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high earth orbit design for lunar-assisted medium class explorer missionsThis study investigates the application of high-Earth orbit (HEO) trajectories to missions requiring long on-target integration times, avoidance of the Earth's radiation belt, and minimal effects of Earth and Lunar shadow periods which could cause thermal/mechanical stresses on the science instruments. As used here, a HEO trajectory is a particular solution to the restricted three-body problem in the Earth-Moon system with the orbit period being either 1/2 of, or 1/4 of, the lunar sidereal period. A primary mission design goal is to find HEO trajectories where, for a five-year mission duration, the minimum perigee radius is greater than seven Earth radii (R(sub E)). This minimum perigee radius is chosen so that, for the duration of the mission, the perigee is always above the relatively heavily populated geosynchronous radius of 6.6 R(sub E). A secondary goal is to maintain as high an ecliptic inclination as possible for the duration of the mission to keep the apsis points well out of the Ecliptic plane. Mission design analysis was completed for launch dates in the month of June 2003, using both direct transfer and phasing loop transfer techniques, to a lunar swingby for final insertion into a HEO. Also provided are analysis results of eclipse patterns for the trajectories studied, as well as the effects of launch vehicle errors and launch delays.
Document ID
20010084964
Document Type
Conference Paper
Authors
McGiffin, Daniel A.
(Computer Sciences Corp. Lanham, MD United States)
Mathews, Michael
(Computer Sciences Corp. Lanham, MD United States)
Cooley, Steven
(NASA Goddard Space Flight Center Greenbelt, MD United States)
Date Acquired
August 20, 2013
Publication Date
June 1, 2001
Publication Information
Publication: 2001 Flight Mechanics Symposium
Subject Category
Astrodynamics
Distribution Limits
Public
Copyright
Work of the US Gov. Public Use Permitted.

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IDRelationTitle20010084958Analytic Primary2001 Flight Mechanics Symposium