•The formation flying of two E-sails in elliptic, heliocentric, displaced orbits is analyzed.•The closed-form solution to the relative motion is parameterized in terms of differential displaced ...orbital elements.•The approximate bounds of relative distances are analytically calculated for near-circular orbits.
We present a geometrical methodology for analyzing the formation flying of electric solar wind sail based spacecraft that operate in heliocentric, elliptic, displaced orbits. The spacecraft orbit is maintained by adjusting its propulsive acceleration modulus, whose value is estimated using a thrust model that takes into account a variation of the propulsive performance with the sail attitude. The properties of the relative motion of the spacecraft are studied in detail and a geometrical solution is obtained in terms of relative displaced orbital elements, assumed to be small quantities. In particular, for the small eccentricity case (i.e. for a near-circular displaced orbit), the bounds characterized by the extreme values of relative distances are analytically calculated, thus providing an useful mathematical tool for preliminary design of the spacecraft formation structure.
The goal of this paper is to analyze the optimal transfer towards the periodic comet 29P/Schwassmann–Wachmann 1 of a solar sail-based spacecraft. This periodic and active comet is an interesting and ...still unexplored small body that has been regarded as an object of the Centaurs group. In this work, a classical (heliocentric) orbit-to-orbit transfer is studied from an optimal viewpoint, by finding the spacecraft trajectories that minimize the flight time for a given value of the solar sail characteristic acceleration, that is, the typical performance parameter of a photonic sail. In particular, the optimal Earth–comet transfer is studied both in a typical three-dimensional mission scenario and with a simplified two-dimensional approach, whose aim is to rapidly obtain an accurate estimation of the minimum flight time with a reduced computation cost. The numerical simulations illustrate the mission performance, in terms of the characteristics of the rapid transfer trajectory, as a function of the typical propulsive parameter and the solar sail thrust model.
Observing the universe in the Ultra-Long Wavelength (ULW) regime has been called the ‘last frontier in astronomy’—real imaging capabilities here are yet to be achieved. Obtaining an image of the sky ...in this frequency band can be done by employing a swarm of satellites that together act as an interferometer and collect the required imaging information pieces throughout the course of their operational life. Meeting the mission objective is challenging for such a swarm, since this imposes restrictions on the operational environment and the relative position and velocity vectors between the swarm elements. This work proposes an orbit solution in a Heliocentric Earth-Leading Orbit (HELO) for an autonomous CubeSat swarm with chemical thrusters. A distributed formation flying algorithm is used to aid the collection of the required imaging information pieces. Furthermore, the estimated total mission launch mass is reduced by optimising cost functions and finding favourable position and velocity at start of operational life, as well as by finding favourable thrust manoeuvre patterns. The results show that the mission objective—obtaining a 3D map of the Universe in ULW—can be achieved with 68 6U spacecraft (S/C). Moreover, the swarm can remain in a Radio Frequency Interference (RFI) quiet zone of >5 × 106 km, whilst not drifting further than ~ 6.6 × 106 km from Earth for an operational life of one year.
•Swarm orbit design based on ultra-long wavelength radio interferometry requirements.•Distributed controller restricted to account for hardware characteristics.•Optimisation of initial state reduces maximum required telemetry range.•Preliminary optimisation of orbital manoeuvres reduces total required swarm size.
Among the most central of Tolkien's myths is the creation of the Sun and Moon as the last fruit and flower of the Two Trees of Valinor. The death of the Trees is central in a long chain of events ...that directly leads to the later battles, kin slayings, and geological upheavals in Middle-earth. It is therefore curious that during the writing of The Lord of the Rings (and continuing into the later 1950s and 1960s), Tolkien began second-guessing himself, and became concerned with what he called "the astronomically absurd business of the making of the Sun and Moon." Beginning with the experimental 1948 "Round World" cosmology of the Ainulindale C* the elder Tolkien explores what his son terms a "radical transformation of the astronomical myth," changes that appear jarring to his son's sensibilities concerning what his father came to call a "primitive" mythology but Christopher defends as "in conception beautiful." As the cosmological writings become further removed from the medievalist geocentric worldview reflected in writings Christopher (himself a medieval scholar) had been carefully collecting and editing for nearly two decades, his commentary seems severely curtailed, mainly limited to philology and drawing a few cursory connections to similar passages within the same volume.
We present the general concept of a telescope with optics and detectors mounted on two separate spacecrafts, in orbit around the telescope’s target (scopocentric or target-centric orbit), and using ...propulsion to maintain the Target-Optics-Detector alignment and Optics-Detector distance. Specifically, we study the case of such a telescope with the Sun as the target, orbiting at
∼
1 AU. We present a simple differential acceleration budget for maintaining Target-Optics-Detector alignment and Optics-Detector distance, backed by simulations of the orbital dynamics, including solar radiation pressure and influence of the planets. Of prime interest are heliocentric orbits (such as Earth-trailing/leading orbits or Distant Retrograde Orbits), where thrust requirement to maintain formation is primarily in a single direction (either sunward or anti-sunward), can be quite minuscule (a few m/s/year), and preferably met by constant-thrust engines such as solar electric propulsion or even by solar sailing via simple extendable and/or orientable flaps or rudders.
A preliminary investigation of the early warning of solar storms caused by Coronal Mass Ejection has been carried out. A long warning time could be obtained with a sailcraft synchronous with the ...Earth-Moon barycenter, and stationed well below the L1 point. In this paper, the theory of heliocentric synchronous sailcraft is set up, its perturbed orbit is analyzed, and a potential solution capable of providing an annual synchrony is carried out. A simple analysis of the response from a low-mass electrochromic actuator for the realization of station-keeping attitude maneuvers is put forwards, and an example of propellantless re-orientation maneuver is studied.
•This paper regards a solution to geomagnetic storms caused by Coronal Mass Ejection.•An early-warning mission concept by photon sailcraft is introduced.•The sailcraft is not constrained around the Sun-Earth L1 point.•Heliocentric synchrony-on-average is computed via small attitude maneuvers.•Sail maneuver propellantless electro-chromic devices are preliminarily investigated.
Solar sails are propellantless propulsion systems that extract momentum from solar radiation pressure. They consist of a large ultrathin membrane, typically aluminized, that reflects incident photons ...from the Sun to generate thrust for space navigation. The purpose of this study is to investigate the optimal performance of a solar sail-based spacecraft in performing two-dimensional heliocentric transfers to inertial points on the ecliptic that lie within an assigned annular region centered in the Sun. Similar to ESA’s Comet Interceptor mission, this type of transfer concept could prove useful for intercepting a potential celestial body, such as a long-period comet, that is passing close to Earth’s orbit. Specifically, it is assumed that the solar sail transfer occurs entirely in the ecliptic plane and, in analogy with recent studies, the flyby points explored are between 0.85au and 1.35au from the Sun. The heliocentric dynamics of the solar sail is described using the classical two-body model, assuming the spacecraft starts from Earth orbit (assumed circular), and an ideal force model to express the sail thrust vector. Finally, no constraint is imposed on the arrival velocity at flyby. Numerical simulation results show that solar sails are an attractive option to realize these specific heliocentric transfers.
In astrodynamics, orbit cranking is usually referred to as an interplanetary transfer strategy that exploits multiple gravity-assist maneuvers to change both the inclination and eccentricity of the ...spacecraft osculating orbit without changing the specific mechanical energy, that is, the semimajor axis. In the context of a solar sail-based mission, however, the concept of orbit cranking is typically referred to as a suitable guidance law that is able to (optimally) change the orbital inclination of a circular orbit of an assigned radius in a general heliocentric three-dimensional scenario. In fact, varying the orbital inclination is a challenging maneuver from the point of view of the velocity change, so orbit cranking is an interesting mission application for a propellantless propulsion system. The aim of this paper is to analyze the performance of a spacecraft equipped with an Electric Solar Wind Sail in a cranking maneuver of a heliocentric circular orbit. The maneuver performance is calculated in an optimal framework considering spacecraft dynamics described by modified equinoctial orbital elements. In this context, the paper presents an analytical version of the three-dimensional optimal guidance laws obtained by using the classical Pontryagin’s maximum principle. The set of (analytical) optimal control laws is a new contribution to the Electric Solar Wind Sail-related literature.
The problem of joint optimization of the trajectory of a spacecraft with an electric propulsion system and for the main parameters of electric propulsion and power supply systems is considered. It is ...well known that for every space transportation operation there is an optimal value of specific impulse of electric propulsion corresponding to the minimum total mass of the system, the power supply system ensuring electric propulsion operation, and the propellant. It is easy to show that there is also an optimal value of electric power of an electric propulsion system, associated with the growth of the required characteristic velocity with the thrust decrease. Optimal specific impulse and electric power can be found only by joint optimization of the trajectory and design parameters of electric propulsion. A simple spacecraft mass budget model and the maximum principle are used for optimization. The necessary optimality conditions for the specific impulse and electric power of electric propulsion system are derived. Numerical examples of the joint optimization of interplanetary trajectory, electric power and specific impulse of electric propulsion systems are presented.
The Solar Wind Ion Focusing Thruster (SWIFT) is a highly-innovative propellantless propulsion concept, recently proposed by Gemmer and Mazzoleni. In its nominal configuration, a SWIFT consists of a ...conically-shaped mesh of positively-charged conducting tethers, with its vertex linked to the spacecraft and its axis oriented towards the Sun. The SWIFT collects and filters the solar wind plasma and suitably directs the positive ions, which are then accelerated by an ion thruster. Such a device is theoretically able to generate a deep-space propulsive acceleration that comes, in part, from the solar wind dynamic pressure impinging on the conical grid and, in part, from the positive ion beam. In particular, the orientation of the ion beam may be chosen in such a way as to set the resultant propulsive acceleration and steer the spacecraft. The aim of this paper is to analyze the performance of a SWIFT-propelled spacecraft in an orbit-to-orbit two-dimensional interplanetary transfer. To that end, some mission scenarios are studied, in an optimal framework, by minimizing the total flight time necessary for the spacecraft to complete the transfer as a function of the propulsion system performance parameters. Numerical simulations are used to compare the optimal flight times calculated in simplified Earth–Venus and Earth–Mars transfers with those obtained by considering other propellantless propulsion systems.
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