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Near-Earth object (NEO) in-situ exploration can provide invaluable information for science, possible future deflection actions and resource utilisation. This is only possible with space missions which approach the asteroid from its vicinity, i.e. rendezvous. This paper explores the use of solar sailing as means of propulsion for NEO rendezvous missions. Given the current state of sail technology, we search for multiple rendezvous missions of up to ten years and characteristic acceleration of up to 0.10 mm/s2. Using a tree-search technique and subsequent trajectory optimisation, we find numerous options of up to three NEO encounters in the launch window 2019–2027. In addition, we explore steerable and throttleable low-thrust (e.g. solar-electric) rendezvous to a particular group of NEOs, the Taurid swarm. We show that an acceleration of 0.23 mm/s2 would suffice for a rendezvous in approximately 2000 days, while shorter transfers are available as the acceleration increases. Finally, we show low-thrust options (0.3 mm/s2) to the fictitious asteroid 2019 PDC, as part of an asteroid deflection exercise. [ABSTRACT FROM AUTHOR]

A major cause of spacecraft orbital variation comes from natural perturbations, which, in close proximity of a body, are dominated by its non-spherical nature. For small bodies, such as asteroids, these effects can be considerable, given their uneven (and uncertain) mass distribution. Solar sail technology is proposed to reduce or eliminate the net secular effects of the irregular gravity field on the orbit. Initially, a sensitivity analysis will be carried out on the system which will show high sensitivity to changes in initial conditions. This presents a challenge for optimisation methods which require an initial guess of the solution. As such, the Genetic Algorithm (GA) is proposed as the preferred optimisation method as this requires no initial guess from the user. A multi-objective optimisation is performed which aims to achieve a periodic orbit whilst also minimising the effort required by the sail to do so. Given the system sensitivity, the control law for one orbit is not necessarily applicable for any subsequent orbit. Therefore, a new method of updating the control law for subsequent orbits is presented, based on linearisation and use of a Control Transition Matrix (CTM). The techniques will later find application in a multiple asteroid rendezvous mission with a solar sail as the primary propulsion system. [ABSTRACT FROM AUTHOR]

This paper investigates the dynamics of an orbital siphon anchored to a rotating ellipsoidal asteroid. The siphon is a chain of tether-connected payload masses arranged vertically from the asteroid surface, envisaged for propellantless delivery of payloads (e.g., mined material) from the asteroid surface to a collecting spacecraft. If the structure is long enough, the centrifugal-induced force can overcome the gravitational force on the payloads, eventually allowing resource payloads to escape. By connecting new payloads at the bottom of this chain while removing upper payloads a net orbital siphon effect is established, which provides a net continuous flow of resources from the asteroid surface to a collecting spacecraft, attached at the top of the siphon. The dynamics of the siphon is investigated in detail by varying a set of relevant parameters, in particular, chain length, anchor location and asteroid shape. It is shown that the system exhibits oscillatory behaviour in the equatorial plane, with decreasing oscillation amplitude over time and that the longest equatorial end is the best anchor location to guarantee proper siphon operation while minimizing the chain length. Eventually, a method is proposed to exploit the equatorial Coriolis-induced oscillations of the siphon to transfer payload masses from the collecting spacecraft to the stable equilibrium points associated with the effective potential of the ellipsoidal asteroid, where a catcher would collect the material. • Dynamics of orbital siphon anchored at ellipsoidal asteroid is studied. • Anchoring at longest equatorial end minimizes siphon length. • Siphon exhibits damped oscillatory behaviour. • Stable equilibrium points can be used as gravitational depots. [ABSTRACT FROM AUTHOR]

A chain of tether-connected payload masses assembled from the surface material of a spherical rotating asteroid is envisaged as a means of delivering a fraction of the asteroid mass into orbit, without the need of external work to be done. Under conditions to be discussed, a net radial force is established on the chain which can be exploited to initialize an orbital siphon effect: new payloads are connected to the chain while top payloads are removed and released into orbit. Adopting simplifying assumptions, the underlying dynamics of the problem is entirely analytical and is investigated in detail. The amount of mass extractable from the asteroid is then discussed, according to a range of strategies. It is proposed that the scheme could in future provide an efficient means of extracting material resources from rotating Near Earth Asteroids. [ABSTRACT FROM AUTHOR]

Hill's approximation models the motion of two small masses gravitationally interacting with each other and perturbed by a large central body. An application of the model is employed in this paper to manoeuvre the relative motion of two asteroids by a small impulse to capture them in bound binary motion in Earth's orbit. The initial conditions prior to the capture manoeuvre are restricted in a parameter space termed the gateway region. The gateway region is produced by applying the constraint that the capture impulse is a real-valued function, and the zero-velocity curve closes, enveloping both asteroids. A full mission scenario is designed with three impulses. The first impulse transfers one asteroid from a far field region to the gateway region, where there is mutual interaction with the second asteroid, which is assumed to be the origin of the relative motion. The second impulse changes the trajectory of the asteroid in the gateway region to a linear drift approaching the second asteroid. The third impulse is calculated by means of increasing the Jacobi integral to reach the critical value required to close the zero-velocity curve. It is demonstrated in principle that the triple impulses enable the candidate asteroids to be captured in a bound pair in Earth's orbit. It is also shown that the thruster on-time for the capture impulse as applied for the different small asteroids considered is short relative to the natural time-scale of the orbital dynamics of the problem, which shows the feasibility of the impulse approximation with the typical thruster forces applied. This strategy could provide a basis for parking small captured near-Earth asteroids in Earth's orbit. • A three-impulse scenario is proposed to form an asteroid pair in bound motion. • Hill's equations are modified with capture impulse to the transverse relative velocity. • An analytical investigation for position and velocity constraints is applied to the capture impulse. • A gateway map is generated to provide possible initial conditions for the capture impulse scenario. • The scenario is illustrated by forming an artificial binary pair in high Earth orbit. [ABSTRACT FROM AUTHOR]

The fascinating idea of shepherding asteroids for science and resource utilization is being considered as a credible concept in a not too distant future. Past studies identified asteroids which could be efficiently injected into manifolds which wind onto periodic orbits around collinear Lagrangian points of the Sun-Earth system. However, the trajectories are unstable, and errors in the capture maneuver would lead to complete mission failure, with potential danger of collision with the Earth, if uncontrolled. This paper investigates the controllability of some asteroids along the transfers and the periodic orbits, assuming the use of a solar-electric low-thrust system shepherding the asteroid. Firstly, an analytical approach is introduced to estimate the stability of the trajectories from a dynamical point of view; then, a numerical control scheme based on a linear quadratic regulator is proposed, where the gains are optimized for each trajectory through a genetic algorithm. A stochastic simulation with a Monte Carlo approach is used to account for different perturbed initial conditions and the epistemic uncertainty on the asteroid mass. Results show that only a small subset of the considered combinations of trajectories/asteroids are reliably controllable, and therefore controllability must be taken into account in the selection of potential targets. [ABSTRACT FROM AUTHOR]

Multi-target missions are an attractive solution to visit multiple bodies, increasing the scientific return and reducing the cost, compared to multiple missions to individual targets. Examples of multi-target missions are multiple active debris removals (MADR) and multiple near-Earth asteroids rendezvous (MNR) missions. MADR missions allow for the disposal of inactive satellites, preventing the build-up of space junk, while MNR missions allow to reduce the expenses of each asteroid observation. Since those missions are long and highly demanding in terms of energy, it is paramount to select the most convenient propulsion system so that the propellant mass and the duration of the mission are minimized. To this end, this paper proposes the use of a multi-objective optimization and artificial neural networks. The methodology is assessed by optimizing trajectories for MADR and MNR sequences with off-the-shelf thrusters. Multiple Pareto-optimal solutions can be identified depending on the propulsion system characteristics, enabling mission designers to trade-off the different options quickly and reliably. [ABSTRACT FROM AUTHOR]

Space resource utilization (SRU) can be a catalyst for a growing space economy. Significant reductions in cost, launch mass and risk for human and robotic activities beyond Earth can be achieved. Commercial entities will have the opportunity to generate revenue through economic development, allowing for a decreased dependency on governmental agencies for the exploration of space. A wide variety of materials can be found on planetary and small solar system bodies, which are potentially of great use to customers in space. This paper conducts a comparison of the economic and commercial potential of SRU on the surface of the Moon, Mars, a representative near-Earth asteroid and a representative main-belt asteroid. In particular, mining of water will be considered which will be electrolysed into hydrogen and oxygen using solar power. Coupled economic and trajectory optimization then results in the minimum specific cost to deliver propellant (LOX/LH 2) from these bodies to a number of customer locations. Customers on the surface of or in orbit around the Earth, Moon, and Mars are investigated, along with Psyche, a large metallic main-belt asteroid. Using a cost model that considers technology maturation to allow current aviation-like costs to be used for development and manufacturing, missions are optimized. Results show that for customers beyond low-Earth orbit, SRU should be used instead of launching water-derived resource directly from the Earth. In particular, near-Earth asteroids are shown to be in an attractive location for mining, processing and delivering a large quantity of LOX/LH 2 at a low specific cost, for customers in the vicinity of the Earth and the Moon, including nearby collinear Lagrange points. Exceptions are customers on the surface of the Earth and the Moon, who, together with customers on the surface of Mars, are better off mining and processing on that same surface. A sensitivity analysis shows that, for all investigated customers except Psyche, there are near-Earth asteroids available which can be used to deliver water-derived resources for a lower specific cost than the most ideal main-belt asteroid. Only once these near-Earth asteroids have been depleted, mining main-belt asteroids and transporting resources back to a customer in the vicinity of the Earth is worth considering. • Economic modelling and trajectory optimization are coupled to minimize specific cost. • A parametric approach allows for unbiased comparison of source/customer-pair results. • The cheapest sources of LOX/LH2 for customers on/near the Earth/Moon/Mars are given. • Using near-Earth asteroids is often more cost-effective than using main-belt asteroids. • It is cost-prohibitive to use Mars-derived propellant elsewhere. [ABSTRACT FROM AUTHOR]

Future space ventures will likely require exploitation of near-Earth asteroid resources. Moreover, it can be envisaged that asteroids may host habitats in their interiors. In fact, a cavern inside an asteroid would be a natural radiation shield against cosmic radiation and may also serve as a confined environment for storage of mined material such as water ice or other processed volatiles such as propellants. To this end, this paper proposes to leverage the asteroid rotational self-energy to remove material from the asteroid interiors and create a spherical cavern, by means of the orbital siphon concept. The siphon is a chain of tether-connected payload masses (the asteroid material), which exploits the rotation of the asteroid for the delivery of mass from the asteroid to escape. Under certain conditions the siphon can be initiated to ensure self-sustained flow of mass from the asteroid to escape. A net orbital siphon effect is generated by connecting new payloads at the bottom of the chain while releasing the upper payloads. Key parameters are discussed, such as the required siphon dimension and the maximum size of the internal cavity that can be excavated, as a function of the asteroid rotational period. Moreover, assuming elastic material behaviour, a closed-form expression for the stress tensor is found and a failure criterion is used to identify regions in the asteroid interiors subjected to the larger stresses. It is shown that the conditions for failure are relaxed as the radius of the internal void increases. [ABSTRACT FROM AUTHOR]

• Integrating economics into trajectory optimization yields more realistic missions. • Asteroid mining can be profitable using both chemical and solar-sail propulsion. • Attainable target asteroids should have Earth-like orbits. • The attainable target region is larger for solar sails than for chemical propulsion. Asteroid mining has the potential to greatly reduce the cost of in-space manufacturing, production of propellant for space transportation and consumables for crewed spacecraft, compared to launching the required resources from the Earth's deep gravity well. This paper discusses the top-level mission architecture and trajectory design for these resource-return missions, comparing high-thrust trajectories with continuous low-thrust solar-sail trajectories. The paper focuses on maximizing the economic Net Present Value, which takes the time-cost of finance into account and therefore balances the returned resource mass and mission duration. The different propulsion methods are compared in terms of maximum economic return and sets of attainable target asteroids. Results for transporting resources to geostationary orbit show that the orbital parameter hyperspace of suitable target asteroids is considerably larger for solar sails, allowing for more flexibility in selecting potential target asteroids. Also, results show that the Net Present Value that can be realized is larger when employing solar sailing instead of chemical propulsion. In addition, it is demonstrated that a higher Net Present Value can be realized when transporting volatiles to the Lunar Gateway instead of geostationary orbit. The paper provides one more step towards making commercial asteroid mining an economically viable reality by integrating trajectory design, propulsion technology and economic modelling. [ABSTRACT FROM AUTHOR]

Abstract This paper investigates the concept of capturing near-Earth asteroids into bound orbits around the Earth by using aerobraking. To guarantee that the candidate asteroids cannot present an impact risk during aerobraking, an initial aerobraking hazard analysis is undertaken and accordingly only asteroids with a diameter less than 30 m are considered as candidates in this paper. Then, two asteroid capture strategies utilizing aerobraking are defined. These are termed single-impulse capture and bi-impulse capture, corresponding to two approaches to raising the perigee height of the captured asteroid's orbit after the aerobraking manoeuvre. A Lambert arc in the Sun-asteroid two-body problem is used as an initial estimate for the transfer trajectory to the Earth and then a global optimisation is undertaken, using the total transfer energy cost and the retrieved asteroid mass ratio (due to ablation) as objective functions. It is shown that the aerobraking can in principle enable candidate asteroids to be captured around the Earth with, in some cases, extremely low energy requirements. Highlights • An aerobraking hazard analysis is undertaken to select candidate asteroids. • Two asteroid capture strategies using aerobraking are proposed. • Maximum mass ratio of captured asteroid and spacecraft is obtained. • Best candidate asteroids which can be captured using aerobraking are found. [ABSTRACT FROM AUTHOR]

This paper investigates two strategies to engineer the capture of one member of a binary asteroid pair by a planetary body after close encounter with that planetary body. It is assumed that the binary pair consists of a smaller minor asteroid in orbit about a larger main asteroid, which encounters a planetary body. In order to develop an engineering model of the problem, first we neglect the mass of the smaller minor asteroid in the binary pair and approximate the model as planar parabolic restricted three-body problem (PPRTBP). Second, the related regularised dynamical equations for the problem are developed. An approximate analytical solution to the problem is then obtained for motion in the vicinity of the main asteroid using the regularised coordinates through a linearized model. This provides insight into the motion of the minor asteroid about the main asteroid, allowing strategies to engineer the capture process to be developed. Based on the topology of the zero velocity curves (ZVCs) for the PPRTBP, we determine the capture region for the problem by developing initial condition maps (ICMs) and investigate the details of the dynamical process for capture. Two capture strategies are then proposed to engineer and extend the possibility for capture of the minor asteroid in binary pair. One is a re-phasing manoeuvre before encounter, which guarantees that the particle is within the capture region of the ICMs. The other is an optimal, single-impulse transfer during encounter to ensure transfer through the ZVC bottleneck and capture of the minor asteroid by the planetary body. The purpose of the paper is to explore such engineering strategies, rather than to provide new insights into natural capture dynamics. [ABSTRACT FROM AUTHOR]

Abstract The scientific interest in near-Earth asteroids (NEAs) and the classification of some of those as potentially hazardous for the Earth stimulated the interest in their exploration. Close-up observations of these objects will drastically increase our knowledge about the overall NEA population. For this reason, a multiple NEA rendezvous mission through solar sailing is investigated, taking advantage of the propellantless nature of this propulsion technology. Considering a spacecraft based on the DLR/ESA Gossamer technology, this work focuses on a method for searching possible sequences of NEA encounters. The effectiveness of the approach is demonstrated through a number of fully-optimised trajectories. The results show that it is possible to visit five NEAs within 10 years with near-term solar-sail technology. Moreover, a study on a reduced NEA database demonstrates the reliability of the approach used, showing that 58% of the sequences found with an approximated trajectory model can be converted into real feasible solar-sail trajectories. Overall, the study shows the effectiveness of the proposed automatic optimisation algorithm, which is able to find solutions for a large number of mission scenarios without any input required from the user. [ABSTRACT FROM AUTHOR]

Near-Earth asteroids have attracted attention for both scientific and commercial mission applications. Due to the fact that the Earth-Moon $$\hbox {L}_{1}$$ and $$\hbox {L}_{2}$$ points are candidates for gateway stations for lunar exploration, and an ideal location for space science, capturing asteroids and inserting them into periodic orbits around these points is of significant interest for the future. In this paper, we define a new type of lunar asteroid capture, termed direct capture. In this capture strategy, the candidate asteroid leaves its heliocentric orbit after an initial impulse, with its dynamics modeled using the Sun-Earth-Moon restricted four-body problem until its insertion, with a second impulse, onto the $$\hbox {L}_{2}$$ stable manifold in the Earth-Moon circular restricted three-body problem. A Lambert arc in the Sun-asteroid two-body problem is used as an initial guess and a differential corrector used to generate the transfer trajectory from the asteroid's initial obit to the stable manifold associated with Earth-Moon $$\hbox {L}_{2}$$ point. Results show that the direct asteroid capture strategy needs a shorter flight time compared to an indirect asteroid capture, which couples capture in the Sun-Earth circular restricted three-body problem and subsequent transfer to the Earth-Moon circular restricted three-body problem. Finally, the direct and indirect asteroid capture strategies are also applied to consider capture of asteroids at the triangular libration points in the Earth-Moon system. [ABSTRACT FROM AUTHOR]

Highly efficient low-thrust propulsion is increasingly applied beyond commercial use, also in mainstream and flagship science missions, in combination with gravity assist propulsion. Another recent development is the growth of small spacecraft solutions, not in size but in numbers and individual capabilities. Just over ten years ago, the DLR-ESTEC G ossamer Roadmap to Solar Sailing was set up to guide technology developments towards a propellant-less and highly efficient class of spacecraft for solar system exploration and applications missions: small spacecraft solar sails designed for carefree handling and equipped with carried application modules. Soon, in three dedicated G ossamer Roadmap Science Working Groups it initiated studies of missions uniquely feasible with solar sails such as Displaced L 1 (DL1) space weather advance warning and monitoring, Solar Polar Orbiter (SPO) delivery to very high inclination heliocentric orbit, and multiple Near-Earth Asteroid (NEA) rendezvous (MNR). Together, they demonstrate the capability of near-term solar sails to achieve at least in the inner solar system almost any kind of heliocentric orbit within 10 years, from the Earth-co-orbital to the extremely inclined, eccentric and even retrograde. Noted as part of the MNR study, sail-propelled head-on retrograde kinetic impactors (RKI) go to this extreme to achieve the highest possible specific kinetic energy for the deflection of hazardous asteroids. At DLR, the experience gained in the development of deployable membrane structures leading up to the successful ground deployment test of a (20 m)2, i.e., 20 m by 20 m square solar sail at DLR Cologne in 1999 was revitalized and directed towards a 3-step small spacecraft development line from as-soon-as-possible sail deployment demonstration (G ossamer -1) via in-flight evaluation of sail attitude control actuators (G ossamer -2) to an envisaged proving-the-principle flight in the Earth-Moon system (G ossamer -3). First, it turned the concept of solar sail deployment on its head by introducing four separable Boom Sail Deployment Units (BSDU) to be discarded after deployment, enabling lightweight 3-axis stabilized sailcraft. By 2015, this effort culminated in the ground-qualified technology of the DLR G ossamer -1 deployment demonstrator Engineering Qualification Model (EQM). For mission types using separable payloads, such as SPO, MNR and RKI, design concepts can be derived from the BSDU characteristic of DLR G ossamer solar sail technology which share elements with the separation systems of asteroid nanolanders like MASCOT. These nano-spacecraft are an ideal match for solar sails in micro-spacecraft format whose launch configurations are compatible with ESPA and ASAP secondary payload platforms. Like any roadmap, this one contained much more than the planned route from departure to destination and the much shorter distance actually travelled. It is full of lanes, narrow and wide, detours and shortcuts, options and decision branches. Some became the path taken on which we previously reported. More were explored along the originally planned path or as new sidings in search of better options when circumstance changed and the project had to take another turn. But none were dead ends, they just faced the inevitable changes when roadmaps face realities and they were no longer part of the road ahead. To us, they were valuable lessons learned or options up our sleeves. But for future sailors they may be on their road ahead. [ABSTRACT FROM AUTHOR]

Abstract Any effort which intends to physically interact with specific asteroids requires understanding at least of the composition and multi-scale structure of the surface layers, sometimes also of the interior. Therefore, it is necessary first to characterize each target object sufficiently by a precursor mission to design the mission which then interacts with the object. In small solar system body (SSSB) science missions, this trend towards landing and sample-return missions is most apparent. It also has led to much interest in Mascot -like landing modules and instrument carriers. They integrate at the instrument level to their mothership and by their size are compatible even with small interplanetary missions. The DLR-ESTEC Gossamer Roadmap NEA Science Working Groups' studies identified Multiple NEA Rendezvous (MNR) as one of the space science missions only feasible with solar sail propulsion. Parallel studies of Solar Polar Orbiter (SPO) and Displaced L 1 (DL1) space weather early warning missions studies outlined very lightweight sailcraft and the use of separable payload modules for operations close to Earth as well as the ability to access any inclination and a wide range of heliocentric distances. These and many other studies outline the unique capability of solar sails to provide access to all SSSB, at least within the orbit of Jupiter. Since the original MNR study, significant progress has been made to explore the performance envelope of near-term solar sails for multiple NEA rendezvous. However, although it is comparatively easy for solar sails to reach and rendezvous with objects in any inclination and in the complete range of semi-major axis and eccentricity relevant to NEOs and PHOs, it remains notoriously difficult for sailcraft to interact physically with a SSSB target object as e.g. the Hayabusa missions do. The German Aerospace Center, DLR, recently brought the Gossamer solar sail deployment technology to qualification status in the Gossamer -1 project. Development of closely related technologies is continued for very large deployable membrane-based photovoltaic arrays in the GoSolAr project. We expand the philosophy of the Gossamer solar sail concept of efficient multiple sub-spacecraft integration to also include landers for one-way in-situ investigations and sample-return missions. These are equally useful for planetary defence scenarios, SSSB science and NEO utilization. We outline the technological concept used to complete such missions and the synergetic integration and operation of sail and lander. We similarly extend the philosophy of Mascot and use its characteristic features as well as the concept of Constraints-Driven Engineering for a wider range of operations. Highlights • Asteroid exploration is key to planetary science, new resources & planetary defense. • Multiple near-Earth asteroid rendezvous (MNR) is uniquely possible by solar sail. • Sail-carried asteroid landers enable direct surface access for science operations. • Small spacecraft solar sails and landers are cost-efficient to build and launch. • Mature technologies were created in the projects Gossamer-1 and MASCOT. [ABSTRACT FROM AUTHOR]