Baoyin Hexi

Lunar Landing Trajectory Design Based on Invariant Manifold

The low-energy lunar landing trajectory design using the invariant manifolds of restricted three-body problem is studied. Considering angle between the ecliptic plane and lunar orbit plane, the four-body problem of sun-earth-moon-spacecraft is divided into two three-body problems, the sun-earth-spacecraft in the ecliptic plane and the earth-moon-spacecraft in the lunar orbit plane. Using the orbit maneuver at the place where the two planes and the invariant manifolds intersect, a general method to design low energy lunar landing trajectory is given. It is found that this method can save the energy about 20% compared to the traditional Hohmann transfer trajectory. The mechanism that the method can save energy is investigated in the point of view of energy and the expression of the amount of energy saved is given. In addition, some rules of selecting parameters with respect to orbit design are provided. The method of energy analysis in the paper can be extended to energy analysis in deep space orbit design.

Comparison of two methods in satellite formation flying

Recently, the research of dynamics and control of the satellite formation flying has been attracting a great deal of attentions of the researchers. The theory of the research was mainly based on Clohessy-Wiltshires (C-Ws) equations, which describe the relative motion between two satellites. But according to some special examples and qualitative analysis, neither the initial parameters nor the period of the solution of C-Ws equations accord with the actual situation, and the conservation of energy is no longer held. A new method developed from orbital element description of single satellite, named relative orbital element method (ROEM), was introduced. This new method, with clear physics conception and wide application range, overcomes the limitation of C-Ws equation, and the periodic solution is a natural conclusion. The simplified equation of the relative motion is obtained when the eccentricity of the main satellite is small. Finally, the results of the two methods (C-Ws equation and ROEM) are compared and the limitations of C-Ws equations are pointed out and explained.

Formation around planetary displaced orbit

The paper investigates the relative motion around the planetary displaced orbit. Several kinds of displaced orbits for geocentric and martian cases were discussed. First, the relative motion was linearized around the displaced orbits. Then, two seminatural control laws were investigated for each kind of orbit and the stable regions were obtained for each case. One of the two control laws is the passive control law that is very attractive for engineering practice. However, the two control laws are not very suitable for the Martian mission. Another special semi-natural control law is designed based on the requirement of the Martian mission. The results show that large stable regions exist for the control law.

Semi-physical Simulation Research of the Antenna Acquiring and Tracking System without Turn Tables

Simulation is an important step during the design of control systems. Based on the series tools of MATLAB\Simulink\xPC, the real-time simulation environment is constructed. By adopting the mixed modeling measures that comprise Simulink, Stateflow and S-function, models for satellite attitude dynamics, all sensors and actuators, control logics of the antenna and etc are all built successfully, then the modes selecting, scheduling and interfaces communication are realized effectively. The tracking and data relay satellite (TDRS) simulator is designed. It simplifies the experiment system. The turn table is eliminated and the experiment system can also reflect the effect of the attitude movement acting on the antenna pointing by using the TDRS simulator. The experiment results indicate that the experiment system can verify the antenna acquiring and tracking processes and is beneficial to the design of the user satellite’s antenna pointing control system for engineers.

Study of zenith pass problem of the inter-satellite linkage antenna

While adopting an elevation-over-azimuth architecture by an inter-satellite linkage antenna of a user satellite, a zenith pass problem always occurs when the antenna is tracing the tracking and data relay satellite (TDRS). This paper deals with this problem by way of, firstly, introducing movement laws of the inter-satellite linkage to predict the movement of the user satellite antenna followed by analyzing the potential pass moment and the actual one of the zenith pass in detail. A number of specific orbit altitudes for the user satellite that can remove the blindness zone are obtained. Finally, on the base of the predicted results from the movement laws of the inter-satellite linkage, the zenith pass tracing strategies for the user satellite antenna are designed under mechanical tilting and program guidance by using a trajectory preprocessor. Simulations have confirmed the reasonability and feasibility of the strategies in dealing with the zenith pass problem.

A new type of aerospace engine to remove space debris

We present a design concept for a space engine that can continuously remove orbital debris by using the debris powder as a propellant. Space debris represents a great potential threat to operational satellites, especially certain high-value spaceships. With the increasing risk of debris, researches reveal that it would reach a phenomenon named the Kessler syndrome that numerous operational satellites being destroyed in a chain reaction in low orbit in the not-far distant future. Initiative removing enough of the debris for the safety of space flight is an inevitable problem for human. Major space organizations are monitoring debris constantly with space debris environment models. At the same time, researchers have proposed a series of methods for removing the debris, such as using electrically conductive tethers to deorbit the debris, using robots with claws, nets or gecko-like suction cups to capture the debris and deploying ground-based or space-based laser to deorbit the debris. Nevertheless, large consumption of fuel for orbital transfer especially noncoplanar, chasing and rendezvousing with the debris is the leading restricting factor. In the concept of debris engine, a robotic space cleaner is used to capture the targeted debris and transfer it to engine. Large debris is firstly disintegrated into small pieces with diameters smaller than one centimeter using a mechanical method. A planetary ball mill is used to grind the pieces into powder with diameters of one micrometer or less. The energy required for this process is obtained from solar power or nuclear. The debris powder is then charged by electron beam. This charged condition can be used to accelerate the movement of the powder in a tandem electrostatic particle accelerator. Continuous thrust is generated by ejecting charged powder from the nozzle of engine because of momentum transfer. This thrust allow the spacecraft and robotic cleaner to perform orbital maneuvers and to rendezvous with other debris. The ejected charged particles will be blown away from the circumterrestrial orbit by solar wind and the effect of Earth magnetic field. Thus, digesting the space debris not only removes the orbital debris but also provides the necessary thrust to propel cleaner. Numerical simulation reveals the specific impulse of debris engine is determined by the accelerating electrostatic potential, charging potential and the charge-to-mass ratio of powder. In addition, the thrust of engine is largely dependent on the mass flow rate of the charged powder at the nozzle and nozzle jet velocity of the powder. With optimal regulation of the specific impulse and thrust, spacecraft can follow an optimal trajectory. In addition, this approach provides a new concept for asteroid exploration and interplanetary flight using a sustainable fuel supplement.

Robust Adaptive Control Method for Inter-satellite Linkage Antenna Acquisition and Tracking

A robust adaptive controller is presented for the control system of Inter-Satellite Linkage (ISL) antenna acquisition and tracking, which faces with the uncertainties of un-modeled dynamics and unbounded disturbance and the control input constraints. The stability is proved by using Lyapunov stabilization theory. The simulation results show that the robust adaptive controller can satisfy all the design requirements, especially the control accuracy is amazingly high contrast to the required accuracy. Furthermore, robust stability of the controller for un-modeled dynamics, uncertainties of parameters and the external disturbance was verified.