SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 43, ISSUE 20 - OCTOBER 07, 2006
18 SPACECRAFT DESIGN, TESTING AND PERFORMANCE
Includes satellites; space platforms; space stations; spacecraft systems and components such as thermal and environmental controls; and spacecraft control and stability characteristics.
For life support systems see 54 Man/System Technology and Life Support.
For related information see also 05 Aircraft Design, Testing and Performance; 39 Structural Mechanics; and 16 Space Transportation and Safety.
20050215027 NASA Glenn Research Center, Cleveland, OH, USA
Multifunctional Inflatable Structure Being Developed for the PowerSphere Concept
Peterson, Todd T.; Research and Technology 2002; March 2003; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
The continuing development of microsatellites and nanosatellites for low Earth orbits requires the collection of sufficient power for instruments onboard a low-weight, low-volume spacecraft. Because the overall surface area of a microsatellite or nanosatellite is small, body-mounted solar cells cannot provide enough power. The deployment of traditional, rigid, solar arrays necessitates larger satellite volumes and weights, and also requires extra apparatus for pointing. One solution to this power choke problem is the deployment of a spherical, inflatable power system. This power system, termed the ‘PowerSphere,' has several advantages, including a high collection area, low weight and stowage volume, and the elimination of solar array pointing mechanisms. Derived from text
Inflatable Structures; Microsatellites; Nanosatellites; Weight (Mass)
20050215040 NASA Langley Research Center, Hampton, VA, USA
Flight Data Analysis of HyShot 2
Hass, Neal E.; Smart, Michael K.; Paull, Alan; [2005]; 17 pp.; In English; 13th AIAA/CIRA International Space Planes and Hypersonic Systems Technologies Conference, 16-20 May 2005, Capua, Italy Contract(s)/Grant(s): 23-745-30-30 Report No.(s): AIAA Paper 2005-3354; Copyright; Avail.: CASI: A03, Hardcopy
The development of scramjet propulsion for alternative launch and payload delivery capabilities has comprised largely of ground experiments for the last 40 years. With the goal of validating the use of short duration ground test facilities, the University of Queensland, supported by a large international contingency, devised a ballistic re-entry vehicle experiment called HyShot to achieve supersonic combustion in flight above Mach 7.5. It consisted of a double wedge intake and two back-to-back constant area combustors; one supplied with hydrogen fuel at an equivalence ratio of 0.33 and the other un-fueled. Following a first launch failure on October 30th 2001, the University of Queensland conducted a successful second launch on July 30th, 2002. Post-flight data analysis of the second launch confirmed the presence of supersonic combustion during the approximately 3 second test window at altitudes between 35 and 29 km. Reasonable correlation between flight and some pre-flight shock tunnel tests was observed. Author
Supersonic Combustion Ramjet Engines; Launching; Payloads; Reentry Vehicles; Supersonic Combustion; Flight Tests
20050215061 NASA Goddard Space Flight Center, Greenbelt, MD, USA
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The Swift Mission and the REM Telescope
Gehrels, N.; Chincarini, G.; Giommi, P.; Mason, K. O.; Nousek, J. A.; Wells, A. A.; White, N. E.; Barthelemy, S. D.; Burrow, D. N.; Hurley, K. C., et al.; International Conference on Advances in Infrastructure for e-Business, e-Education, e-Science, e-Medicine and Mobile Technologies on the Internet (SSGRR 2003s); [2003]; 5 pp.; In English; See also 20050215042; Original contains color illustrations; Copyright; Avail.: CASI: A01, Hardcopy; Available from CASI on CD-ROM only as part of the entire parent document
Following a description of the science drive which originated the Swift Mission, this is US NASA midEX Mission with the collaboration of Italy and the UK, we will describe the status of the hardware and the observing strategy. The telemetry is carried out via the TDRSS satellite for those communications that need immediate response. The data transfer and the scheduled uploading of routine commands will be done through the ASI Malindi station in Kenia. Both in the US and in Europe a large effort will be done to follow the bursts with the maximum of efficiency and as soon as possible after the alert. We will describe how the ESO VLT telescopes are able to respond to the alert. To address the problematic of the dark bursts and to immediately follow up all of the bursts also in the Near Infrared we designed and built a 60 cm NIR Robotic telescope, REM, to be located on the ESO ground at Cerro La Silla. The instrumentation includes also a low dispersion spectrograph with the capability of multi wavelength optical photometry. Author
X Ray Telescopes; Space Missions; Spacecraft Design; TDR Satellites; Robotics
20050215093 NASA Langley Research Center, Hampton, VA, USA
A Design Comparison of Atmospheric Flight Vehicles for the Exploration of Titan
Gasbarre, Joseph F.; Wright, Henry S.; Lewis, Mark J.; [2005]; 18 pp.; In English; AIAA Atmospheric Flight Mechanics Conference and Exhibit, 15-18 Aug. 2005, San Francisco, CA, USA Contract(s)/Grant(s): 23-752-50-20 Report No.(s): AIAA Paper 2005-6235; Copyright; Avail.: CASI: A03, Hardcopy
Titan, the largest moon of Saturn, is one of the most scientifically interesting locations in the Solar System. With a very cold atmosphere that is five times as dense as Earth s, and one and a half times the surface pressure, it also provides one of the most aeronautically fascinating environments known to humankind. While this may seem the ideal place to attempt atmospheric flight, many challenges await any vehicle attempting to navigate through it. In addition to these physical challenges, any scientific exploration mission to Titan will most likely have several operational constraints. One difficult constraint is the desire for a global survey of the planet and thus, a long duration flight within the atmosphere. Since many of the scientific measurements that would be unique to a vehicle flying through the atmosphere (as opposed to an orbiting spacecraft) desire near-surface positioning of their associated instruments, the vehicle must also be able to fly within the first scale height of the atmosphere. Another difficult constraint is that interaction with the surface, whether by landing or dropped probe, is also highly desirable from a scientific perspective. Two common atmospheric flight platforms that might be used for this mission are the airplane and airship. Under the assumption of a mission architecture that would involve an orbiting relay spacecraft delivered via aerocapture and an atmospheric flight vehicle delivered via direct entry, designs were developed for both platforms that are unique to Titan. Consequently, after a viable design was achieved for each platform, their advantages and disadvantages were compared. This comparison included such factors as deployment risk, surface interaction capability, mass, and design heritage. When considering all factors, the preferred candidate platform for a global survey of Titan is an airship. Author
Titan; Aerospace Vehicles; Deployment; Surface Reactions
20050215102 NASA Marshall Space Flight Center, Huntsville, AL, USA
MIT-NASA Workshop: Transformational Technologies
Mankins, J. C., Editor; Christensen, C. B.; Gresham, E. C.; Simmons, A.; Mullins, C. A.; March 2005; 48 pp.; In English; MIT-NASA Workshop: Transformational Technologies, 11-12 Dec. 2003, Cambridge, MA, USA Report No.(s): NASA/CP-2005-213741; M-1134; No Copyright; Avail.: CASI: A03, Hardcopy
As a space faring nation, we are at a critical juncture in the evolution of space exploration. NASA has announced its Vision for Space Exploration, a vision of returning humans to the Moon, sending robots and eventually humans to Mars, and exploring the outer solar system via automated spacecraft. However, mission concepts have become increasingly complex, with the potential to yield a wealth of scientific knowledge. Meanwhile, there are significant resource challenges to be met. Launch costs remain a barrier to routine space flight; the ever-changing fiscal and political environments can wreak havoc on mission planning; and technologies are constantly improving, and systems that were state of the art when a program began can quickly become outmoded before a mission is even launched. This Conference Publication describes the workshop and featured presentations by world-class experts presenting leading-edge technologies and applications in the areas of power and propulsion; communications; automation, robotics, computing, and intelligent systems; and transformational techniques for space activities. Workshops such as this one provide an excellent medium for capturing the broadest possible array of insights and expertise, learning from researchers in universities, national laboratories, NASA field Centers, and industry to help better our future in space. Author
Robotics; Propulsion; Automatic Control; Computers; Space Exploration; Spacecraft Design
20050215115 NASA Langley Research Center, Hampton, VA, USA
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Radiation Protection for Lunar Mission Scenarios
Clowdsley, Martha S.; Nealy, John E.; Wilson, JohnW.; Anderson, Brooke M.; Anderson, Mark S.; Krizan, Shawn A.; [2005]; 11 pp.; In English; AIAA SPACE 2005 Conference and Exposition, 30 Aug. - 1 Sep. 2005, Long Beach, CA, USA Contract(s)/Grant(s): 23-612-10-81-A1; Copyright; Avail.: CASI: A03, Hardcopy
Preliminary analyses of shielding requirements to protect astronauts from the harmful effects of radiation on both short-term and long-term lunar missions have been performed. Shielding needs for both solar particle events (SPEs) and galactic cosmic ray (GCR) exposure are discussed for transit vehicles and surface habitats. This work was performed under the aegis of two NASA initiatives. The first study was an architecture trade study led by Langley Research Center (LaRC) in which a broad range of vehicle types and mission scenarios were compared. The radiation analysis for this study primarily focused on the additional shielding mass required to protect astronauts from the rare occurrence of a large SPE. The second study, led by Johnson Space Center (JSC), involved the design of lunar habitats. Researchers at LaRC were asked to evaluate the changes to mission architecture that would be needed if the surface stay were lengthened from a shorter mission duration of 30 to 90 days to a longer stay of 500 days. Here, the primary radiation concern was GCR exposure. The methods used for these studies as well as the resulting shielding recommendations are discussed. Recommendations are also made for more detailed analyses to minimize shielding mass, once preliminary vehicle and habitat designs have been completed. Here, methodologies are mapped out and available radiation analysis tools are described. Since, as yet, no dosimetric limits have been adopted for missions beyond low earth orbit (LEO), radiation exposures are compared to LEO limits. Uncertainties associated with the LEO career effective dose limits and the effects of lowering these limits on shielding mass are also discussed. Author
Radiation Protection; Astronauts; Exposure; Radiation Dosage; Radiation Effects; Shielding
20050215301 Christian Brothers Univ., Memphis, TN, USA
Integration Of Launch Vehicle Simulation/Analysis Tools And Lunar Cargo Lander Design, Part 2/2
DeJean, George Brian; Shiue, Yeu-Sheng Paul; King, Jeffrey; The 2004 NASA Faculty Fellowship Program Research Reports; January 2005, pp. X-1 - X-8; In English; See also 20050215300; Original contains color illustrations; No Copyright; Avail.: CASI: A02, Hardcopy
Part 2, which will be discussed in this report, will discuss the development of a Lunar Cargo Lander (unmanned launch vehicle) that will transport usable payload from Trans- Lunar Injection to the moon. The Delta IV-Heavy was originally used to transport the Lunar Cargo Lander to TLI, but other launch vehicles have been studied. In order to uncover how much payload is possible to land on the moon, research was needed in order to design the sub-systems of the spacecraft. The report will discuss and compare the use of a hypergolic and cryogenic system for its main propulsion system. The guidance, navigation, control, telecommunications, thermal, propulsion, structure, mechanisms, landing gear, command, data handling, and electrical power sub-systems were designed by scaling off other flown orbiters and moon landers. Once all data was collected, an excel spreadsheet was created to accurately calculate the usable payload that will land on the moon along with detailed mass and volume estimating relations. As designed, The Lunar Cargo Lander can plant 5,400 lbm of usable payload on the moon using a hypergolic system and 7,400 lbm of usable payload on the moon using a cryogenic system. Author
Computerized SIMulation; Control Systems Design; Launch Vehicles; Navigation; Payloads; Telecommunication
20050215310 Middle Tennessee State Univ., Murfreesboro, TN, USA
Reliability Assessment Of Conceptual Launch Vehicles
Bloomer, Lisa A.; The 2004 NASA Faculty Fellowship Program Research Reports; January 2005, pp. V-1 - V-5; In English; See also 20050215300; No Copyright; Avail.: CASI: A01, Hardcopy
Planning is underway for new NASA missions to the moon and to MARS. These missions carry a great deal of risk, as the Challenger and Columbia accidents demonstrate. In order to minimize the risks to the crew and the mission, risk reduction must be done at every stage, not only in quality manufacturing, but also in design. It is necessary, therefore, to be able to compare the risks posed in different launch vehicle designs. Further, these designs have not yet been implemented, so it is necessary to compare these risks without being able to test the vehicles themselves. This paper will discuss some of the issues involved in this type of comparison. It will start with a general discussion of reliability estimation. It will continue with a short look at some software designed to make this estimation easier and faster. It will conclude with a few recommendations for future tools. Derived from text
Manufacturing; NASA Programs; Reliability; Test Vehicles
20050215317 Tennessee State Univ., Nashville, TN, USA
Study of a 30-M Boom For Solar Sail-Craft: Model Extendibility and Control Strategy
Keel, Leehyun; The 2004 NASA Faculty Fellowship Program Research Reports; January 2005, pp. XXIII-1 - XXIII-5; In English; See also 20050215300; Original contains color illustrations; No Copyright; Avail.: CASI: A01, Hardcopy
Space travel propelled by solar sails is motivated by the fact that the momentum exchange that occurs when photons are reflected and/or absorbed by a large solar sail generates a small but constant acceleration. This acceleration can induce a constant thrust in very large sails that is sufficient to maintain a polar observing satellite in a constant position relative to the Sun or Earth. For long distance propulsion, square sails (with side length greater than 150 meters) can reach Jupiter in two years and Pluto in less than ten years.
Converting such design concepts to real-world systems will require accurate analytical models and model parameters. This requires extensive structural dynamics tests. However, the low mass and high flexibility of large and light weight structures such as solar sails makes them unsuitable for ground testing. As a result, validating analytical models is an extremely difficult problem.
On the other hand, a fundamental question can be asked. That is whether an analytical model that represents a small-scale version of a solar-sail boom can be extended to much larger versions of the same boom. To answer this question, we considered a long deployable boom that will be used to support the solar sails of the sail-craft.
The length of fully deployed booms of the actual solar sail-craft will exceed 100 meters. However, the test-bed we used in our study is a 30 meter retractable boom at MSFC. We first develop analytical models based on Lagrange s equations and the standard Euler-Bernoulli beam. Then the response of the models will be compared with test data of the 30 meter boom at various deployed lengths. For this stage of study, our analysis was limited to experimental data obtained at 12ft and 18ft deployment lengths. The comparison results are positive but speculative. To observe properly validate the analytic model, experiments at longer deployment lengths, up to the full 30 meter, have been requested.
We expect the study to answer the extendibility question of the analytical models. In operation, rapid temperature changes can be induced in solar sails as they transition from day to night and vice versa. This generates time dependent thermally induced forces, which may in turn create oscillation in structural members such as booms. Such oscillations have an adverse effect on system operations, precise pointing of instruments and antennas and can lead to self excited vibrations of increasing amplitude. The latter phenomenon is known as thermal flutter and can lead to the catastrophic failure of structural systems.
To remedy this problem, an active vibration suppression system has been developed. It was shown that piezoelectric actuators used in conjunction with a Proportional Feedback Control (PFC) law (or Velocity Feedback Control (VFC) law) can induce moments that can suppress structural vibrations and prevent flutter instability in spacecraft booms. In this study, we will investigate control strategies using piezoelectric transducers in active, passive, and/or hybrid control configurations. Advantages and disadvantages of each configuration will be studied and experiments to determine their capabilities and limitations will be planned. In particular, special attention will be given to the hybrid control, also known as energy recycling, configuration due to its unique characteristics. Derived from text
Feedback Control; Mathematical Models; Solar Sails; Booms (Equipment); Retractable Equipment
20050215330 University of Northern Arizona, Flagstaff, AZ, USA
Rack Distribution Effects on MPLM Center of Mass
Tester, John T.; The 2004 NASA Faculty Fellowship Program Research Reports; January 2005, pp. XLIV-1 - XLIV-5; In English; See also 20050215300; Original contains color and black and white illustrations; No Copyright; Avail.: CASI: A01, Hardcopy
This research was in support of exploring the need for more flexible ‘center of gravity (CG) specifications than those currently established by NASA for the Multi-Purpose Logistics Module (MPLM). The MPLM is the cargo carrier for International Space Station (ISS) missions. The MPLM provides locations for 16 standard racks, as shown in Figure 1; not all positions need to be filled in any given flight. The MPLM coordinate system (X(sub M), Y(sub M), Z(sub M)) is illustrated as well. For this project, the primary missions of interest were those which supply the ISS and remove excess materials on the return flights. These flights use a predominate number of ‘Resupply Stowage Racks' (RSR) and ‘Resupply Stowage Platforms' (RSP). In these two types of racks, various smaller items are stowed. Hence, these racks will exhibit a considerable range of mass values as well as a range as to where their individual CG are located. Derived from text
Center of Mass; Multi-Purpose Logistics Modules; Gravitation
20050215337 Clark-Atlanta Univ., GA, USA
Development of a Modeling Framework to Support Control Investigations of Sailcraft Missions A First Cut: ABLE Sailcraft Dynamics Model
Sarathy, Sriprakash; The 2004 NASA Faculty Fellowship Program Research Reports; January 2005, pp. XXXVII-1 -XXXVII-6; In English; See also 20050215300; No Copyright; Avail.: CASI: A02, Hardcopy
Solar Sailcraft, the stuff of dreams of the H.G. Wells generation, is now a rapidly maturing reality. The promise of unlimited propulsive power by harnessing stellar radiation is close to realization. Currently, efforts are underway to build, prototype and test two configurations. These sails are designed to meet a 20m sail requirement, under guidance of the In-Space Propulsion (ISP) technology program office at MSFC. While these sails will not fly , they are the first steps in improving our understanding of the processes and phenomena at work. As part of the New Millennium Program (NMP) the ST9 technology validation mission hopes to launch and fly a solar sail by 2010 or sooner. Though the Solar Sail community has been studying and validating various concepts over two decades, it was not until recent breakthroughs in structural and material technology, has made possible to build sails that could be launched. With real sails that can be tested (albeit under earth conditions), the real task of engineering a viable spacecraft has finally commenced. Since it is not possible to accurately or practically recreate the actual operating conditions of the sailcraft (zero-G, vacuum and extremely low temperatures), much of the work has focused on developing accurate models that can be used to predict behavior in space, and for sails that are 6-10 times the size of currently existing sails. Since these models can be validated only with real test data under ‘earth' conditions, the process of modeling and the identification of uncertainty due to model assumptions and scope need to be closely considered. Sailcraft models that exist currently, are primarily focused on detailed physical representations at the component level, these are intended to support prototyping efforts. System level models that cut across different sail configurations and control concepts while maintaining a consistent approach are non-existent. Much effort has been focused on the areas of thrust performance, solar radiation prediction, and sail membrane behavior vis-a-vis their reflective geometry, such as wrinkling/folding/furling as it pertains to thrust prediction. A parallel effort has been conducted on developing usable models for developing attitude control systems (ACS), for different sail configurations in different regimes. There has been very little by way of a system wide exploration of the impact of the various control schemes, thrust prediction models for different sail configurations being considered. Author
Solar Sails; Stellar Radiation; Attitude Control; Performance Prediction; Propulsion; Thrust
20050215340 Mississippi State Univ., MS, USA
On Structural Design of a Mobile Lunar Habitat with Multi-Layered Environmental Shielding
Rais-Rohani, Masoud; The 2004 NASA Faculty Fellowship Program Research Reports; January 2005, pp. XXXIV-1 -XXXIV-5; In English; See also 20050215300; Original contains color and black and white illustrations; No Copyright; Avail.: CASI: A01, Hardcopy
The future human lunar missions are expected to undertake far more ambitious activities than those of the Apollo program with the possibility of some missions lasting up to several months. Such extended missions require the use of large-size lunar outposts to accommodate living quarters for the astronauts as well as indoor laboratory facilities. The greatest obstacle to the prolonged human presence on the Moon is the threat posed by the harsh lunar environment that is plagued with multi-source high-energy radiation exposure as well as frequent barrage of meteoroids. Hence, for such extended missions to succeed, it is vital that the future lunar outposts be designed to provide a safe habitat for the astronauts. Over the past few years, a variety of ideas and concepts for future lunar outposts and bases have been proposed. With shielding as the primary concern, some have suggested the use of natural structures such as lava tubes while others have taken a more industrial approach and suggested the construction of fixed structures in the form of inflatable, inflatable with rigid elements, and tent-style membrane. For evaluation of these structural design concepts, Drake and Richter1 have proposed a rating system based on such factors as effectiveness, importance, and timing. While all of these designs, in general, benefit from in-situ resource utilization (i.e., lunar regolith) for shielding, they share a common disadvantage of being fixed to one particular location that would limit exploration to the region in close proximity of the outpost. Derived from text
Structural Design; Space Capsules; Lunar Exploration; Radiation Dosage
20050215414 NASA Glenn Research Center, Cleveland, OH, USA
Multifunctional Inflatable Structure Being Developed for the PowerSphere Concept
Peterson, Todd T.; Research and Technology 2003; May 2004; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
NASA has funded a collaborative team of The Aerospace Corporation, ILC Dover, Lockheed Martin, and NASA Glenn Research Center to develop the Multifunctional Inflatable Structure (MIS) for a ‘PowerSphere' concept through a NASA Research Announcement. This power system concept has several advantages, including a high collection area, low weight and stowage volume, and the elimination of all solar array pointing mechanisms. The current 3-year effort will culminate with the fabrication and testing of a fully functional engineering development unit. The baseline design of the Power-Sphere consists of two opposing semispherical domes connected to a central spacecraft. Each semispherical dome consists of hexagonal and pentagonal solar cell panels that together form a geodetic sphere. Inflatable ultraviolet (UV) rigidizable tubular hinges between the solar cell panels and UV rigidizable isogrid center columns with imbedded flex circuitry form the MIS. The reference configuration for the PowerSphere is a 0.6-m-diameter (fully deployed) spacecraft with a total mass budget of 4 kg (1 kg for PowerSphere, 3 kg for spacecraft) capable of producing 29 W of electricity with 10-percent-efficient thin-film solar cells. In a stowed configuration, the solar cell panels will be folded sequentially to the outside of the instrument decks. The center column will be z-folded between the instrument decks and the spacecraft housing for packaging. The instrument panel will secure the z-folded stack with launch ties. After launch, once the release tie is triggered, the center column and hinge tubes will inflate and be rigidized in their final configurations by ultraviolet radiation. The overall PowerSphere deployment sequence is shown pictorially in the following illustration. Author
Inflatable Structures; Ultraviolet Radiation; Solar Arrays; Product Development; Fabrication
20050215461 Missouri Univ., Rolla, MO, USA
Overview of Microwave and MillimeterWave Testing Activities for the Inspection of the Space Shuttle SOH and Heat Tiles
Zoughi, R.; [2005]; 1 pp.; In English; 32nd Annual Review of Progress in Quantitative Nondestructive Evaluation (WNDE) Conference, 31 Jul. - 5 Aug. 2005, Brunswick, ME, USA Contract(s)/Grant(s): NNM04AA15A; No Copyright; Avail.: Other Sources; Abstract Only
Microwave and millimeter wave nondestructive testing and evaluation methods, have shown great potential for inspecting the Space Shuttle s external tank spray on foam insulation (SOFI) and acreage heat tiles. These methods are capable of producing high-resolution images of et interior of these structures. To this end, several different microwave and millimeter wave nondestructive testing methods have been investigated for this purpose. These methods have included near-field as well as focused approaches ranging in frequency from 10 GHz to beyond 100 GHz. Additionally, synthetic aperture focusing methods have also been developed in this regime for obtaining high-resolution images of the interior of these critical structures. These methods possess the potential for producing 3D images of these structures in a relatively short amount of time. This paper presents a summary of these activities in addition to providing examples of images produced using these diverse methods. Author
Millimeter Waves; Microwaves; Nondestructive Tests; External Tanks; Insulation; Foams
20050215563 NASA Marshall Space Flight Center, Huntsville, AL, USA
Low Earth Orbit Environmental Effects on Space Tether Materials
Finckernor, Miria M.; Gitlemeier, Keith A.; Hawk, Clark W.; Watts, Ed; [2005]; 23 pp.; In English Contract(s)/Grant(s): 159-89-10; Copyright; Avail.: CASI: A03, Hardcopy
Atomic oxygen (AO) and ultraviolet (UV) radiation erode and embrittle most polymeric materials. This research was designed to test several different materials and coatings under consideration for their application to space tethers, for resistance to these effects. The samples were vacuum dehydrated, weighed and then exposed to various levels of AO or UV radiation at the NASA Marshall Space Flight Center. They were then re-weighed to determine mass loss due to atomic oxygen erosion, inspected for damage and tensile tested to determine strength loss. The experiments determined that the Photosil coating process, while affording some protection, damaged the tether materials worse than the AO exposure. TOR-LM also failed to fully protect the materials, especially from UV radiation. The POSS and nickel coatings did provide some protection to the tethers, which survived the entire test regime. M5 was tested, uncoated, and survived AO exposure, though its brittleness prevented any tensile testing. Author
Ultraviolet Radiation; Environment Effects; Low Earth Orbits; Spacecraft Construction Materials; Nickel Coatings; Oxygen Atoms
20050215570 Georgia Inst. of Tech., Atlanta, GA, USA
Command Generation and Control of Momentum Exchange Electrodynamic Reboost Tethered Satellite
Robertson, Michael J.; 2005; 17 pp.; In English Contract(s)/Grant(s): NGT8-52943; No Copyright; Avail.: CASI: A03, Hardcopy
The research completed for this NASA Graduate Student Research Program Fellowship sought to enhance the current state-of-the-art dynamic models and control laws for Momentum Exchange Electrodynamic Reboost satellite systems by utilizing command generation, specifically Input Shaping. The precise control of tethered spacecraft with flexible appendages is extremely difficult. The complexity is magnified many times when the satellite must interact with other satellites as in a momentum exchange via a tether. The Momentum Exchange Electronic Reboost Tether (MXER) concept encapsulates all of these challenging tasks [l]. Input Shaping is a command generation technique that allows flexible spacecraft to move without inducing residual vibration [2], limit transient deflection [3] and utilize fuel-efficient actuation [4]. Input shaping is implemented by convolving a sequence of impulses, known as the input shaper, with a desired system command to produce a shaped input that is then used to drive the system. This process is demonstrated in Figure 1. The shaped command is then use to drive the system without residual vibration while meeting many other performance specifications. The completed work developed tether control algorithms for retrieval. A simple model of the tether response has been developed and command shaping was implemented to minimize unwanted dynamics. A model of a flexible electrodynamic tether has been developed to investigate the tether s response during reboost. Command shaping techniques have been developed to eliminate the tether oscillations and reduce the tether s deflection to pre-specified levels during reboost. Additionally, a model for the spin-up of a tethered system was developed. This model was used in determining the parameters for optimization the resulting angular velocity. Author
Command and Control; Control Theory; Tethered Satellites; Spacecraft Control; Functional Design Specifications; Flexible Spacecraft
20050215584 NASA Kennedy Space Center, Cocoa Beach, FL, USA
STS-114: Discovery Day 5 Post MMT Briefing July 30, 2005
3 pp.; In English; 38 min. playing time, in color, with sound; No Copyright; Avail.: CASI: V03, Videotape-VHS: B03, Videotape-Beta
Wayne Hill, Space Shuttle Deputy Program Manager, and Chair of the Mission Management Team, reports the following: the testing of the thermal protection system in space, on orbit was successful; the MMT meeting formally approved the mission extension for one day; image analysis from the launch phase has been completed; tile and blankets had been formally cleared by the Engineering team; additional inspection of the reinforced carbon-carbon (RCC); gap filler, heat shield and small black spot phenomenon are carefully inspected and evaluated to prepare the safe return of the Discovery. Bill Gerstenmaier, ISS Program Manager reports on the extra vehicular activities (EVA): replaced the GPS antenna; prepared the airlock to attach the ESP2; re-powered CMG2; transfer of CWCs; and the MISSE packages were wrapped around thru at the end of the air lock. He mentioned the high currents problem seen on CMG3, which they will have to take off line to check and understand the problem. Consumables, engineering performance of the three sensors (LCS, LDRI, ITVC), CMG removal and replacement, EVA2 and EVA3, gap fillers, RCC, hydrogen tank pre-press cycles, thermal protection system, inspection, ISS supply maintenance, and projection of next flight are topics covered with the News Media. CASI
Space Transportation System; Space Shuttles; Launching
20050215604
NASA Marshall Space Flight Center, Huntsville, AL, USA
Solar Sail Roadmap Mission GN and C Challenges
Heaton, Andrew F.; [2005]; 10 pp.; In English; AIAA GN&C Conference, 15-19 Aug. 2005, San Francisco, CA, USA; Original contains black and white illustrations; No Copyright; Avail.: CASI: A02, Hardcopy
The NASA In-Space Propulsion program is funding development work for solar sails to enhance future scientific opportunities. Key to this effort are scientific solar sail roadmap missions identified by peer review. The two near-term missions of interest are L1 Diamond and Solar Polar Imager. Additionally, the New Millennium Program is sponsoring the Space Technology 9 (ST9) demonstration mission. Solar sails are one of five technologies competing for the ST9 flight demonstration. Two candidate solar sail missions have been identified for a potential ST9 flight. All the roadmap missions and candidate flight demonstration missions face various GN&C challenges. A variety of efforts are underway to address these challenges. These include control actuator design and testing, low thrust optimization studies, attitude control system design and modeling, control-structure interaction studies, trajectory control design, and solar radiation pressure model development. Here we survey the various efforts underway and identify a few of specific recent interest and focus. Author
Aerospace Engineering; Solar Sails; Trajectory Control; Design Optimization; Flight Tests; Actuators; Attitude Control
20050215611 NASA Marshall Space Flight Center, Huntsville, AL, USA
Applications of the Electrodynamic Tether to Interstellar Travel
Matloff, Gregory L.; Johnson, Les; June 21, 2005; 11 pp.; In English Contract(s)/Grant(s): SAIC-440055739; Copyright; Avail.: CASI: A03, Hardcopy
After considering relevant properties of the local interstellar medium and defining a sample interstellar mission, this paper considers possible interstellar applications of the electrodynamic tether, or EDT. These include use of the EDT to provide on-board power and affect trajectory modifications and direct application of the EDT to starship acceleration. It is demonstrated that comparatively modest EDTs can provide substantial quantities of on-board power, if combined with a large-area electron-collection device such as the Cassenti toroidal-field ramscoop. More substantial tethers can be used to accomplish large-radius thrustless turns. Direct application of the EDT to starship acceleration is apparently infeasible. Author
Interstellar Travel; Tethering; Electrodynamics
Source: NASA.
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