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  • 1
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    Polar Spin Axis Anomaly (1997)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: The Global Geospace Science (GGS) Polar Plasma Laboratory (POLAR) spacecraft was launched on February 24, 1996, by a Delta 2. The spacecraft, a major axis spinner, appeared to function nominally throughout the early mission phase, which included several deployments, and orbit and attitude maneuvers. Of particular interest is the fact that the spacecraft was launched with a deliberate dynamic imbalance. During a segment of early orbit operations, a pair of Lanyard Deployed Booms (LDB) were extended. These booms were not identical; the intent was that the spacecraft would be nearly dynamically balanced after they were deployed. The spacecraft contained two dynamic balance mechanisms intended to fine tune the balance on orbit. However, subsequent images taken by the science instruments on the Despun Platform during the dynamic balancing segment indicated an offset of the principal spin axis from the geometric axis. This offset produced a sinusoidal blurring of the science images sufficiently large to degrade science data below mission requirement specifications. In the end, the imbalance encountered in flight was significantly outside the correction capability of the balances. The purpose of this paper is to examine the flight data during the various deployment and maneuver stages of the early orbit operations coupled with analytical simulations to discuss some of the potential causes of the resultant imbalance.
    Keywords: Spacecraft Design, Testing and Performance
    Type: Flight Mechanics Symposium 1997; 17-31; NASA-CP-3345
    Format: text
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  • 2
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    Backup Attitude Control Algorithms for the MAP Spacecraft (1999)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: The Microwave Anisotropy Probe (MAP) is a follow-on to the Differential Microwave Radiometer (DMR) instrument on the Cosmic Background Explorer (COBE) spacecraft. The MAP spacecraft will perform its mission, studying the early origins of the universe, in a Lissajous orbit around the Earth-Sun L(sub 2) Lagrange point. Due to limited mass, power, and financial resources, a traditional reliability concept involving fully redundant components was not feasible. This paper will discuss the redundancy philosophy used on MAP, describe the hardware redundancy selected (and why), and present backup modes and algorithms that were designed in lieu of additional attitude control hardware redundancy to improve the odds of mission success. Three of these modes have been implemented in the spacecraft flight software. The first onboard mode allows the MAP Kalman filter to be used with digital sun sensor (DSS) derived rates, in case of the failure of one of MAP's two two-axis inertial reference units. Similarly, the second onboard mode allows a star tracker only mode, using attitude and derived rate from one or both of MAP's star trackers for onboard attitude determination and control. The last backup mode onboard allows a sun-line angle offset to be commanded that will allow solar radiation pressure to be used for momentum management and orbit stationkeeping. In addition to the backup modes implemented on the spacecraft, two backup algorithms have been developed in the event of less likely contingencies. One of these is an algorithm for implementing an alternative scan pattern to MAP's nominal dual-spin science mode using only one or two reaction wheels and thrusters. Finally, an algorithm has been developed that uses thruster one shots while in science mode for momentum management. This algorithm has been developed in case system momentum builds up faster than anticipated, to allow adequate momentum management while minimizing interruptions to science. In this paper, each mode and algorithm will be discussed, and simulation results presented.
    Keywords: Spacecraft Design, Testing and Performance
    Type: 1999 Flight Mechanics Symposium; 391-405; NASA/CP-1999-209235
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  • 3
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    High-Temperature, High-Flux Multifoil Shield Developed for Space Applications (1998)
    In:  CASI
    Details
    Publication Date: 2018-06-02
    Description: Spacecraft employing solar dynamic power systems typically use parabolic, point focus concentrators to collect solar power and direct it to the aperture of a heat receiver. Solar fluxes several thousand times the intensity of one solar constant are typically produced in the focal plane of such concentrators. Under heat loading this severe, passively cooled surfaces constructed of most engineering materials would rapidly melt. Therefore, high-temperature shielding is required to protect heat receiver surfaces and other spacecraft surfaces that may be exposed to high flux. To meet this challenge for the joint U.S./Russian Solar Dynamic Flight Demonstration Program, AlliedSignal Aerospace and the NASA Lewis Research Center developed a high-temperature, high-flux multifoil shield tolerant of extreme heat loading conditions in a vacuum environment. The shield is passively cooled, obviating the need for pumped fluid loops and/or heat pipe cooling systems with their attendant cost, mass, complexity, and reliability issues.
    Keywords: Spacecraft Design, Testing and Performance
    Type: Research and Technology 1997; NASA/TM-1998-206312
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  • 4
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    Dark Forward Electrical Testing of the Mir Cooperative Solar Array (1997)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: During July 11 to 13, 1995, a team from the NASA Lewis Research Center performed dark forward electrical testing on the Mir Cooperative Solar Array (MCSA) flight unit in the Space Station Processing Facility at the NASA Kennedy Space Center. The MCSA was jointly designed and built by the United States and Russia to supply approximately 6 kW of electricity to the Russian Mir space station. The primary objective of testing was to assess the overall electrical performance of the flight array after handling and shipment from Russia to NASA Kennedy. This objective was achieved without the high cost and difficulties of deploying and illuminating the MCSA as is usually done with large-area solar arrays. The data obtained provided U.S. and Russian program managers with a high level of confidence in the MCSA electrical performance prior to the array's launch on shuttle mission STS-74 in November 1995 and its deployment on Mir in May 1996.
    Keywords: Spacecraft Design, Testing and Performance
    Type: Research and Technology 1996; NASA-TM-107350
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  • 5
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    CM-2 Environmental/Modal Testing of SPACEHAB Racks (2001)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: Combined environmental/modal vibration testing has been implemented at the NASA Glenn Research Center's Structural Dynamics Laboratory. The benefits of combined vibration testing are that it facilitates test article modal characterization and vibration qualification testing. The Combustion Module-2 (CM-2) is a space experiment that will launch on shuttle mission STS-107 in the SPACEHAB Research Double Module. The CM-2 flight hardware is integrated into a SPACEHAB single and double rack. CM-2 rack-level combined vibration testing was recently completed on a shaker table to characterize the structure's modal response and verify the random vibration response. Control accelerometers and limit force gauges, located between the fixture and rack interface, were used to verify the input excitation. Results of the testing were used to verify the loads and environments for flight on the shuttles.
    Keywords: Spacecraft Design, Testing and Performance
    Type: Research and Technology 2000; NASA/TM-2001-210605
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  • 6
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    An Innovative Manufacturing of CCC Ion Thruster Grids by North Carolina A&T's RTM Carbon/Carbon Process (2003)
    In:  Other Sources
    Details
    Publication Date: 2019-07-18
    Description: Electric ion thrusters are the preferred engines for deep space missions, because of very high specific impulse. The ion engine consists of screen and accelerator grids containing thousands of concentric very small holes. The xenon gas accelerates between the two grids, thus developing the impulse force. The dominant life-limiting mechanism in the state-of-the-art molybdenum thrusters is the xenon ion sputter erosion of the accelerator grid. Carbon/carbon composites (CCC) have shown to be have less than 1/7 the erosion rates than the molybdenum, thus for interplanetary missions CCC engines are inevitable. Early effort to develop CCC composite thrusters had a limited success because of limitations of the drilling technology and the damage caused by drilling. The proposed is an in-situ manufacturing of holes while the CCC is made. Special low CTE molds will be used along with the NC A&T s patented resin transfer molding (RTM) technology to manufacture the CCC grids. First, a manufacture process for 10-cm diameter thruster grids will be developed and verified. Quality of holes, density, CTE, tension, flexure, transverse fatigue and sputter yield properties will be measured. After establishing the acceptable quality and properties, the process will be scaled to manufacture 30-cm diameter grids. The properties of the two grid sizes are compared with each other.
    Keywords: Spacecraft Design, Testing and Performance
    Type: HBCUs/OMUs Research Conference Agenda and Abstracts; 20; NASA/TM-2003-212207
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  • 7
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    Concept Design of High Power Solar Electric Propulsion Vehicles for Human Exploration (2011)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Human exploration beyond low Earth orbit will require enabling capabilities that are efficient, affordable and reliable. Solar electric propulsion (SEP) has been proposed by NASA s Human Exploration Framework Team as one option to achieve human exploration missions beyond Earth orbit because of its favorable mass efficiency compared to traditional chemical propulsion systems. This paper describes the unique challenges associated with developing a large-scale high-power (300-kWe class) SEP vehicle and design concepts that have potential to meet those challenges. An assessment of factors at the subsystem level that must be considered in developing an SEP vehicle for future exploration missions is presented. Overall concepts, design tradeoffs and pathways to achieve development readiness are discussed.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NASA/TM-2011-217281 , IAC-11-D2.3.5 , E-18036 , 62nd International Astronautical Congress; Oct 03, 2011 - Oct 07, 2011; Cape Town; South Africa
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  • 8
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    Vibration Modal Characterization of a Stirling Convertor via Base-Shake Excitation (2003)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The U.S. Department of Energy (DOE), Lockheed Martin (LM), Stirling Technology Company (STC), and NASA John H. Glenn Research Center (GRC) are currently developing a high-efficiency Stirling convertor for use in a Stirling Radioisotope Generator (SRG). NASA and DOE have identified the SRG for potential use as an advanced power system for future NASA Space Science missions, providing spacecraft onboard electric power for deep space missions and power for unmanned Mars rovers. Low-level, baseshake sine vibration tests were conducted on the Stirling Technology Demonstration Convertor (TDC), at NASA GRC's Structural Dynamics Laboratory, in February 2001, as part of the development of this Stirling technology. The purpose of these tests was to provide a better understanding of the TDC's internal dynamic response to external vibratory base excitations. The knowledge obtained can therein be used to help explain the success that the TDC enjoyed in its previous random vibration qualification tests (December 1999). This explanation focuses on the TDC s internal dynamic characteristics in the 50 to 250 Hz frequency range, which corresponds to the maximum input levels of its qualification random vibration test specification. The internal dynamic structural characteristics of the TDC have now been measured in two separate tests under different motoring and dynamic loading conditions: (1) with the convertor being electrically motored, under a vibratory base-shake excitation load, and (2) with the convertor turned off, and its alternator internals undergoing dynamic excitation via hammer impact loading. This paper addresses the test setup, procedure and results of the base-shake vibration testing conducted on the motored TDC, and will compare these results with those results obtained from the dynamic impact tests (May 2001) on the nonmotored TDC.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NASA/TM-2003-212479 , E-14017 , AIAA Paper 2003-6096 , First International Energy Conversion Engineering Conference; Aug 17, 2003 - Aug 21, 2003; Portsmouth, VA; United States
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  • 9
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    Lessons Learned From CM-2 Modal Testing and Analysis (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Combustion Module-2 (CM-2) is a space experiment that launches on Shuttle mission STS-107 in the SPACEHAB Double Research Module. The CM-2 flight hardware is installed into SPACEHAB single and double racks. The CM-2 flight hardware was vibration tested in the launch configuration to characterize the structure's modal response. Cross-orthogonality between test and analysis mode shapes were used to assess model correlation. Lessons learned for pre-test planning and model verification are discussed.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NASA/TM-2002-211692 , NAS 1.15:211692 , E-13422 , Ninth International Congress on Sound and Vibration; Jul 08, 2002 - Jul 11, 2002; Orlando, FL; United States
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  • 10
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    Off-Nominal Performance of the International Space Station Solar Array Wings Under Orbital Eclipse Lighting Scenarios (2005)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper documents testing and analyses to quantify International Space Station (ISS) Solar Array Wing (SAW) string electrical performance under highly off-nominal, low-temperature-low-intensity (LILT) operating conditions with nonsolar light sources. This work is relevant for assessing feasibility and risks associated with a Sequential Shunt Unit (SSU) remove and replace (R&R) Extravehicular Activity (EVA). During eclipse, SAW strings can be energized by moonlight, EVA suit helmet lights or video camera lights. To quantify SAW performance under these off-nominal conditions, solar cell performance testing was performed using full moon, solar simulator and Video Camera Luminaire (VCL) light sources. Test conditions included 25 to 110 C temperatures and 1- to 0.0001-Sun illumination intensities. Electrical performance data and calculated eclipse lighting intensities were combined to predict SAW current-voltage output for comparison with electrical hazard thresholds. Worst case predictions show there is no connector pin molten metal hazard but crew shock hazard limits are exceeded due to VCL illumination. Assessment uncertainties and limitations are discussed along with operational solutions to mitigate SAW electrical hazards from VCL illumination. Results from a preliminary assessment of SAW arcing are also discussed. The authors recommend further analyses once SSU, R&R, and EVA procedures are better defined.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NASA/TM-2005-213988 , AIAA Paper 2005-5671 , E-15311 , Third International Energy Conversion Engineering Conference American Institute of Aeronautics and Astronautics, Inc.; Aug 15, 2005 - Aug 18, 2005; San Francisco, CA.; United States
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