Satellite Deorbit Research Articles

Electric sail, photonic sail and deorbiting applications of the freely guided photonic blade

We consider a freely guided photonic blade (FGPB) which is a centrifugally stretched sheet of photonic sail membrane that can be tilted by changing the centre of mass or by other means. The FGPB can be installed at the tip of each main tether of an electric solar wind sail (E-sail) so that one can actively manage the tethers to avoid their mutual collisions and to modify the spin rate of the sail if needed. This enables a more scalable and modular E-sail than the baseline approach where auxiliary tethers are used for collision avoidance. For purely photonic sail applications one can remove the tethers and increase the size of the blades to obtain a novel variant of the heliogyro that can have a significantly higher packing density than the traditional heliogyro. For satellite deorbiting in low Earth orbit (LEO) conditions, analogous designs exist where the E-sail effect is replaced by the negative polarity plasma brake effect and the photonic pressure by atmospheric drag. We conclude that the FGPB appears to be an enabling technique for diverse applications. We also outline a way of demonstrating it on ground and in LEO at low cost.

Performance Analysis of Bipropellant Propulsion System on Special Working Conditions

At the end of a test satellite, oxidant of the propulsion system was exhausted ahead of fuel. Long-term monopropellant working condition without combustion, cool gas flow working condition and interim working condition were experienced, which trouble the de-orbit control of the satellite. In order to evaluate the performance of the propulsion system on the special working conditions during the satellite de-orbit, theoretic analysis of the specific thrusters is introduced. And effective calculation method is built. Based on the method and obtained remote data, the thrust and specific impulse of the propulsion system on monopropellant working condition and cool gas flow working condition are calculated. It is indicated that, on monopropellant working condition, the thrust of 10N thruster declines 83%, the specific impulse declines about 82%. On cool gas flow working condition, the thrust of the 10N thrusters is less than 0.12N, and declines with the decrease of the tank pressure. The specific impulse is less than 146s. These results can be consulted by other satellites confronting similar conditions.

Libration dynamics and stability of electrodynamic tethers in satellite deorbit

This paper studies libration dynamics and stability of deorbiting nano-satellites by short and bare electrodynamic tethers. A critical aspect of satellite deorbit by an electrodynamic tether is to maintain the tether aligned with the local vertical and stable while subjected to external perturbations. The dynamics of electrodynamic tether system in deorbit application is divided into the orbital motion of the center of system’s mass and the tether libration motion relative to that center. Major space environmental perturbations including the current-induced electrodynamic force, atmospheric drag, oblateness effect of the Earth, irregularity of geomagnetic field, variable plasma density, solar radiation pressure, and lunisolar gravitational attractions are considered in the dynamic analysis. Quantitative analyses are provided in order to characterize the order of the perturbative torques during the deorbit process. A single index is derived from the libration energy to stabilize the libration motion by regulating the current in the tether through simple on-off switching. Numerical results show that the libration dynamics of an electrodynamic tether has significant impacts on the deorbit process and the electrodynamic tether cannot effectively deorbit satellites without libration stability control. The proposed current regulation strategy is simple and very effective in stabilizing libration motion of an electrodynamic tether.

Survival Probability of Round and Tape Tethers Against Debris Impact

The current space environment, consisting of manmade debris and micrometeoroids, poses a risk to safe operations in space, and the situation is continuously deteriorating due to in-orbit debris collisions and to new satellite launches. Bare electrodynamic tethers can provide an efficient mechanism for rapid deorbiting of satellites from low Earth orbit at end of life. Because of its particular geometry (length very much larger than cross-sectional dimensions), a tether may have a relatively high risk of being severed by the single impact of small debris. The rates of fatal impact of orbital debris on round and tape tethers of equal length and mass, evaluated with an analytical approximation to debris flux modeled by NASA’s ORDEM2000, shows much higher survival probability for tapes. A comparative numerical analysis using debris flux model ORDEM2000 and ESA’s MASTER2005 validates the analytical result and shows that, for a given time in orbit, a tape has a probability of survival of about one and a half orders of magnitude higher than a round tether of equal mass and length. Because deorbiting from a given altitude is much faster for the tape due to its larger perimeter, its probability of survival in a practical sense is quite high.

A passive satellite deorbiting strategy for medium earth orbit using solar radiation pressure and the J2 effect

The growing population of space debris poses a serious risk to the future of space flight. To effectively manage the increase of debris in orbit, end-of life disposal has become a key requirement for future missions. This poses a challenge for Medium Earth Orbit (MEO) spacecraft which require a large Δv to re-enter the atmosphere or reach the geostationary graveyard orbit. This paper further explores a passive strategy based on the joint effects of solar radiation pressure and the Earth's oblateness acting on a high area-to-mass-ratio object. The concept was previously presented as an analytical planar model. This paper uses a full 3D model to validate the analytical results numerically for equatorial circular orbits first, then investigating higher inclinations. It is shown that for higher inclinations the initial position of the Sun and right ascension of the ascending node become increasingly important. A region of very low required area-to-mass-ratio is identified in the parameter space of semi-major axis and inclination which occurs for altitudes below 10,000km.

Deorbiting Dynamics of Electrodynamic Tether

rzhong@yorku.caAccepted 28 August 2011This paper studies satellite deorbit using electrodynamic tether (EDT) propulsion. Gaus-sian perturbation equations are used to model the orbital dynamics of EDT with envi-ronmental perturbations of electrodynamic force, aerodynamic drag and the effect ofEarth’s oblateness. Differential equations for the induced voltage-current across EDTare derived and solved with boundary conditions determined by mission objectives andhardware devices. A simplified analytical method for solving the equations is proposed toimprove computational efficiency. Simulations find that the orbit of a satellite deorbitedby EDT will become elliptical in near polar orbits due to the higher-order perturbationof Earth’s magnetic field. This is beneficial for the near polar orbits where the electro-dynamic force is less effective, because the atmosphere at a lower perigee will providelarger air drag to dissipate the orbital kinetic energy of satellite faster. Moreover, weproved that the polarity reverse of the induced voltage/current across EDT in near polarorbits does not affect the kinetic energy dissipation by the current induced Lorentz force.Compared the decay by air drag only, the orbit decaying time of a satellite with EDTwill be reduced by three and two orders of magnitudes in the equatorial and polar orbitsrespectively.Keywords: Electrtodynamic tether; deorbit; dynamics; tethered satellite; magnetic field.

Nonlinear Control of a Double Pendulum Electrodynamic Tether System

T HE electrodynamic tether (EDT) system uses the Lorentz force as a thruster generated by the interference between the Earth’s magnetic field and the electric current along the tether. The main advantage of this system is that its required electric current can be produced by solar power and electric particles such as Xenon, the mass of which is much smaller than those of traditional chemical propellants. As a result, this system is expected to be employed as a new space technology that will be useful in many types of space applications [1] such as reboosting of the international space station [2]; deorbiting of dysfunction satellites [3] (space debris removal system); the momentum exchange electrodynamic reboost (MXER) [4] system, which could be used to boost payloads from a low Earth orbit (LEO) to a geosynchronous transfer orbit (GTO); and scientific missions that could include the observation of meteors, the highaltitude atmosphere, and the aurora. The motion of the EDT system consists of librations that are similar to themotion of a pendulum and vibrations of the tether that are similar to the vibrations of a string. The librational motion of the tether system in an elliptic orbit is known to be chaotic [5,6]. The librational and vibrational motion of the tethered satellite system must be stabilized to successfully perform space operations such as observation of the atmospheric environment at high altitude. For the purpose of stabilization, a great number of control schemes have been proposed for various aspects of control [7], such as stationkeeping, deployment and retrieval control, generation of optimal trajectories [8], suppression of tether vibrations by electric current variation [9], periodic motion of the electrodynamic tether systemon inclined orbits [10], and effect of the electrodynamic force on orbit [11]. However, to the best knowledge of the authors, except delayed feedback controls, application of nonlinear control schemes to the multipendulum type of electrodynamic tether systems for achieving constellation missions has not yet been widely studied. The tether system treated in this Note consists of two subsatellites and a mother satellite, such as the space shuttle, connected in series via conductive massless tethers. Although this model seems simple, because the mass and flexibility of the tether are ignored, it can be used to investigate the behavior of the in-plane motion of an electrodynamic tether system in an elliptic orbit without significant computational effort and to investigate the applicability of the nonlinear controllers to the simple model. In this study, two nonlinear control methods to stabilize the inplane librational motion of an electrodynamic tether system in an elliptic orbit are investigated. The nonlinear controllers treated in this Note are a decoupling control method [6,12] and amodel-followingdecoupling control method [6,12]. The decoupling control can be used to control each tether attitude independently, and this independent motion is suitable for achieving satellite constellations. If this motion is realized for a three-mass tethered satellite system, various scientific missions will be possible, including the observation of planets with a magnetic field and the observation of the aurora from more than two directions simultaneously. The model-following-decoupling control method is introduced to achieve a periodic motion by which the Earth atmosphere at the specific altitude can be studied periodically. In this study, the periodic in-plane motion of a tethered satellite system in a circular orbit is employed as the reference trajectory for tracking by the actual tethered satellite system in an elliptic orbit. The results of numerical simulations show that the nonlinear control schemes considered herein can stabilize the in-plane librational motion of a double pendulum electrodynamic tether system in an elliptic orbit.

Collision avoidance during deorbiting of satellites at end-of-life

The deorbiting of satellites at end-of-life is the only effective means to keep low-Earth orbit clean. With the success of solar-electric propulsion systems deorbiting now becomes economically feasible also for spacecraft above 1000 km altitude. The deorbiting in this case will take several months. The collision risk with 10-cm objects during this time is only of the order of 10 −6 per square meter, however, it may be used as an argument to reorbit an aging satellite above LEO rather than to deorbit it. With a proper collision avoidance concept this already low risk of an on-orbit collision with an operational satellite can almost be excluded. A 40 mN thrust system is sufficient for a 1000 kg spacecraft to keep a safe radial distance of 200 m from any target orbit.

Analysis of Bare-Tether Systems for Deorbiting Low-Earth-Orbit Satellites

Performances, design criteria, and system mass of bare tethers for satellite deorbiting missions are analyzed. Orbital conditionsand tethercrosssection dee nea tetherlength, suchthat1 )shortertethers areelectron collecting practically in their whole extension and 2 ) longer tethers collect practically the short-circuit current in a e xed segment length. Long tethers have a higher drag efe ciency (dee ned as the drag force vs the tether mass ) and are better adapted to adverse plasma densities. Dragging efe ciency and mission-related costs are used to dee ne design criteria for tether geometry. A comparative analysis with electric thrusters shows that bare tethers have much lower costs for low- and midinclination orbits and remain an attractive option up to 70 deg.

Satellite de-orbiting by means of electrodynamic tethers part i: general concepts and requirements

It has been recently suggested that conductive tethers, besides other applications, may be effective for the de-orbiting of satellites at the end of their operational life. In this paper we give a preliminary evaluation of this capability with respect to a well-defined tether system and using the expertise and results from the recently flown TSS-1 project. The system considered is a long conducting tether, covered by an insulator, with a passive electron collector at the positive termination and a hollow-cathode electron emitter at the opposite end. In analyzing this system, we point out the crucial importance of the contact impedances associated to the coupling between the tether terminations and the ionospheric plasma. We give a first evaluation of the de-orbiting time for a typical case and conclude about feasibility of a tether system to de-orbit satellites.

Tethers and debris mitigation

In recent years, the use of tethers has been proposed for reduction of space debris either through momentum transfer or use of electrodynamic effects. Tethers have been shown to at least theoretically allow for quick, elegant and cost-effective deorbit of defunct satellites or spent stages. On the other hand, the large risk that tethers themselves may pose to other satellites in orbit has been recognized as well. The large collision area of tethers, combined with operational hazards and meteoroid risk may result in a large orbital exposure. For example, in 1997, the ESA/Dutch 35-km tether deployment of YES from TEAMSAT was inhibited after an analysis of the collision risk for the case the tether operation would fail. The question rises how these two points of view compare to eachother. This paper intends to highlight a representative selection of the proposed tether applications while taking into account the added risks caused by the tethers themselves. Typical applications from recent literature will be briefly described, such as an Ariane 502 spent stage re-entry from GTO and the concept of deboost of defunct satellites by interaction of a conductive tether with the Earth magnetic field. Mass savings of the tethered sytems versus conventional equivalents will be evaluated. Based on a crude risk analysis, involving elements such as mission complexity, dynamic stability, meteoroid risk and orbital life time, a general outline of limiting factors can be given for the various applications. Special attention is reserved for implementation of mechanisms that help reduce this tether risk, such as the DUtether (Tether Degradable by Ultraviolet), utilization of airdrag and solar pressure, the effect of residual current in bare tethers, tether retrieval etc. It is proposed how a net tether-induced mitigation can be compared to that of conventional alternatives, i.e. deboost by rocket engine or a completely passive approach. This comparison is put in the perspective of an ever-increasing occupation of the space environment. It is concluded that tethers can in fact help mitigate the debris risk and that for each application a useful niche can be defined. It is argued that eliminating pollution directly after use of the precious resource of space is not only good custom, but also an important way to make the risk of debris controllable and independent of future trends. Although tethers may have large exposure in terms of area-time product, they deliver a quick cleaning service that may be appreciated by the future users of space.