Orbiting solar reflectors (OSRs) are flat, thin and lightweight reflective structures that are proposed to enhance terrestrial solar energy generation by illuminating large terrestrial solar power plants locally around dawn/dusk and during night hours. The incorporation of OSRs into terrestrial energy systems may offset the daylight-only limitation of terrestrial solar energy. However, the quantity of solar energy delivered to the Earth's surface remains low due to short duration of orbital passes and the low density of the reflected solar power due to large slant ranges. To compensate for these, this paper proposes a constellation of multiple reflectors in low-Earth orbit for scalable enhancement of the quantity of energy delivered. Circular near-polar orbits of 1000 km altitude in the terminator region are considered in a Walker-type constellation for a preliminary analysis. Starting from a simplified approach, the equations of Walker constellations describing the distribution of the reflectors are first modified by introducing a phasing parameter to ensure repeating pass-geometry over solar power farms. This approach allows for a single groundtrack optimisation to define the constellation, which was carried out by a genetic algorithm for single and two reflectors per orbit with an objective function defined as the total quantity of energy delivered per day, to existing and hypothetical solar power projects around the Earth. When full-scale constellations are considered with a number of reflectors, the quantity of solar energy delivered is substantial in the broader context of global terrestrial solar energy generation.

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Many proposals are being made for cleaner, more sustainable forms of energy production. Terrestrial solar photovoltaic farms (SPFs) could potentially deliver large quantities of energy to the grid, although these are limited to daytime use. The output from these SPFs could be enhanced, particularly around dawn and dusk, by the use of orbiting solar reflectors (OSRs) in near-polar orbit. These would reflect an image of the solar disk, or solar image (SI), onto the SPFs to augment their energy output. Pointing requirements are therefore to ensure that reflected sunlight is delivered to the terrestrial SPF, avoiding the losses incurred by an offset of the SI and the SPF itself. The SI would typically be of order 10 km for a reflector in a 1000 km orbit. Given the potentially large size of the reflectors, this presents a challenge for the attitude determination and control system (ADCS) to ensure that the maximum quantity of energy can be delivered to a SPF, typically requiring large control moment gyro actuators. In addition, there exist numerous sources of error in the ADCS which can cause further degradation in the quantity of energy delivered to the SPF. These errors can manifest in the resolution of the various sensors, flexible structural modes, manufacturing inaccuracies, and misalignments due to vibration during launch. This paper will investigate the effects of pointing error sources (PES) on the reflector ADCS and so on the quantity of energy delivered to the SPF. With the application of a PD controller with feedforward compensation, and the set of noise characteristics defined in this paper, numerical simulations will show the typical losses in energy delivered to SPFs of 0.015% when the model accounts for PES in onboard sensors, actuator uncertainty and flexible structural modes.

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Many proposals are being made for cleaner, more sustainable forms of energy production. Terrestrial solar photovoltaic farms (SPFs) could potentially deliver large quantities of energy to the grid, although these are limited to daytime use. The output from these SPFs could be enhanced, particularly around dawn and dusk, by the use of orbiting solar reflectors (OSRs) in near-polar orbit. These would reflect an image of the solar disk, or solar image (SI), onto the SPFs to augment their energy output. Pointing requirements are therefore to ensure that reflected sunlight is delivered to the terrestrial SPF, avoiding the losses incurred by an offset of the SI and the SPF itself. The SI would typically be of order 10 km for a reflector in a 1000 km orbit. Given the potentially large size of the reflectors, this presents a challenge for the attitude determination and control system (ADCS) to ensure that the maximum quantity of energy can be delivered to a SPF, typically requiring large control moment gyro actuators. In addition, there exist numerous sources of error in the ADCS which can cause further degradation in the quantity of energy delivered to the SPF. These errors can manifest in the resolution of the various sensors, flexible structural modes, manufacturing inaccuracies, and misalignments due to vibration during launch. This paper will investigate the effects of pointing error sources on the reflector ADCS and so on the quantity of energy delivered to the SPF. With the application of a PD controller with feedforward compensation, and the set of noise characteristics defined in this paper, numerical simulations will show the typical losses in energy delivered to SPFs of 37.50% when the model accounts for pointing error sources in onboard sensors, actuator uncertainty and flexible structural modes. Improvements in sensor noise characteristics, as well as application of a more robust controller, could improve upon this result.

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Orbiting solar reflectors may be a useful assets to illuminate solar power farms to enhance their utility when direct sunlight is not available. The assessment of their feasibility for a variety of applications requires accurate calculations of how much solar energy can be delivered from a variety of orbits. This paper presents a generic, three-dimensional semi-analytical model that outputs the quantity of solar energy for a given circular orbit and solar power farm position at the beginning of a pass. The model extends previous studies by including new phenomena such as the Earth's oblateness, rotation, shadow on the reflector and orbit around the Sun, in addition to time-dependent geometric and atmospheric losses. These additions provide new analytical insights into the delivery of reflected solar energy delivery and demonstrate the importance of high-fidelity modelling. The strengths of the model are illustrated for a 1000 km altitude Sun-synchronous orbit throughout, as well as a range of other orbits and solar power farms located at different latitudes and longitudes.

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Orbiting reflectors offer the possibility of illuminating large terrestrial solar power plants to enhance their output, particularly at dawn and dusk when their output is low but energy spot prices can be high. While the concept of orbiting solar reflectors has been considered in various forms in the past, there is now a timely overlap of rapidly growing global demand for clean energy services, falling launch costs through reusability and the emergence of in-orbit manufacturing technologies to enable the fabrication of large, ultra-lightweight space structures. This paper provides an end-to-end analysis of a possible minimum initial architecture to deliver such global clean energy services. The analysis will cover orbit selection, attitude control requirements, structural analysis and economical viability, followed by a discussion on regulatory issues, future improvements and further applications.

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Orbiting solar reflectors (OSRs) are flat, thin and lightweight reflective structures that are proposed to enhance terrestrial solar energy generation by illuminating large terrestrial solar power plants locally around dawn/dusk and in night hours. The incorporation of OSRs into terrestrial energy systems may offset the daylight-only limitation of terrestrial solar energy. However, the quantity of solar energy delivered to the Earth's surface remains low due to short duration of orbital passes and the low density of the reflected solar power due to large slant ranges and the projected image of the solar disk. To compensate for these, this paper proposes a constellation of multiple reflectors in low-Earth orbit to enhance the quantity of energy delivered and extend pass durations. Circular near-polar orbits of 1000 km altitude in the terminator region are considered in a Walker-type constellation as a preliminary analysis to avoid eclipses and to include Sun-synchronous orbits. Starting from a simplified approach, equations of Walker constellations describing the distribution of the constellation reflectors is modified to ensure scalable expansion of the constellation by a phasing parameter between the reflectors. Thanks to this approach, a single groundtrack optimisation is sufficient to describe the constellation. This optimisation was carried out for single and two reflectors per orbit, with an objective function defined as the total quantity of energy delivered per day to existing and hypothetical solar power projects around the Earth. When full-scale constellations are considered with 5, 10 and 20 reflectors, the quantity of solar energy delivered linearly scales with the number of reflectors and considerable in the broader context of global solar energy.

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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.

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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.

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Orbiting solar reflectors can be employed to redirect incoming sunlight to provide/reject additional solar energy to/from a planetary body. The concept has been studied in the past for a variety of applications, among which enhancing terrestrial solar power generation, supporting lunar exploration and terraforming Mars are the most prominent. Despite the potential of the concept, previous studies have assumed a perfect reflector, and have only relied on simplified geometric analyses to calculate the quantity of energy delivered with minimal consideration of geometric and physical losses. In this paper, an analytical model is developed for a reflector in a circular orbit to calculate the total energy delivered to a stationary ground-target, such as a solar power farm. A perfectly ideal flat reflector is also assumed to avoid material specific considerations but atmospheric transmission losses, fixed ground-target size and solar panel orientation are included. Case studies demonstrate the significance of high-fidelity modelling, provide new insights into the scalability of the results to test the effectiveness of the concept at different solar system objects, including delivering solar energy at the Earth, Moon and Mars.@en

The many small-body exploration missions that have occurred over the last few decades have shown that small solar system objects are covered with granular material of varying depth. These missions have also observed that granular materials are mobilized from the surfaces at speeds on the order of escape speed by different contributing mechanisms. Those result in various outcomes including escape and reimpact. The latter contributes to further impact-driven evolution. Despite the long history of the research in the field of planetary cratering, low-speed impacts have not been studied extensively under gravity levels relevant to small-bodies. Earth-based low-gravity platforms lack the ability to probe microgravity impact physics for a sufficiently long duration to collect meaningful data from experiments. In order to overcome these challenges, this study uses discrete-element method (DEM) simulations to test low-speed cratering at 5–50 cm/s in granular materials in microgravity. The study first presents a procedure for post-processing the raw simulation data to extract the information relevant to the crater-scaling relationships and demonstrates their applicability for crater sizes, ejecta properties and crater formation time. The implications of the results are discussed in the light of results from recent small-body exploration missions.@en

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.

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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.

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DESTINY+ is a medium-class interplanetary mission, selected by the Japan Aerospace Exploration Agency for potential launch windows in the first half of 2020s. The mission will demonstrate innovative spacecraft subsystem technologies, including a new type of ion engine for future missions. The mission will also collect scientific data through high-speed flyby observations and dust measurements from asteroid (3200) Phaethon and its related body (155140) 2005 UD, to understand their origin and reveal the content of extraterrestrial dust in the context of origin of life. The limited control authority on the spacecraft, the orbits of the target asteroids, and the specific mission requirements pose a challenging task for the trajectory design of DESTINY+. Multiple-target low-thrust optimal trajectories are explored in this paper to fulfill the goals of the DESTINY+ mission. An effective methodology is presented to convert feasible impulsive transfer solutions into low-thrust initial guesses and combine with gravity-assist maneuvers to reveal new high-fidelity optimal trajectories in real ephemeris models. The early mission analysis results demonstrate multitudes of flyby opportunities that provide robustness against programmatic and operational delays in the mission schedule.@en

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.

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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.

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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.

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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.

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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⁺,.

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