Orbiting Solar Reflectors (OSRs) can be used to reflect sunlight locally to terrestrial solar power plants to enhance solar energy generation. Displaced polar orbits can, in principle, change the geometry of passes of OSR over terrestrial solar power plants. Such non-Keplerian orbits can be obtained by orienting the reflector at a fixed pitch angle with respect to the Sun-line, such that the solar radiation pressure (SRP) induced force would shift the orbit plane in the anti-Sun line. This, in principle, would allow extending night-time or high-latitude solar energy delivery without eclipses. This paper investigates a range of displaced highly non-Keplerian orbits for OSRs and assesses their operational use. Displaced polar orbits are generated in the two-body problem using a rotating reference frame considering the Earth's oblateness up to J₂ and the SRP force. Their stability is reviewed and an optimal control scheme is presented with reflector area control. As a novel application, a compound reflector system is proposed, which consists of a large Sun-facing parabolic collector in a polar orbit displaced in the anti-Sun direction and a smaller free-flying flat director placed near the focus of the parabolic collector, displaced by the reflected SRP in the Sun direction. The conditions for the synchronized motion and the sizing of the reflectors are investigated. The quantity of solar energy delivered to the Earth is calculated for both the compound and single reflector systems and it is shown that the displaced polar orbits could enhance solar energy delivery significantly.

Constellations of orbiting solar reflectors on Sun-synchronous repeat ground track orbits can in principle illuminate large terrestrial solar power plants after sunset or before sunrise. This will enhance the number of hours per day during which such solar power plants can deliver clean energy to the grid. In order to develop and deploy such large-scale in-orbit infrastructure, a number of technology demonstrations will be required to de-risk the technology and build confidence for investment. This paper considers potential technology demonstration activities for orbiting solar reflectors, from laboratory-scale testing to high altitude balloon flight and sub-scale in-orbit demonstration. The key technological requirements for orbiting solar reflectors are identified and the utility of each demonstration step evaluated. An integrated technology development, technology demonstration and investment roadmap is then presented.

The delivery of global clean energy services represents a key challenge for the 21st century. In order to deliver such services, it is clear that large-scale solar power farms will continue to grow both in number and size. In principle, ultralight membrane orbiting solar reflectors can illuminate large-scale solar power farms during the critical dawn/dusk hours of the day, enhancing the utility of terrestrial solar power. The key advantage is that only a relatively modest mass needs to be delivered to Earth orbit. This paper discusses the technical challenges associated with the development, deployment and operation of such a space-based energy service. Business development models are discussed along with regulatory issues and finally an integrated technology demonstration roadmap is presented.

DESTINY⁺ is an upcoming JAXA Epsilon medium-class mission to fly by the Geminids meteor shower parent body (3200) Phaethon. It will be the world's first spacecraft to escape from a near-geostationary transfer orbit into deep space using a low-thrust propulsion system. In doing so, DESTINY⁺ will demonstrate a number of technologies that include a highly efficient ion engine system, lightweight solar array panels, and advanced asteroid flyby observation instruments. These demonstrations will pave the way for JAXA's envisioned low-cost, high-frequency space exploration plans. Following the Phaethon flyby observation, DESTINY⁺ will visit additional asteroids as its extended mission. The mission design is divided into three phases: a spiral-shaped apogee-raising phase, a multi-lunar-flyby phase to escape Earth, and an interplanetary and asteroids flyby phase. The main challenges include the optimization of the many-revolution low-thrust spiral phase under operational constraints; the design of a multi-lunar-flyby sequence in a multi-body environment; and the design of multiple asteroid flybys connected via Earth gravity assists. This paper shows a novel, practical approach to tackle these complex problems, and presents feasible solutions found within the mass budget and mission constraints. Among them, the baseline solution is shown and discussed in depth; DESTINY⁺ will spend two years raising its apogee with ion engines, followed by four lunar gravity assists, and a flyby of asteroids (3200) Phaethon and (155140) 2005 UD. Finally, the flight operations plan for the spiral phase and the asteroid flyby phase are presented in detail.

The OSIRIS-REx mission collected a sample from the surface of the asteroid (101955) Bennu in 2020 October. Here, we study the impact of the OSIRIS-REx Touch-and-Go Sampling Acquisition Mechanism (TAGSAM) interacting with the surface of an asteroid in the framework of granular physics. Traditional approaches to estimating the penetration depth of a projectile into a granular medium include force laws and scaling relationships formulated from laboratory experiments in terrestrial-gravity conditions. However, it is unclear that these formulations extend to the OSIRIS-REx scenario of a 1300-kg spacecraft interacting with regolith in a microgravity environment. We studied the TAGSAM interaction with Bennu through numerical simulations using two collisional codes, pkdgrav and gdc-i. We validated their accuracy by reproducing the results of laboratory impact experiments in terrestrial gravity. We then performed TAGSAM penetration simulations varying the following geotechnical properties of the regolith: packing fraction (P), bulk density, inter-particle cohesion (σc), and angle of friction (φ). We find that the outcome of a spacecraft-regolith impact has a non-linear dependence on packing fraction. Closely packed regolith (P ? 0.6) can effectively resist the penetration of TAGSAM if φ ? 28° and/or σc ? 50 Pa. For loosely packed regolith (P ? 0.5), the penetration depth is governed by a drag force that scales with impact velocity to the 4/3 power, consistent with energy conservation. We discuss the importance of low-speed impact studies for predicting and interpreting spacecraft-surface interactions. We show that these low-energy events also provide a framework for interpreting the burial depths of large boulders in asteroidal regolith.

Ballistic landers enable orbiting asteroid missions to perform surface science at limited additional cost and risk. Due to asteroids’ weak gravity and irregular terrain, lander deployment trajectories will consist of several chaotic bounces. Although impacts on regolith-covered asteroids are numerically expensive to model, impacts on rocky asteroids can be modeled with simpler, impulsive contact models. One such model is that by Stronge, which was successfully used in large-scale Monte Carlo studies of asteroid lander deployment. This model parameterizes impacts with (fixed) material restitution and friction coefficients, but has not been validated for the low-velocity regime of an assembled, nonspherical body. This paper uses an air-bearing setup to perform 2D experiments of a rectangular floating assembly impacting a concrete block with V⩽25 cm/s. The impact velocity, assembly attitude, and block attitude are varied across 2,400 experimental runs of both normal and tangential impacts. Optical tracking is used to extract the pre- and post-impact velocities of the assembly. In a majority of cases, Stronge's model can be fit to the experiments to extract the corresponding restitution and friction coefficients. We find that the coefficients are not fixed with respect to the impact velocity and attitude, but that their variation is seemingly random. In some tangential impact cases, the model even fails to reproduce the observed behavior althogether. This suggests that there may not be a simple way to reconcile Stronge's fixed-material-coefficient model with reality, although it may retain practical use if the coefficients are randomly varied in each impact of a simulation.

Solar electric propulsion is a key enabling technology that has improved the efficiency of space transport. With specific impulses that are typically ten times higher than the chemical counterpart, electric motors allow a considerable saving in propellant mass at the expense of longer times of flight. However, the length of the transfer process and the specific operational needs require to develop a different operational concept for the navigation and orbit control that can be sustained during the different phases of the mission. In this paper, a trade-off is performed among several operational concepts and solutions for multi-revolutions SEP transfers with application to the DESTINY⁺ mission. The GTO-to-Moon low-thrust transfer is first computed in a high-fidelity model with a tangential thrust strategy and later optimized with a five-order Legendre-Gauss-Lobatto collocation method. The impact of eclipses, radiation, thrust outages and misfires, and orbit tracking is analyzed in detailed and included in the transcript optimal problem as algebraic constraints where possible. Numerical results show that the driving factors for the optimal trajectory are related to the operations of the spacecraft rather than the final mass or time of flight.

We present a small-scale mission concept to characterize the permanently shadowed regions of the lunar south pole. MARAUDERS aims to measure in situ for the first time the presence, distribution, and state of volatiles in one permanently shaded crater at a greater resolution than existing orbital measurements using up to 12 deployed impactors. A total of 15 permanently shadowed regions have been characterized as potential landing sites candidates for the probes. The science principle is based on penetrometry, that has proven in the past to be an efficient technique to estimate regolith properties from acceleration profiles. We demonstrate this concept by numerically simulating the surface interaction between our probes and the lunar regolith, thereby demonstrating how deceleration profiles can elucidate information on key regolith properties and help discriminate between two ice-regolith end-members. The preliminary payload design indicates that a good baseline for the impactors would be a spherical shell of 30–40 ​mm in size and ~90 ​g in mass per impactor, including electronics and the communication system. This would sum up to an overall payload of ~1 ​kg contained in a volume of ~15.10⁻⁴ ​m³, which is in agreement with a small-scale payload. Preliminary landing trajectory design enabled the computing of a nominal deployment scenario (with constraint on altitude, ejection velocity and spin rate) that would provide dispersions of the probes from ~250 ​m down to ~20 ​m if deployed from orbit, and down to ~10 ​m if deployed from a carrier lander/rover. Both scenarios will be able to comply with the MARAUDERS’ objectives to assess: (1) the presence (2) the distribution and (3) the surface strength heterogeneity (that can be traced back to the state of volatiles through lab experiments) of water-ice volatiles in permanently shadowed regions at a resolution ​< ​10 ​s ​m via ground-truth measurements. Future work will be dedicated to experimental work to validate the modelling as a proof of concept for the MARAUDERS, as well as the development of the payload.

DESTINY⁺ (Demonstration and Experiment of Space Technology for INterplanetary voYage, Phaethon fLyby and dUSt analysis) is a small-sized high-performance deep space vehicle proposed at ISAS/JAXA. The trajectory design of DESTINY⁺ is divided into several phases. First phase is an orbit injection into an extended elliptical orbit launched by the Epsilon rocket with the additional solid kick motor. Second phase is many revolutions transfer to raise apogee altitude by low thrust propulsion system to the moon orbit nearby. And at third phase, the distant flyby and the swing-by around the moon is designed to give DESTINY⁺ momentum to escape Earth gravitational field. At an interplanetary phase, DESTINY⁺ goes to an Asteroid Phaethon for flyby observation. After the Phaethon flyby, DESTINY⁺ is planned to go back toward Earth for gravity assist and go to another asteroid 2005UD which thought to have split from Phaethon. This paper discusses DESTINY⁺'s low-thrust trajectory design. As for the many revolution transfer phase, the low-thrust trajectory is optimized by the multi-objective optimization using genetic algorithm. In this phase, we minimize the time of flight, the passage of time of radiation belt, the work time of low thrust propulsion system and the maximum eclipse period. After the spacecraft reaches to the moon's orbit, it utilizes the moon swing-by several times to connect to the transfer trajectory for Asteroid Phaethon. From these studies, we can show the feasibility of the mission design of DESTINY⁺,.

Landing on Phobos and bringing samples from its surface would settle the debate on the origin of the Martian moons and support future manned exploration to Mars. To fulfill these scientific objectives, JAXA is planning to send a sample return probe to Phobos by the first half of the next decade, named the Martian Moons eXploration (MMX) mission. In order to explore scientifically interesting regions of Phobos, as well as to support the sampling operations of MMX, a number of Deployable CAMera 5 payloads are proposed to be deployed from quasi-satellite orbits (QSOs) around the Martian moon. This paper explores the feasibility of ballistic deployments from QSOs under realistic dynamical environment and surface constraints in order to guarantee surface settlement within the lifespan of DCAM5. First, we analyze the dynamical environment and escape speeds from Phobos by means of the Circular Hill Problem. Then, the surface coefficient of restitution is estimated by generic impacts onto Phobos regolith via discrete element method simulations. By combining these two analyses, maximum allowable impact velocities for surface settling are calculated and applied to downselect the number of feasible ballistic landings from QSOs. It is found that access to Phobos surface is possible especially along the leading and trailing sides of the Martian moon and in agreement with the engineering requirements of DCAM5.

Small body missions can significantly benefit from deploying small landing systems onto the surface of the visited object. Despite the potential benefit that they may bring, deployments of landers in small body environments may entail significant mission design challenges. This paper thus addresses the potential of ballistic landing opportunities in binary asteroid moons from a mission design perspective, particularly focusing on reliability aspects of the trajectories. Two binaries that were previously identified as target bodies in several missions/proposals, Didymos and 1996 FG3, are considered in this paper. The dynamics near them are modeled by means of the Circular Restricted Three Body Problem (CR3BP), which provides a reasonable representation of a standard binary system. Natural landing trajectories that allow both minimum-velocity local-vertical touchdown and deployment from a safe distance are investigated. Coefficient of restitution values are used as a design parameter to compute the first touchdown speeds that ensure sufficient reliability of landing trajectories. A simple reliability index, which is derived via uncertainty ellipsoid from covariance analysis, is introduced to create a global reliability map across the asteroid surfaces. Assuming 3σ deployment errors on the order of 90 m and 2 cm/s, the results show that ballistic landing operations are likely to be successful for larger binary moons if the deployments target near equatorial regions within longitude range 320 ^o –20°. It has also been shown that the deployments to smaller binary moons may require higher accuracy in navigation and deployment systems in their mothership, and/or closer deployment distances.

This paper aims to assess the feasibility of an interstellar mission to reach the Alpha Centauri star system and delivering scientific data, using current and near-future technology. The mission baseline uses 100 GW of laser power to accelerate a spacecraft of 2750 kg to 5% the speed of light with light sail technology, resulting in a travel time of about a century. This paper explores several aspects of the mission: Possible locations of the laser infrastructure and different sail materials to achieve the required acceleration are discussed. Deceleration using a magnetic sail in the interstellar medium and in the heliosphere of the target star, taking into account mathematical models from Zubrin [6], Freeman [22] and Gros [21], is studied. Potential orbits in the star system are considered for observation and data collection. Finally, a multi-spacecraft mission architecture is presented, as it would allow for the spacecraft to be launched sequentially, thus exploiting the possibility of continuously operating the laser infrastructure.