Automation And Optimization Of Spacecraft Rendezvous Phasing In A Near-Circular Orbit

The objective of phasing phase of a rendezvous or an on orbit servicing mission is to reduce the phase angle between the chaser and target spacecraft.The chaser is commonly controlled by the ground during the phasing phase in the previous researches.To improve reliability of the orbital maneuver,when an impulse is applied on the chaser,the chaser should have a direct communication with ground stations.The maneuvers have to take place at a fixed point in time to meet communication conditions.Therefore,for the purpose of reducing the mission cost,simplifying the ground control center,and minimizing the phasing time and the fuel consumption for the rendezvous and the on orbit servicing mission,this dissertation studies the automation and optimization of the spacecraft rendezvous phasing in a near-circular orbit.An improved method based on special-point-based maneuvers is presented to realize onboard autonomous rendezvous phasing in a near-circular orbit.The phasing segment is divided into three subphases(the initial subphase,the natural phasing subphase,and the adjusting subphase)according to the relationship of orbital phase angle and semimajor axis difference between the chaser and the target.The maneuver to natural phasing orbit is done in the initial subphase.In the natural phasing subphase,the phase angle between two spacecraft is reduced naturally due to the semimajor axis difference,and then the maneuver to the adjusting orbit is implemented.In the adjusting subphase,the phase angle is reduced step by step to satisfy the close rendezvous approaches requirement.In addition to the inplane maneuvers,the coplanar orbital corrections to eliminate the crosstrack difference are done in the corresponding subphase.All orbital maneuvers are given as guidance impulses and calculated onboard autonomously.Precise orbital simulation results show that the autonomous guidance algorithm for rendezvous-phasing maneuvers is feasible.The minimum-fuel transfer needed for servicing client geostationary satellites is found while considering perturbations that have been neglected in previous studies.The effect of Earth’s triaxiality on the semimajor axis and longitude is derived by the method of averaging and used for designing a two-impulse planar phasing maneuver.The phasing maneuver is then extended to a three-impulse planar maneuver for matching the eccentricity.The inclination variation due to lunisolar perturbations is modeled to determine the maneuver needed for eliminating the normal direction excursion.Then,an optimal rendezvous model is built with Earth’s triaxiality and lunisolar perturbations included and applied for servicing a sparsely distributed geostationary constellation.The results show that the duration of each rendezvous of the optimized solution is mainly determined by third-body effects.The proposed maneuver strategy with a tesseral-term correction can achieve relatively high accuracy in long duration orbital transfers under the effects of Earth’s actual gravitational field,lunisolar attraction,and solar radiation pressure.An analytic initial costates estimation method is presented for the minimum-time station change of geostationary satellites with low thrust.Different from the previous research,which formulated it as a time-constrained station change maximization problem,this method models the practical station change mission as a minimum-time problem in a more straightforward way.Furthermore,an analytic method is presented to reduce the number of the unknown initial costates from five to two,which could consequently reduce the number of iterations and improve the robustness of the shooting procedure.We present two approximation techniques to generate a solution of the final time,as well as the initial costates associated with the semimajor axis,longitude,and mass.The two techniques rely on an approximation of the optimal control by a tangential control and removal of the short-periodic terms.Numerous simulations show that the initial costates of semimajor axis,longitude,and mass of the optimal control strategy match well with the guess values generated by the presented analytic estimation method.Hence,the minimum-time station change solution can be obtained rapidly by employing the estimated initial costates.Minimum-fuel station change of geostationary satellites with low thrust is studied while taking into account of the significant perturbation forces for geostationary Earth orbit(GEO).The effect of Earth’s triaxiality,lunisolar perturbations,and solar radiation pressure on terminal conditions of long duration GEO transfers is derived and used for establishing the station-change model with consideration of significant perturbation forces.A method is presented for analytically evaluating the effect of Earth’s triaxiality on the semimajor axis and longitude during a station-change.The minimum-fuel problem is solved by the indirect optimization method.The easier and related minimum-energy problem is first addressed and then the energy-to-fuel homotopy is employed to finally obtain the solution of the minimum-fuel problem.Several effective techniques are employed in solving the two-point boundary-value problem with a shooting method to overcome the problem of the small convergence radius and the sensitivity of the initial costate variables.These methods include normalization of the initial costate vector,computation of the analytic Jacobians matrix,and switching detection.The simulation results show that the solution of the minimum-fuel station-change with low-thrust considering significant perturbation forces can be obtained rapidly by applying these preceding techniques.