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| 1 | Survey of convex optimization for aerospace applications显示文摘Convex optimization is a class of mathematical programming problems with polynomial complexity for which state-of-the-art, highly efficient numerical algorithms with predeterminable computational bounds exist. Computational efficiency and tractability in aerospace engineering, especially in guidance, navigation, and control (GN&C), are of paramount importance. With theoretical guarantees on solutions and computational efficiency, convex optimization lends itself as a very appealing tool. Coinciding the strong drive toward autonomous operations of aerospace vehicles, convex optimization has seen rapidly increasing utility in solving aerospace GN&C problems with the potential for onboard real-time applications. This paper attempts to provide an overview on the problems to date in aerospace guidance, path planning, and control where convex optimization has been applied. Various convexification techniques are reviewed that have been used to convexify the originally nonconvex aerospace problems. Discussions on how to ensure the validity of the convexification process are provided. Some related implementation issues will be introduced as well. | Xinfu Liu Ping Lu Binfeng Pan | 2017 | Astrodynamics2017，1，1： | 32 |
| 2 | Angles-only initial relative orbit determination algorithm for noncooperative spacecraft proximity operations显示文摘This research furthers the development of a closed-form solution to the angles-only initial relative orbit determination problem for non-cooperative target close-in proximity operations when the camera offset from the vehicle center-of-mass allows for range observability.In previous work,the solution to this problem had been shown to be non-global optimal in the sense of least square and had only been discussed in the context of Clohessy–Wiltshire.In this paper,the emphasis is placed on developing a more compact and improved solution to the problem by using state augmentation least square method in the context of the Clohessy–Wiltshire and Tschauner–Hempel dynamics,derivation of corresponding error covariance,and performance analysis for typical rendezvous missions.A two-body Monte Carlo simulation system is used to evaluate the performance of the solution.The sensitivity of the solution accuracy to camera offset,observation period,and the number of observations are presented and discussed. | Baichun Gong Wendan Li Shuang Li Weihua Ma Lili Zheng | 2018 | Astrodynamics2018，2，3： | 20 |
| 3 | Overview of China’s 2020 Mars mission design and navigation显示文摘Scheduled for an Earth-to-Mars launch opportunity in 2020,the China’s Mars probe will arrive on Mars in 2021 with the primary objective of injecting an orbiter and placing a lander and a rover on the surface of the Red Planet.For China’s 2020 Mars exploration mission to achieve success,many key technologies must be realized.In this paper,China’s 2020 Mars mission and the spacecraft architecture are first introduced.Then,the preliminary launch opportunity,Earth–Mars transfer,Mars capture,and mission orbits are described.Finally,the main navigation schemes are summarized. | Xiuqiang Jiang Bin Yang Shuang Li | 2018 | Astrodynamics2018，2，1： | 17 |
| 4 | A survey on artificial intelligence trends in spacecraft guidance dynamics and control显示文摘The rapid developments of artificial intelligence in the last decade are influencing aerospace engineering to a great extent and research in this context is proliferating.We share our observations on the recent developments in the area of spacecraft guidance dynamics and control,giving selected examples on success stories that have been motivated by mission designs.Our focus is on evolutionary optimisation,tree searches and machine learning,including deep learning and reinforcement learning as the key technologies and drivers for current and future research in the field.From a high-level perspective,we survey various scenarios for which these approaches have been successfully applied or are under strong scientific investigation.Whenever possible,we highlight the relations and synergies that can be obtained by combining different techniques and projects towards future domains for which newly emerging artificial intelligence techniques are expected to become game changers. | Dario Izzo Marcus Martens Binfeng Pan | 2019 | Astrodynamics2019，3，4： | 15 |
| 5 | Planetary defense mission concepts for disrupting/pulverizing hazardous asteroids with short warning time显示文摘This paper presents an overview of space mission concepts for disrupting or pulverizing hazardous asteroids, especially with warning time shorter than approximately 10 years. An innovative mission concept, referred to as a nuclear hypervelocity asteroid intercept vehicle (HAIV) system, employs both a kinetic-energy impactor and nuclear explosive devices. A new mission concept of exploiting a multiple kinetic-energy impactor vehicle (MKIV) system that doesn’t employ nuclear explosives is proposed in this paper, especially for asteroids smaller than approximately 150 m in diameter. The multiple shock wave interaction effect on disrupting or pulverizing a small asteroid is discussed using hydrodynamic simulation results. A multi-target terminal guidance problem and a planetary defense mission design employing a heavy-lift launch vehicle are also brie y discussed in support of the new non-nuclear MKIV mission concept. The nuclear HAIV and non-nuclear MKIV systems complement to each other to effectively mitigate the various asteroid impact threats with short warning time. | Bong Wie Ben Zimmerman Joshua Lyzhoft George Vardaxis | 2017 | Astrodynamics2017，1，1： | 10 |
| 6 | Real-time control for fuel-optimal Moon landing based on an interactive deep reinforcement learning algorithm显示文摘In this study,a real-time optimal control approach is proposed using an interactive deep reinforcement learning algorithm for the Moon fuel-optimal landing problem.Considering the remote communication restrictions and environmental uncertainties,advanced landing control techniques are demanded to meet the high requirements of real-time performance and autonomy in the Moon landing missions.Deep reinforcement learning(DRL)algorithms have been recently developed for real-time optimal control but suffer the obstacles of slow convergence and difficult reward function design.To address these problems,a DRL algorithm is developed using an actor-indirect method architecture to achieve the optimal control of the Moon landing mission.In this DRL algorithm,an indirect method is employed to generate the optimal control actions for the deep neural network(DNN)learning,while the trained DNNs provide good initial guesses for the indirect method to promote the efficiency of training data generation.Through sufficient learning of the state-action relationship,the trained DNNs can approximate the optimal actions and steer the spacecraft to the target in real time.Additionally,a nonlinear feedback controller is developed to improve the terminal landing accuracy.Numerical simulations are given to verify the effectiveness of the proposed DRL algorithm and demonstrate the performance of the developed optimal landing controller. | Lin Cheng Zhenbo Wang Fanghua Jiang | 2019 | Astrodynamics2019，3，4： | 8 |
| 7 | Formation ight design for a LISA-like gravitational wave observatory via Cascade optimization显示文摘Laser Interferometer Space Antenna(LISA)is a project to detect and measure gravitational waves.The project has three spacecraft ying in a formation of near equilateral triangle in a heliocentric orbit trailing Earth.Many sources of perturbations cause the con guration to deviate from the nominal.This paper studies the formation design problem for a LISA-like mission by considering ephemeris-based dynamics.This type of mission is well-known for addressing several strict mission requirements under the realistic dynamics.The problem is formulated as optimizing multiple mission performance indices.It is observed that some indices are correlated with each other,whereas some indices have di erent sensitivities with respect to the semi-major axis.Therefore,the problem is transformed into a two-step cascade single-objective optimization,in which the optimal solution of the rst optimization problem is fed to the second optimization as initial value.In addition,the major perturbing celestial bodies are picked up to make a simpli ed but accurate enough dynamics to speed up the optimization.Numerical examples verify the analysis and show the e ectiveness of the optimization procedure.The in uences of mission lifetime and spatial scales on the optimal solutions are also presented. | Chihang Yang Hao Zhang | 2019 | Astrodynamics2019，3，2： | 8 |
| 8 | Neural-network-based terminal sliding-mode control for thrust regulation of a tethered space-tug显示文摘This paper studies the thrust regulation of the tethered space-tug in order to stabilize the target towed by a flexible tether.To compromise between model accuracy and simplicity,a rigid-flexible coupling multi-body model is proposed as the full model of the tethered space-tug.More specifically,the tug and the towed target are assumed as rigid bodies,whereas the flexible tether is approximated as a series of hinged rods.The rods are assumed extensible but incompressible.Then the equations of motion of the multi-body system are derived based on the recursive dynamics algorithm.The attitude motion of the towed target is stabilized by regulating the thrust on the tug,whereas the tether-tension-caused perturbation to the tug's attitude motion is eliminated by the control torque on the tug.The regulated thrust is achieved by first designing an optimal control trajectory considering the simplified system model with constraints for both state variables and control input.Then the trajectory is tracked using a neural-network based terminal sliding-mode controller.The radial basis function neural network is used to estimate the unknown nonlinear difference between the simple model and the full model,while the terminal sliding mode controller ensures the rapid tracking control of the target's attitude motion.Thrust saturation and tether slackness avoidance are also considered.Finally,numerical simulations prove the effectiveness of the proposed controller using the regulated thrust.Without disturbing orbital motion much,the attitude motion of the tug and the target are well stabilized and the tether slackness is avoided. | Rui Zhong Shijie Xu | 2018 | Astrodynamics2018，2，2： | 7 |
| 9 | Target sequence optimization for multiple debris rendezvous using low thrust based on characteristics of SSO显示文摘A method is proposed to select the target sequence for a J 2-perturbed multiple debris rendezvous mission aimed at removing dozens of debris from several thousand debris candidates running on sun-synchronous orbits(SSO).The solving methodology proceeds in two steps:Firstly,the variance of the right ascension of ascending node(RAAN)of the debris group is used for narrowing down the potential debris candidate;secondly,the debris of the candidate group that has closest RAAN to the current debris is chosen as the next debris.The low thrust near-minimum-fuel trajectories of each rendezvous leg are obtained by the indirect optimization method.The proposed approach is demonstrated for the problem of the 8th China Trajectory Optimization Competition(CTOC).The radar cross section(RCS)of the debris is also considered in the first step since the primary performance index of the competition is to maximize the total RCS of the debris visited.The results show that the proposed approach achieves better performance within a competition period.Of the many rendezvous sequences found,the best one submitted for the competition obtained a total RCS of 184 by accomplishing rendezvous with 70 debris within a transfer duration of one year. | Shuge Zhao Jingrui Zhang Kaiheng Xiang Rui Qi | 2017 | Astrodynamics2017，1，1： | 6 |
| 10 | Optimization of observing sequence based on nominal trajectories of symmetric observing con guration显示文摘This paper presents the crucial method for obtaining our team's results in the 8th Global Trajectory Optimization Competition(GTOC8).Because the positions and velocities of spacecraft cannot be completely determined by one observation on one radio source,the branch and bound method for sequence optimization of multi-asteroid exploration cannot be directly applied here.To overcome this diculty,an optimization method for searching the observing sequence based on nominal low-thrust trajectories of the symmetric observing con guration is proposed.With the symmetric observing con guration,the normal vector of the triangle plane formed by the three spacecraft rotates in the ecliptic plane periodically and approximately points to the radio sources which are close to the ecliptic plane.All possible observing opportunities are selected and ranked according to the nominal trajectories designed by the symmetric observing con guration.First,the branch and bound method is employed to nd the optimal sequence of the radio source with thrice observations.Second,this method is also used to nd the optimal sequence of the left radio sources.The nominal trajectories are then corrected for accurate observations.The performance index of our result is 128,286,317.0 km which ranks the second place in GTOC8. | Hongwei Yang Gao Tang Fanghua Jiang | 2018 | Astrodynamics2018，2，1： | 6 |
| 11 | Comparative evaluation of three machine learning algorithms on improving orbit prediction accuracy显示文摘In this paper,the recently developed machine learning(ML)approach to improve orbit prediction accuracy is systematically investigated using three ML algorithms,including support vector machine(SVM),artificial neural network(ANN),and Gaussian processes(GPs).In a simulation environment consisting of orbit propagation,measurement,estimation,and prediction processes,totally 12 resident space objects(RSOs)in solar-synchronous orbit(SSO),low Earth orbit(LEO),and medium Earth orbit(MEO)are simulated to compare the performance of three ML algorithms.The results in this paper show that ANN usually has the best approximation capability but is easiest to overfit data;SVM is the least likely to overfit but the performance usually cannot surpass ANN and GPs.Additionally,the ML approach with all the three algorithms is observed to be robust with respect to the measurement noise. | Hao Peng Xiaoli Bai | 2019 | Astrodynamics2019，3，4： | 4 |
| 12 | Safe rendezvous scenario design for geostationary satellites with collocation constraints显示文摘Rendezvous on the geostationary orbit(GEO)is much more complex than that on the low earth orbit and has a higher critical requirement for safety performance.This paper presents a safe scenario design method for GEO rendezvous proximity missions where the safety constraint of a collocated satellite is considered.A recently proposed quantitative index considering trajectory uncertainty is introduced to analyze the safety performance of the scenario parameters including the V-bar keeping positions and the fly-by trajectory radius.Furthermore,an exhaustive analysis is performed to find the dangerous regions of the V-bar keeping positions and the appropriate semi-major axis of the fly-by ellipse,considering the safety requirements of both the target and the collocated satellite.A geometry method is then developed for designing a feasible and suboptimal safe rendezvous scenario.The method is tested by designing four rendezvous scenarios with±V-bar approach directions respectively in the situations with and without one collocated satellite.Safety performance and velocity increments of the scenarios are compared and a conclusion is reached that the collocated satellite has a significant influence on the scenario design. | Ya-Zhong Luo Zhen-Jiang Sun | 2017 | Astrodynamics2017，1，1： | 4 |
| 13 | Dust in the Jupiter system outside the rings显示文摘Jupiter is one of the major targets for planetary exploration,and dust in the Jovian system is of great interest to researchers in the field of planetary science.In this paper,we review the five dust populations outside the ring system:grains in the region of the Galilean moons,potential dust from plumes on Europa,Jovian stream particles,particles in the outer region of the Jovian system ejected from the irregular satellites,and dust in the region of the Trojan asteroids.The physical environment for the dust dynamics is described,including the gravity,the magnetic field and the plasma environment.For each population,the dust sources are described,and the relevant perturbation forces are discussed.Observations and results from modeling are reviewed,and the distributions of the individual dust populations are shown.The understanding of the Jovian dust environment allows to assess the dust hazard to spacecraft,and to characterize the material exchange between the Jovian moons,their surface properties and distribution of non-icy constituents. | Xiaodong Liu Jürgen Schmidt | 2019 | Astrodynamics2019，3，1： | 4 |
| 14 | Performance analysis of three-dimensional differential geometric guidance law against low-speed maneuvering targets显示文摘The performance of the three-dimensional differential geometric guidance law with proportional navigation formation against a target maneuvering arbitrarily with time-varying normal acceleration is thoroughly analyzed using the Lyapunov-like approach.The validation of this guidance law is firstly proved,and then the performance issues such as capturability,heading error control efficiency,line of sight rate convergence,and commanded acceleration requirement are analyzed,under the condition that the missile is initially flying toward the target with a speed advantage.It is proved that an intercept can occur and the line of sight rate and missile commanded acceleration can be limited in certain ranges,if the initial heading error is small and the navigation gain is sufficiently large.The nonlinear relative dynamics between the missile and the target is taken into full account,and the analysis process is simple and intuitive,due to the use of a convenient line of sight rotating coordinate system.Finally,the new theoretical findings are validated by numerical simulations. | Kebo Li Wenshan Su Lei Chen | 2018 | Astrodynamics2018，2，3： | 3 |
| 15 | Reachable domain for spacecraft with ellipsoidal Delta-V distribution显示文摘Conventional reachable domain(RD)problem with an admissible velocity increment,Δv,in an isotropic distribution,was extended to the general case withΔv in an anisotropic ellipsoidal distribution.Such an extension enables RD to describe the effect of initial velocity uncertainty because a Gaussian form of velocity uncertainty can be regarded as possible velocity deviations that are confined within an error ellipsoid.To specify RD in space,the boundary surface of RD,also known as the envelope,should be determined.In this study,the envelope is divided into two parts:inner and outer envelopes.Thus,the problem of solving the RD envelope is formulated into an optimization problem.The inner and outer reachable boundaries that are closest to and farthest away from the center of the Earth,respectively,were found in each direction.An optimal control policy is then formulated by using the necessary condition for an optimum;that is,the first-order derivative of the performance function with respect to the control variable becomes zero.Mathematical properties regarding the optimal control policy is discussed.Finally,an algorithm to solve the RD envelope is proposed.In general,the proposed algorithm does not require any iteration,and therefore benefits from quick computation.Numerical examples,including two coplanar cases and two 3D cases,are provided,which demonstrate that the proposed algorithm works efficiently. | Changxuan Wen Chao Peng Yang Gao | 2018 | Astrodynamics2018，2，3： | 3 |
| 16 | Nonlinear semi-analytical uncertainty propagation of trajectory under impulsive maneuvers显示文摘The usage of state transition tensors(STTs)was proved as an effective method for orbital uncertainty propagation.However,orbital maneuvers and their uncertainties are not considered in current STT-based methods.Uncertainty propagation of spacecraft trajectory with maneuvers plays an important role in spaceflight missions,e.g.,the rendezvous phasing mission.Under the effects of impulsive maneuvers,the nominal trajectory of a spacecraft will be divided into several segments.If the uncertainty is piecewise propagated using the STTs one after another,large approximation errors will be introduced.To overcome this challenge,a set of modified STTs is derived,which connects the segmented trajectories together and allows for directly propagating uncertainty from the initial time to the final time.These modified STTs are then applied to analytically propagate the statistical moments of navigation and impulsive maneuver uncertainties.The probability density function is obtained by combining STTs with the Gaussian mixture model.The proposed uncertainty propagator is shown to be efficient and affords good agreement with Monte Carlo simulations.It also has no dimensionality problem for high-dimensional uncertainty propagation. | Zhen Yang Ya-Zhong Luo Jin Zhang | 2019 | Astrodynamics2019，3，1： | 3 |
| 17 | Review on solar sail technology显示文摘This paper reviews solar sail trajectory design and dynamics,attitude control,and structural dynamics.Within the area of orbital dynamics,methods relevant to transfer trajectory design and non-Keplerian orbit generation are discussed.Within the area of attitude control,di erent control strategies,including utilisation of solar radiation pressure and conventional actuators,are discussed.Finally,the methods of modelling structural dynamics during and after deployment are discussed,before considering possible future trends in developing of solar sailing missions. | Shengping Gong Malcolm Macdonald | 2019 | Astrodynamics2019，3，2： | 3 |
| 18 | Optimal circle-to-rectilinear orbit transfer with circumferential thrust显示文摘This paper investigates the optimal transfer trajectories from a circular parking orbit towards the apocenter of a rectilinear ellipse,where the spacecraft reaches a quasi-stationary condition relative to an inertial reference frame.The spacecraft is equipped with a propulsion system that provides a circumferential continuous propulsive acceleration,that is,an acceleration whose direction is perpendicular to the primary body-spacecraft line.The performance index to minimize is the total flight time,and an indirect method is used to analyze the transfer trajectories.In this context,the optimal transfer performance is obtained as a function of the spacecraft propulsive acceleration magnitude through an interpolation procedure of numerical simulations.The results obtained with a continuous thrust propulsion system are also compared with those derived from a multi-impulse transfer.Finally,the paper investigates a heliocentric mission scenario in which the spacecraft minimizes the flight time required to reach a rectilinear ellipse with a given value of the aphelion radius. | Alessandro A.Quarta Giovanni Mengali Andrea Caruso | 2019 | Astrodynamics2019，3，1： | 2 |
| 19 | Image-based autonomous navigation of Hayabusa2 using artificial landmarks: The design and brief in-flight results of the first landing on asteroid Ryugu显示文摘Hayabusa2 is an asteroid sample return mission carried out by the Japan Aerospace Exploration Agency.The spacecraft was launched in 2014 and arrived at the target asteroid Ryugu on June 27,2018.During the 1.5-year proximity phase,several critical operations(including two landing/sampling operations)were successfully performed.They were based on autonomous image-based descent and landing techniques.This paper describes an imagebased autonomous navigation scheme of the Hayabusa2 mission using artificial landmarks named target markers(TMs).Its basic algorithm,and the in-flight results of the first touchdown and its rehearsal,are shown. | Naoko Ogawa Fuyuto Terui Yuya Mimasu Kent Yoshikawa Go Ono Seiji Yasuda Kota Matsushima Tetsuya Masuda Hiroki Hihara Junpei Sano Takashi Matsuhisa Satoshi Danno Manabu Yamada Yasuhiro Yokota Yuto Takei Takanao Saiki Yuichi Tsuda | 2020 | Astrodynamics2020，4，2： | 2 |
| 20 | Orbital rendezvous performance comparison of differential geometric and ZEM/ZEV feedback guidance algorithms显示文摘In this paper,the performance of two distinct classes of feedback guidance algorithms is evaluated for a spacecraft rendezvous problem utilizing a continuous low-thrust propulsion system.They are the DG(Differential Geometric)and ZEM/ZEV(Zero-Effort-Miss/Zero-Effort-Velocity)feedback guidance algorithms.Even though these two guidance algorithms do not attempt to minimize the onboard fuel consumption orΔV directly,theΔV requirement is used as a measure of their orbital rendezvous performance for various initial conditions and a wide range of the rendezvous time(within less than one orbital period of the target vehicle).For the DG guidance,the effects of its guidance parameter and terminal time on the closed-loop performance are evaluated by numerical simulations.For the ZEM/ZEV guidance,its nearfuel-optimality is further demonstrated for a rapid,short-range orbital rendezvous,in comparison with the corresponding open-loop optimal solutions.Furthermore,the poorΔV performance of the ZEM/ZEV guidance for a slow,long-range orbital rendezvous is remedied by simply adding an initial drift phase.The ZEM/ZEV feedback guidance algorithm and its appropriate variants are then shown to be a simple practical solution to a non-impulsive rendezvous problem,in comparison with the DG guidance as well as the open-loop optimal guidance. | Pengyu Wang Yanning Guo Bong Wie | 2019 | Astrodynamics2019，3，1： | 2 |
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