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A special issue of Aerospace (ISSN 2226-4310). This special issue belongs to the section "Astronautics & Space Science".

Deadline for manuscript submissions: closed (30 November 2023) | Viewed by 21370

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Dr. Andris Slavinskis

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Space Technology Department, UT Tartu Observatory, Observatooriumi 1, 61602 Tõravere, Tartu Maakond, Estonia
Interests: nanospacecraft; CubeSat; nanosatellite; interplanetary propulsion; electric solar wind sail; space debris; deorbiting; plasma brake; mission design; deep-space missions; spacecraft control; optical imaging; celestial navigation
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Dr. Pekka Janhunen

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Finnish Meteorological Institute, Erik Palménin aukio 1, 00560 Helsinki, Finland
Interests: space plasma physics; sustainable space; electric solar wind sail; plasma brake; propulsion; space settlements
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Dear Colleagues,

Spacecraft size and propulsion are major limiting factors in space mission design. Chemical and electric propulsion require the spacecraft size to be several orders of magnitude larger than CubeSats. The CubeSat Standard in conjunction with the New Space movement has revolutionized the space industry and scientific exploration. CubeSats consist of one or multiple 10×10×10 cm units stacked together in order to achieve the desired mission objectives. With a typical CubeSat mass in the range of 1–10 kg, their propellant storage capabilities are extremely limited if available at all.

Propellantless propulsion systems use an external force to propel the spacecraft, instead of an onboard propellant. This can be photon pressure and solar wind originating in the Sun, as well as magnetic field originating in a planet’s core or atmospheric particles dragging the spacecraft to a lower altitude. We can employ physical light sails to reflect photons and travel the Solar System. A similar drag sail can be used in low Earth orbit (LEO) for orbital debris mitigation with deorbiting. Virtual electromagnetic sails can also be generated: the electric sail deflects solar wind particles using the Coulomb drag force to travel sunward and away from the star, electrodynamic tethers use Lorentz force to increase and lower a satellite’s altitude, and the plasma brake employs the Coulomb drag interaction with the ionosphere for deorbiting. We invite you to submit papers on topics covering CubeSat sails and tethers – fundamental aspects, simulations, designs, optimization, operations, applications in Earth orbit and deep space as well as in-orbit results.

Prof. Dr. Andris Slavinskis
Dr. Pekka Janhunen
Guest Editors

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Keywords

propellantless CubeSat propulsion
lightsails (photon pressure propulsion) and dragsails (atmospheric drag)
electric solar wind sail and plasma brake (Coulomb drag propulsion)
electrodynamics tethers (Lorentz force propulsion)
earth orbit as well as interplanetary CubeSats
orbital debris
on-board orbit and attitude control

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Advances in CubeSat Sails and Tethers (2nd Edition) in Aerospace (1 article)

Published Papers (12 papers)

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Research

29 pages, 7672 KiB

Open AccessArticle

Electric Sail Test Cube–Lunar Nanospacecraft, ESTCube-LuNa: Solar Wind Propulsion Demonstration Mission Concept

by Andris Slavinskis, Mario F. Palos, Janis Dalbins, Pekka Janhunen, Martin Tajmar, Nickolay Ivchenko, Agnes Rohtsalu, Aldo Micciani, Nicola Orsini, Karl Mattias Moor, Sergei Kuzmin, Marcis Bleiders, Marcis Donerblics, Ikechukwu Ofodile, Johan Kütt, Tõnis Eenmäe, Viljo Allik, Jaan Viru, Pätris Halapuu, Katriin Kristmann, Janis Sate, Endija Briede, Marius Anger, Katarina Aas, Gustavs Plonis, Hans Teras, Kristo Allaje, Andris Vaivads, Lorenzo Niccolai, Marco Bassetto, Giovanni Mengali, Petri Toivanen, Iaroslav Iakubivskyi, Mihkel Pajusalu and Antti Tamm Show full author list Hide full author list

Aerospace 2024, 11(3), 230; https://doi.org/10.3390/aerospace11030230 - 14 Mar 2024

Viewed by 1142

Abstract

The electric solar wind sail, or E-sail, is a propellantless interplanetary propulsion system concept. By deflecting solar wind particles off their original course, it can generate a propulsive effect with nothing more than an electric charge. The high-voltage charge is applied to one or multiple centrifugally deployed hair-thin tethers, around which an electrostatic sheath is created. Electron emitters are required to compensate for the electron current gathered by the tether. The electric sail can also be utilised in low Earth orbit, or LEO, when passing through the ionosphere, where it serves as a plasma brake for deorbiting—several missions have been dedicated to LEO demonstration. In this article, we propose the ESTCube-LuNa mission concept and the preliminary cubesat design to be launched into the Moon’s orbit, where the solar wind is uninterrupted, except for the lunar wake and when the Moon is in the Earth’s magnetosphere. This article introduces E-sail demonstration experiments and the preliminary payload design, along with E-sail thrust validation and environment characterisation methods, a cis-lunar cubesat platform solution and an early concept of operations. The proposed lunar nanospacecraft concept is designed without a deep space network, typically used for lunar and deep space operations. Instead, radio telescopes are being repurposed for communications and radio frequency ranging, and celestial optical navigation is developed for on-board orbit determination. Full article

(This article belongs to the Special Issue Advances in CubeSat Sails and Tethers)

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16 pages, 26466 KiB

Open AccessArticle

Robust Flight Tether for In-Orbit Demonstrations of Coulomb Drag Propulsion

by Petri Toivanen, Pekka Janhunen, Jarmo Kivekäs and Meri Mäkelä

Aerospace 2024, 11(1), 62; https://doi.org/10.3390/aerospace11010062 - 09 Jan 2024

Viewed by 944

Abstract

A new method of producing robust multi-wire tethers for Coulomb drag applications was developed. The multi-wire structure required for redundancy against the micrometeoroid flux of the space environment is realised through the method of wire twist bonding traditionally used for chicken wire. In the case of the Coulomb drag tether, the diameter of the individual wires is 50 μm, which introduces the main technological challenge. To manufacture the tether, a manually driven tether machine was designed and built. Two multi-wire tethers for Coulomb drag applications were produced for two in-orbit demonstrations of the FORESAIL-1 and ESTCube-2 CubeSat missions. The flight tethers were both 60 m long as produced, clearly demonstrating beyond the level of proof of concept the applicability of both the method and the manually driven tether machine. Altogether, 6480 twist bonds were produced without a single wire cut. In this paper, the requirements for the tether are listed and justified. The production method is reviewed, and the 4-wire tether produced is evaluated against the requirements. Finally, the test procedures of the tether are described, and on the basis of the results, it is concluded that the tether can tolerate a tension of 14 g without the twist bonds slipping or the tether structure collectively collapsing. Furthermore, the tether can be reeled from the production reel to the flight reel, which simplifies the final integration of the tether reeling system with the Coulomb drag propulsion device. Full article

(This article belongs to the Special Issue Advances in CubeSat Sails and Tethers)

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19 pages, 4481 KiB

Open AccessArticle

Decentralized Differential Aerodynamic Control of Microsatellites Formation with Sunlight Reflectors

by Kirill Chernov, Uliana Monakhova, Yaroslav Mashtakov, Shamil Biktimirov, Dmitry Pritykin and Danil Ivanov

Aerospace 2023, 10(10), 840; https://doi.org/10.3390/aerospace10100840 - 26 Sep 2023

Viewed by 721

Abstract

The paper presents a study of decentralized control for a satellite formation flying mission that uses differential lift and drag to enforce the relative positioning requirements. All spacecraft are equipped with large sunlight reflectors so that, given the appropriate lighting conditions, the formation as a whole can be made visible from the Earth as a configurable pixel image in the sky. The paper analyzes the possibility of achieving a pre-defined lineup of the formation by implementing decentralized aerodynamic-based control through the orientation of sunlight reflectors relative to the incoming airflow. The required relative trajectories are so-called projected circular orbits which ensure the rotation of the image with the orbital period. The choice of the reference trajectory for each satellite is obtained by minimizing the total sum of relative trajectory residuals. The control law is based on the linear-quadratic regulator with the decentralized objective function of reducing the mean deviation of each satellite’s trajectory relative to the other satellites. The accuracy of the required image construction and convergence time depending on the initial conditions and orbit altitude are studied in the paper. Full article

(This article belongs to the Special Issue Advances in CubeSat Sails and Tethers)

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27 pages, 8625 KiB

Open AccessArticle

Electric Sail Mission Expeditor, ESME: Software Architecture and Initial ESTCube Lunar Cubesat E-Sail Experiment Design

by Mario F. Palos, Pekka Janhunen, Petri Toivanen, Martin Tajmar, Iaroslav Iakubivskyi, Aldo Micciani, Nicola Orsini, Johan Kütt, Agnes Rohtsalu, Janis Dalbins, Hans Teras, Kristo Allaje, Mihkel Pajusalu, Lorenzo Niccolai, Marco Bassetto, Giovanni Mengali, Alessandro A. Quarta, Nickolay Ivchenko, Joan Stude, Andris Vaivads, Antti Tamm and Andris Slavinskis Show full author list Hide full author list

Aerospace 2023, 10(8), 694; https://doi.org/10.3390/aerospace10080694 - 05 Aug 2023

Cited by 3 | Viewed by 1316

Abstract

The electric solar wind sail, or E-sail, is a novel deep space propulsion concept which has not been demonstrated in space yet. While the solar wind is the authentic operational environment of the electric sail, its fundamentals can be demonstrated in the ionosphere where the E-sail can be used as a plasma brake for deorbiting. Two missions to be launched in 2023, Foresail-1p and ESTCube-2, will attempt to demonstrate Coulomb drag propulsion (an umbrella term for the E-sail and plasma brake) in low Earth orbit. This paper presents the next step of bringing the E-sail to deep space—we provide the initial modelling and trajectory analysis of demonstrating the E-sail in solar wind. The preliminary analysis assumes a six-unit cubesat being inserted in the lunar orbit where it deploys several hundred meters of the E-sail tether and charges the tether at 10–20 kV. The spacecraft will experience acceleration due to the solar wind particles being deflected by the electrostatic sheath around the charged tether. The paper includes two new concepts: the software architecture of a new mission design tool, the Electric Sail Mission Expeditor (ESME), and the initial E-sail experiment design for the lunar orbit. Our solar-wind simulation places the Electric Sail Test Cube (ESTCube) lunar cubesat with the E-sail tether in average solar wind conditions and we estimate a force of

1.51 \times 10^{- 4}

N produced by the Coulomb drag on a 2 km tether charged to 20 kV. Our trajectory analysis takes the 15 kg cubesat from the lunar back to the Earth orbit in under three years assuming a 2 km long tether and 20 kV. The results of this paper are used to set scientific requirements for the conceptional ESTCube lunar nanospacecraft mission design to be published subsequently in the Special Issue “Advances in CubeSat Sails and Tethers”. Full article

(This article belongs to the Special Issue Advances in CubeSat Sails and Tethers)

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21 pages, 2806 KiB

Open AccessArticle

LightSail 2 Solar Sail Control and Orbit Evolution

by Justin R. Mansell, John M. Bellardo, Bruce Betts, Barbara Plante and David A. Spencer

Aerospace 2023, 10(7), 579; https://doi.org/10.3390/aerospace10070579 - 22 Jun 2023

Cited by 4 | Viewed by 2303

Abstract

The propellantless thrust of solar sails makes them capable of entirely new classes of missions compared to conventional or electric engines. Initiated in 2010, the Planetary Society’s LightSail program sought to advance solar sail technology with the flights of LightSail 1 and 2. From launch in 2019 to deorbit in late 2022, LightSail 2 demonstrated the first controlled solar sailing in Earth’s orbit using a CubeSat. By adjusting the orientation of the sail relative to the sun twice per orbit, LightSail 2 controlled solar radiation pressure on the sail to offset losses in orbital energy from atmospheric drag. Previous papers analyzed early mission results to show the effect this had on reducing the spacecraft’s orbital decay rate. Subsequent refinements to the spacecraft’s attitude control made throughout the mission eventually enabled it to achieve sustained net increases in orbital energy. This paper presents an analysis of the orbit changes and attitude control performance over the entire mission. Methods of assessing and improving the sail control are described. Activities and attitude behavior during the final deorbit phase are also analyzed, with results relevant to future drag sails as well as solar sail missions. Full article

(This article belongs to the Special Issue Advances in CubeSat Sails and Tethers)

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22 pages, 6955 KiB

Open AccessArticle

Feasibility Study of the Bare-Photovoltaic-Tether Concept: Prototypes and Experimental Performance Evaluation of the Photovoltaic Tether Segment

by Leo Peiffer, Christian Perfler and Martin Tajmar

Aerospace 2023, 10(4), 386; https://doi.org/10.3390/aerospace10040386 - 21 Apr 2023

Viewed by 1508

Abstract

Consumable-free electron emitters are presently not feasible for autonomous tether-based deorbit devices such as E.T.PACK due to their power requirement. The bare-photovoltaic-tether (BPT) concept combines the bare tether electron collection with a tether segment, coated with thin film Copper Indium Gallium Selenide solar cells to harvest additional power for the cathodic contact, potentially enabling propellant-less operation. This thesis presents the first prototype of the photovoltaic tether segment, its architecture, its electrical characteristics, major challenges of the system and possible solutions. Photovoltaic tether segments of up to 3 m in length were manufactured, consisting of parallelized submodules of 25 cm in length. Due to space limitations, only the I-V-characteristics of these submodules were measured under a self-built Class BCA LED Solar-Simulator inside a vacuum chamber and at varying temperatures between −100 °C and 100 °C. In addition, the suitability of the concept for a low Earth orbit environment was assessed by performing atomic oxygen exposure tests using a microwave-based low pressure plasma atomic oxygen source. Based on the experimental data, a model is provided for predicting the performance of the photovoltaic segment in orbit, highlighting the main problems of the BPT: temperature, orientation and partial shading. Full article

(This article belongs to the Special Issue Advances in CubeSat Sails and Tethers)

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13 pages, 2462 KiB

Open AccessArticle

E-Sail Option for Plunging a Spacecraft into the Sun’s Atmosphere

by Giovanni Mengali and Alessandro A. Quarta

Aerospace 2023, 10(4), 340; https://doi.org/10.3390/aerospace10040340 - 01 Apr 2023

Cited by 2 | Viewed by 1041

Abstract

A close observation of the near-Sun region, with in situ measurements, requires that a scientific probe be placed in a heliocentric orbit with a perihelion distance on the order of a few solar radii only. This is the solution adopted by the Parker Solar Probe (PSP), whose mission design uses a very complex transfer trajectory with seven Venus gravity assists to reach a perihelion radius of roughly

9.9

solar radii in about seven years. This paper aims to discuss the capability of an Electric Solar-Wind Sail (E-sail), i.e., a propellantless propulsion system that exploits the solar wind as a deep-space thrust source using a grid of long and artificially charged tethers, to drive a scientific probe toward a heliocentric orbit with characteristics similar to that considered during the initial design of the PSP mission. The two-dimensional trajectory analysis of an E-sail-based spacecraft is performed in an optimal framework, by considering the physical constraints induced by the thermal loads acting on the propellantless propulsion system when the spacecraft approaches the inner Sun regions. This means that, during the transfer trajectory, the E-sail-based spacecraft must avoid a spherical region around the Sun whose radius depends on the mechanical characteristics of the charged tethers. The paper shows that feasible solutions, in terms of optimal transfer trajectories, are possible even when a medium-performance E-sail is considered in the spacecraft design. In that context, the obtained trajectory can drive a scientific probe on the target (high elliptic) orbit in less than two years, without the use of any intermediate flyby maneuver. Full article

(This article belongs to the Special Issue Advances in CubeSat Sails and Tethers)

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16 pages, 1613 KiB

Open AccessArticle

E-Sail Optimal Trajectories to Heliostationary Points

by Alessandro A. Quarta and Giovanni Mengali

Aerospace 2023, 10(2), 194; https://doi.org/10.3390/aerospace10020194 - 17 Feb 2023

Cited by 1 | Viewed by 1338

Abstract

The aim of this paper is to investigate the performance of a robotic spacecraft, whose primary propulsion system is an electric solar wind sail (E-sail), in a mission to a heliostationary point (HP)—that is, a static equilibrium point in a heliocentric and inertial reference frame. A spacecraft placed at a given HP with zero inertial velocity maintains that heliocentric position provided the on-board thrust is able to counterbalance the Sun’s gravitational force. Due to the finite amount of storable propellant mass, a prolonged mission toward an HP may be considered as a typical application of a propellantless propulsion system. In this respect, previous research has been concentrated on the capability of high-performance (photonic) solar sails to reach and maintain such a static equilibrium condition. However, in the case of a solar-sail-based spacecraft, an HP mission requires a sail design with propulsive characteristics that are well beyond the capability of current or near-future technology. This paper shows that a medium-performance E-sail is able to offer a viable alternative to the use of photonic solar sails. To that end, we discuss a typical HP mission from an optimal viewpoint, by looking for the minimum time trajectory necessary for a spacecraft to reach a given HP. In particular, both two- and three-dimensional scenarios are considered, and the time-optimal mission performance is analyzed parametrically as a function of the HP heliocentric position. The paper also illustrates a potential mission application involving the observation of the Sun’s poles from such a static inertial position. Full article

(This article belongs to the Special Issue Advances in CubeSat Sails and Tethers)

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11 pages, 6812 KiB

Open AccessArticle

Optimal Earth Gravity-Assist Maneuvers with an Electric Solar Wind Sail

by Lorenzo Niccolai, Marco Bassetto, Alessandro A. Quarta and Giovanni Mengali

Aerospace 2022, 9(11), 717; https://doi.org/10.3390/aerospace9110717 - 14 Nov 2022

Viewed by 1725

Abstract

Propellantless propulsive systems such as Electric Solar Wind Sails are capable of accelerating a deep-space probe, only requiring a small amount of propellant for attitude and spin-rate control. However, the generated thrust magnitude is usually small when compared with the local Sun’s gravitational attraction. Therefore, the total velocity change necessary for the mission is often obtained at the expense of long flight times. A possible strategy to overcome this issue is offered by an Earth gravity-assist maneuver, in which a spacecraft departs from the Earth’s sphere of influence, moves in the interplanetary space, and then re-encounters the Earth with an increased hyperbolic excess velocity with respect to the starting planet. An Electric Solar Wind Sail could effectively drive the spacecraft in the interplanetary space to perform such a particular maneuver, taking advantage of an augmented thrust magnitude in the vicinity of the Sun due to the increased solar wind ion density. This work analyzes Earth gravity-assist maneuvers performed with an Electric Solar Wind Sail based probe within an optimal framework, in which the final hyperbolic excess velocity with respect to the Earth is maximized for a given interplanetary flight time. Numerical simulations highlight the effectiveness of this maneuver in obtaining a final heliocentric orbit with high energy. Full article

(This article belongs to the Special Issue Advances in CubeSat Sails and Tethers)

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14 pages, 985 KiB

Open AccessArticle

Trajectory Approximation of a Coulomb Drag-Based Deorbiting

by Lorenzo Niccolai, Marco Bassetto, Alessandro A. Quarta and Giovanni Mengali

Aerospace 2022, 9(11), 680; https://doi.org/10.3390/aerospace9110680 - 02 Nov 2022

Cited by 2 | Viewed by 1652

Abstract

The presence of a number of space debris in low Earth orbits poses a serious threat for current spacecraft operations and future space missions. To mitigate this critical problem, international guidelines suggest that an artificial satellite should decay (or be transferred to a graveyard orbit) within a time interval of 25 years after the end of its operative life. To that end, in recent years deorbiting technologies are acquiring an increasing importance both in terms of academic research and industrial efforts. In this context, the plasma brake concept may represent a promising and fascinating innovation. The plasma brake is a propellantless device, whose working principle consists of generating an electrostatic Coulomb drag between the planet’s ionosphere ions and a charged tether deployed from a satellite in a low Earth orbit. This paper discusses an analytical method to approximate the deorbiting trajectory of a small satellite equipped with a plasma brake device. In particular, the proposed approach allows the deorbiting time to be estimated through an analytical equation as a function of the design characteristics of the plasma brake and of the satellite initial orbital elements. Full article

(This article belongs to the Special Issue Advances in CubeSat Sails and Tethers)

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19 pages, 1412 KiB

Open AccessArticle

Optimal Circle-to-Ellipse Orbit Transfer for Sun-Facing E-Sail

by Alessandro A. Quarta, Giovanni Mengali, Marco Bassetto and Lorenzo Niccolai

Aerospace 2022, 9(11), 671; https://doi.org/10.3390/aerospace9110671 - 29 Oct 2022

Cited by 4 | Viewed by 1657

Abstract

The transfer between two coplanar Keplerian orbits of a spacecraft with a continuous-thrust propulsion system is a classical problem of astrodynamics, in which a numerical procedure is usually employed to find the transfer trajectory that optimizes (i.e., maximizes or minimizes) a given performance index such as, for example, the delivered payload mass, the propellant mass, the total flight time, or a suitable combination of them. In the last decade, this class of problem has been thoroughly analyzed in the context of heliocentric mission scenarios of a spacecraft equipped with an Electric Solar Wind Sail as primary propulsion system. The aim of this paper is to further extend the existing related literature by analyzing the optimal transfer of an Electric Solar Wind Sail-based spacecraft with a Sun-facing attitude, a particular configuration in which the sail nominal plane is perpendicular to the Sun-spacecraft (i.e., radial) direction, so that the propulsion system is able to produce its maximum propulsive acceleration magnitude. The problem consists in transferring the spacecraft, which initially traces a heliocentric circular orbit, into an elliptic coplanar orbit of given eccentricity with a minimum-time trajectory. Using a classical indirect approach for trajectory optimization, the paper shows that a simplified version of the optimal control problem can be obtained by enforcing the typical transfer constraints. The numerical simulations show that the proposed approach is able to quantify the transfer performance in a parametric and general form, with a simple and efficient algorithm. Full article

(This article belongs to the Special Issue Advances in CubeSat Sails and Tethers)

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13 pages, 1687 KiB

Open AccessArticle

Rapid Evaluation of the Decay Time of a Plasma Brake-Based CubeSat

by Marco Bassetto, Lorenzo Niccolai, Alessandro A. Quarta and Giovanni Mengali

Aerospace 2022, 9(11), 636; https://doi.org/10.3390/aerospace9110636 - 23 Oct 2022

Cited by 1 | Viewed by 1809

Abstract

The plasma brake is a propellantless device conceived for de-orbiting purposes. It consists of an electrically charged thin tether that generates a Coulomb drag by interacting with the ionosphere. In essence, a plasma brake may be used to decelerate an out-of-service satellite and to ensure its atmospheric re-entry within the time limits established by the Inter-Agency Space Debris Coordination Committee. Moreover, since it only needs a small amount of electric power to work properly, the plasma brake is one of the most cost-effective systems for space debris mitigation. This paper exploits a recent plasma brake acceleration model to construct an iterative algorithm for the rapid evaluation of the decay time of a plasma-braked CubeSat, which initially traced a circular low Earth orbit. The altitude loss at the end of each iterative step was calculated using the linearized Hill–Clohessy–Wiltshire equations. It showed that the proposed algorithm, which was validated by comparing the approximate solution with the results from numerically integrating the nonlinear equations of motion, reduced computational time by up to four orders of magnitude with negligible errors in CubeSat position. Full article

(This article belongs to the Special Issue Advances in CubeSat Sails and Tethers)

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