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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. Author Microwave Anisotropy Probe; Observatories; Thermal Vacuum Tests; Power Supplies; Control Equipment; Heating Equipment; Spacecraft Temperature 20030022776 NASA Glenn Research Center, Cleveland, OH, USA Catalyzed Ignition of Bipropellants in Microtubes Schneider, Steven J.; Boyarko, George A.; Sung, Chih-Jen; February 2003; 17 pp.; In English; 41st Aerospace Sciences Meeting and Exhibit, 6-9 Jan. 2003, Reno, NV, USA; Original contains color illustrations Contract(s)/Grant(s): WBS-22-755-04-08 Report No.(s): NASA/TM-2003-212206; NAS 1.15:212206; E-13810; AIAA Paper 2003?0674; Copyright; Avail: CASI; A03, Hardcopy This paper addresses the need to understand the physics and chemistry involved in propellant combustion processes in micro-scale combustors for propulsion systems on micro-spacecraft. These spacecraft are planned to have a mass less than 50 kilograms with attitude control estimated to be in the 10 milli-Newton thrust class. These combustors are anticipated to be manufactured using Micro Electrical Mechanical Systems (MEMS) technology and are expected to have diameters approaching the quenching diameter of the propellants. Combustors of this size are expected to benefit significantly from surface catalysis processes. Miniature flame tube apparatus is chosen for this study because microtubes can be easily fabricated from known catalyst materials and their simplicity in geometry can be used in fundamental simulations to more carefully characterize the measured heat transfer and pressure losses for validation purposes. Experimentally, we investigate the role of catalytically active surfaces within 0.4 and 0.8 mm internal diameter micro-tubes, with special emphases on ignition and extinction processes in fuel rich gaseous hydrogen and gaseous oxygen. Flame thickness and reaction zone thickness calculations predict that the diameters of our test apparatus are below the quenching diameter of the propellants in sub-atmospheric tests. Temperature and pressure rises in resistively heated platinum and palladium micro-tubes are used as an indication of exothermic reactions. Specific data on mass flow versus preheat temperature required to achieve ignition are presented. Author Chemical Propulsion; Spacecraft Propulsion; Rocket Engines; Low Thrust Propulsion; Liquid Rocket Propellants; Combustion Physics; Combustion Chemistry; Microstructure 20030025254 NASA Goddard Space Flight Center, Greenbelt, MD, USA A Comparison of Trajectory Optimization Methods for the Impulsive Minimum Fuel Rendezvous Problem Hughes, Steven P.; Mailhe, Laurie M.; Guzman, Jose J.; February 2003; 20 pp.; In English; 26th Annual Guidance and Control Conference, Feb. 2003, Breckenridge, CO., USA; Copyright; Avail: CASI; A03, Hardcopy In this paper we present, a comparison of trajectory optimization approaches for 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 and Nelder-Meade Simplex. Several cost function parameterizations are considered for the direct approach. We choose one direct approach that appears to be the most flexible. Both the direct and indirect methods are applied to a variety of test cases which are chosen to demonstrate the performance of each method in different flight regimes. The first test case is a simple circular-to-circular coplanar rendezvous. The second test case is an elliptic-to-elliptic line of apsides rotation. The final test case is an orbit phasing maneuver sequence in a highly elliptic orbit. For each test case we present a comparison of the performance of all methods we consider in this paper. Author Trajectory Optimization; Parameterization; Fuel Control 20030025383 NASA Goddard Space Flight Center, Greenbelt, MD, USA Development of a Thrust Stand to Meet LISA Mission Requirements Willis, William D., III; Zakrzwski, Charles M.; Merkowitz, Stephen M.; [2002]; 9 pp.; In English; AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 7-10 Jul. 2002; Original contains black and white illustrations; No Copyright; Avail: CASI; A02, Hardcopy 29
A thrust stand has been built to measure the force-noise produced by electrostatic micro-Newton (muN) thrusters. The LISA mission’s Disturbance Reduction System (DRS) requires thrusters that are capable of producing continuous thrust levels between 1-100 muN with a resolution of 0.1 muN. The stationary force-noise produced by these thrusters must not exceed 0.1 muN/dHz in the measurement bandwidth 10(exp -4) to 1 Hz. The LISA Thrust Stand (LTS) is a torsion-balance type thrust stand designed to meet the following requirements: stationary force-noise measurements from l0( -4) to 1 Hz with 0.1 muN/dHz sensitivity, absolute thrust measurements from 1-100 muN with better than 0.1 muN resolution, and dynamic thruster response from to 10 Hz. The LTS employs a unique vertical configuration, autocollimator for angular position measurements, and electrostatic actuators that are used for dynamic pendulum control and null-mode measurements. Force-noise levels are measured indirectly by characterizing the thrust stand as a spring-mass system. The LTS was initially designed to test the indium FEEP thruster developed by the Austrian Research Center in Seibersdorf (ARCS), but can be modified for testing other thrusters of this type. Author Thrust Measurement; Thrust; Dynamic Response; Dynamic Control; Bandwidth; LISA (Observatory) 20030025704 NASA Langley Research Center, Hampton, VA, USA Mach 4 Performance of a Fixed-Geometry Hypersonic Inlet with Rectangular-to-Elliptical Shape Transition Smart, Michael K.; Trexler, Carl A.; [2003]; 12 pp.; In English; 41st AIAA Aerospace Sciences Meeting and Exhibit, 6-9 Jan. 2003, Reno, NV, USA; Original contains color illustrations Report No.(s): AIAA Paper 2003-0012; No Copyright; Avail: CASI; A03, Hardcopy Wind-tunnel testing of a hypersonic inlet with rectangular-to-elliptical shape transition has been conducted at Mach 4.0. These tests were performed to investigate the starting and back-pressure limits of this fixed-geometry inlet at conditions well below the Mach 5.7 design point. Results showed that the inlet required side spillage holes in order to self-start at Mach 4.0. Once started, the inlet generated a compression ratio of 12.6, captured almost 80\% of available air and withstood a back-pressure ratio of 30.3 relative to tunnel static pressure. The spillage penalty for self-starting was estimated to be 4\% of available air. These experimental results, along with previous experimental results at Mach 6.2 indicate that fixed-geometry inlets with rectangular-to-elliptical shape transition are a viable configuration for airframe-integrated scramjets that operate over a significant Mach number range. Nomenclature Author Hypersonic Inlets; Mach Number; Wind Tunnel Tests; Shapes; Pressure Ratio 20030027869 NASA Glenn Research Center, Cleveland, OH, USA Space Nuclear Power and Propulsion Systems Technology: Enabling Future Planetary Exploration. Segment 7: Power Management and Distribution Soltis, James V.; February 1, 2003; 52 pp.; In English; Space Technology and Application International Forum, 1-2 Feb. 2003, Albuquerque, NM, USA; Original contains black and white illustrations; No Copyright; Avail: CASI; A04, Hardcopy Control and integration of energy conversion in space nuclear reactor power systems is discussed. Specific topics include: Systems Engineering of the PMAD subsystem and interference with other subsystems, including environmental and mission unique considerations, and integration issues with payloads. Author Nuclear Propulsion; Energy Conversion; Systems Engineering; Spacecraft Control; Spacecraft Power Supplies; Nuclear Reactors 20030028448 NASA Glenn Research Center, Cleveland, OH, USA Innovative Airbreathing Propulsion Concepts for High-speed Applications Whitlow, Woodrow, Jr.; [2002]; 22 pp.; In English; 23rd Congress of International Council of the Aeronautical Sciences, 9-12 Sep. 2002; Original contains black and white illustrations; No Copyright; Avail: CASI; A03, Hardcopy The current cost to launch payloads to low earth orbit (LEO) is approximately loo00 U.S. dollars ($) per pound ($22000 per kilogram). This high cost limits our ability to pursue space science and hinders the development of new markets and a productive space enterprise. This enterprise includes NASA’s space launch needs and those of industry, universities, the military, and other U.S. government agencies. NASA’s Advanced Space Transportation Program (ASTP) proposes a vision of the future where space travel is as routine as in today’s commercial air transportation systems. Dramatically lower launch costs will be required to make this vision a reality. In order to provide more affordable access to space, NASA has established new goals in its Aeronautics and Space Transportation plan. These goals target a reduction in the cost of launching payloads 30
Page 2 and 3: NASA STI Program ... in Profile Sin
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A new approach is proposed to asses
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dance in the Arabian Sea and Bay of
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Operational Vertical Sounder) Verti
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The Moderate Resolution Imaging Spe
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droughts of the 1930s and 1950s. Th
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20030031380 NASA Goddard Space Flig
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Data from the Tropical Rainfall Mea
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Mid-latitude forest ecosystems have
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Understanding functional associatio
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cyanobacterial mats, oxygenic photo
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as CIAO/SDP software developers, we
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Exploration Rover (MER) expeditions
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62 COMPUTER SYSTEMS Includes comput
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data in guard cells leads to the pr
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71 ACOUSTICS Includes sound generat
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20030025751 NASA Langley Research C
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analyzed using quantum kinetic theo
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Science Through ARts (STAR) is an e
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85 TECHNOLOGY UTILIZATION AND SURFA
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20030022798 California Inst. of Tec
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mostly at intermediate velocities.
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close to the spacecraft. We then us
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This source had been a bright persi
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the X-ray properties of the cluster
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We revise the physics of primary el
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size of the gullies in relation to
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92 SOLAR PHYSICS Includes solar act
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society, GISS is part of the Earth
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Initial Operation and Checkout of S
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ATMOSPHERIC EFFECTS The Sun’s Imp
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Effect of a BN Interphase that Debo
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Scalability of a Low-Cost Multi-Ter
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DATA TRANSMISSION Reliable Transpor
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Atomic Data for the K-Vacancy State
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FORMATION FLYING A Nonlinear, Six-D
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HIGH TEMPERATURE Alumina Volatility
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Atomic Data for the K-Vacancy State
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MATRICES (MATHEMATICS) A Time Integ
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NASA’s Sentinels Monitoring Weath
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PACIFIC OCEAN Tropospheric Ozone Ov
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PROVING Experimental Verification o
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SACCHAROMYCES Redundancies in Large
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SOLAR CORONA Are CME ‘interaction
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STANDARDS NASA GSFC Mechanical Engi
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TENSILE STRESS Temperature Dependen
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TYPE 3 BURSTS Energetic Particle Pr
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Rayleigh-Taylor Gravity Waves and Q
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Balagadde, Frederick The Electrical
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Casey, Carolyn A. Systems Engineeri
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Donnelly, Michael Performance of th
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Gliebe, Philip Recent Progress in E
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Hughes, S. P. Primer Vector Optimiz
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Kirk, C. Laurence Effective Frequen
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Luckring, J. M. Transonic Reynolds
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Naasz, Bo J. Integrated Orbit and A
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Raffaelle, Ryne Future Photovoltaic
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Shirron, Peter J. A Cryogenic, Insu
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Urban and Rural Ozone Collect over
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Xu, Yun-Jun Coupled Attitude and Or
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