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Analytical solution to three-body problems is applied to the problem of predicting the orbit of a lunar satellite and determining the orbit of a near-earth satellite. Using this solution, the state vector may be generated at any time without intermediate numerical extrapolation.
First order short and long perturbations of lunar satellite orbit by mascon
A method was examined for joint construction of a selenocentric fundamental system which can be realized by a coordinate catalog of reference contour points uniformly positioned over the entire lunar surface, and determination of the parameters characterizing the gravitational field, rotation, and orbital motion of the moon. Characteristic of the problem formulation is the introduction of a new complex of inconometric measurements which can be made using pictures obtained from an artificial lunar satellite. The proposed method can be used to solve similar problems on any other planet for which surface images can be obtained from a spacecraft. Characteristic of the proposed technique for solving the problem is the joint statistical analysis of all forms of measurements: orbital iconometric, earth-based trajectory, and also a priori information on the parameters in question which is known from earth-based astronomical studies.
Nearside tracking data from Apollo subsatellites and Lunar Orbiter 5 are used to construct a model of the farside lunar gravity field. An indirect method of calculating the average effect of all the gravity anomalies which affect spacecraft orbits is reduced to a linear estimation problem by eliminating the integration of spacecraft dynamics. Doppler frequency shifts between the lunar satellite and earth-based tracking station are used to measure the line of sight velocity component between the orbiter and the tracker; gravity determination is then performed in a weighted least squares estimator which employs the orbital rates as data and formulates the normal equations using the associated Keplerian state. Simulations performed demonstrate that farside gravity features can be economically determined by use of gravity estimates obtained from long-period models of nearside orbiting tracks.
An analysis is given of how satellite, lunar laser, and very long base interferometry stations available during the MERIT Campaign in 1983/84 can contribute to the detection of short periodic variations in the rotational parameters of the earth, as well as the determination of the differences between the various Conventional Terrestrial and Inertial Reference Frames inherent in the above systems. Specific observational requirements are given both by objective and by country.
Four areas of research in optical communications in support of future manned space missions being carried out at Johnson Space Center are discussed. These are the Space Station Freedom proximity operations, direct LEO-to-ground communications, IR voice communications inside manned spacecraft, and deep space and lunar satellite operations. The background, requirements, and scenario for each of these areas of research are briefly described.
We present SpeX 0.8-2.5 micron spectra and Br-gamma or J H or K-band images of the Moon on 2006 July 10 UT, 2006 Sep 1 UT, and 2006 Sep 3 UT. The first two dates are data taken in preparation (near Full Moon and 1 st quarter Moon) for observing the impact of ESA's SMART-1 Lunar satellite on 2006 Sep 3 UT. We hope to be presenting images (about 1 arc minute by 1 arc minute), and low resolution (Prism mode) spectra of the SMART-1 impact to be taken with the 60 arc second long slit.
As power levels of advanced spacecraft climb above 25 kW, higher solar array operating voltages become attractive. Even in today s satellites, operating spacecraft buses at 100 V and above has led to arcing in GEO communications satellites, so the issue of spacecraft charging and solar array arcing remains a design problem. In addition, micrometeoroid impacts on all of these arrays can also lead to arcing if the spacecraft is at an elevated potential. For example, tests on space station hardware disclosed arcing at 75V on anodized A1 structures that were struck with hypervelocity particles in Low Earth Orbit (LEO) plasmas. Thus an understanding of these effects is necessary to design reliable high voltage solar arrays of the future, especially in light of the Vision for Space Exploration of NASA. In the future, large GEO communication satellites, lunar bases, solar electric propulsion missions, high power communication systems around Mars can lead to power levels well above 100 kW. As noted above, it will be essential to increase operating voltages of the solar arrays well above 80 V to keep the mass of cabling needed to carry the high currents to an acceptable level. Thus, the purpose of this paper is to discuss various solar array approaches, to discuss the results of testing them at high voltages, in the presence of simulated space plasma and under hypervelocity impact. Three different types of arrays will be considered. One will be a planar array using thin film cells, the second will use planar single or multijunction cells and the last will use the Stretched Lens Array (SLA - 8-fold concentration). Each of these has different approaches for protection from the space environment. The thin film cell based arrays have minimal covering due to their inherent radiation tolerance, conventional GaAs and multijunction cells have the traditional cerium-doped microsheet glasses (of appropriate thickness) that are usually attached with Dow Corning DC 93-500 silicone adhesive. In practice, these cover glasses and adhesive do not cover the cell edges. Finally, in the SLA, the entire cell and cell edges are fully encapsulated by a cover glass that overhangs the cell perimeter and the silicone adhesive covers the cell edges providing a sealed environment. These three types of blanket technology have been tested at GRC and Auburn. The results of these tests will be described. For example, 15 modules composed of four state-of-the-art 2x4 cm GaAs solar cells with 150 pm cover glasses connected in two-cell series strings were tested at high voltage, in plasma under hypervelocity impact. A picture of one of the modules is shown in figure 1. These were prepared by standard industry practice from a major supplier and had efficiencies above 18%. The test results and other fabrication factors that influenced the tests will be presented. In addition, results for SLA segments tested under the same conditions will be presented. Testing of thin film blankets at GRC will also be presented. Figure 1 : Typical GaAs Solar Cell Module These results will show significant differences in resistance to arcing that are directly related to array design and manufacturing procedures. Finally, the approaches for mitigating the problems uncovered by these tests will be described. These will lay the foundation for future higher voltage array operation, even including voltages above 300-600 V for direct drive SEP applications.
The purpose of this staff study, made at the request of the chairman, is to serve members of the Committee on Aeronautical and Space Sciences as a source of basic information on Project Mercury, the man-in-space program of the National Aeronautics and Space Administration. The study is largely derived from unclassified information released by the National Aeronautics and Space Administration and testimony concerning Project Mercury given during hearings before this committee. The program descriptions are based upon current program planning. Since this is a highly advanced research and development program, the project is obviously subject to changes that may result from future developments and accomplishments characteristic of such research activities. Certain information with respect to revised schedules, obtained on a classified basis by the committee during inspection trips, is necessarily omitted. The appendixes to the study include information that may prove helpful on various aspects of space flight and exploration. Included are unofficial comments and observations relating to Russia's manned space flight activities and also a complete chronology of all satellites, lunar probes, and space probes up to the present.
Satellite orbit perturbation analysis for lunar gravitational field determination using von Zeipel transformation
Data relay and tracking capability of lunar communication satellites above lunar far side surface
The Lunar CRater Observation and Sensing Satellite (LCROSS) launched with the Lunar Reconnaissance Orbiter (LRO) on June 18, 2009. While the science function of the LCROSS mission was to determine the presence of water-ice in a permanently-shadowed crater on the moon, the operational purpose was to be a pioneer for future low-cost, risk-tolerant small satellite NASA missions. Recent strategic changes at the Agency level have only furthered the importance of small satellite missions. NASA Ames Research Center and its industry partner, Northrop-Grumman, initiated this spacecraft project two-years after its co-manifest mission had started, with less than one-fifth the budget. With a $79M total cost cap (including operations and reserves) and 31-months until launch, LCROSS needed a game-changing approach to be successful. At the LCROSS Confirmation Review, the ESMD Associate Administrator asked the Project team to keep a close record of lessons learned through the course of the mission and share their findings with the Agency at the end of the mission. This paper summarizes the Project, the mission, its risk position, and some of the more notable lessons learned.
The Lunar CRater Observation and Sensing Satellite (LCROSS) launched with the Lunar Reconnaissance Orbiter (LRO) on June 18, 2009. While the science purpose of the LCROSS mission was to determine the presence of water-ice in a permanently-shadowed crater on the moon, the functional purpose was to be a pioneer for future low-cost, risk-tolerant small satellite NASA missions. Recent strategic changes at the Agency level have only furthered the importance of small satellite missions. NASA Ames Research Center and its industry partner, Northrop-Grumman, initiated this spacecraft project two-years after its co-manifest mission had started, with less than one-fifth the budget. With a $79M total cost cap (including operations and reserves) and 31-months until launch, LCROSS needed a game-changing approach to be successful. At the LCROSS Confirmation Review, the ESMD Associate Administrator asked the Project team to keep a close record of lessons learned through the course of the mission and share their findings with the Agency at the end of the mission. This paper summarizes the Project, the mission, its risk position, and some of the more notable lessons learned.
Orbiting lunar relay satellite networks for manned landing missions on lunar far side
Data relay and tracking capability of lunar communication satellites
Two design options were explored to address the requirement to provide lunar piloted missions with continuous communications for outpost and sortie missions. Two unique orbits were assessed, along with the appropriate spacecraft (S/C) to address these requirements. Both constellations (with only two S/C each) provide full time coverage (24 hr/7 d) for a south polar base and also provide continuous 7 day coverage for sorties for specified sites and periodic windows. Thus a two-satellite system can provide full coverage for sorties for selected windows of opportunity without reconfiguring the constellation.
Application of double-averaged equations of artificial satellite motion to lunar orbiter
Generalized von Zeipel treatment of lunar and artificial satellite theories, generating single canonical transformation by variable separation technique