Simulation Results for the Lunar Orbiter

Previous Next
Stable Manifolds
This plot shows the x, y, and z components of position error for the low lunar orbiter in a 50x95 km altitude polar orbit. This orbit was chosen because of its reduced station keeping costs, but work is currently underway to generate a frozen orbit with an even smaller station keeping budget. The lines plotted in black show the error in the lunar orbiter's position estimate computed on-board the halo spacecraft. The dashed red lines indicate the ±2σ limits from the estimation error covariance. It takes a day or two for the EKF to converge for the lunar orbiter. The "true" low lunar orbit used to generate the observations was propagated using a variable step 7-8th order Runge Kutta integrator, a 20x20 portion of the lunar gravity field LP100K, a solar radiation pressure (SRP) model, and the point-mass gravitational acceleration of the Earth, Moon, Sun, Venus, Mars, Jupiter, and Saturn. The positions of the planets were computed using the JPL DE405 planetary ephemeris. Observations were generated every 60 seconds during nine 1-hour tracking periods each day. The orbit propagator in the EKF used a gravity field generated by taking a statistical sample of the LP100K gravity field using its covariance. The EKF SRP model was forced to generate SRP acceleration errors of about 1x10-9 m/s2. Using the on-board orbit estimates, the lunar orbiter computed and executed stationkeeping maneuvers about every two weeks. The stationkeeping maneuvers had a 5% execution error, which was also estimated by the EKF. The resulting 1σ position error RSS was 6.87 meters.

Geryon Home Start Previous [Slide 23] Next