No full-text available
To read the full-text of this research,
you can request a copy directly from the author.
Coulomb drag devices: electric solar wind sail propulsion and ionospheric deorbiting
Abstract
A charged tether or wire experiences Coulomb drag when inserted into flowing
plasma. In the solar wind the Coulomb drag can be utilised as efficient
propellantless interplanetary propulsion as the electric solar wind sail
(electric sail, E-sail). In low Earth orbit (LEO) the same plasma physical
effect can be utilised for efficient low-thrust deorbiting of space debris
objects (the plasma brake). The E-sail is rotationally stabilised while the
deorbiting Coulomb drag devices
According to numerical estimates, Coulomb drag devices have very promising
performance figures, both for interplanetary propulsion and for deorbiting in
LEO. Much of the technology is common to both applications. E-sail technology
development was carried out in ESAIL FP7 project (2011-2013) which achieved TRL
4-5 for key hardware components that can enable 1 N class interplanetary E-sail
weighing less than 200 kg. The thrust of the E-sail scales as inverse solar
distance and its power consumption (nominally 700 W/N at 1 au) scales as the
inverse distance squared. As part of the ESAIL project, a continuous 1 km
sample of E-sail tether was produced by an automatic and scalable "tether
factory". The manufacturing method uses ultrasonic wire to wire bonding which
was developed from ordinary wire to plate bonding for the E-sail purpose. Also
a "Remote Unit" device which takes care of deployment and spin rate control was
prototyped and successfully environmentally tested. Our Remote Unit prototype
is operable in the solar distance range of 0.9-4 au.
The 1-U CubeSat ESTCube-1 was launched in May 2013 and it will try to measure
the Coulomb drag acting on a 10 m long tether in LEO when charged to 500 V
positive or negative. A more advanced version of the experiment with 100 m
tether is under preparation and will be launched in 2015 with the Aalto-1 3-U
CubeSat to polar LEO.
... At a given time instant t 0 0, when the spacecraft altitude is h(t 0 ) h 0 (i.e., the orbital radius is r(t 0 ) r 0 = h 0 + R ⊕ , R ⊕ being the Earth's mean radius), and the true anomaly is ν(t 0 ) ν 0 , the PB conducting tether is unreeled and charged by a suitable electric voltage source. Note that if the tether polarity is negative, the spacecraft itself can act as an electric power supply, since it acquires a negative charge due to the high thermal mobility of the electrons, thus removing the necessity of a power source [21,35]. In the rest of the work, a negative voltage is accordingly assumed. ...
... The mathematical method used for the trajectory analysis must be completed with a thrust model for the PB-induced drag. According to Ref. [35], the magnitude of the Coulomb drag D c generated by a PB conducting tether can be written as where L t is the tether length, m i is the ions mass, n is the plasma bulk number density, v is the relative velocity of the ions relative to the spacecraft, el is the elementary charge, and 0 is the vacuum permittivity. In Equation (30), the auxiliary voltage V a is defined as ...
... where V t is the tether voltage, which is assumed negative according to Refs. [21,35]. ...
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.
... The plasma brake [1,2] is an innovative technology capable of supplying a propulsive acceleration to a spacecraft on a low Earth orbit (LEO) without any propellant consumption, by exploiting the (electrostatic) Coulomb collisions between a long space tether and the charged particles in a plasma stream. In a typical configuration [2,3,4] a single charged tether deployed by a spacecraft, see Fig. 1, interacts with the ionized upper stages of Earth's atmosphere (ionosphere), and provides a decelerating thrust (Coulomb drag) orthogonal to the tether line. The idea of plasma brake is a consequence of the Electric Solar Wind Sail (E-sail) propulsive concept, which can be traced back to 2004 [5]. ...
... Previous works [1,3,24] on plasma brake system analysis have shown that a negatively-charged tether is more convenient compared to a positively-charged one in terms of design simplicity, for a LEO mission scenario. In fact, in the positive polarity case, the plasma brake system requires a voltage source and an electron gun to maintain the necessary voltage by expelling the accumulated electrons. ...
... According to Refs. [3,22], the thrust per unit length dl of a single negatively-charged tether can be expressed as ...
Plasma brake is an innovative propellantless propulsion system concept that exploits the Coulomb collisions between a charged tether and the ions in the surrounding environment (typically, the ionosphere) to generate an electrostatic force orthogonal to the tether direction. Previous studies on the plasma brake effect have emphasized the existence of a number of different parameters necessary to obtain an accurate description of the propulsive acceleration from a physical viewpoint. The aim of this work is to discuss an analytical model capable of estimating, with the accuracy required by a preliminary mission analysis, the performance of a spacecraft equipped with a plasma brake in a (near-circular) low Earth orbit. The simplified mathematical model is first validated through numerical simulations, and is then used to evaluate the plasma brake performance in some typical mission scenarios, in order to quantify the influence of the system parameters on the mission performance index.
... However, the solar activity level influences the plasma density, the presence of atomic oxygen, and the particle temperature, thus affecting the Coulomb drag generation through the terms n 0 , m i , and T i ; see Eq. (52). In fact, the effects of the solar activity on the drag expression were evaluated by Janhunen [149,150], who estimated that the ratio of the drag corresponding to a solar maximum to the that generated at a solar minimum amounts to about 3.5. Although this difference is relevant and must be considered when estimating the decay profile, the deorbiting times provided by Refs. ...
The Electric Solar Wind Sail (E-sail) is an innovative propellantless propulsion system conceived by Pekka Janhunen in 2004 for use in interplanetary space. An E-sail consists of a network of electrically charged tethers maintained at a high voltage level by an electron emitter. The electrostatic field surrounding the E-sail extracts momentum from the incoming solar wind ions, thus giving rise to the generation of a continuous thrust. In a geocentric context, the same physical principle is also exploited by the plasma brake, a promising option for reducing the decay time of satellites in low Earth orbits after the end of their operational life. This paper discusses the scientific advances of both E-sail and plasma brake concepts from their first design to the current state of the art. A general description of the E-sail architecture is first presented with particular emphasis on the proposed tether deployment mechanisms and thermo-structural analyses that have been carried out over the recent years. The working principle of an E-sail is then illustrated and the evolution of the thrust and torque vector models is retraced to emphasize the subsequent refinements that these models have encountered. The dynamic behavior of an E-sail is also analyzed by illustrating the mathematical tools that have been proposed and developed for both orbital dynamics and attitude control. A particular effort is devoted to reviewing the numerous mission scenarios that have been studied to date. In fact, the extensive literature about E-sail-based mission scenarios demonstrates the versatility of such an innovative propulsion system in an interplanetary framework. Credit is given to the very recent studies on environmental uncertainties, which highlight the importance of using suitable control strategies for the compensation of solar wind fluctuations. Finally, the applications of the plasma brake are thoroughly reviewed.
... The Electric Solar Wind Sail (E-sail) is a recent, propellantless, propulsion system concept [1,2,3] that consists of a number of thin, long, and conducting tethers, which are deployed and stretched out using the centrifugal force obtained by spinning the spacecraft around its symmetry axis [4]. The tethers, which are kept at a high positive potential by an onboard electron gun, experience Coulomb drag [5] by interacting with the solar wind plasma stream and generate a propulsive thrust [6,7]. The E-sail is theoretically able to allow a spacecraft to reach and maintain highly non-Keplerian orbits, which are difficult, or even impossible, to be generated with more conventional propulsion systems such as chemical or electrical thrusters [8]. ...
This paper analyzes the performance of an Electric Solar Wind Sail for generating and maintaining an elliptic, heliocentric, displaced non-Keplerian orbit. In this sense, this paper extends and completes recent studies regarding the performances of an Electric Solar Wind Sail that covers a circular, heliocentric, displaced orbit of given characteristics. The paper presents the general equations that describe the elliptic orbit maintenance in terms of both spacecraft attitude and performance requirements, when a refined thrust model (recently proposed for the preliminary mission design) is taken into account. In particular, the paper also discusses some practical applications on particular mission scenarios in which an analytic solution of the governing equations has been found.
... Part of the material in this paper is taken from the Space Propulsion 2014 proceedings paper. 9 ...
A voltage-biased metallic tether creates an electrostatic potential structure around itself. If put into flowing plasma, the electrostatic field deflects the trajectories of plasma ions. The resulting momentum transfer from plasma to the tether can be used for propulsive purposes. The electric solar wind sail (E-sail) is one such application. The E-sail has a set of centrifugally stabilised positively biased tethers in the solar wind. It enables propellantless propulsion in the solar wind, much in the same way as the photonic solar sail, but typically with order of magnitude higher thrust per propulsion system mass. The E-sail’s high efficiency stems from the fact that the tether’s electrostatic field penetrates a significant distance (some 100-200 m distance at 1 au) into the solar wind plasma while the tether that creates the electric field can be made very thin. As a result, the E-sail’s virtual sail area can be millions of times larger than the physical area of the wires of which the tether is made. Thus, even though the solar wind’s dynamic pressure is 5000 times less than the photonic pressure, the E-sail can achieve high thrust per mass. If using 20 kV voltage, the achieved thrust is some 0.5 mN/km while a kilometre of tether weighs about 10 grams.
Even though E-sail’s thrust source - the solar wind - is highly variable, the E-sail’s thrust varies much less, due to some nontrivial plasma feedback mechanisms. Furthermore, because the thrust is independently controllable in direction (by inclining the sail) and magnitude (by changing the voltage), the E-sail is accurately navigable and also able to tack sunward.
The E-sail works everywhere in the solar system except inside Earth’s magnetosphere, because inside the magnetosphere there is no solar wind. However, in low Earth orbit (LEO), a closely related device called the plasma brake can be used for satellite orbit lowering and deorbiting. The plasma brake is a negatively biased Coulomb drag tether with moderate voltage. It uses the relative velocity between the orbiting satellite and nearly stationary ionospheric plasma to provide controllable orbit-lowering drag. As the E-sail, the plasma brake is also very efficient in terms of thrust per mass. Another important benefit of the plasma brake is that the tether is so thin that it is safe to other space assets in case of accidental collision.
... So far, the soundness of the E-sail concept has been checked either by means of laboratory tests, or through accurate plasmadynamic simulations [23][24][25]. A 100 m long E-sail tether is scheduled to fly onboard the Aalto-1 CubeSat, which is planned to be launched into a LEO orbit in late spring 2016 [26,27]. ...
The Electric Solar Wind Sail is an advanced propulsion system concept that, similar to the more conventional solar sail, is able to generate a propulsive thrust without any propellant. The main performances of such a propulsion system have been studied in different mission scenarios and are reported in the literature. However, the analyses available so far are based on a simplified thrust model that neglects the effect of the spacecraft attitude on both the thrust modulus and its direction. The recent availability of a refined thrust model requires a critical reappraisal of the simulation results and a new analysis of the optimal trajectories of a spacecraft equipped with such a propulsion system. The aim of this paper is to review the different thrust models used over the last years for mission analysis purposes, and to illustrate the optimal control law and the corresponding minimum-time trajectories that can be obtained with the new, refined, thrust model. The study highlights new analytical relations for the propulsive thrust as a function of the spacecraft attitude, whereas simple and accurate closed-form equations are also proposed for the study of a classical circle-to-circle coplanar heliocentric orbit transfer.
... Even though the strength of the E-sail effect (Coulomb drag effect on charged tether or wire) has not yet been measured in space, laboratory measurements by Siguier et al. [4] around a charged wire in a flowing plasma resembling LEO conditions indicate [5] that the size of the forming electron sheath is in good agreement with earlier theoretical predictions [6]. On a technical side, a 1 km long sample of E-sail tether has been already produced [7], a lightweight Remote Unit compatible with a solar distance of 0.9-4 au is at TRL 4-5 [8] and a 100 m long E-sail tether will fly onboard Aalto-1 CubeSat, which is scheduled to be launched into a LEO orbit in 2015 [9]. ...
The recent successes of the European Rosetta mission have shown the possibility of a close observation with one of the most evasive celestial bodies in the Solar System, the comets, and the practical feasibility of a comet rendezvous to obtain detailed information and in situ measurements. This paper discusses a preliminary study of the transfer trajectory toward the comet 67P/Churyumov-Gerasimenko (the same target used by Rosetta) for a spacecraft whose primary propulsion system is an electric solar wind sail. The use of a propellantless propulsion system with a continuous thrust is theoretically able to simplify the transfer trajectory by avoiding the need of intermediate flyby maneuvers. The problem is addressed in a parametric way, by looking for the possible optimal launch windows as a function of the propulsion system performance. The study is completed by a mass breakdown analysis of the spacecraft, for some mission scenarios of practical interest, based on the actual payload mass of the spacecraft Rosetta.
... Part of the material in this paper is taken from the Space Propulsion 2014 proceedings paper [9]. ...
In Coulomb drag propulsion, a long high voltage tether or system of tethers gathers momentum from a natural plasma stream such as solar wind or ionospheric plasma ram flow. A positively polarised tether in the solar wind can be used for efficient general-purpose interplanetary propellantless propulsion (the electric solar wind sail or E-sail), whereas a negatively polarised tether in LEO can be used for efficient deorbiting of satellites (the plasma brake). Aalto-1 is a 3-U cubesat to be launched in May 2016. The satellite carries three scientific experiments including 100 m long Coulomb drag tether experiment. The tether is made of four 25 and 50 micrometre diameter aluminium wires that are ultrasonically bonded together every few centimetre intervals. The tether can be charged by an onboard voltage source up to one kilovolt positive and negative. The Coulomb drag is measured by monitoring the spin rate.
... Independent laboratory measurements of the width of the electron sheath around a positively biased tether placed in streaming plasma were made by Siguier et al. (2013). The laboratory results were produced in a plasma mimicking conditions in LEO and they are in good agreement with PIC thrust predictions (Janhunen, 2014b). ...
The solar wind electric sail (E-sail) is a planned in-space propulsion device
that uses the natural solar wind momentum flux for spacecraft propulsion with
the help of long, charged, centrifugally stretched tethers. The problem of
accurately predicting the E-sail thrust is still somewhat open, however, due to
a possible electron population trapped by the tether. Here we develop a new
type of particle-in-cell (PIC) simulation for predicting E-sail thrust. In the
new simulation, electrons are modelled as a fluid, hence resembling hydrid
simulation, but in contrast to normal hybrid simulation, the Poisson equation
is used as in normal PIC to calculate the self-consistent electrostatic field.
For electron-repulsive parts of the potential, the Boltzmann relation is used.
For electron-attractive parts of the potential we employ a power law which
contains a parameter that can be used to control the number of trapped
electrons. We perform a set of runs varying the parameter and select the one
with the smallest number of trapped electrons which still behaves in a
physically meaningful way in the sense of producing not more than one solar
wind ion deflection shock upstream of the tether. By this prescription we
obtain thrust per tether length values that are in line with earlier estimates,
although somewhat smaller. We conclude that the Boltzmann PIC simulation is a
new tool for simulating the E-sail thrust. This tool enables us to calculate
solutions rapidly and allows to easily study different scenarios for trapped
electrons.
... This is based on numerical simulations [7]. When these simulations are run for plasma representing LEO conditions [8], their predicted electron sheath width (which is a proxy for E-sail thrust per length) is in good agreement with laboratory measurements of the sheath width in LEO-like conditions [9]. The thrust of an E-sail is inversely proportional to the distance from the Sun as F $ 1=r [7]. ...
http://www.sciencedirect.com/science/article/pii/S0094576515001290
The novel propellantless electric solar wind sail (E-sail) concept promises
efficient low thrust transportation in the solar system outside Earth's
magnetosphere. Combined with asteroid mining to provide water and synthetic
cryogenic rocket fuel in orbits of Earth and Mars, possibilities for affordable
continuous manned presence on Mars open up. Orbital fuel and water eliminate
the exponential nature of the rocket equation and also enable reusable
bidirectional Earth-Mars vehicles for continuous manned presence on Mars. Water
can also be used as radiation shielding of the manned compartment, thus
reducing the launch mass further. In addition, the presence of fuel in Mars
orbit provides the option for an all-propulsive landing, thus potentially
eliminating issues of heavy heat shields and augmenting the capability of
pinpoint landing. With this E-sail enabled scheme, the recurrent cost of
continuous bidirectional traffic between Earth and Mars might ultimately
approach the recurrent cost of running the International Space Station, ISS.
Electric sail-based propulsion is an innovative propellant-less propulsion technology that generates continuous thrust through the interaction between an artificial electric field and the solar wind. In an electric sail, the propulsive acceleration is adjusted by controlling the attitude of its normal plane and the coefficient determining the maximum thrust. However, the attitude adjustment speed of the electric sail is relatively small and the direction of the normal vector is constrained. Consequently, the electric sail is required to maintain a continuous propulsive acceleration vector when flying by the intermediate targets in multitarget interplanetary exploration. Therefore, an indirect optimization of three-dimensional optimal continuous interplanetary trajectory for electric sails with refined thrust model is investigated in this article. First, the optimal propulsive angles and thrust adjustment coefficient of electric sails with a refined thrust model are derived using Pontryagin's minimum principle. Second, a homotopy function is introduced in the process of trajectory optimization with an indirect method to approximate the step of the thrust adjustment coefficient to improve the accuracy of numerical integration. Additionally, the initial costates of the electric sail are transformed from the numerical simulation results of the Bezier shaping approach (BSA) and integrative BSA (IBSA) using the Karush–Kuhn–Tucker condition. The numerical simulation results for the Earth–Mars rendezvous mission and the multitarget flyby mission reveal that the initial values of the costates are effective to implement an indirect optimization process of the optimal continuous trajectory for the rapid convergence of the electric sail. According to the numerical simulation results for multitarget mission, the indirect method is on average 2.5% better than the Gauss pseudospectral method (GPM) in overall index, and the calculation time is only 1.06% of GPM.
As an innovative spacecraft propellantless propulsion technology, the electric sail can generate electric fields that interact with the solar wind, generating continuous propulsive thrust. However, unlike conventional electric thrusters, the thrust vector of the electric sail is constrained, and the attitude cannot be adjusted quickly. Consequently, the position and velocity of each flight trajectory at the connecting node and the direction of propulsion acceleration must be continuous. In this study, the continuous trajectory optimization of the electric sail–based probe in multi-target exploration is investigated. An integrative Bezier shaping approach (IBSA) is proposed to realize the rapid optimization of a multi-target exploration trajectory with continuous acceleration. The relationship between boundary constraints, intermediate constraints, and Bezier basis function coefficients is derived, transforming the original continuous optimal control problem of multiple stages into a problem of nonlinear programming that naturally satisfies the boundary and intermediate constraints. Numerical simulation results demonstrate that the IBSA can obtain a better interplanetary trajectory with a shorter calculation time than that obtained with the traditional Bezier shaping approach using piecewise optimization. Furthermore, the simulation results are successfully applied to the initial value estimation of the Gauss pseudospectral method (GPM). Compared with the GPM, the proposed shaping approach differs from the objective function by approximately 2%–6%, but requires approximately 1.5% of the computational time. During the preliminary mission design stage, the IBSA can be used for rapid feasibility assessments of different electric sail–based probe mission profiles.
Space exploration has been long constrained by the efficiency and capability of modern chemical rocket. Propellant-less propulsion has been proposed as a solution to expand the boundary of space exploration. In this paper, we examine the possibilities of a propellantless propulsion scheme through the interaction between the spacecraft and ambient plasma. The spacecraft is charged to high electric potential by constantly shooting electrons away. The high voltage spacecraft will deplete the surrounding electrons, thus interact with a wide range of the background plasma (solar wind) and thus effectively extract momentum from the plasma. By taking advantage of the exploitable ambient plasma, a spacecraft can reach very high speed, thus considerably reducing the travel time. The scheme is also applicable for braking, which is helpful in the exploration of inner planets like Venus and Mercury, and the stopping at the destination planets or stars.
The electric sail is an innovative concept for spacecraft propulsion, which can generate continuous thrust without propellant by reflecting solar wind ions. In previous studies, the thrust of an electric sail is described by a classical model that neglects the effects of the electric sail attitude on the propulsive thrust modulus and direction. This study reappraised the performance of the electric sail in the Vesta and Ceres exploration mission with an advanced thrust model that considers the effect of the spacecraft attitude on both the thrust modulus and direction. By using a hybrid optimization method, the trajectory optimization of the electric-sail-based spacecraft from Earth to Vesta and Ceres is implemented in an optimization framework. Numerical results show that the minimal flight time with the advanced thrust model is longer than that with the classical model. The difference in performance between the classical and advanced models is attributable to over-estimation of the maximum thrust cone angle and the thrust modulus by the classical model.
We analyse the potential of the electric solar wind sail for solar system space missions. Applications studied include fly-by missions to terrestrial planets (Venus, Mars and Phobos, Mercury) and asteroids, missions based on non-Keplerian orbits, one-way boosting to outer solar system, off-Lagrange point space weather forecasting and low-cost impactor probes for added science value to other missions. We also discuss the generic idea of data clippers (returning large volumes of high resolution scientific data from distant targets packed in memory chips) and possible exploitation of asteroid resources. We support our analysis by orbit calculations assuming circular and coplanar orbits for planets. Some particular challenge areas requiring further research work and related to some more ambitious mission scenarios are also identified and discussed.
This work presents the outline and so far completed design of the Aalto-1 science mission. Aalto-1 is a multi-payload remote sensing nanosatellite, built almost entirely by students. The satellite aims for a 500–900 km sun-synchronous orbit, and includes an accurate attitude dynamics and control unit, a UHF/VHF housekeeping and S-band data links, and a GPS unit for positioning (radio positioning and NORAD TLE's are planned to be used as backups). It has three specific payloads: a spectral imager based on piezo-actuated Fabry–Perot interferometry, designed and built by The Technical Research Center of Finland (VTT); a miniaturized radiation monitor (RADMON) jointly designed and built by Universities of Helsinki and Turku ; and an electrostatic plasma brake designed and built by the Finnish Meteorological Institute (FMI), derived from the concept of the e-sail, also originating from FMI. Two phases are important for the payloads, the technology demonstration and the science phase. Emphasis is placed on technological demonstration of the spectral imager and RADMON, and suitable targets have already been chosen to be completed during that phase, while the plasma brake will start operation in the latter part of the science phase. The technology demonstration will be over in relatively short time, while the science phase is planned to last two years. The science phase is divided into two smaller phases: the science observations phase, during which only the spectral imager and RADMON will be operated for 6–12 months, and the plasma brake demonstration phase, which is dedicated to the plasma brake experiment for at least a year. These smaller phases are necessary due to the drastically different power, communication and attitude requirements of the payloads. The spectral imager will be by far the most demanding instrument on board, as it requires most of the downlink bandwidth, has a high peak power and attitude performance. It will acquire images in a series up to at least 20 spectral bands within the 500–900 nm spectral range, forming the desired spectral data cube product. Shortly before an image is acquired, the parallel visual spectrum camera will take a broader picture for comparison. Also stereoscopic imaging is planned. The amount of data collected by the spectral imager is adjustable, and ranges anywhere from 10 to 500 MB. The RADMON will be on 80% of an orbit period in average and together with housekeeping data will gather around 2 MB of data in 24 h. An operational limitation is formed due to the S-band downlink capability of 29–49 MB per 24 h for a 500 900 km orbit altitude, as only one ground station is planned to be available to the satellite. This will limit both type and quantity of spectral imager images taken during the science phase. The plasma brake will in turn be on within an angle of 20° over the poles for efficient use of the Earth's magnetic field and ionosphere during its spin-up and operation.
The electric solar wind sail is a propulsion system that uses long centrifugally spanned and electrically charged tethers to extract the solar wind momentum for spacecraft thrust. The sail angle with respect to the sun direction can be controlled by modulating the voltage of each tether separately to produce net torque for attitude control and thrust vectoring. A solution for the voltage modulation that maintains any realistic sail angle under constant solar wind is obtained. Together with the adiabatic invariance of the angular momentum, the tether spin rate and coning angle is solved as functions of temporal changes in the solar wind dynamic pressure, the tether length, or the sail angle. The obtained modulation also gives an estimate for the fraction of sail performance (electron gun power) to be reserved for sail control. We also show that orbiting around the Sun with a xed sail angle leads to a gradual increase (decrease) in the sail spin rate when spiraling outward (inward). This eect arises from the fact that the modulation of the electric sail force can only partially cancel the Coriolis eect, and the remaining component lays in the spin plane having a cumulative eect on the spin rate.
The electric solar wind sail (E-sail) is an innovative propellantless concept for interplanetary space propulsion that uses the natural solar wind as a thrust source with the help of long, artificially charged tethers. The characteristic property of an E-sail based spacecraft is that the propulsive acceleration scales as the inverse Sun-spacecraft distance, and the thrust vector can be varied within about 30 deg away from radial direction. The aim of this paper is to estimate the transfer times required to fulfill a mission toward the near-Earth asteroid 1998 KY26. In doing so the propulsive acceleration of the E-sail, at a reference distance from the Sun, is used as a performance parameter so that the numerical results are applicable to E-sails of different sizes and different payload masses. The paper shows that the flight time scales nearly linearly with the inverse of the spacecraft maximum propulsive acceleration at 1 astronomical unit from the Sun, when the acceleration is greater than 0.3 mm=s 2 . For smaller propulsive accelerations the relationship for the flight time is more involved, because the transfer trajectory is complex and more than one revolution around the Sun is necessary to accomplish the mission. The numerical analysis involves a sample return mission in which the total flight time is parametrically correlated with the starting date for a given E-sail propulsion system.
The scientific mission of ESTCube-1, launched in May 2013, is to measure the
Electric solar wind sail (E-sail) force in orbit. The experiment is planned to
push forward the development of E-sail, a propulsion method recently invented
at the Finnish Meteorological Institute. E-sail is based on extracting momentum
from the solar wind plasma flow by using long thin electrically charged
tethers. ESTCube-1 is equipped with one such tether, together with hardware
capable of deploying and charging it. At the orbital altitude of ESTCube-1
(660--680~km) there is no solar wind present. Instead, ESTCube-1 shall observe
the interaction between the charged tether and the ionospheric plasma. The
ESTCube-1 payload uses a 10-meter, partly two-filament E-sail tether and a
motorized reel on which it is stored. The tether shall be deployed from a
spinning satellite with the help of centrifugal force. An additional mass is
added at the tip of the tether to assist with the deployment. During E-sail
experiment the tether shall be charged to 500~V potential. Both positive and
negative voltages shall be experimented with. The voltage is provided by a
dedicated high voltage source and delivered to the tether through a slip ring
contact. When the negative voltage is applied to the tether, the satellite body
is expected to attract electron flow capable of compensating for the ion flow,
which runs to the tether from the surrounding plasma. With the positive voltage
applied, onboard cold cathode electron guns are used to remove excess electrons
to maintain the positive voltage of the tether. In this paper we present the
design and structure of the tether payload of ESTCube-1.
We analyse the potential of the electric solar wind sail for solar system
space missions. Applications studied include fly-by missions to terrestrial
planets (Venus, Mars and Phobos, Mercury) and asteroids, missions based on
non-Keplerian orbits (orbits that can be maintained only by applying continuous
propulsive force), one-way boosting to outer solar system, off-Lagrange point
space weather forecasting and low-cost impactor probes for added science value
to other missions. We also discuss the generic idea of data clippers (returning
large volumes of high resolution scientific data from distant targets packed in
memory chips) and possible exploitation of asteroid resources. Possible orbits
were estimated by orbit calculations assuming circular and coplanar orbits for
planets. Some particular challenge areas requiring further research work and
related to some more ambitious mission scenarios are also identified and
discussed.
The plasma brake is a thin negatively biased tether which has been proposed
as an efficient concept for deorbiting satellites and debris objects from low
Earth orbit. We simulate the interaction with the ionospheric plasma ram flow
with the plasma brake tether by a high performance electrostatic particle in
cell code to evaluate the thrust. The tether is assumed to be perpendicular to
the flow. We perform runs for different tether voltage, magnetic field
orientation and plasma ion mass. We show that a simple analytical thrust
formula reproduces most of the simulation results well. The interaction with
the tether and the plasma flow is laminar when the magnetic field is
perpendicular to the tether and the flow. If the magnetic field is parallel to
the tether, the behaviour is unstable and thrust is reduced by a modest factor.
The case when the magnetic field is aligned with the flow can also be unstable,
but does not result in notable thrust reduction. We also fix an error in an
earlier reference. According to the simulations the predicted thrust of the
plasma brake is large enough to make the method promising for satellite
deorbiting. As a numerical example we estimate that a 5 km long plasma brake
tether weighing 0.055 kg could produce 0.43 mN breaking force which is enough
to reduce the orbital altitude of a 260 kg debris mass by 100 km during one
year.
The solar wind electric sail is a novel propellantless space propulsion
concept. According to numerical estimates, the electric sail can produce a
large total impulse per propulsion system mass. Here we consider using a 0.5 N
electric sail for boosting a 550 kg spacecraft to Uranus in less than 6 years.
The spacecraft is a stack consisting of the electric sail module which is
jettisoned at Saturn distance, a carrier module and a probe for Uranus
atmospheric entry. The carrier module has a chemical propulsion ability for
orbital corrections and it uses its antenna for picking up the probe's data
transmission and later relaying it to Earth. The scientific output of the
mission is similar to what the Galileo Probe did at Jupiter. Measurement of the
chemical and isotope composition of the Uranian atmosphere can give key
constraints for different formation theories of the solar system. A similar
method could also be applied to other giant planets and Titan by using a fleet
of more or less identical electric sail equipped probes.
The electric solar wind sail (E-sail) is a new type of propellantless
propulsion system for Solar System transportation, which uses the
natural solar wind for producing spacecraft propulsion. This paper
discusses a mass breakdown and a performance model for an E-sail
spacecraft that hosts a scientific payload of prescribed mass. In
particular, the model is able to estimate the total spacecraft mass and
its propulsive acceleration as a function of various design parameters
such as the tethers number and their length. A number of subsystem
masses are calculated assuming existing or near-term E-sail technology.
In light of the obtained performance estimates, an E-sail represents a
promising propulsion system for a variety of transportation needs in the
Solar System.
Space debris in the form of abandoned satellites is a growing concern, especially at the heavily populated 600-1000 km altitude orbits. To prevent new space junk from forming, new satellites should be equipped with a deorbiting mechanism. The problem is especially tricky for the emerging class of very small satellites for which using a braking rocket as a deorbit mechanism may have a prohibitively high relative cost impact. We describe a novel type of deorbiting mechanism that is suitable for small satellites with a mass of upto a few hundred kilograms. The method is a plasma brake device based on coulomb drag interaction between the ionospheric plasma and a negatively charged thin tether. The method resembles the well-known electrodynamic tether deorbit mechanism, but the underlying physical mechanism is different and the new method has an order of magnitude smaller mass and power consumption. The new method uses the same physical principle (coulomb drag) as the recently invented electric solar wind sail propulsion method. Furthermore, the tether required by the plasma brake is so thin that, if accidentally cut, the loose fragments of it pose no threat to other spacecraft and will rapidly descend into the atmosphere. The electrostatic plasma brake could enable an extended use of small satellites by resolving their associated space debris problem. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.
Outgoing optimal (minimum time) trajectories for electric sail spacecraft are calculated. The study includes trajectories for reaching a distance of 100 AU from the Sun, escape trajectories, and missions aimed at obtaining a flyby with Uranus or Neptune. The results are parameterized as a function of the electric sail acceleration at 1 AU . Using an electric sail of modest complexity, the attainable flight-times are quite attractive. Because no gravity assists are used, the mission trajectories investigated do not suffer of complications such as rare launch windows. Missions with coast arcs (in which the propulsion is switched off at some point) are also analyzed, because they might be needed for outgoing missions which include the Pioneer anomaly study.
This work presents the outline and so far completed design of the Aalto-1 science mission. Aalto-1 is a multi-payload remote-sensing nanosatellite, built almost entirely by students. The satellite aims for a 500–900 km sun-synchronous orbit and includes an accurate attitude dynamics and control unit, a UHF/VHF housekeeping and S-band data links, and a GPS unit for positioning (radio positioning and NORAD TLE's are planned to be used as backup). It has three specific payloads: a spectral imager based on piezo-actuated Fabry–Perot interferometry, designed and built by The Technical Research Centre of Finland (VTT); a miniaturised radiation monitor (RADMON) jointly designed and built by Universities of Helsinki and Turku; and an electrostatic plasma brake designed and built by the Finnish Meteorological Institute (FMI), derived from the concept of the e-sail, also originating from FMI. Two phases are important for the payloads, the technology demonstration and the science phase. The emphasis is placed on technological demonstration of the spectral imager and RADMON, and suitable targets have already been chosen to be completed during that phase, while the plasma brake will start operation in the latter part of the science phase. The technology demonstration will be over in a relatively short time, while the science phase is planned to last two years. The science phase is divided into two smaller phases: the science observations phase, during which only the spectral imager and RADMON will be operated for 6–12 months and the plasma brake demonstration phase, which is dedicated to the plasma brake experiment for at least a year. These smaller phases are necessary due to the drastically different power, communication and attitude requirements of the payloads. The spectral imager will be by far the most demanding instrument on board, as it requires most of the downlink bandwidth, has a high peak power and attitude performance. It will acquire images in a series up to at least 20 spectral bands within the 500–900 nm spectral range, forming the desired spectral data cube product. Shortly before an image is acquired, the parallel visual spectrum camera will take a broader picture for comparison. Also stereoscopic imaging is planned. The amount of data collected by the spectral imager is adjustable, and ranges anywhere from 10 to 500 MB. The RADMON will be on 80% of an orbit period on average and together with housekeeping data will gather around 2 MB of data in 24 h. An operational limitation is formed due to the S-band downlink capability of 29–49 MB per 24 h for a 500 900 km orbit altitude, as only one ground station is planned to be available for the satellite. This will limit both type and quantity of spectral imager images taken during the science phase. The plasma brake will in turn be within an angle of 20° over the poles for efficient use of the Earth's magnetic field and ionosphere during its spin-up and operation.
Missions towards the boundaries of the Solar System require long transfer times and advanced propulsion systems. An interesting option is offered by electric sails, a new propulsion concept that uses the solar wind dynamic pressure for generating a continuous thrust without the need for reaction mass. The aim of this paper is to investigate the performance of such a propulsion system for obtaining escape conditions from the Solar System and planning a mission to reach the heliosphere boundaries. The problem is studied in an optimal framework, by minimizing the time to reach a given solar distance or a given hyperbolic excess speed. Depending on the value of the sail characteristic acceleration, it is possible that, in an initial mission phase, the sailcraft may approach the Sun to exploit the increased available thrust due to the growing solar wind electron density. The corresponding optimal trajectory is constrained to not pass inside a heliocentric sphere whose admissible radius is established by thermal constraints. Once the escape condition is met, the sail is jettisoned and the payload alone continues its journey without any propulsion system. A medium performance electric sail is shown to have the potentialities to reach the heliosheath, at a distance of 100 AU, in about fifteen years. Finally, the Interstellar Heliopause Probe mission is used as a reference mission to further quantify the electric sail capabilities for an optimal transfer towards the heliopause nose (200 AU).
An electric sail is capable of guaranteeing the fulfilment of a class of trajectories that would be otherwise unfeasible through
conventional propulsion systems. In particular, the aim of this paper is to analyze the electric sail capabilities of generating
a class of displaced non-Keplerian orbits, useful for the observation of the Sun’s polar regions. These orbits are characterized
through their physical parameters (orbital period and solar distance) and the spacecraft propulsion capabilities. A comparison
with a solar sail is made to highlight which of the two systems is more convenient for a given mission scenario. The optimal
(minimum time) transfer trajectories towards the displaced orbits are found with an indirect approach.
The electric solar wind sail (E-Sail) is a new propulsion method for interplanetary travel which was invented in 2006 and is currently under development. The E-Sail uses charged tethers to extract momentum from the solar wind particles to obtain propulsive thrust. According to current estimates, the E-Sail is 2-3 orders of magnitude better than traditional propulsion methods (chemical rockets and ion engines) in terms of produced lifetime-integrated impulse per propulsion system mass. Here we analyze the problem of using the E-Sail for directly deflecting an Earth-threatening asteroid. The problem then culminates into how to attach the E-Sail device to the asteroid. We assess alternative attachment strategies, namely straightforward direct towing with a cable and the gravity tractor method which works for a wider variety of situations. We also consider possible techniques to scale up the E-Sail force beyond the baseline one Newton level to deal with more imminent or larger asteroid or cometary threats. As a baseline case we consider an asteroid of effective diameter of 140 m and mass of 3 million tons, which can be deflected with a baseline 1 N E-Sail within 10 years. With a 5 N E-Sail the deflection could be achieved in 5 years. Once developed, the E-Sail would appear to provide a safe and reasonably low-cost way of deflecting dangerous asteroids and other heavenly bodies in cases where the collision threat becomes known several years in advance.
An electric solar wind sail is a recently introduced propellantless space propulsion method whose technical development has also started. The electric sail consists of a set of long, thin, centrifugally stretched and conducting tethers which are charged positively and kept in a high positive potential of order 20 kV by an onboard electron gun. The positively charged tethers deflect solar wind protons, thus tapping momentum from the solar wind stream and producing thrust. The amount of obtained propulsive thrust depends on how many electrons are trapped by the potential structures of the tethers, because the trapped electrons tend to shield the charged tether and reduce its effect on the solar wind. Here we present physical arguments and test particle calculations indicating that in a realistic three-dimensional electric sail spacecraft there exist a natural mechanism which tends to remove the trapped electrons by chaotising their orbits and causing them to eventually collide with the conducting tethers. We present calculations which indicate that if these mechanisms were able to remove trapped electrons nearly completely, the electric sail performance could be about five times higher than previously estimated, about 500 nN/m, corresponding to 1 N thrust for a baseline construction with 2000 km total tether length.
An electric solar wind sail is a recently introduced propellantless space propulsion method whose technical development has also started. In its original version, the electric sail consists of a set of long, thin, centrifugally stretched and conducting tethers which are charged positively and kept in a high positive potential of 20 kV by an onboard electron gun. The positively charged tethers deflect solar wind protons, thus tapping momentum from the solar wind stream and producing thrust. Here we consider a variant of the idea with negatively charged tethers. The negative polarity electric sail seems to be more complex to implement than the positive polarity variant since it needs an ion gun instead of an electron gun as well as a more complex tether structure to keep the electron field emission current in check with the tether surface. However, since this first study of the negative polarity electric sail does not reveal any fundamental issues, more detailed studies would be warranted.
In this paper, sailing and navigation in the solar wind with a spacecraft powered by an electric sail is addressed. The electric sail is a novel propellantless spacecraft propulsion concept based on positively charged tethers that are centrifugally uncoiled and stabilised to extract the solar wind momentum by repelling the solar wind protons. Steering of such a sail ship is realised either by changing the tether voltage or the sail spin plane. To model the solar wind, we use spacecraft observations for the density and wind speed at 1 AU and assume that the speed is constant and density decreases in square of the distance from the Sun. Using the electric sail thrust formula, we describe the sail response to the solar wind variations, especially, the self-reefing effect leading to a smooth spacecraft acceleration even during periods of large densities or fast winds. As a result, the variations of the acceleration are statistically small relative to the density and wind speed variations. Considering the navigation, we adopt an optimal transfer orbit to Mars originally obtained for constant solar wind speed and density. The orbit and associated sail operations including a coasting phase are then used as the navigation plan to Mars. We show that passive navigation based only on the statistical results is far too inaccurate for planetary missions and active navigation is required. We assume a simple active navigation system that monitors only the actual orbital speed with an onboard accelerometer and matches it with the optimal orbital speed by altering the tether voltage independently from the future solar wind conditions. We launch 100 test spacecraft with a random launch date and show that with the active navigation 85% (100%) of the spacecraft reach a distance relative to Mars less than about 10 (70) Mars radii with a residual speed less than 20 m/s (80 m/s). As a conclusion, the electric sail is highly navigable and it suits for targeting planets and asteroids, in addition to broad targets such as the heliopause.
One possibility for propellantless propulsion in space is to use the momentum flux of the solar wind. A way to set up a solar wind sail is to have a set of thin long wires which are kept at high positive potential by an onboard electron gun so that the wires repel and deflect incident solar wind protons. The efficiency of this so-called electric sail depends on how large force a given solar wind exerts on a wire segment and how large electron current the wire segment draws from the solar wind plasma when kept at a given potential. We use 1-D and 2-D electrostatic plasma simulations to calculate the force and present a semitheoretical formula which captures the simulation results. We find that under average solar wind conditions at 1 AU the force per unit length is (5±1×10?8 N/m for 15 kV potential and that the electron current is accurately given by the well-known orbital motion limited (OML) theory cylindrical Langmuir probe formula. Although the force may appear small, an analysis shows that because of the very low weight of a thin wire per unit length, quite high final speeds (over 50 km/s) could be achieved by an electric sailing spacecraft using today's flight-proved components. It is possible that artificial electron heating of the plasma in the interaction region could increase the propulsive effect even further.
The electric solar wind sail (E-sail) is a space propulsion concept that uses the natural solar wind dynamic pressure for producing spacecraft thrust. In its baseline form, the E-sail consists of a number of long, thin, conducting, and centrifugally stretched tethers, which are kept in a high positive potential by an onboard electron gun. The concept gains its efficiency from the fact that the effective sail area, i.e., the potential structure of the tethers, can be millions of times larger than the physical area of the thin tethers wires, which offsets the fact that the dynamic pressure of the solar wind is very weak. Indeed, according to the most recent published estimates, an E-sail of 1 N thrust and 100 kg mass could be built in the rather near future, providing a revolutionary level of propulsive performance (specific acceleration) for travel in the solar system. Here we give a review of the ongoing technical development work of the E-sail, covering tether construction, overall mechanical design alternatives, guidance and navigation strategies, and dynamical and orbital simulations.
BETs is a three-year project financed by the Space Program of the European Commission, aimed at developing an efficient deorbit system that could be carried on board any future satellite launched into Low Earth Orbit (LEO). The operational system involves a conductive tape-tether left bare to establish anodic contact with the ambient plasma as a giant Langmuir probe. As a part of this project, we are carrying out both numerical and experimental approaches to estimate the collected current by the positive part of the tether. This paper deals with experimental measurements performed in the IONospheric Atmosphere Simulator (JONAS) plasma chamber of the Onera-Space Environment Department. The JONAS facility is a 9- vacuum chamber equipped with a plasma source providing drifting plasma simulating LEO conditions in terms of density and temperature. A thin metallic cylinder, simulating the tether, is set inside the chamber and polarized up to 1000 V. The Earth's magnetic field is neutralized inside the chamber. In a first time, tether collected current versus tether polarization is measured for different plasma source energies and densities. In complement, several types of Langmuir probes are used at the same location to allow the extraction of both ion densities and electron parameters by computer modeling (classical Langmuir probe characteristics are not accurate enough in the present situation). These two measurements permit estimation of the discrepancies between the theoretical collection laws, orbital motion limited law in particular, and the experimental data in LEO-like conditions without magnetic fields. In a second time, the spatial variations and the time evolutions of the plasma properties around the tether are investigated. Spherical and emissive Langmuir probes are also used for a more extensive characterization of the plasma in space and time dependent analysis. Results show the- ion depletion because of the wake effect and the accumulation of ions upstream of the tether. In some regimes (at large positive potential), oscillations are observed on the tether collected current and on Langmuir probe collected current in specific sites.
The solar wind electric sail (electric sail, E-sail) is a new and
potentially revolutionarily efficient method of interplanetary
propulsion. The E-sail taps the momentum flux of the solar wind with the
help of long, thin, highly positively charged and centrifugally
stretched tethers. According to estimations, a full-scale E-sail could
weigh 100-200 kg and produce 1 N thrust at 1 AU. The thrust would scale
as 1/r where r is the distance from the sun and the thrust direction
could be vectored by ˜ 30° away from radial by inclining the
sail. Here we present new particle in cell (PIC) simulation results of
the E-sail thrust. The challenge in modeling the E-sail from first
principles is the possible existence of trapped electrons. We sketch a
way by which the PIC simulations might be possible to extrapolate to the
natural limit in the future.
We produced a 1 km continuous piece of multifilament electric solar wind sail tether of μm-diameter aluminum wires using a custom made automatic tether factory. The tether comprising 90 704 bonds between 25 and 50 μm diameter wires is reeled onto a metal reel. The total mass of 1 km tether is 10 g. We reached a production rate of 70 m∕24 h and a quality level of 1[per thousand] loose bonds and 2[per thousand] rebonded ones. We thus demonstrated that production of long electric solar wind sail tethers is possible and practical.
Missions towards potentially hazardous asteroids require considerable propellant-mass consumption and complex flybys maneuvers with conventional propulsion systems. A very promising option is offered by an electric sail, an innovative propulsion concept, that uses the solar-wind dynamic pressure for generating a continuous and nearly radial thrust without the need for reaction mass. The aim of this paper is to investigate the performance of such a propulsion system for performing rendezvous missions towards all the currently known potentially hazardous asteroids, a total of 1025 missions. The problem is studied in an optimal framework by minimizing the total flight time. Assuming a canonical value of sail characteristic acceleration, we show that about 67% of the potentially hazardous asteroids may be reached within one year of mission time, with 137 rendezvous in the first six months. A detailed study towards asteroid 99942 Apophis is reported, and a comparison with the corresponding performance achievable with a flat solar sail is discussed.
Electric sail for extracting momentum from the solar wind was described. Another way to use the solar wind momentum is to use a magnetic sail, which requires a superconducting one-dimensional loop wire of length smaller than the linear extent of the electric wire mesh. It was found that the electric sail does not need any magnetic fields or superconductors and its power requirements are modest. Accurate maneuvering is difficult but the method is expected to be well suited for missions whose purpose is just to travel fast out of the solar system across the heliosphere to interstellar space.
We present a novel ultrasonic wire-to-wire bonding method for bonding two micrometer-thick metal
wires together. A special jig and an industrial wire bonder perform the bonding. This wire-to-wire bonding
is the core unit process to produce space tether for the Electric Solar Wind Sail. The proposed method
was validated experimentally with 38 bonds where a 25 um and a 50 um by diameter Al wires that were
first bonded together after which the bond was pull tested. The measured average pull force was
(74 ± 15) mN whereas the lowest pull force value was 40 mN. The results show that wire-to-wire bonds
of sufficient strength can be produced for the Electric Solar Wind Sail tether application. Tether manufacturing
was demonstrated with a separate test where a 1.4 m long tether was produced featuring more
than 100 wire-to-wire bonds.
The aim of this paper is to study, from a mission analysis point of view, the performance of a hybrid propulsion concept for a two-dimensional transfer towards a planet of the Solar System. The propulsion system is obtained by combining a chemical thruster, used for the phases of Earth escape and/or target planet capture, with an electric sail, which provides a continuous thrust during the heliocentric transfer. Two possible mission scenarios are investigated: in the first case the sailcraft reaches the target planet with zero hyperbolic excess velocity, thus performing a classical rendezvous mission in a heliocentric framework. In the second mission scenario, a given final hyperbolic excess velocity relative to the planet is tolerated in order to decrease the mission flight time. The amount of final hyperbolic excess velocity is used as a simulation parameter for a tradeoff study in which the minimum flight time is related to the total velocity variation required by the chemical thruster to accomplish the mission, that is, for Earth escape and planetary capture.
The electric solar wind sail (E-sail) is a novel, efficient propellantless
propulsion concept which utilises the natural solar wind for spacecraft
propulsion with the help of long centrifugally stretched charged tethers. The
E-sail requires auxiliary propulsion applied to the tips of the main tethers
for creating the initial angular momentum and possibly for modifying the
spinrate later during flight to counteract the orbital Coriolis effect and
possibly for mission specific reasons. We introduce the possibility of
implementing the required auxiliary propulsion by small photonic blades (small
radiation pressure solar sails). The blades would be stretched centrifugally.
We look into two concepts, one with and one without auxiliary tethers. The use
of photonic blades has the benefit of providing sufficient spin modification
capability for any E-sail mission while keeping the technology fully
propellantless. We conclude that the photonic blades appear to be a feasible
and attractive solution to E-sail spinrate control.
Design description of the Remote Unit
- S Wagner
- J Sundqvist
- G Thornell
Wagner, S., J. Sundqvist and G. Thornell, Design description of the Remote Unit, ESAIL
FP7 project deliverable D41.2, 2012 (http://www.
electric-sailing.fi/fp7/docs/D412.pdf).
Summary of orbit calculations supporting WP61, ESAIL FP7 project deliverable D62
- G Mengali
- A A Quarta
- G Aliasi
Mengali, G., A.A. Quarta and G. Aliasi, Summary
of orbit calculations supporting WP61, ESAIL
FP7 project deliverable D62.1, 2013 (http://www.
electric-sailing.fi/fp7/docs/D62.1.pdf).
Remote Unit test results
- G Thornell
- J Sundqvist
Thornell, G. and J. Sundqvist, Remote Unit
test results, ESAIL FP7 project deliverable
D41.4, 2013 (http://www.electric-sailing.
fi/fp7/docs/D414.pdf).
Auxiliary tether report, ESAIL FP7 project deliverable D24
- J Polkko
Polkko, J., Auxiliary tether report, ESAIL
FP7
project
deliverable
D24.1,
2012
(http://www.electric-sailing.fi/fp7/
docs/D241_auxtether.pdf).
Simulator user guide, ESAIL FP7 project deliverable D51
- P Janhunen
Janhunen, P., Simulator user guide, ESAIL FP7
project deliverable D51.1, 2013 (http://www.
electric-sailing.fi/fp7/docs/D511.pdf).
Report of performed runs
- P Janhunen
Janhunen, P., Report of performed runs, ESAIL
FP7 project deliverable D51.2, 2013 (http://www.
electric-sailing.fi/fp7/docs/D51.2.pdf).
Failure mode and recovery strategy analysis, ESAIL FP7 project deliverable D53
- P Janhunen
- P Toivanen
- P Pergola
Janhunen, P., P. Toivanen and P. Pergola, Failure mode and recovery strategy analysis, ESAIL
FP7 project deliverable D53.1, 2013 (http://www.
electric-sailing.fi/fp7/docs/D531.pdf).