India’s Lunar Mission (Chandrayaan- 1) Orbit Determination System
- Paper number
IAC-06-C1.5.06
- Author
Dr. Vighnesam Narayanasetti Venkata, Indian Space Research Organisation (ISRO), India
- Coauthor
Mr. Anatta Sonney, Indian Space Research Organisation (ISRO), India
- Coauthor
Mr. Pramod Kumar Soni, Indian Space Research Organisation (ISRO), India
- Year
2006
- Abstract
India’s first mission to the moon named Chandrayaan –1 is proposed to be a lunar polar circular orbiter at an altitude of about 100km. PSLV-C4 has been identified as a possible candidate launch vehicle for Chandrayaan - 1. The launcher is capable of placing a spacecraft of about 1000kg mass in an elliptic parking orbit (240 X 36000)km with 18 deg inclination. The tracking of the Launch phase, GTO phase, Earth Transfer Orbit phase and LTT (Lunar Transfer Trajectory) phase up to a slant range of 1,00,000km will be carried out by the existing 10/11/12-m dish antennae ISRO’s network stations. The DSN (Deep Space Network) will take over the tracking requirements once the above limit of slant range is approached. It is proposed to provide S-band range, range rate and angle information in the early phase of the mission to derive the spacecraft position. During the normal operation phase (Lunar orbiting phase), the proposed Indian DSN station at Bangalore will provide the tracking data. During the initial phase the tracking support will be taken from Berslake (DSN) and Saskatoon along with ISRO’s network of stations. The determined daily orbital estimates are used for spacecraft navigation, mission planning, and science processing. ISRO’s operational GEO missions Orbit Determination System (ODS) software was updated to cater the need aspect of meeting Lunar Mission ODS requirement during its different phases of the mission. This paper describes the Orbit Determination System (ODS) for Chandrayaan-1 during different phases of the mission. ODS methodology, software validation especially trajectory generation and achievable orbit determination accuracies during different phases of the mission are studied in detail. The main computation process of the orbit determination system is trajectory generation and estimation. Trajectory generation is performed through numerical integration of differential equations of motion of satellite. The force model mainly includes earth's gravitation/moon’s gravitation, aerodynamic drag, third body gravitation and solar radiation pressure. Cowell's method is used for trajectory generation. Weighted least squares technique and iterative differential correction process is used to obtain the refined state. Software was evaluated for trajectory generation with “Lunar Prospector Mission” ephemeris obtained from GSFC. Achievable orbit determination accuracy during different phases of the mission was carried out with simulated tracking data from the network of possible stations considered for the Lunar Mission. One day orbit prediction difference between ODS ephemeris and GSFC’s ephemeris for LP mission is about 25m and 4 cm/sec in position and velocity respectively during Lunar mapping phase. Orbit determination accuracy during lunar mapping phase is about 500m with Bangalore and Bearslake DSN ground stations tracking data.
- Abstract document
- Manuscript document
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