A two-phase approach is proposed to deal with the integration problem in theloop layout. Tabu search is applied to cell construction in phase 1 to minimize the inter-cell flow,and the heuristic for layout design is us...A two-phase approach is proposed to deal with the integration problem in theloop layout. Tabu search is applied to cell construction in phase 1 to minimize the inter-cell flow,and the heuristic for layout design is used as phase 2 to optimize the sum of intra-cell andinter-cell transportation cost. The final computational results demonstrate the validation of thetwo-phase approach.展开更多
This letter proposes a method for designing a specific formation of satellites where the flying motion only exists in a circle orbit plane of the reference satellite, which means that the orbit eccen- tricity is zero....This letter proposes a method for designing a specific formation of satellites where the flying motion only exists in a circle orbit plane of the reference satellite, which means that the orbit eccen- tricity is zero. This method combines the Hill equation, the Kepler equation, and the geometrical meaning of orbit elements. It creates the redundancy condition to simplify the deducing process, utilizes multiple conditions to solve the orbit elements for the satellite formation, and obtains the analytical relationship of the orbit elements for the formation satellites with the formation parameters and the orbit elements of the reference satellite. Using these formulations, the orbit elements and formation parameters for the formation satellites can be solved for the given orbit elements of the reference satellite. The letter describes the proposed double-ellipse formation for both GMTI and InSAR, and the validity of the formation is demonstrated via simulation.展开更多
Spacecraft formation flying is an attractive new concept in international aeronautic fields because of its powerful functions and low cost. In this paper, the formation design and PD closed-loop control of spacecraft ...Spacecraft formation flying is an attractive new concept in international aeronautic fields because of its powerful functions and low cost. In this paper, the formation design and PD closed-loop control of spacecraft formation flying in elliptical orbits are discussed. Based on two-body relative dynamics, the true anomaly is applied as independent variable instead of the variable of time. Since the apogee is considered as the starting point, the six integrating constants are calculated. Therefore, the algebraic solution is obtained for the relative motion in elliptical orbits. Moreover, the formation design is presented and both circular formation and line formation are provided in terms of an algebraic solution. This paper also discusses the PD-closed loop control for precise formation control in elliptical orbits. In this part, the error-type state equation is put forward and the linear quadratic regulator （LQR） method is used to calculate PD parameters. Though the gain matrix calculated from LQR is time-variable because the error-type state equation is time variable, the PD parameters are also considered as constants because of their small changes in simulation. Finally, taking circular formation as an example, the initial orbital elements are achieved for three secondary spacecraft. And the numerical simulation is analyzed under PD formation control with initial errors and J2 perturbation. The simulation results demonstrate the validity of PD closed-loop control scheme.展开更多
This paper presents the methods and results for the trajectory design and optimization for the low earth orbit (LEO) satellites in formation to observe the geostationary orbit (GEO) satellites’ beams. The background ...This paper presents the methods and results for the trajectory design and optimization for the low earth orbit (LEO) satellites in formation to observe the geostationary orbit (GEO) satellites’ beams. The background of the trajectory design mission is the 9th China Trajectory Optimization Competition (CTOC9). The formation is designed according to the observation demands. The flying sequence is determined by a reference satellite using a proposed improved ephemeris matching method (IEMM). The formation is changed, maintained and transferred following the reference satellite employing a multi-impulse control method (MICM). Then the total observation value is computed by propagating the orbits of the satellites according to the sequence and transfer strategies. Based on the above methods, we have obtained a fourth prize in the CTOC9. The proposed methods are not only fit for this competition, but can also be used to fulfill the trajectory design missions for similar multi-object explorations.展开更多
基金This project is supported by National Natural Science Foundation of China (No.59889505, No.70071017).
文摘A two-phase approach is proposed to deal with the integration problem in theloop layout. Tabu search is applied to cell construction in phase 1 to minimize the inter-cell flow,and the heuristic for layout design is used as phase 2 to optimize the sum of intra-cell andinter-cell transportation cost. The final computational results demonstrate the validation of thetwo-phase approach.
文摘This letter proposes a method for designing a specific formation of satellites where the flying motion only exists in a circle orbit plane of the reference satellite, which means that the orbit eccen- tricity is zero. This method combines the Hill equation, the Kepler equation, and the geometrical meaning of orbit elements. It creates the redundancy condition to simplify the deducing process, utilizes multiple conditions to solve the orbit elements for the satellite formation, and obtains the analytical relationship of the orbit elements for the formation satellites with the formation parameters and the orbit elements of the reference satellite. Using these formulations, the orbit elements and formation parameters for the formation satellites can be solved for the given orbit elements of the reference satellite. The letter describes the proposed double-ellipse formation for both GMTI and InSAR, and the validity of the formation is demonstrated via simulation.
文摘Spacecraft formation flying is an attractive new concept in international aeronautic fields because of its powerful functions and low cost. In this paper, the formation design and PD closed-loop control of spacecraft formation flying in elliptical orbits are discussed. Based on two-body relative dynamics, the true anomaly is applied as independent variable instead of the variable of time. Since the apogee is considered as the starting point, the six integrating constants are calculated. Therefore, the algebraic solution is obtained for the relative motion in elliptical orbits. Moreover, the formation design is presented and both circular formation and line formation are provided in terms of an algebraic solution. This paper also discusses the PD-closed loop control for precise formation control in elliptical orbits. In this part, the error-type state equation is put forward and the linear quadratic regulator （LQR） method is used to calculate PD parameters. Though the gain matrix calculated from LQR is time-variable because the error-type state equation is time variable, the PD parameters are also considered as constants because of their small changes in simulation. Finally, taking circular formation as an example, the initial orbital elements are achieved for three secondary spacecraft. And the numerical simulation is analyzed under PD formation control with initial errors and J2 perturbation. The simulation results demonstrate the validity of PD closed-loop control scheme.
文摘This paper presents the methods and results for the trajectory design and optimization for the low earth orbit (LEO) satellites in formation to observe the geostationary orbit (GEO) satellites’ beams. The background of the trajectory design mission is the 9th China Trajectory Optimization Competition (CTOC9). The formation is designed according to the observation demands. The flying sequence is determined by a reference satellite using a proposed improved ephemeris matching method (IEMM). The formation is changed, maintained and transferred following the reference satellite employing a multi-impulse control method (MICM). Then the total observation value is computed by propagating the orbits of the satellites according to the sequence and transfer strategies. Based on the above methods, we have obtained a fourth prize in the CTOC9. The proposed methods are not only fit for this competition, but can also be used to fulfill the trajectory design missions for similar multi-object explorations.
