Precise Orbit Determination Of Small Body Spacecraft And Gravitational Mass Estimation

Spacecraft precise orbit determination and GM estimate play a key role in small body missions.In terms of Rosetta mission and China's future small body mission,I thoroughly studied the basic theories,methods,models,algorithms and technical details of them.First,I developed my own software based on LUGREAS by code rewriting and function extension.All the works in this thesis were done using this software.Then I developed two updated Earth-based two-way Doppler models and processed two-way Doppler data collected in Rosetta mission.At the end,I designed the mission timeline,the nominal spacecraft orbit and simulation experiments for China's future asteroid mission.The works of this thesis are as following:(1)I summarized the theory,methods,models,algorithms and technical details of small body spacecraft precise orbit determination and GM estimate.The key points about time-space coordinate reference system were first described.Basic force models,measurement models and related partial derivatives were described in detail and corresponding calculation steps were also given.The measurement calibration was also described.(2)Based on LUGREAS,I developed my own software—ASGREAS(ASteroid Gravity REcovery and Analysis System)by code rewriting and function extension.The software can be used for spacecraft precise orbit determination,GM and ephemeris estimate.The software consists of three subsystems—subsystems for the estimate of arc-independent parameters,global parameters and small body ephemeris,respectively.These three subsystems communicate with each other by defined interface file and script functions.Specially,global parameters can be estimated parallelly using Open MP.The computation speed of the software is thus improved by 16?20 times.(3)I presented two updated two-way Doppler model—UTD model and UID model,aiming at impairing the effects of numerical errors in the computation of traditional two-way Doppler model.UTD model contains a new algorithm of spacecraft position difference when the central body is the Sun.UID model can be used when the form of the central body ephemeris is not in Chebyshev.The numerical noise can be reduced by two orders of magnitude using these two models.Moreover,an updated formulation of computing the Chebyshev time was given at last for analyzing their contribution to numerical errors,as an example of SPICE SPK type-2/3 kernel.The numerical noise can be reduced by two orders of magnitude using this formulation.(4)I processed two-way Doppler data collected during Rosetta Lutetia flyby using ASGREAS.First,I developed a model for the effect of Rosetta HGA slew on two-way Doppler observables.Second,based on parameter analysis,two different strategies were given for Rosetta precise orbit and Lutetia GM determination.In this thesis,Lutetia GM were determined to(0.1115±0.0028)km~3/s~2(2.5%,3).This result is consistent with which is determined by Rosetta radio science team.(5)I designed the mission timeline and spacecraft accompanying orbit of four different altitudes for China's future asteroid mission.The geometry between 2016HO3,the Sun and the Earth was analyzed.During the three-month escort phase,the spacecraft will accompany with the target along the line of“Earth-2016HO3”.(6)Combining Earth-based two-way range,two-way Doppler and onboard one-way range,the simulation experiments are conducted in two different scenarios.In the first scenario,the spacecraft accompanied the asteroid at a constant altitude for three months.In the second scenario,the spacecraft gradually approaches the asteroid with several orbit controls.The average accuracy of spacecraft position is about 200 m.The accuracy of the asteroid GM can be determined to 2%?20%when the spacecraft accompanied the asteroid at a constant altitude.When the spacecraft gradually approaches the asteroid,the GM accuracy can be determined to3%?11%.(7)I analyzed the evolution of 2016HO3 ephemeris error and the effect on single-arc precise orbit determination and GM estimate.2016HO3 ephemeris,as well as spacecraft orbit and 2016HO3 GM,were then solved simultaneously in different scenarios.The ephemeris error at the end of the escort phase was 4?5 times larger than which at the beginning of the escort phase.The error of spacecraft orbit was several tens of kilometer and the error in GM estimate is larger than the magnitude of GM in half of scenarios.The accuracy of spacecraft orbit and GM estimate can be improved by 10 times through solving for 2016HO3 ephemeris.If the range measurement is unbiased,the ephemeris accuracy is several hundred meters.If the range measurement is biased,the ephemeris error is several kilometers.In this case,the asteroid ephemeris should be refined by combining ground-based telescopes and onboard images.