit (SSTO) vehicle that delivers a small payload to low earth orbit. The small payload class was chosen in order to minimize the risk and cost of development of this revolutionary system. A preliminary design study by the GTX team has resulted in the current configuration that offers considerable promise for meeting the stated goal. The size and gross lift-off weight resulting from scaling the current design to closure however may be considered impractical for the small payload. In lieu of evolving the project's reference vehicle to a large-payload class, this paper offers the alternative of using solid-rocket motors in order to close the vehicle at a practical scale. This approach offers a near-term, quasi-reusable system that easily evolves to reusable SSTO following subsequent development and optimization. This paper presents an overview of the impact of the addition of SRM's to the GTX reference vehicle's performance and trajectory. The overall methods of vehicle modeling and trajectory optimization will also be presented. A key element in the trajectory optimization is the use of the program OTIS 3.10 that provides rapid convergence and a great deal of flexibility to the user. This paper will also present the methods used to implement GTX requirements into OTIS modeling.
    Keywords: Spacecraft Propulsion and Power
    Type: 26th JANNAF Airbreathing Propulsion Subcommittee Meeting; Volume 1; 305-320; CPIA-Publ-713-Vol-1
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  • 20
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    JPL micro-thrust propulsion activities (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Description: Formation flying and microspacecraft constellation missions pose new propulsion requirements. Formationflying spacecraft, due to the tight positioning and pointing control requirements, may need thrust control within 1- 20 uN to an accuracy of 0.1 uN for LISA and ST-7, for example. Future missions may have extended thrust ranges into the sub - mN range. However, all do require high specific impulses (〉500 sec) due to long required thruster firings.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA Nanotech 2002; Houston, TX; United States
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  • 21
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    Piezoelectrically actuated microvalves for micropropulsion applications (2002)
    In:  Other Sources
    Details
    Publication Date: 2018-06-08
    Description: This paper presents a piezoelectric microvalve technology with a high pressure handling capability for micropropulsion applications.
    Keywords: Spacecraft Propulsion and Power
    Type: ASME International Mechanical Congree and Exposition; New Orleans, LA; United States
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    A piezoelectric microvalve for micropropulsion (2002)
    In:  Other Sources
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    Publication Date: 2018-06-11
    Description: This paper presents a piezoelectric microvalve technology with a high pressure handling capability for micropropulsion applications.
    Keywords: Spacecraft Propulsion and Power
    Type: Nanotech 2002 At the Edge of Revolution; Houston, TX; United States
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    Stirling Research Laboratory Providing Independent Performance Verification of Convertors for a Stirling Radioisotope Generator (2002)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: The Department of Energy (DOE), Germantown, Maryland, Stirling Technology Company (STC), Kennewick, Washington, and NASA Glenn Research Center are developing a free-piston Stirling convertor for a high-efficiency Stirling Radioisotope Generator for NASA Space Science missions. This generator is being developed for multimission use, including providing electric power for unmanned Mars rovers and for deep space missions. STC is developing the 55-W Technology Demonstration Convertor (TDC) under contract to DOE. Glenn is conducting an in-house technology project to assist in developing the convertor for readiness for space qualification and mission implementation. As part of this effort, a Stirling Research Laboratory was established to test the TDC's and related technologies. A key task is providing an independent verification and validation of the TDC performance. Four TDC's are now being tested at Glenn. Acceptance testing has been completed for all convertors, and in general, performance agreed well with that achieved by STC prior to the delivery of the convertors. Performance mapping has also been completed on two of the convertors over a range of hot-end temperatures (450 to 650 C), cold-end temperatures (80 to 120 C), and piston amplitudes (5.2 to 6.2 mm). These test data are available online at http://www.grc.nasa.gov/WWW/tmsb/. The TDC's can be tested in either a horizontal orientation with dual-opposed convertors or in a vertical orientation with a single convertor. Synchronized dual-opposed pairs are used for dynamically balanced operation that results in very low levels of vibration. The Stirling Research Laboratory also supports launch environment testing of the TDC's in Glenn's Structural Dynamics Laboratory and electromagnetic interference and electromagnetic compatibility characterization and reduction efforts. In addition, the TDC's will be used for long-term endurance testing, and preparations are underway for unattended operation.
    Keywords: Spacecraft Propulsion and Power
    Type: Research and Technology 2001; NASA/Tm-2002-211333
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    Solar Electric Propulsion Technologies Being Designed for Orbit Transfer Vehicle Applications (2002)
    In:  CASI
    Details
    Publication Date: 2018-06-02
    Description: There is increasing interest in employing Solar Electric Propulsion (SEP) for new missions requiring transfer from low Earth orbit to the Earth-Moon Lagrange point, L1. Mission architecture plans place the Gateway Habitat at L1 in the 2011 to 2016 timeframe. The Gateway Habitat is envisioned to be used for Lunar exploration, space telescopes, and planetary mission staging. In these scenarios, an SEP stage, or "tug," is used to transport payloads to L1--such as the habitat module, lunar excursion and return vehicles, and chemical propellant for return crew trips. SEP tugs are attractive because they are able to efficiently transport large (less than 10,000 kg) payloads while minimizing propellant requirements. To meet the needs of these missions, a preliminary conceptual design for a general-purpose SEP tug was developed that incorporates several of the advanced space power and in-space propulsion technologies (such as high-power gridded ion and Hall thrusters, high-performance thin-film photovoltaics, lithium-ion batteries, and advanced high-voltage power processing) being developed at the NASA Glenn Research Center. A spreadsheet-based vehicle system model was developed for component sizing and is currently being used for mission planning. This model incorporates a low-thrust orbit transfer algorithm to make preliminary determinations of transfer times and propellant requirements. Results from this combined tug mass estimation and orbit transfer model will be used in a higher fidelity trajectory model to refine the analysis.
    Keywords: Spacecraft Propulsion and Power
    Type: Research and Technology 2001; NASA/TM-2002-211333
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    Refractive Secondary Solar Concentrator Demonstrated High-Temperature Operation (2002)
    In:  CASI
    Details
    Publication Date: 2018-06-02
    Description: Space applications that utilize solar thermal energy--such as electric power conversion systems, thermal propulsion systems, and furnaces--require highly efficient solar concentration systems. The NASA Glenn Research Center is developing the refractive secondary concentrator, which uses refraction and total internal reflection to efficiently concentrate and direct solar energy. When used in combination with advanced lightweight primary concentrators, such as inflatable thin films, the refractive secondary concentrator enables very high system concentration ratios and very high temperatures. Last year, Glenn successfully demonstrated a secondary concentrator throughput efficiency of 87 percent, with a projected efficiency of 93 percent using an antireflective coating. Building on this achievement, Glenn recently successfully demonstrated high-temperature operation of the secondary concentrator when it was used to heat a rhenium receiver to 2330 F. The high-temperature demonstration of the concentrator was conducted in Glenn's 68-ft long Tank 6 thermal vacuum facility equipped with a solar simulator. The facility has a rigid panel primary concentrator that was used to concentrate the light from the solar simulator onto the refractive secondary concentrator. NASA Marshall Space Flight Center provided a rhenium cavity, part of a solar thermal propulsion engine, to serve as the high-temperature receiver. The prototype refractive secondary concentrator, measuring 3.5 in. in diameter and 11.2 in. long, is made of single-crystal sapphire. A water-cooled splash shield absorbs spillage light outside of the 3.5-in. concentrator aperture. Multilayer foil insulation composed of tungsten, molybdenum, and niobium is used to minimize heat loss from the hightemperature receiver. A liquid-cooled canister calorimeter is used to measure the heat loss through the multilayer foil insulation.
    Keywords: Spacecraft Propulsion and Power
    Type: Research and Technology 2001; NASA/TM-2002-211333
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    A piezoelectric microvalve for micropropulsions (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: Nanotech 2002: At the Edge of Revolution; Houston, TX; United States
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  • 27
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    Plasma generation near an ion engine discharge chamber hollow cathode (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Huntsville, AL; United States
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  • 28
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    Numerical simulations of ion thruster accelerator grid erosion (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Description: The highly successful demonstration of ion propulsion on Deep Space 1 has stimulated the study of more demanding applications of ion propulsion. These future applications require ion thrusters capable of providing significantly greater specific impulses and total impulses than the current state-of-the-art Higher specific impulses aggravate the known wear out mechanisms of the ion accelerator system.
    Keywords: Spacecraft Propulsion and Power
    Type: Joint Propulsion Conference; Indianapolis, IN; United States
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  • 29
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    Minimum impulse thruster valve design and development (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA JPC and E 2002; Indianapolis, IN; United States
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    Initial Cassini propulsion system in-flight characterization (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: Joint Propulsion Conference and Exhibit; Indianapolis, IN; United States
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    Electric propulsion at the Jet Propulsion Laboratory (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Huntsville, AL; United States
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  • 32
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    Performance of gas fed pulsed plasma thrusters using water vapor propellant (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Description: The performance of two gas-fed pulsed plasma thrusters (GFPPTs) using both argon and water vapor for propellant has been measured using the JPL microthrust stand.
    Keywords: Spacecraft Propulsion and Power
    Type: 38th AIAA Joint Propulsion Conference; Indianapolis, IN; United States
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    Experimental and computational investigation of the performance of a micro-ion thruster (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Description: A micro-ion thruster assembly with a characteristic diameter of 3-cm has been developed at JPL for testing and optimization of various system parameters.
    Keywords: Spacecraft Propulsion and Power
    Type: Joint Propulsion Conference; Indianapolis, IN; United States
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    Performance characteristics of the deep space 1 flight spare ion thruster long duration test, the first 21,300 hours of operation (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Description: A long duration test of the DSl flight spare ion thruster (FT2) is presently being conducted at the Jet Propulsion Laboratory. To, date the thruster has accumulated over 23,500 hours of operation, and 190 kg of Xenon propellant, over 230% of the initial design life. The primary objectives of the test include the processing of 200 kg of Xenon propellant, the identification of unknown failure modes, the characterization and drivers of these failure modes, and to measure performance degradation as the thruster wears. The test is fitted with an extensive array of diagnostics to measure engine wear and performance degradation. To date the most notable erosion processes include severe discharge cathode keeper erosion, accelerator grid erosion, reduction in electrical isolation of the neutralizer assembly, and deposit formation within the neutralizer orifice, reducing margin from plume mode. Over the past 23,500 hours of operation, performance degradation has been minimal, and it is anticipated that the above erosion processes will not preclude the thruster from processing over 200 kg of Xenon.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA/ASME/SAE/ASEE 38th Joint Propulsion Conference and Exhibit; Indianapolis, IN; United States
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    High efficiency, high power ion thrusters for operation at specific impulses greater than 7000 s (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: Advanced Space Propulsion Workshop; Pasadena, CA; United States
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    Candidate thruster technologies for NEP (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: Advanced Space Propulsion Workshop; Pasadena, CA; United States
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    Long life cathodes for high power NEP missions (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: Advanced Space Propulsion Workshop; Pasadena, CA; United States
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    NEP mission performance technology drivers (2002)
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: Advanced Space Propulsion Workshop; Pasadena, CA; United States
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  • 39
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    NEP mission performance technology drivers (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: Advanced Space Propulsion Workshop; Pasadena, CA; United States
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    Long life hollow cathodes for high power NEP missions (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: Advanced Space Propulsion Workshop; Pasadena, CA; United States
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  • 41
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    Laser ablation microthruster technology (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Description: A review of Laser Albation Microthruster (LAmuT) technology along with recent experimental results are presented.
    Keywords: Spacecraft Propulsion and Power
    Type: 33rd AIAA Plasmadynamics and Lasers Conference; Maui, HI; United States
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  • 42
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    Advanced electric propulsion for outer planet exploration (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: 6th International Symposium on Space Propulsion of the 21st Century; Paris; France
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  • 43
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    Ion thrusters and LFAs for outer planet exploration (2002)
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    Publication Date: 2018-06-08
    Description: Given the recent advancements in power generation, waste heat rejection systems and electric propulsion, a reassessment of the benefits of Nuclear Electric Propulsion (NEP) is provided.
    Keywords: Spacecraft Propulsion and Power
    Type: AAAF 6th International Symposium; Versailles; France
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  • 44
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    Initial Cassini propulsion system in-flight characterization (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Description: The Cassini Propulsion Module Subsystem has performed excellently throughout the first four years of mission operations. The PMS is the most complex interplanetary propulsion subsystem ever flown, with separate monopropellant and bipropellant propulsion modules, each replete with many redundant components.
    Keywords: Spacecraft Propulsion and Power
    Type: 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibition; Indianapolis, IN; United States
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    Present status and future directions in micropropulsion (2002)
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: International Workshop on Rocket Propulsion: Present and Future; Pozzuoli; Italy
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  • 46
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    Field emission array cathode material selection for compatibility with electric propulsion applications (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Description: In the investigation discussed in this paper, the response of FEA cathodes and FEA material candidates to simulated electric propulsion system environments are presented.
    Keywords: Spacecraft Propulsion and Power
    Type: 201st Meeting Electrochemical Society; Philadelphia, PA; United States
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    Flywheel Energy Storage System Designed for the International Space Station (2002)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: Following successful operation of a developmental flywheel energy storage system in fiscal year 2000, researchers at the NASA Glenn Research Center began developing a flight design of a flywheel system for the International Space Station (ISS). In such an application, a two-flywheel system can replace one of the nickel-hydrogen battery strings in the ISS power system. The development unit, sized at approximately one-eighth the size needed for ISS was run at 60,000 rpm. The design point for the flight unit is a larger composite flywheel, approximately 17 in. long and 13 in. in diameter, running at 53,000 rpm when fully charged. A single flywheel system stores 2.8 kW-hr of useable energy, enough to light a 100-W light bulb for over 24 hr. When housed in an ISS orbital replacement unit, the flywheel would provide energy storage with approximately 3 times the service life of the nickel-hydrogen battery currently in use.
    Keywords: Spacecraft Propulsion and Power
    Type: Research and Technology 2001; NASA/TM-2002-211333
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    Magnetohydrodynamic MACH Code Used to Simulate Magnetoplasmadynamic Thrusters (2002)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: The On-Board Propulsion program at the NASA Glenn Research Center is utilizing a state of-the-art numerical simulation to model the performance of high-power electromagnetic plasma thrusters. Such thrusters are envisioned for use in lunar and Mars cargo transport, piloted interplanetary expeditions, and deep-space robotic exploration of the solar system. The experimental portion of this program is described in reference 1. This article describes the numerical modeling program used to guide the experimental research. The synergistic use of numerical simulations and experimental research has spurred the rapid advancement of high-power thruster technologies for a variety of bold new NASA missions. From its inception as a U.S. Department of Defense code in the mid-1980's, the Multiblock Arbitrary Coordinate Hydromagnetic (MACH) simulation tool has been used by the plasma physics community to model a diverse range of plasma problems--including plasma opening switches, inertial confinement fusion concepts, compact toroid formation and acceleration, z-pinch implosion physics, laser-target interactions, and a variety of plasma thrusters. The MACH2 code used at Glenn is a time-dependent, two-dimensional, axisymmetric, multimaterial code with a multiblock structure. MACH3, a more recent three-dimensional version of the code, is currently undergoing beta tests. The MACH computational mesh moves in an arbitrary Lagrangian-Eulerian (ALE) fashion that allows the simulation of diffusive-dominated and dispersive-dominated problems, and the mesh can be refined via a variety of adaptive schemes to capture regions of varying characteristic scale. The mass continuity and momentum equations model a compressible viscous fluid, and three energy equations are used to simulate nonthermal equilibrium between electrons, ions, and the radiation field. Magnetic fields are modeled by an induction equation that includes resistive diffusion, the Hall effect, and a thermal source for magnetic fields. Various models of plasma resistivity are included, along with ablation models and multiport circuit solvers. The set of equations is closed using either an ideal gas or real equation of state.
    Keywords: Spacecraft Propulsion and Power
    Type: Research and Technology 2001; NASA/TM-2002-211333
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    Active Control of Combustor Instability Shown to Help Lower Emissions (2002)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: In a quest to reduce the environmental impact of aerospace propulsion systems, extensive research is being done in the development of lean-burning (low fuel-to-air ratio) combustors that can reduce emissions throughout the mission cycle. However, these lean-burning combustors have an increased susceptibility to thermoacoustic instabilities, or high-pressure oscillations much like sound waves, that can cause severe high-frequency vibrations in the combustor. These pressure waves can fatigue the combustor components and even the downstream turbine blades. This can significantly decrease the safe operating life of the combustor and turbine. Thus, suppression of the thermoacoustic combustor instabilities is an enabling technology for lean, low-emissions combustors. Under the Aerospace Propulsion and Power Base Research and Technology Program, the NASA Glenn Research Center, in partnership with Pratt & Whitney and United Technologies Research Center, is developing technologies for the active control of combustion instabilities. With active combustion control, the fuel is pulsed to put pressure oscillations into the system. This cancels out the pressure oscillations being produced by the instabilities. Thus, the engine can have lower pollutant emissions and long life.The use of active combustion instability control to reduce thermo-acoustic-driven combustor pressure oscillations was demonstrated on a single-nozzle combustor rig at United Technologies. This rig has many of the complexities of a real engine combustor (i.e., an actual fuel nozzle and swirler, dilution cooling, etc.). Control was demonstrated through modeling, developing, and testing a fuel-delivery system able to the 280-Hz instability frequency. The preceding figure shows the capability of this system to provide high-frequency fuel modulations. Because of the high-shear contrarotating airflow in the fuel injector, there was some concern that the fuel pulses would be attenuated to the point where they would not be effective for control. Testing in the combustor rig showed that open-loop pulsing of the fuel was, in fact, able to effectively modulate the combustor pressure. To suppress the combustor pressure oscillations due to thermoacoustic instabilities, it is desirable to time the injection of the fuel so that it interferes with the instability. A closed-loop control scheme was developed that uses combustion pressure feedback and a phase-shifting controller to time the fuel-injection pulses. Some suppression of the pressure oscillations at the 280-Hz instability frequency was demonstrated (see the next figure). However, the overall peak-to- peak pressure oscillations in the combustor were only mildly reduced. Improvements to control hardware and control methods are being continued to gain improved closed-loop reduction of the pressure oscillations.pulse the fuel at
    Keywords: Spacecraft Propulsion and Power
    Type: Research and Technology 2001; NASA/TM-2002-211333
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    Power and propulsion - the keys to space exploration (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Keywords: Spacecraft Propulsion and Power
    Type: Continuing Learning Experience - Frontiers of Space; Fullerton, CA; United States
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    In-FEEP ion beam neutralization with thermionic and field emission cathodes (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Description: Charge neutralization of an In-FEEP thruster was demonstrated with three different electron sources by zeroing the floating potential of the thruster and neutralizer system. The three cathodes used in the investigation include a mixed metal thermionic cathode, a carbon nanotube field emission cathode, and a Spindt-type Mo field emission array cathode.
    Keywords: Spacecraft Propulsion and Power
    Type: International Electric Propulsion Technology; Pasadena, CA; United States
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    The DS1 hyper-extended mission (2002)
    In:  Other Sources
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    Publication Date: 2018-06-08
    Description: The Primary Mission for Deep Space (DS1) was to demonstrate 12 new technologies one of which was the Ion Propulsion System (IPS). After successfully completing its primary mission, DS1 was given a new mission. The objective of this Extended Mission was to fly by the Comet Borrelly. After the successful Borrelly encounter, the ion thruster on DS1 had to be operated for more than 14,000 hours in space. This provided a unique opportunity to investigate the condition of the thruster after long-term operation in space and determine, to the extent possible, if the thruster wear in space is consistent with that observed in long-duration ground tests.
    Keywords: Spacecraft Propulsion and Power
    Type: Joint Propulsion Conference; Indianapolis, IN; United States
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    Space Solar Power Concepts: Demonstrations to Pilot Plants (2002)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: The availability of abundant, affordable power where needed is a key to the future exploration and development of space as well as future sources of clean terrestrial power. One innovative approach to providing such power is the use of wireless power transmission (WPT). There are at least two possible WPT methods that appear feasible; microwave and laser. Microwave concepts have been generated, analyzed and demonstrated. Technologies required to provide an end-to-end system have been identified and roadmaps generated to guide technology development requirements. Recently, laser W T approaches have gained an increased interest. These approaches appear to be very promising and will possibly solve some of the major challenges that exist with the microwave option. Therefore, emphasis is currently being placed on the laser WPT activity. This paper will discuss the technology requirements, technology roadmaps and technology flight experiments demonstrations required to lead toward a pilot plant demonstration. Concepts will be discussed along with the modeling techniques that are used in developing them. Feasibility will be addressed along with the technology needs, issues and capabilities for particular concepts. Flight experiments and demonstrations will be identified that will pave the road from demonstrations to pilot plants and beyond.
    Keywords: Spacecraft Propulsion and Power
    Type: 53rd IAF; Oct 10, 2002 - Oct 19, 2002; Houston, TX; United States
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    The FRC Acceleration Space Thruster (FAST) Experiment (2002)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: The objective of the FRC (Field Reversed Configuration) Acceleration Space Thruster (FAST) Experiment is to investigate the use of a repetitive FRC source as a thruster, specifically for an NEP (nuclear electric propulsion) system. The Field Reversed Configuration is a plasmoid with a closed poloidal field line structure, and has been extensively studied as a fusion reactor core. An FRC thruster works by repetitively producing FRCs and accelerating them to high velocity. An FRC thruster should be capable of I(sub sp)'s in the range of 5,000 - 25,000 seconds and efficiencies in the range of 60 - 80 %. In addition, they can have thrust densities as high as 10(exp 6) N/m2, and as they are inductively formed, they do not suffer from electrode erosion. The jet-power should be scalable from the low to the high power regime. The FAST experiment consists of a theta-pinch formation chamber, followed by an acceleration stage. Initially, we will produce and accelerate single FRCs. The initial focus of the experiment will be on the ionization, formation and acceleration of a single plasmoid, so as to determine the likely efficiency and I(sub sp). Subsequently, we will modify the device for repetitive burst-mode operation (5-10 shots). A variety of diagnostics are or will be available for this work, including a HeNe interferometer, high-speed cameras, and a Thomson-scattering system. The status of the experiment will be described.
    Keywords: Spacecraft Propulsion and Power
    Type: Advanced Space Propulsion Workshop; Jun 04, 2002 - Jun 06, 2002; Pasadena, CA; United States
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    Fundamentals of Plasma Sails Propulsion Concept (2002)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: The Mini-Magnetospheric Plasma Propulsion (M2P2), originally proposed by Winglee et al. [2000], is based on the two-fluid plasma model and requires a 15-km frontal standoff distance (or 20-km cross-sectional diameter) in order for the magnetic bubble to absorb sufficient momentum from the SW to accelerate a spacecraft to the unprecedented speeds of 50-80 km/s after an acceleration period of about three months. Winglee et al. [2000] derived the above size requirement based on an extrapolation of their simulated results in which a system much smaller than a M2P2 was used (p. 21,074 of their study). We submit, however, that a fluid model has no validity for such a small scale size-even in the region near the plasma source! It is assumed in the MHD fluid model, normally applied to the magnetosphere, that the characteristic scale-size is much greater than the Larmor radius and ion skin depth of the SW. In the case of the M2P2, however, the size of the magnetic bubble is actually less than or, at best, comparable to, the scale of these characteristic parameters and, therefore, a kinetic approach, which addresses the smallscale physical mechanisms involved, must be used. A fully three-dimensional version of the hybrid code is used in our M2P2 (Plasma Sails) studies was originally developed by Delamere et al. [1999]. The M2P2 plasma sail is an excellent application for this hybrid code. The primary advantage of this code is the seamless interface between fluid and kinetic descriptions of the ion populations. A kinetic description is not necessary for the dense inner regions of the magnetic bubble and tremendous computational savings can be realized by treating this dense, magnetized ion population with the fluid description. It is essential, however, that the outer bubble regions be treated kinetically as well as the SW protons. Comparison of full size M2P2 simulation based on 3D MHD and kinetic models show that kinetic treatment introduces much more asymmetry to the considering problem and the possibility of kinetic instabilities development.
    Keywords: Spacecraft Propulsion and Power
    Type: 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit; Jul 20, 2003 - Jul 23, 2003; Huntsville, AL; United States
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    A Comparison of Trajectory Optimization Methods for the Impulsive Minimum Fuel Rendezvous Problem (2002)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: In this paper we present a comparison of optimization approaches to the minimum fuel rendezvous problem. Both indirect and direct methods are compared for a variety of test cases. The indirect approach is based on primer vector theory. The direct approaches are implemented numerically and include Sequential Quadratic Programming (SQP), Quasi-Newton, Simplex, Genetic Algorithms, and Simulated Annealing. Each method is applied to a variety of test cases including, circular to circular coplanar orbits, LEO to GEO, and orbit phasing in highly elliptic orbits. We also compare different constrained optimization routines on complex orbit rendezvous problems with complicated, highly nonlinear constraints.
    Keywords: Spacecraft Propulsion and Power
    Type: 26th Annual Guidance and Control Conference; Feb 01, 2003 - Feb 28, 2003; Breckenridge, CO; United States
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    LWS/SET Technology Experiment Carrier (2002)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: This paper examines the approach taken to building a low-cost, modular spacecraft bus that can be used to support a variety of technology experiments in different space environments. It describes the techniques used and design drivers considered to ensure experiment independence from as yet selected host spacecraft. It describes the technology experiment carriers that will support NASA's Living With a Star Space Environment Testbed space missions. NASA has initiated the Living With a Star (LWS) Program to develop a better scientific understanding to address the aspects of the connected Sun-Earth system that affect life and society. A principal goal of the program is to bridge the gap between science, engineering, and user application communities. The Space Environment Testbed (SET) Project is one element of LWS. The Project will enable future science, operational, and commercial objectives in space and atmospheric environments by improving engineering approaches to the accommodation and/or mitigation of the effects of solar variability on technological systems. The SET Project is highly budget constrained and must seek to take advantage of as yet undetermined partnering opportunities for access to space. SET will conduct technology validation experiments hosted on available flight opportunities. The SET Testbeds will be developed in a manner that minimizes the requirements for accommodation, and will be flown as flight opportunities become available. To access the widest range of flight opportunities, two key development requirements are to maintain flexibility with respect to accommodation constraints and to have the capability to respond quickly to flight opportunities. Experiments, already developed to the technology readiness level of needing flight validation in the variable Sun-Earth environment, will be selected on the basis of the need for the subject technology, readiness for flight, need for flight resources and particular orbit. Experiments will be accumulated by the Project and manifested for specific flight opportunities as they become available. The SET Carrier is designed to present a standard set of interfaces to SET technology experiments and to be modular and flexible enough to interface to a variety of possible host spacecraft. The Carrier will have core components and mission unique components. Once the core carrier elements have been developed, only the mission unique components need to be defined and developed for any particular mission. This approach will minimize the mission specific cost and development schedule for a given flight opportunity. The standard set of interfaces provided by SET to experiments allows them to be developed independent of the particulars of a host spacecraft. The Carrier will provide the power, communication, and the necessary monitoring features to operate experiments. The Carrier will also provide all of the mechanical assemblies and harnesses required to adapt experiments to a particular host. Experiments may be hosted locally with the Carrier or remotely on the host spacecraft. The Carrier design will allow a single Carrier to support a variable number of experiments and will include features that support the ability to incrementally add experiments without disturbing the core architecture.
    Keywords: Spacecraft Propulsion and Power
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    System Description of the Gasdynamic Mirror Propulsion System (2002)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: Nuclear fusion appears to be a most promising concept for producing extremely high specific impulse rocket engines. One particular fusion concept which seems to be very well suited for fusion propulsion applications is the gasdynamic mirror (GDM). An experimental GDM device has been constructed at the NASA Marshall Space Flight Center to provide an initial assessment of the feasibility of this type of propulsion system. A systems shakedown of the device is currently underway with initial experiments slated to occur in early 2002. The device has been constructed so as to allow a considerable degree of flexibility in its configuration thus permitting the experiment to easily grow over time without necessitating a great deal of additional fabrication. This flexibility is due in large part to the modular nature of the machine wherein additional modules may be added as needed to meet varying experimental objectives. Figure 1 shows the current configuration of the Gasdynamic Mirror Experiment, and Table 1 describes the main features of the device.
    Keywords: Spacecraft Propulsion and Power
    Type: Track-53392 , American Nuclear Society Annual Meeting; Hollywood, FL; United States
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    Trajectories for High Specific Impulse High Specific Power Deep Space Exploration (2002)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: Flight times and deliverable masses for electric and fusion propulsion systems are difficult to approximate. Numerical integration is required for these continuous thrust systems. Many scientists are not equipped with the tools and expertise to conduct interplanetary and interstellar trajectory analysis for their concepts. Several charts plotting the results of well-known trajectory simulation codes were developed and are contained in this paper. These charts illustrate the dependence of time of flight and payload ratio on jet power, initial mass, specific impulse and specific power. These charts are intended to be a tool by which people in the propulsion community can explore the possibilities of their propulsion system concepts. Trajectories were simulated using the tools VARITOP and IPOST. VARITOP is a well known trajectory optimization code that involves numerical integration based on calculus of variations. IPOST has several methods of trajectory simulation; the one used in this paper is Cowell's method for full integration of the equations of motion. An analytical method derived in the companion paper was also evaluated. The accuracy of this method is discussed in the paper.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA Paper 2002-4233 , 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Jul 07, 2002 - Jul 10, 2002; Indianapolis, IN; United States
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    Conceptual Design of an MTF Space Propulsion System (2002)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: Most fusion propulsion concepts that have been investigated in the past employ some form of inertial or magnetic confinement separately, and are encumbered by the need for advanced drivers (e.g. laser) or steady-state magnetic confinement systems (e.g. superconductors) that have historically resulted in large, massive spacecraft designs. Here we present a comparatively new approach, Magnetized Target Fusion (MTF), which offers a nearer-term avenue for realizing the tremendous performance benefits of fusion propulsion. MTF attempts to combine the favorable attributes of both inertially and magnetically confined fusion to achieve both efficient and low-cost compressional plasma heating and energy confinement. The key advantage of MTF is its less demanding requirements for driver energy and power processing. Additional features include: 1) very low system masses and volumes, 2) relatively low waste heat, 3) substantial utilization of energy from product neutrons, 4) efficient, low peak-power drivers based on existing pulsed power technology, 5) very high Isp , specific power and thrust, and 6) relatively affordable R&D pathways. MTF overcomes many of the problems associated with traditional fusion techniques, thus making it particularly attractive for space applications. Isp greater than 50,000 seconds and specific powers greater than 20 kilowatts/kilogram appear feasible using relatively near-term pulse power and plasma gun technology.
    Keywords: Spacecraft Propulsion and Power
    Type: DOE Innovative Confinement Concepts; Jan 22, 2002 - Jan 24, 2002; College Park, MD; United States
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    Plasma Liner Research for MTF at NASA Marshall Space Flight Center (2002)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: The current research effort at NASA Marshall Space Flight Center (MSFC) in MTF is directed towards exploring the critical physics issues of potential embodiments of MTF for propulsion, especially standoff drivers involving plasma liners for MTF. There are several possible approaches for forming plasma liners. One approach consists of using a spherical array of plasma jets to form a spherical plasma shell imploding towards the center of a magnetized plasma, a compact toroid. Current experimental plan and status to explore the physics of forming a 2-D plasma liner (shell) by merging plasma jets are described. A first-generation coaxial plasma guns (Mark-1) to launch the required plasma jets have been built and tested. Plasma jets have been launched reproducibly with a low jitter, and velocities in excess of 50 km/s for the leading edge of the plasma jet. Some further refinements are being explored for the plasma gun, Successful completion of these single-gun tests will be followed by an experimental exploration of the problems of launching a multiple number of these jets simultaneously to form a cylindrical plasma liner.
    Keywords: Spacecraft Propulsion and Power
    Type: DOE Innovative Confinement Concepts; Jan 22, 2002 - Jan 24, 2002; College Park, MD; United States
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    Robust Low Cost Liquid Rocket Combustion Chamber by Advanced Vacuum Plasma Process (2002)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: Next-generation, regeneratively cooled rocket engines require materials that can meet high temperatures while resisting the corrosive oxidation-reduction reaction of combustion known as blanching, the main cause of engine failure. A project was initiated at NASA-Marshal Space Flight Center (MSFC) to combine three existing technologies to build and demonstrate an advanced liquid rocket engine combustion chamber that would provide a 100 mission life. Technology developed in microgravity research to build cartridges for space furnaces was utilized to vacuum plasma spray (VPS) a functional gradient coating on the hot wall of the combustion liner as one continuous operation, eliminating any bondline between the coating and the liner. The coating was NiCrAlY, developed previously as durable protective coatings on space shuttle high pressure fuel turbopump (HPFTP) turbine blades. A thermal model showed that 0.03 in. NiCrAlY applied to the hot wall of the combustion liner would reduce the hot wall temperature 200 F, a 20% reduction, for longer life. Cu-8Cr-4Nb alloy, which was developed by NASA-Glenn Research Center (GRC), and which possesses excellent high temperature strength, creep resistance, and low cycle fatigue behavior combined with exceptional thermal stability, was utilized as the liner material in place of NARloy-Z. The Cu-8Cr-4Nb material exhibits better mechanical properties at 650 C (1200 F) than NARloy-Z does at 538 C (1000 F). VPS formed Cu-8Cr-4Nb combustion chamber liners with a protective NiCrAlY functional gradient coating have been hot fire tested, successfully demonstrating a durable coating for the first time. Hot fire tests along with tensile and low cycle fatigue properties of the VPS formed combustion chamber liners and witness panel specimens are discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: 39th Space Congress; Apr 29, 2002 - May 02, 2002; Cocoa Beach, FL; United States
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    Advancing Sensor Technology for Aerospace Propulsion (2002)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: NASA's Stennis Space Center (SSC) and Glenn Research Center (GRC) participate in the development of technologies for propulsion testing and propulsion applications in air and space transportation. Future transportation systems and the test facilities needed to develop and sustain them are becoming increasingly complex. Sensor technology is a fundamental pillar that makes possible development of complex systems that must operate in automatic mode (closed loop systems), or even in assisted-autonomous mode (highly self-sufficient systems such as planetary exploration spacecraft). Hence, a great deal of effort is dedicated to develop new sensors and related technologies to be used in research facilities, test facilities, and in vehicles and equipment. This paper describes sensor technologies being developed and in use at SSC and GRC, including new technologies in integrated health management involving sensors, components, processes, and vehicles.
    Keywords: Spacecraft Propulsion and Power
    Type: SE-2002-05-00044-SSC , International Mechanical Engineering Congress and Exposition; Nov 17, 2002 - Nov 22, 2002; New Orleans, LA; United States
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    Evaluation and Characterization Study of Dual Pulse Laser-Induced Spark (DPLIS) For Rocket Engine Ignition System Application (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Space Flight Center (MSFC). Initial hot-fire tests in a small-scale rocket chamber at MSFC have demonstrated the DPLIS concept having two main advantages over existing laser ignition concepts. First, the DPLIS can be potentially optimized its laser pulse format to maximize the initial plasma volume, the plasma lifetime, as well as the flame kernel growth rate. Characterization studies of the laser pulse format are now underway with experiments of igniting gaseous hydrogen/air in a Hencken burner. Once ignition is achieved, the flame is open to the atmosphere. This open environment allows easy access for diagnostics of the ignition phenomenon. The quick turn-around time of conducting experiments on this burner make it more amenable for conducting a large number of experiments for statistical analysis of the sensitivity of the test parameters. The results from these experiments will help optimize the laser format for future testing in an H2/O2 subscale rocket chamber.
    Keywords: Spacecraft Propulsion and Power
    Type: 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Jul 20, 2003 - Jul 23, 2003; Huntsville, AL; United States
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    Quasi 1-D Study of Pulse Detonation Rocket Engine Blowdown Gasdynamics and Performance (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This viewgraph representation provides an overview of research which develops a quasi one dimensional chemistry computational fluid dynamics code to study the effect of nozzle design on the performance of pulse detonation rocket engines (PDREs). Topics considered include: PDREs vs. steady-state rocket engines (SSREs), PDRE cycles, numerical models of idealized PDRE performance, thrust determination of PDRE, specific geometries, and nozzle design and geometry.
    Keywords: Spacecraft Propulsion and Power
    Type: Propulsion Engineering Research Center 14th Annual Symposium on Propulsion; Dec 10, 2002 - Dec 11, 2002; State College, PA; United States
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    Characterization of Space Environmental Effects on Candidate Solar Sail Material (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The National Aeronautics and Space Administration's (NASA) Marshall Space Flight Center (MSFC) is concentrating research into the utilization of photonic materials for spacecraft propulsion. Spacecraft propulsion, using photonic materials, will be achieved using a solar sail. A sail operates on the principle that photons, originating from the sun, impart pressure and provide a source of spacecraft propulsion. The pressure can be increased, by a factor of two if the sun-facing surface is perfectly reflective. Solar sails are generally composed of a highly reflective metallic front layer, a thin polymeric substrate, and occasionally a highly emissive back surface. The Space Environmental Effects Team at MSFC is actively characterizing candidate solar sail materials to evaluate the thermo-optical and mechanical properties after exposure to a simulated Geosynchronous Transfer Orbit (GTO) radiation environment. The technique of radiation dose verses material depth profiling was used to determine the orbital equivalent exposure doses. The solar sail exposure procedures and results of the material characterization will be discussed.
    Keywords: Spacecraft Propulsion and Power
    Type: Propulsion Engineering Research Center 14th Annual Symposium on Propulsion; Dec 10, 2002 - Dec 11, 2002; University Park, PA; United States
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    JANNAF Rocket Nozzle Technology Subcommittee Executive Committee Report (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This viewgraph presentation provides information on the structure and activities of the panels of the Joint Army Navy NASA Air Force (JANNAF) Rocket Nozzle Technology Subcommittee. The panels profiled are the Processing Science and Materials Panel, the Nozzle Design, Test, and Evaluation Panel, the Nozzle Analysis and Modeling Panel, and the Nozzle Control Systems Panel. The presentation also lists meetings, workshops, and publications in which the subcommittee participated during the reporting period.
    Keywords: Spacecraft Propulsion and Power
    Type: 51st JANNAF Propulsion Meeting; Nov 18, 2002 - Nov 21, 2002; Lake Buena Vista, FL; United States
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    Assessment of Various Flow Solvers Used to Predict the Thermal Environment inside Space Shuttle Solid Rocket Motor Joints (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: It is very important to accurately predict the gas pressure, gas and solid temperature, as well as the amount of O-ring erosion inside the space shuttle Reusable Solid Rocket Motor (RSRM) joints in the event of a leak path. The scenarios considered are typically hot combustion gas rapid pressurization events of small volumes through narrow and restricted flow paths. The ideal method for this prediction is a transient three-dimensional computational fluid dynamics (CFD) simulation with a computational domain including both combustion gas and surrounding solid regions. However, this has not yet been demonstrated to be economical for this application due to the enormous amount of CPU time and memory resulting from the relatively long fill time as well as the large pressure and temperature rising rate. Consequently, all CFD applications in RSRM joints so far are steady-state simulations with solid regions being excluded from the computational domain by assuming either a constant wall temperature or no heat transfer between the hot combustion gas and cool solid walls.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA Paper 2002-4300 , 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit; Jul 07, 2002 - Jul 10, 2002; Indianapolis, IN; United States
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    Photovoltaic Power for Future NASA Missions (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Recent advances in crystalline solar cell technology are reviewed. Dual-junction and triple-junction solar cells are presently available from several U. S. vendors. Commercially available triple-junction cells consisting of GaInP, GaAs, and Ge layers can produce up to 27% conversion efficiency in production lots. Technology status and performance figures of merit for currently available photovoltaic arrays are discussed. Three specific NASA mission applications are discussed in detail: Mars surface applications, high temperature solar cell applications, and integrated microelectronic power supplies for nanosatellites.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA Paper 2002-0718 , AIAA 40th Aerospace Sciences Meeting and Exhibit; Jan 14, 2002 - Jan 17, 2002; Reno, NV; United States
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    Extending Ion Engine Technology to NEXT and Beyond (2002)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Extending ion engine technology beyond the current state-of-the art primary interplanetary electric propulsion system, the 2.3-kW NASA Solar Electric Propulsion Technology and Applications Readiness (NSTAR) system, will require thrusters with improved propellant throughput and total impulse capability. Many of the design choices that culminated in the NSTAR thrusters must be revisited, and their application to next generation ion engine technology must be evaluated. The concept of derating, which was successfully employed in NSTAR, has been applied to the 40 cm NASA Evolutionary Xenon Thruster (NEXT) currently under development at NASA Glenn Research Center (GRC). At 5-kW, NEXT operates with the same average beam current density as NSTAR, and at 10-kW, the peak beam current density is only ten percent greater than NSTAR. The result is that similar Ion optics technology is expected to yield comparable lifetime. Thick-accelerator- grid ion optics are also being tested to realize additional lifetime benefits. A 40-A discharge cathode is being developed for NEXT based on scaling the NSTAR design. Nevertheless, the experiences of the NSTAR ground tests and the thruster on the Deep Space One spacecraft indicate that the discharge cathode wear must be studied experimentally and theoretically to ensure that it meets the lifetime requirements. Although NEXT is in its infancy, investigations have already begun to examine possible modifications to engine design for even higher-power and higher-specific impulse engines. Ion optics using alternate materials such as titanium, graphite, or carbon-carbon composite are currently being investigated due to their low sputter yields at high voltage. To avoid the difficulties encountered using electrodes at high-currents, the use of a microwave-based ion thruster is under investigation for potential high-power ion thruster systems requiring long lifetimes. Additionally, alternative propellants are being considered for applications requiring high-specific impulse (〉〉 5000 s) and extremely long-life (〉〉 15,000 hr). Testing requirements make condensable propellants attractive for high-power engines. Although the NSTAR ion engine demonstrated the flight maturity of ion thruster technology, many challenges remain for the development of thrusters with improved propellant throughput and power handling capabilities.
    Keywords: Spacecraft Propulsion and Power
    Type: 29th International Conference on Plasma Science; May 27, 2002 - May 30, 2002; Banff; Canada
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    A Systems Model for Power Technology Assessment (2002)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: A computer model is under continuing development at NASA Glenn Research Center that enables first-order assessments of space power technology. The model, an evolution of NASA Glenn's Array Design Assessment Model (ADAM), is an Excel workbook that consists of numerous spreadsheets containing power technology performance data and sizing algorithms. Underlying the model is a number of databases that contain default values for various power generation, energy storage and power management and distribution component parameters. These databases are actively maintained by a team of systems analysts so that they contain state-of-art data as well as the most recent technology performance projections. Sizing of the power subsystems can be accomplished either by using an assumed mass specific power (W/kg) or energy (Wh/kg) or by a bottoms-up calculation that accounts for individual component performance and masses. The power generation, energy storage and power management and distribution subsystems are sized for given mission requirements for a baseline case and up to three alternatives. This allows four different power systems to be sized and compared using consistent assumptions and sizing algorithms. The component sizing models contained in the workbook are modular so that they can be easily maintained and updated. All significant input values have default values loaded from the databases that can be over-written by the user. The default data and sizing algorithms for each of the power subsystems are described in some detail. The user interface and workbook navigational features are also discussed. Finally, an example study case that illustrates the model's capability is presented.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211728 , E-13482 , NAS 1.15:211728 , IECEC-2002-20038 , 37th Intersociety Energy Conversion Engineering Conference; Jul 28, 2002 - Aug 02, 2002; Washington, DC; United States
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    Battery Reinitialization of the Photovoltaic Module of the International Space Station (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The photovoltaic (PV) module on the International Space Station (ISS) has been operating since November 2000 and supporting electric power demands of the ISS and its crew of three. The PV module contains photovoltaic arrays that convert solar energy to electrical power and an integrated equipment assembly (IEA) that houses electrical hardware and batteries for electric power regulation and storage. Each PV module contains two independent power channels for fault tolerance. Each power channel contains three batteries in parallel to meet its performance requirements and for fault tolerance. Each battery consists of 76 Ni-Hydrogen (Ni-H2) cells in series. These 76 cells are contained in two orbital replaceable units (ORU) that are connected in series. On-orbit data are monitored and trended to ensure that all hardware is operating normally. Review of on-orbit data showed that while five batteries are operating very well, one is showing signs of mismatched ORUs. The cell pressure in the two ORUs differs by an amount that exceeds the recommended range. The reason for this abnormal behavior may be that the two ORUs have different use history. An assessment was performed and it was determined that capacity of this battery would be limited by the lower pressure ORU. Steps are being taken to reduce this pressure differential before battery capacity drops to the point of affecting its ability to meet performance requirements. As a first step, a battery reinitialization procedure was developed to reduce this pressure differential. The procedure was successfully carried out on-orbit and the pressure differential was reduced to the recommended range. This paper describes the battery performance and the consequences of mismatched ORUs that make a battery. The paper also describes the reinitialization procedure, how it was performed on orbit, and battery performance after the reinitialization. On-orbit data monitoring and trending is an ongoing activity and it will continue as ISS assembly progresses.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211713 , E-13464 , NAS 1.15:211713 , IECEC-2002-20033 , 37th Intersociety Energy Conversion Engineering Conference; Jul 28, 2002 - Aug 02, 2002; Washington, DC; United States
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    Phase 1 Space Fission Propulsion Energy Source Design (2002)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Fission technology can enable rapid, affordable access to any point in the solar system. If fission propulsion systems are to be developed to their full potential; however, near-term customers must be identified and initial fission systems successfully developed, launched, and operated. Studies conducted in fiscal year 2001 (IISTP, 2001) show that fission electric propulsion (FEP) systems with a specific mass at or below 50 kg/kWjet could enhance or enable numerous robotic outer solar system missions of interest. At the required specific mass, it is possible to develop safe, affordable systems that meet mission requirements. To help select the system design to pursue, eight evaluation criteria were identified: system integration, safety, reliability, testability, specific mass, cost, schedule, and programmatic risk. A top-level comparison of four potential concepts was performed: a Testable, Passive, Redundant Reactor (TPRR), a Testable Multi-Cell In-Core Thermionic Reactor (TMCT), a Direct Gas Cooled Reactor (DGCR), and a Pumped Liquid Metal Reactor.(PLMR). Development of any of the four systems appears feasible. However, for power levels up to at least 500 kWt (enabling electric power levels of 125-175 kWe, given 25-35% power conversion efficiency) the TPRR has advantages related to several criteria and is competitive with respect to all. Hardware-based research and development has further increased confidence in the TPRR approach. Successful development and utilization of a "Phase I" fission electric propulsion system will enable advanced Phase 2 and Phase 3 systems capable of providing rapid, affordable access to any point in the solar system.
    Keywords: Spacecraft Propulsion and Power
    Type: 2002 American Nuclear Society Meeting: International Congress on Advanced Nuclear Power Plants; Jun 09, 2002 - Jun 13, 2002; Hollywood, FL; United States
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    Space Shuttle Five-Segment Booster (Short Course) (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: NASA is considering upgrading the Space Shuttle by adding a fifth segment (FSB) to the current four-segment solid rocket booster. Course materials cover design and engineering issues related to the Reusable Solid Rocket Motor (RSRM) raised by the addition of a fifth segment to the rocket booster. Topics cover include: four segment vs. five segment booster, abort modes, FSB grain design, erosive burning, enhanced propellant burn rate, FSB erosive burning model development and hardware configuration.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA RSRM Short Course; Jul 07, 2002 - Jul 10, 2002; Indianapolis, IN; United States
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    The MAP Propulsion Subsystem (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper describes the requirements, design, integration, test, performance, and lessons learned of NASA's Microwave Anisotropy Probe (MAP) propulsion subsystem. MAP was launched on a Delta-II launch vehicle from NASA's Kennedy Space Center on June 30, 2001. Due to instrument thermal stability requirements, the Earth-Sun L2 Lagrange point was selected for the mission orbit. The L2 trajectory incorporated phasing loops and a lunar gravity assist. The propulsion subsystem's requirements are to manage momentum, perform maneuvers during the phasing loops to set up the lunar swingby, and perform stationkeeping at L2 for 2 years. MAP's propulsion subsystem uses 8 thrusters which are located and oriented to provide attitude control and momentum management about all axes, and delta-V in any direction without exposing the instrument to the sun. The propellant tank holds 72 kg of hydrazine, which is expelled by unregulated blowdown pressurization. Thermal management is complex because no heater cycling is allowed at L2. Several technical challenges presented themselves during I and T, such as in-situ weld repairs and in-situ bending of thruster tubes to accommodate late changes in the observatory CG. On-orbit performance has been nominal, and all phasing loop, mid-course correction, and stationkeeping maneuvers have been successfully performed to date.
    Keywords: Spacecraft Propulsion and Power
    Type: AIAA-2002-4156 , AIAA/ASME/SAE/ASEE 38th Joint Propulsion Conference; Jul 07, 2002 - Jul 10, 2002; Indianapolis, IN; United States
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    Ablative Laser Propulsion Using Multi-Layered Material Systems (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Experimental investigations are ongoing to study the force imparted to materials when subjected to laser ablation. When a laser pulse of sufficient energy density impacts a material, a small amount of the material is ablated. A torsion balance is used to measure the momentum produced by the ablation process. The balance consists of a thin metal wire with a rotating pendulum suspended in the middle. The wire is fixed at both ends. Recently, multi-layered material systems were investigated. These multi-layered materials were composed of a transparent front surface and opaque sub surface. The laser pulse penetrates the transparent outer surface with minimum photon loss and vaporizes the underlying opaque layer.
    Keywords: Spacecraft Propulsion and Power
    Type: 33rd AIAA Plasmadynamics and Lasers Conference; May 20, 2002 - May 23, 2002; Maui, HI; United States
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    On-Orbit Propulsion System Project Overview (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This conference presentation reports on the progress on NASA's On-Orbit Propulsion System Project which aims to support the development of second generation reusable launch vehicles (RLV) through advanced research and development and risk reduction activities. Topics covered include: project goals, project accomplishments, risk reduction activities, thruster design and development initiatives, and Aerojet LOX/Ethanol engine development and testing.
    Keywords: Spacecraft Propulsion and Power
    Type: 1st AIAA/IAF Symposium on Future Reusable Launch Vehicles; Apr 11, 2002 - Apr 12, 2002; Huntsville, AL; United States
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    Integrated System Test of an Airbreathing Rocket (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This viewgraph presentation provides information on NASA's attempts to develop an air-breathing propulsion in an effort to make future space transportation safer, more reliable and significantly less expensive than today's missions. Spacecraft powered by air-breathing rocket engines would be completely reusable, able to take off and land at airport runways and ready to fly again within days. A radical new engine project is called the Integrated System Tests of an Air-breathing Rocket, or ISTAR.
    Keywords: Spacecraft Propulsion and Power
    Type: 6th International Symposium on Propulsion for Space Transportation for the 21st Century; May 14, 2002 - May 16, 2002; Versailles; France
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    Selection and Prioritization of Advanced Propulsion Technologies for Future Space Missions (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The exploration of our solar system will require spacecraft with much greater capability than spacecraft which have been launched in the past. This is particularly true for exploration of the outer planets. Outer planet exploration requires shorter trip times, increased payload mass, and ability to orbit or land on outer planets. Increased capability requires better propulsion systems, including increased specific impulse. Chemical propulsion systems are not capable of delivering the performance required for exploration of the solar system. Future propulsion systems will be applied to a wide variety of missions with a diverse set of mission requirements. Many candidate propulsion technologies have been proposed but NASA resources do not permit development of a] of them. Therefore, we need to rationally select a few propulsion technologies for advancement, for application to future space missions. An effort was initiated to select and prioritize candidate propulsion technologies for development investment. The results of the study identified Aerocapture, 5 - 10 KW Solar Electric Ion, and Nuclear Electric Propulsion as high priority technologies. Solar Sails, 100 Kw Solar Electric Hall Thrusters, Electric Propulsion, and Advanced Chemical were identified as medium priority technologies. Plasma sails, momentum exchange tethers, and low density solar sails were identified as high risk/high payoff technologies.
    Keywords: Spacecraft Propulsion and Power
    Type: 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit; Jul 07, 2002 - Jul 10, 2002; Indianapolis, IN; United States
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    Intelligent Systems for Power Management and Distribution (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The motivation behind an advanced technology program to develop intelligent power management and distribution (PMAD) systems is described. The program concentrates on developing digital control and distributed processing algorithms for PMAD components and systems to improve their size, weight, efficiency, and reliability. Specific areas of research in developing intelligent DC-DC converters and distributed switchgear are described. Results from recent development efforts are presented along with expected future benefits to the overall PMAD system performance.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211370/REV1 , NAS 1.15:211370/REV1 , E-13192-1/REV1 , IECEC2001-AT-40 , 36th Intersociety Energy Conversion Engineering Conference; Jul 29, 2001 - Aug 02, 2001; Savannah, GA; United States
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    Single Axis Attitude Control and DC Bus Regulation with Two Flywheels (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A computer simulation of a flywheel energy storage single axis attitude control system is described. The simulation models hardware which will be experimentally tested in the future. This hardware consists of two counter rotating flywheels mounted to an air table. The air table allows one axis of rotational motion. An inertia DC bus coordinator is set forth that allows the two control problems, bus regulation and attitude control, to be separated. Simulation results are presented with a previously derived flywheel bus regulator and a simple PID attitude controller.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211812 , NAS 1.15:211812 , E-13507 , IECEC-2002-20078 , 37th Intersociety Energy Conversion Engineering Conference; Jun 28, 2002 - Aug 02, 2002; Washington, DC; United States
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    Comparison of ISS Power System Telemetry with Analytically Derived Data for Shadowed Cases (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Accurate International Space Station (ISS) power prediction requires the quantification of solar array shadowing. Prior papers have discussed the NASA Glenn Research Center (GRC) ISS power system tool SPACE (System Power Analysis for Capability Evaluation) and its integrated shadowing algorithms. On-orbit telemetry has become available that permits the correlation of theoretical shadowing predictions with actual data. This paper documents the comparison of a shadowing metric (total solar array current) as derived from SPACE predictions and on-orbit flight telemetry data for representative significant shadowing cases. Images from flight video recordings and the SPACE computer program graphical output are used to illustrate the comparison. The accuracy of the SPACE shadowing capability is demonstrated for the cases examined.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211715 , E-13467 , IECEC-2002-20113 , NAS 1.15:211715 , 37th Intersociety Energy Conversion Engineering Conference; Jul 28, 2002 - Aug 02, 2002; Washington, DC; United States
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    A Parametric Assessment of the Mission Applicability of Thin-film Solar Arrays (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Results are presented from a parametric assessment of the applicability and spacecraft-level impacts of very lightweight thin-film solar arrays with relatively large deployed areas for representative space missions. The most and least attractive features of thin-film solar arrays are briefly discussed. A calculation is then presented illustrating that from a solar array alone mass perspective, larger arrays with less efficient but lighter thin-film solar cells can weigh less than smaller arrays with more efficient but heavier crystalline cells. However, a spacecraft-level systems assessment must take into account the additional mass associated with solar array deployed area: the propellant needed to desaturate the momentum accumulated from area-related disturbance torques and to perform aerodynamic drag makeup reboost. The results for such an assessment are presented for a representative low Earth orbit (LEO) mission, as a function of altitude and mission life, and a geostationary Earth orbit (GEO) mission. Discussion of the results includes a list of specific mission types most likely to benefit from using thin-film arrays. The presentation concludes with a list of issues to be addressed prior to use of thin-film solar arrays in space and the observation that with their unique characteristics, very lightweight arrays using efficient, thin film cells on flexible substrates may become the best array option for a subset of Earth orbiting and deep space missions.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211720 , E-13471 , NAS 1.15:211720 , Space Power Workshop 2002; Apr 22, 2002 - Apr 25, 2002; Redondo Beach, CA; United States
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    Reinventing the Solar Power Satellite (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-19
    Description: Economy of scale is inherent in the microwave power transmission aperture/spot-size trade-off, resulting in a requirement for large space systems in the existing design concepts. Unfortunately, this large size means that the initial investment required before the first return, and the price of amortization of this initial investment, is a daunting (and perhaps insurmountable) barrier to economic viability. As the growth of ground-based solar power applications will fund the development of the PV technology required for space solar power and will also create the demand for space solar power by manufacturing a ready-made market, space power systems must be designed with an understanding that ground-based solar technologies will be implemented as a precursor to space-based solar. for low initial cost, (3) operation in synergy with ground solar systems, and (4) power production profile tailored to peak rates. A key to simplicity of design is to maximize the integration of the system components. Microwave, millimeter-wave, and laser systems are analyzed. A new solar power satellite design concept with no sun-tracking and no moving parts is proposed to reduce the required cost to initial operational capability.
    Keywords: Spacecraft Propulsion and Power
    Type: Paper IAC-02-R.1.06 , 53rd International Astronautical Congress; Oct 10, 2002 - Oct 19, 2002; Houston, TX; United States|Second World Space Congress; Oct 10, 2002 - Oct 19, 2002; Houston, TX; United States|34th COSPAR Scientific Assembly; Oct 10, 2002 - Oct 19, 2002; Houston, TX; United States
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    Momentum and Heat Flux Measurements in the Exhaust of VASIMR Using Helium Propellant (2002)
    In:  Other Sources
    Details
    Publication Date: 2019-07-18
    Description: Electromagnetic thrusters typically use electric and magnetic fields to accelerate and exhaust plasma through interactions with the charged particles in the plasma. The energy required to create the plasma, i.e. ionization energy, is potential energy between the electron and ion. This potential energy is typically lost since it is not recovered as the plasma is exhausted and is known as frozen flow loss. If the frozen flow energy is a small fraction of the total plasma energy, this frozen flow loss may be negligible. However, if the frozen flow energy is a major fraction of the total plasma energy, this loss can severely reduce the energy efficiency of the thruster. Recovery and utilization of this frozen flow energy can improve the energy efficiency of a thruster during low specific impulse operating regimes when the ionization energy is a large fraction of the total plasma energy. This paper quantifies the recovery of the frozen flow energy, i.e. recombination energy, via the process of surface recombination for helium. To accomplish this task the momentum flux and heat flux of the plasma flow were measured and compared to calculated values from electrostatic probe data. This information was used to deduce the contribution of recombination energy to the total heat flux on a flat plate as well as to characterize the plasma conditions. Helium propellant was investigated initially due to its high ionization potential and hence available recombination energy.
    Keywords: Spacecraft Propulsion and Power
    Type: International Electric Propulsion Conference; Mar 17, 2003 - Mar 21, 2003; Toulouse; France
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    Real-Time Embedded Software Verification and Validation 2001 (2002)
    In:  Other Sources
    Details
    Publication Date: 2019-07-18
    Description: As the space applications become more complex and timing constraints on control actions are more stringent, the task of integrating and testing NASA's real-time systems (such as X-38 Crew Return Vehicle, and certain International Space Station autonomous systems) has become a great challenge. A testing environment where can preserve consistent temporal behaviors as in the target execution must be established for system-level verification and software quality assurance. Our goal is to develop an analysis suite for validation and verification of real-time systems that are used to perform human- in-the-loop control operations during safety-critical missions. The suite will be able to carry out quantitative approaches of coverage diagnostic and temporal behavior evaluation in order to measure test coverage, to optimize test utilization, and to verify timing correctness.
    Keywords: Spacecraft Propulsion and Power
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    Analysis of Parallel Burn Without Crossfeed TSTO RLV Architectures and Comparison to Parallel Burn With Crossfeed and Series Burn Architectures (2002)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: There are currently three dominant TSTO class architectures. These are Series Burn (SB), Parallel Burn with crossfeed (PBw/cf), and Parallel Burn without crossfeed (PBncf). The goal of this study was to determine what factors uniquely affect PBncf architectures, how each of these factors interact, and to determine from a performance perspective whether a PBncf vehicle could be competitive with a PBw/cf or SB vehicle using equivalent technology and assumptions. In all cases, performance was evaluated on a relative basis for a fixed payload and mission by comparing gross and dry vehicle masses of a closed vehicle. Propellant combinations studied were LOX: LH2 propelled orbiter and booster (HH) and LOX: Kerosene booster with LOX: LH2 orbiter (KH). The study conclusions were: 1) a PBncf orbiter should be throttled as deeply as possible after launch until the staging point. 2) a detailed structural model is essential to accurate architecture analysis and evaluation. 3) a PBncf TSTO architecture is feasible for systems that stage at mach 7. 3a) HH architectures can achieve a mass growth relative to PBw/cf of 〈 20%. 3b) KH architectures can achieve a mass growth relative to Series Burn of 〈 20%. 4) center of gravity (CG) control will be a major issue for a PBncf vehicle, due to the low orbiter specific thrust to weight ratio and to the position of the orbiter required to align the nozzle heights at liftoff. 5 ) thrust to weight ratios of 1.3 at liftoff and between 1.0 and 0.9 when staging at mach 7 appear to be close to ideal for PBncf vehicles. 6) performance for all vehicles studied is better when staged at mach 7 instead of mach 5. The study showed that a Series Burn architecture has the lowest gross mass for HH cases, and has the lowest dry mass for KH cases. The potential disadvantages of SB are the required use of an air-start for the orbiter engines and potential CG control issues. A Parallel Burn with crossfeed architecture solves both these problems, but the mechanics of a large bipropellant crossfeed system pose significant technical difficulties. Parallel Burn without crossfeed vehicles start both booster and orbiter engines on the ground and thus avoid both the risk of orbiter air-start and the complexity of a crossfeed system. The drawback is that the orbiter must use 20% to 35% of its propellant before reaching the staging point. This induces a weight penalty in the orbiter in order to carry additional propellant, which causes a further weight penalty in the booster to achieve the same staging point. One way to reduce the orbiter propellant consumption during the first stage is to throttle down the orbiter engines as much as possible. Another possibility is to use smaller or fewer engines. Throttling the orbiter engines soon after liftoff minimizes CG control problems due to a low orbiter liftoff thrust, but may result in an unnecessarily high orbiter thrust after staging. Reducing the number or size of engines size may cause CG control problems and drift at launch. The study suggested possible methods to maximize performance of PBncf vehicle architectures in order to meet mission design requirements.
    Keywords: Spacecraft Propulsion and Power
    Type: 39th Joint Propulsion Conference; Jul 20, 2003 - Jul 23, 2003; Huntsville, AL; United States
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    Fusion for Space Propulsion (2002)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: The need for fusion propulsion for interplanetary flights is discussed. For a propulsion system, there are three important system attributes: (1) The absolute amount of energy available, (2) the propellant exhaust velocity, and (3) the jet power per unit mass of the propulsion system (specific power). For efficient and affordable human exploration of the solar system, propellant exhaust velocity in excess of 100 km/s and specific power in excess of 10 kW/kg are required. Chemical combustion obviously cannot meet the requirement in propellant exhaust velocity. Nuclear fission processes typically result in producing energy in the form of heat that needs to be manipulated at temperatures limited by materials to about 2,800 K. Using the fission energy to heat a low atomic weight propellant produces propellant velocity of the order of 10 kinds. Alternatively the fission energy can be converted into electricity that is used to accelerate particles to high exhaust velocity. However, the necessary power conversion and conditioning equipment greatly increases the mass of the propulsion system. Fundamental considerations in waste heat rejection and power conditioning in a fission electric propulsion system place a limit on its jet specific power to the order of about 0.2 kW/kg. If fusion can be developed for propulsion, it appears to have the best of all worlds - it can provide the largest absolute amount of energy, the propellant exhaust velocity (〉 100 km/s), and the high specific jet power (〉 10 kW/kg). An intermediate step towards fusion propulsion might be a bimodal system in which a fission reactor is used to provide some of the energy to drive a fusion propulsion unit. There are similarities as well as differences between applying fusion to propulsion and to terrestrial electrical power generation. The similarities are the underlying plasma and fusion physics, the enabling component technologies, the computational and the diagnostics capabilities. These physics and engineering capabilities have been demonstrated for a fusion reactor gain (Q) of the order of unity (TFTR: 0.25, JET: 0.65, JT-60: Q(sub eq) approx. 1.25). These technological advances made it compelling for considering fusion for propulsion.
    Keywords: Spacecraft Propulsion and Power
    Type: American Nuclear Society (ANS); Jun 03, 2002 - Jun 09, 2002; Hollywood, FL; United States
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    Power System for Venus Surface Exploration (2002)
    In:  Other Sources
    Details
    Publication Date: 2019-07-18
    Description: A radioisotope power and cooling system is designed to provide electrical power for a probe operating on the surface of Venus. Most foreseeable electronics devices and sensors cannot operate at the 450 C ambient surface temperature of Venus. Because the mission duration is substantially long and the use of thermal mass to maintain an operable temperature range is likely impractical, some type of active refrigeration may be required to keep electronic components at a temperature below ambient. The fundamental cooling parameters are the cold sink temperature, the hot sink temperature, and the amount of heat to be removed. In this instance, it is anticipated that electronics would have a nominal operating temperature of 300 C. Due to the highly thermal convective nature of the high-density (90 bar CO2) atmosphere, the hot sink temperature was assumed to be 50 C, which provided a 500 C temperature of the cooler's heat rejecter to the ambient atmosphere. The majority of the heat load on the cooler is from the high temperature ambient surface environment on Venus, with a small contribution of heat generation from electronics and sensors. Both thermoelectric (RTG) and dynamic power conversion systems were analyzed, based on use of a standard isotope (General-purpose heat source, or GPHS) brick. For the radioisotope Stirling power converter configuration designed, the Sage model predicts a thermodynamic power output capacity of 478.1 watts, which slightly exceeds the required 469.1 watts. The hot sink temperature is 1200 C, and the cold sink temperature is 500 C. The required heat input is 1740 watts. This gives a thermodynamic efficiency of 27.48 %. It is estimated that the mechanical efficiency of the power converter design is on the order of 85 %, based on experimental measurements taken from 500-watt power class, laboratory-tested Stirling engines. The overall efficiency is calculated to be 23.36 %. The mass of the power converter is estimated at approximately 21.6 kg. Additional information is included in the original extended abstract.
    Keywords: Spacecraft Propulsion and Power
    Type: Seventh European Space Power Conference; May 09, 2005 - May 13, 2005; Como; Italy
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    Stennis Space Center Propulsion Test Customer Day 2002 Presentation (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: An overview and breakdown of the 2001 workforce at the Stennis Space Center is provided.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/IS-2002-05-2003-SSC
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    Sputtering Erosion Measurement on Boron Nitride as a Hall Thruster Material (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: The durability of a high-powered Hall thruster may be limited by the sputter erosion resistance of its components. During normal operation, a small fraction of the accelerated ions will impact the interior of the main discharge channel, causing its gradual erosion. A laboratory experiment was conducted to simulate the sputter erosion of a Hall thruster. Tests of sputter etch rate were carried out using 300 to 1000 eV Xenon ions impinging on boron nitride substrates with angles of attack ranging from 30 to 75 degrees from horizontal. The erosion rates varied from 3.41 to 14.37 Angstroms/[sec(mA/sq cm)] and were found to depend on the ion energy and angle of attack, which is consistent with the behavior of other materials.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211837 , NAS 1.15:211837 , E-13537
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    Rocket Motor Joint Construction Including Thermal Barrier (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: A thermal barrier for extremely high temperature applications consists of a carbon fiber core and one or more layers of braided carbon fibers surrounding the core. The thermal barrier is preferably a large diameter ring, having a relatively small cross-section. The thermal barrier is particularly suited for use as part of a joint structure in solid rocket motor casings to protect low temperature elements such as the primary and secondary elastomeric O-ring seals therein from high temperature gases of the rocket motor. The thermal barrier exhibits adequate porosity to allow pressure to reach the radially outward disposed O-ring seals allowing them to seat and perform the primary sealing function. The thermal barrier is disposed in a cavity or groove in the casing joint, between the hot propulsion gases interior of the rocket motor and primary and secondary O-ring seals. The characteristics of the thermal barrier may be enhanced in different applications by the inclusion of certain compounds in the casing joint, by the inclusion of RTV sealant or similar materials at the site of the thermal barrier, and/or by the incorporation of a metal core or plurality of metal braids within the carbon braid in the thermal barrier structure.
    Keywords: Spacecraft Propulsion and Power
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    Comprehensive Evaluation of Power Supplies at Cryogenic Temperatures for Deep Space Applications (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: The operation of power electronic systems at cryogenic temperatures is anticipated in many future space missions such as planetary exploration and deep space probes. In addition to surviving the space hostile environments, electronics capable of low temperature operation would contribute to improving circuit performance, increasing system efficiency, and reducing development and launch costs. DC/DC converters are widely used in space power systems in the areas of power management, conditioning, and control. As part of the on-going Low Temperature Electronics Program at NASA, several commercial-off-the-shelf (COTS) DC/DC converters, with specifications that might fit the requirements of specific future space missions have been selected for investigation at cryogenic temperatures. The converters have been characterized in terms of their performance as a function of temperature in the range of 20 C to - 180 C. These converters ranged in electrical power from 8 W to 13 W, input voltage from 9 V to 72 V and an output voltage of 3.3 V. The experimental set-up and procedures along with the results obtained on the converters' steady state and dynamic characteristics are presented and discussed.
    Keywords: Spacecraft Propulsion and Power
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    Low-Current, Xenon Orificed Hollow Cathode Performance for In-Space Applications (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: An experimental investigation of the operating characteristics of 3.2-mm diameter orificed hollow cathodes was conducted to examine low current and low flow rate operation. Cathode power was minimized with an orifice aspect ratio of approximately one and the use of an enclosed keeper. Cathode flow rate requirements were proportional to orifice diameter and the inverse of the orifice length. The minimum power consumption in diode mode was 10-W, and the minimum mass flow rate required for spot-mode emission was approximately 0.08-mg/s. Cathode temperature profiles were obtained using an imaging radiometer and conduction was found to be the dominant heat transfer mechanism from the cathode tube. Orifice plate temperatures were found to be weakly dependent upon the flow rate and strongly dependent upon the current.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211574 , NAS 1.15:211574 , E-13358
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    Atomic-Based Combined Cycle Propulsion System and Method (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: A method and system are provided for propelling an aerodynamic vehicle into space. The aerodynamic vehicle uses a nuclear-based thermal rocket (NTR) propulsion system capable of producing a hydrogen exhaust. A flow of air is introduced into the hydrogen exhaust to augment the thrust force at speeds of the vehicle up to approximately Mach 6. When the speed of the vehicle is approximately Mach 6 and the altitude of the vehicle is approximately 40 kilometers, the flow of air is stopped and the vehicle is propelled into space using only the NTR.
    Keywords: Spacecraft Propulsion and Power
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    Rocket Propellants Engine Design/Operations/Validation (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Lockheed Martin Astronautics Operations (LMA) was competitively awarded a contract May 21, 2001 for next generation launch system architecture definition and technology maturation. The Second Generation Launch Vehicle Program objectives include reducing the technical and programmatic risk of proceeding to full scale development of the system by establishing requirements for the next generation launch system and maturing critical technologies needed by the system. LMA will conduct analyses and trades to optimize the architecture ETO elements including configuration, conceptual designs, and preliminary operations definition. To fully understand the engine and propellant trades were conducted by LMA to yield the optimized architecture system from the operability, reliability, safety, and cost perspectives. A government/industry team addressed the required trade studies, the parameters and weighting factors, and the most critical trades were addressed. This report summarizes the participation of JCM Consulting, Inc. in the propellant trade study.
    Keywords: Spacecraft Propulsion and Power
    Type: NONP-NASA-CD-2002032679
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    Power Control and Monitoring Requirements for Thermal Vacuum/Thermal Balance Testing of the MAP Observatory (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The specific heater control requirements for the thermal vacuum and thermal balance testing of the Microwave Anisotropy Probe (MAP) Observatory at the Goddard Space Flight Center (GSFC) in Greenbelt, Maryland are described. The testing was conducted in the 10m wide x 18.3m high Space Environment Simulator (SES) Thermal Vacuum Facility. The MAP thermal testing required accurate quantification of spacecraft and fixture power levels while minimizing heater electrical emissions. The special requirements of the MAP test necessitated construction of five (5) new heater racks.
    Keywords: Spacecraft Propulsion and Power
    Type: 22nd IEST-NASA/ASTM/AIAA/CSA Space Simulation Conference; Oct 21, 2002 - Oct 24, 2002
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    Turbulent Mixing of Primary and Secondary Flow Streams in a Rocket-Based Combined Cycle Engine (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: This viewgraph presentation gives an overview of the turbulent mixing of primary and secondary flow streams in a rocket-based combined cycle (RBCC) engine. A significant RBCC ejector mode database has been generated, detailing single and twin thruster configurations and global and local measurements. On-going analysis and correlation efforts include Marshall Space Flight Center computational fluid dynamics modeling and turbulent shear layer analysis. Potential follow-on activities include detailed measurements of air flow static pressure and velocity profiles, investigations into other thruster spacing configurations, performing a fundamental shear layer mixing study, and demonstrating single-shot Raman measurements.
    Keywords: Spacecraft Propulsion and Power
    Type: JANNAF 38th Combustion Subcommittee Meeting; Apr 08, 2002 - Apr 12, 2002; Destin, FL; United States
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    Design and Study of a LOX/GH2 Throttleable Swirl Injector for Rocket Applications (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: A LOX/GH2 swirl injector was designed for a 10:1 propellant throttling range. To accomplish this, a dual LOX (liquid oxygen) manifold was used feeding a single common vortex chamber of the swirl element. Hot-fire experiments were conducting for rocket chamber pressures from 80 to 800 psia at a mixture ratio of nominally 6.0 using steady flow, single-point-per-firing cases as well as dynamic throttling conditions. Low frequency (mean) and high frequency (fluctuating) pressure transducer data, flow meter measurements, and Raman spectroscopy images for mixing information were obtained. The injector design, experimental setup, low frequency pressure data, and injector performance analysis will be presented. C efficiency was very high (approximately 100%) at the middle of the throttle-able range with somewhat lower performance at the high and low ends. From the analysis of discreet steady state operating conditions, injector pressure drop was slightly higher than predicted with an inviscid analysis, but otherwise agreed well across the design throttling range. Analysis of the dynamic throttling data indicates that the injector may experience transient conditions that effect pressure drop and performance when compared to steady state results.
    Keywords: Spacecraft Propulsion and Power
    Type: 2002 JANNAF 38th Combustion Subcommittee Meeting; Apr 08, 2002 - Apr 12, 2002; Destin, FL; United States
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    Measurement of Doubly Charged Ions in Ion Thruster Plumes (2002)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The ratio of doubly to singly charged ions was measured in the plumes of a 30 cm and of a 40 cm ion thruster. The measured ratio was correlated with observed erosion rates and thruster operating conditions. The measured and calculated erosion rates paralleled variation in the j(sup ++)/j(sup +) ratio and indicated that the erosion was dominated by Xe III. Simple models of cathode potential surfaces which were developed in support of this work were in agreement with this conclusion and provided a predictive capability of the erosion given the ratio of doubly to singly charged ion currents.
    Keywords: Spacecraft Propulsion and Power
    Type: NASA/TM-2002-211295 , E-13098 , NAS 1.15:211295 , IEPC-01-310 , 27th International Electric Propulsion Conference; Oct 14, 2001 - Oct 19, 2001; Pasadena, CA; United States
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