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The paper aims to introduce a trade-off method for selecting a mission concept for an asteroid mining mission. In particular, the method is applied to the KaNaRiA mission concept selection. After introducing the KaNaRiA project, the KaNaRiA mission concept selection and reference scenario are described in detail.

The paper introduces past relevant asteroid missions in general and the previous studies on asteroid mining in particular. Based on the review of past mission concepts to minor planets, the paper discusses the operational phases of a potential industrial and commercial space mining mission. The methodology for selecting a mission reference scenario is explained and the selected KaNaRiA mission scenario is described.

The key technology driver for a space mining mission is the autonomous on-board capability related to navigation, guidance and handling of hardware/software anomalies or unexpected events. With the methodology presented here, it is possible to derive a mission concept which provides an adequate test-bed for the validation and verification of algorithms for enhanced spacecraft autonomy. This is the primary scientific and engineering goal of the KaNaRiA project.

The mission concept selection method presented here can be used as a generalized approach for mining missions targeting asteroids in the solar system.

The availability and usage of space resources is seen as a possible solution for the imminent problem of diminishing terrestrial materials in the foreseen future. This paper explains a methodology to select mission concepts for asteroid mining missions.

The purpose of this paper is to present a methodology and experimental results on using global optimization algorithms to determine the optimal orbit, based on the mission requirements, for a set of spacecraft flying in formation with an asteroid.

A behavioral‐based hybrid global optimization approach is used to first characterize the solution space and find families of orbits that are a fixed distance away from the asteroid. The same optimization approach is then used to find the set of Pareto optimal solutions that minimize both the distance from the asteroid and the variation of the Sun‐spacecraft‐asteroid angle. Two sample missions to asteroids, representing constrained single and multi‐objective problems, were selected to test the applicability of using an in‐house hybrid stochastic‐deterministic global optimization algorithm (Evolutionary Programming and Interval Computation (EPIC)) to find optimal orbits for a spacecraft flying in formation with an orbit. The Near Earth Asteroid 99942 Apophis (2004 MN₄) is used as the case study due to a fly‐by of Earth in 2029 leading to two potential impacts in 2036 or 2037. Two black‐box optimization problems that model the orbital dynamics of the spacecraft were developed.

It was found for the two missions under test, that the optimized orbits fall into various distinct families, which can be used to design multi‐spacecraft missions with similar orbital characteristics.

The global optimization software, EPIC, was very effective at finding sets of orbits which met the required mission objectives and constraints for a formation of spacecraft in proximity of an asteroid. The hybridization of the stochastic search with the deterministic domain decomposition can greatly improve the intrinsic stochastic nature of the multi‐agent search process without the excessive computational cost of a full grid search. The stability of the discovered families of formation orbit is subject to the gravity perturbation of the asteroid and to the solar pressure. Their control, therefore, requires further investigation.

This paper contributes to both the field of space mission design for close‐proximity orbits and to the field of global optimization. In particular, suggests a common formulation for single and multi‐objective problems and a robust and effective hybrid search method based on behaviorism. This approach provides an effective way to identify families of optimal formation orbits.

The purpose of this paper is to assess the orbital perturbation caused by the gravitational orbit–attitude coupling of spacecraft in the proximity of asteroids.

The gravitational orbit–attitude coupling perturbation (GOACP), which has been neglected before in the close-proximity orbital dynamics about asteroids, is investigated and compared with other orbital perturbations. The GOACP has its origin in the fact that the gravity acting on a non-spherical extended body is actually different from that acting on a point mass located at the body’s center of mass, which is the approximated model in the orbital dynamics. Besides, a case study of a tethered satellite system is given by numerical simulations.

It is found that the ratio of GOACP to the asteroid’s non-spherical gravity is the order of ρ/a_e, where ρ is the spacecraft’s characteristic dimension and a_e is the asteroid’s mean radius. It can also be seen that as ρ increases, GOACP will also increase but the solar radiation pressure (SRP) will decrease due to the decreasing area-to-mass ratio. The GOACP will be more significant than SRP at small orbital radii for a large-sized spacecraft. Based on the results by analyses and simulations, it can be concluded that GOACP needs to be considered in the orbital dynamics for a large-sized spacecraft in the proximity of a small asteroid.

This study is of great importance for the future asteroids missions for scientific explorations and near-Earth objects mitigation.

The GOACP, which has been neglected before, is revealed and studied.

The purpose of this paper is to present a full fourth‐order model of the gravity gradient torque of spacecraft around asteroids by taking into consideration of the inertia integrals of the spacecraft up to the fourth order, which is an improvement of the previous fourth‐order model of the gravity gradient torque.

The fourth‐order gravitational potential of the spacecraft is derived based on Taylor expansion. Then the expression of the gravity gradient torque in terms of gravitational potential derivatives is derived. By using the formulation of the gravitational potential, explicit formulations of the full fourth‐order gravity gradient torque are obtained. Then a numerical simulation is carried out to verify the model.

It is found that the model is more sound and precise than the previous fourth‐order model due to the consideration of higher‐order inertia integrals of the spacecraft. Numerical simulation results show that the motion of the previous fourth‐order model is quite different from the exact motion, while the full fourth‐order model fits the exact motion very well. The full fourth‐order model is precise enough for high‐precision attitude dynamics and control around asteroids.

This high‐precision model is of importance for the future asteroids missions for scientific explorations and near‐Earth objects (NEOs) mitigation.

In comparison with previous models, a gravity gradient torque model around asteroids that is more sound and precise is established. This model is valuable for high‐precision attitude dynamics and control around asteroids.

This paper aims to report the first iteration on the Light Detection and Ranging (LIDAR) Engineering Model altimeter named HELENA. HELENA is a Time of Flight (TOF) altimeter that provides time-tagged distances and velocity measurements. The LIDAR can be used for support near asteroid navigation and provides scientific information. The HELENA design comprises two types of technologies: a microchip laser and low noise sensor. The synergies between these two technologies enable developing a compact instrument for range measurements of up to 14 km. Thermal-mechanical and radiometric simulations of the HELENA telescope are reported in this paper. The design is subjected to vibrational, static and thermal conditions, and it was possible to conclude by the results that the telescope is compliant with the random vibration levels, the static load and the operating temperatures.

The Asteroid Impact & Deflection Assessment (AIDA) is a collaboration between the NASA DART mission and ESA Hera mission. The aim scope is to study the asteroid deflection through a kinetic collision. DART spacecraft will collide with Didymos-B, while ground stations monitor the orbit change. HERA spacecraft will study the post-impact scenario. The HERA spacecraft is composed by a main spacecraft and two small CubeSats. HERA will monitor the asteroid through cameras, radar, satellite-to-satellite doppler tracking, LIDAR, seismometry and gravimetry.

The HELENA design comprises two types of technologies: a microchip laser and low noise sensor. The synergies between these two technologies enable developing a compact instrument for range measurements of up to 14 km.

In this paper is reported the first iteration on the LIDAR Engineering Model altimeter named HELENA. HELENA is a TOF altimeter that provides time-tagged distances and velocity measurements. The LIDAR can be used for support near asteroid navigation and provides scientific information. The HELENA design comprises two types of technologies: a microchip laser and low noise sensor. The synergies between these two technologies enable developing a compact instrument for range measurements of up to 14 km.

Describes the technological developments which are establishing the foundation for an exciting era of in situ exploration missions to planets, comets and asteroids with advanced robotic systems. Also outlines important concurrent terrestrial applications and spin offs of the space robotics technology. These include high‐precision robotic manipulators for microsurgical operations and dexterous arm control systems.

The purpose of this paper is to study the application of the planetary aerogravity‐assist (AGA) technique to the interplanetary transfer mission with low‐thrust engine, and the design and optimization approach of low‐thrust AGA trajectory.

In the research, the transfer trajectory with planetary AGA maneuver is analyzed first, the maximum atmospheric turn angle and the matching condition for AGA trajectory is derived out, which is the significant principle for AGA trajectory design and studies. Then, a design and optimization approach for interplanetary low‐thrust trajectory with AGA maneuver is developed. The complicated design problem is transformed into a parameter optimization problem with multiple nonlinear constraints by using calculus of variations and the matching condition associated with AGA trajectory. Furthermore, since the optimization problem is very sensitive to the launch date and AGA maneuver parameters, three ordinal sub‐problems are reformulated to reduce the sensitivity. Finally, a direct/indirect hybrid approach is utilized to solve these sub‐problems.

The planetary AGA maneuver is feasible and effective in decreasing the propellant consumption and flight time for interplanetary low‐thrust mission and provides better performance than pure planetary gravity assist. Moreover, the proposed approach is effective to design and optimize the low‐thrust transfer mission with AGA maneuver.

In further research, some simple preliminary design approaches for interplanetary low‐thrust trajectory with AGA maneuver are required to developed, which can provide a good initial conjecture for a hybrid optimization algorithm.

The paper provides the matching condition for interplanetary AGA transfer trajectory by analyzing some characteristics of planetary AGA maneuver, and presents an effective approach to design and optimize interplanetary low‐thrust AGA trajectory.

A severe breakage problem has been solved at Precision Propeller, Indianapolis, by using AerMet 100 alloy to make a shaft that holds propellers in place for straightening. The alloy is a high‐strength, high‐toughness steel developed by Carpenter Technology Corp., Reading, Pennsylvania.

The paper aims to address the combined attitude control and Sun tracking problem in a flexible spacecraft in the presence of external and internal disturbances. The attitude stabilization of a flexible satellite is generally a challenging control problem, because of the facts that satellite kinematic and dynamic equations are inherently nonlinear, the rigid–flexible coupling dynamical effect, as well as the uncertainty that arises from the effect of actuator anomalies.

To deal with these issues in the combined attitude and Sun tracking system, a novel control scheme is proposed based on the adaptive fuzzy Jacobian approach. The augmented spacecraft model is then analyzed and the Lyapunov-based backstepping method is applied to develop a nonlinear three-axis attitude pointing control law and the adaptation law.

Numerical results show the effectiveness of the proposed adaptive control scheme in simultaneously tracking the desired attitude and the Sun.

Reaction wheels are commonly used in many spacecraft systems for the three-axis attitude control by delivering precise torques. If a reaction wheel suffers from an irreversible mechanical breakdown, then it is likely going to interrupt the mission, or even leading to a catastrophic loss. The pitch-axis mounted solar array drive assemblies (SADAs) can be exploited to anticipate such situation to generate a differential torque. As the solar panels are rotated by the SADAs to be orientated relative to the Sun, the pitch-axis wheel control torque demand can be compensated by the differential torque.

The proposed Jacobian control scheme is inspired by the knowledge of Jacobian matrix in the trajectory tracking of robotic manipulators.