Research On Guidance Problem Of Tail-Case Intercept Using Small Satellites

Small satellites have several advantages including lower mission cost, smaller size, built and launched more guickly, and so on, which make small satellites predominant to be the platform of space weapons. In this paper, the tail-case intercept mission scenario utilized a smallsatellite as an interceptor is the research background. The key technologies including the optimal trajectory planning and precision guidance technology are systematically studied. The long-range guidance scheme, midcourse and terminal guidance law are designed and analyzed based on the characteristic of small satellites.Firstly, the operation functions of small satellites worked as the platforms of space weapons are introduced. Then, the operation mission based on small satellites is planned with system view, and the components and key technologies are defined, and also the flight phases of the whole rendezvous trajectory are demonstrated. At last, the mission model is established. This research work is the basis of the following research.The long-range guidance law considers the situation that the fuel the smallsatellite holds and the transfer time are limited. The relation of characteristic velocity, transfer time, and initial angle for Lambert's transfer problem between two circular orbits is analyzed by both analytical and numerical methods. The solution methods of the optimal intercept problem under the constraints of fuel, transfer time, and rendezvous angle are investigated, at the same time, the method decreasing the fuel expenditure and transfer time is given. The simulation results prove the long-range guidance law is effective.For the problem of small satellite co-orbital maneuver, the analytical guidance method with J2 perturbation based on the homogenious central force field is derived. The iterative guidance arithmetic to solve the tail-case intercept problem is present, in which the analytical solutions are substituted by the orbit papameters of Keplerian orbit. An approach to solve multiple-thrust transfer problem is investigated by use of the design idea of every thrust eliminating a part of the position miss at the terminal time. Genetic algorithm is utilized to optimize the ignition time, thrust time and thrust direction. Simulation results test that the fuel expenditure is decreased.The quantitative precision analysis for several classic zero-effort-intercept manifolds is made, and the necessary conditions at the burnoff time of the midcourse phase are deduced from the simulation results. The optimal midcourse guidance with the terminal constraints being the zero-effort-intercept manifold is derived. A midcourse guidance scheme by regulating the relative distance, the first order derivative of the relative distance, and the sightline angle speed is given, which can satisfy the initial requirety of the terminal guidance phase.Fuzzy Model Reference Learning Control (FMRLC) is designed. In the algorithm, dynamically focused learning (DFL) strategy based fuzzy control and adaptive learning of fuzzy rules are adopted. The problem of too small learning region in the learning is overcomed. The ability to adapt the disturbance and the control precision are improved.