Review of Deorbiting Technology

Due to the historical practice of abandoning in orbit decommissioned spacecraft, payload fairings, and upper stages of launch vehicles there is currently a shell of synthetic debris around the Earth. As of October 2013 the U.S. Space Surveillance Network is tracking over 13000 Earth-orbiting space debris objects larger than 10 cm [NASA, 2013]. The threat this poses to active and future spacecraft sent into orbit is an ever growing problem of special concern to LEO altitude regions where the accumulation of debris is more severe.

Reviews by panels of international experts have repeatedly failed to identify an universal concise plan which is both, technically feasible in the near-term, and economically viable. Hence, there are currently no international laws that control the generation of new space debris and enact sanctions. However, a number of orbit debris mitigation strategies and guidelines have been proposed over the years by NASA [NASA, 1995, NASA, 2000] and ESA [ESA, 1999, ESA, 2004]. These include mission planning, satellite hardware design strategies, explosion prevention guidelines, collision avoidance techniques, as well as the well-known 25-year post-mission lifetime non-regulated requirement. Targeted active debris removal (ADR) of high risk objects is also seen as an action needed for a sustainable Earth orbit environment [Liou, 2010, Liou, 2011]. In 2007, all the space-faring nations that form the Inter-Agency Space Debris Coordination Committee (IADC) drew up a voluntary code of best practice, that established mitigation guidelines to be followed in the use of outer space [IADC, 2007].

Several studies, such as the ones summarised in [Petro, 1992, Mayer and Chao, 2000, Macdonald et al., 2013], have been funded in the last two decades to evaluate techniques for end-of-life disposal of space assets of a wide range of masses and initial orbit altitudes. In the case of atmospheric reentry of the spacecraft or deorbiting, the techniques more widely accepted can be categorized in four groups: chemical propulsion maneuvers; low-thrust propulsion transfer; drag enhancement; and electrodynamic tethers. Nevertheless, some of the most recent studies that have actual target missions only focus on evaluating propulsion technologies as means of deorbiting in LEO, as these as seen as more readily available solutions. For example, in [Burkhardt et al., 2002], traditional mono or bi-propellant systems, cold gas, solid propulsion, and electrical propulsion were compared and traded-off with respect to many different parameters.

According to [Burkhardt et al., 2002], there are three optimal options for accelerating the nat- ural altitude decay of low altitude Earth-bound spacecraft: progressively take energy from the orbit by for example letting the thin atmosphere deorbit the object with a drag-augmentation device (un-controlled deorbiting); direct retrieval and deorbiting from an initial altitude (controlled deor- biting); or maneuver to an orbit from which atmospheric drag will remove the satellite within a given time frame, e.g. 25 years (hybrid-controlled deorbiting). Depending on the propulsion system carried onboard or the lack of one, the initial altitude and mass of the object, and the likelihood to survive reentry that would risk life on ground, one of these options will be more attractive or necessary to mission designers.

consideration given their low TRL, high risks involved in using them, and more limited applicability. These include: ground and space based, laser and solar ablation [Bekey, 1997, Bondarenko et al., 1997, Campbell, 2000, Schall, 2002]; foam-based ADR [Andrenucci et al., 2011]; the Ion Beam Shepherd [Combardelli, 2011]; space tugs [Starke et al., 2011]; the Catchers’s Mitt [Phillips, 2010]; and Tungsten dust [Ganguli et al., 2012]. Also, from the traditional chemical propulsion systems, cold gas propulsion is not considered efficient for low-thrust propulsion transfer giving its extremely low specific impulse, and solid propulsion is discarded due to the risk of generating further debris from propellant slag.

The leading and more mature deorbiting technologies and their applicability in LEO are briefly summarized next:

Mono and bi-propellant propulsion. It is the traditional technology used for spacecraft orbit manoeuvring and station-keeping, and therefore, a great part of the satellites launched would already carry on-board one of these systems. Chemical motors can offer thrust levels of several hundred to thousands of Newtons. Of the down-selected deorbiting options, it is the only that can offer direct controlled atmospheric re-entry if the spacecraft is still operational, so as to point the motors in the right direction. However, the mass of the required additional propellant can still be quite significant counting for more than 5-10% and 10-20% of the total mass of the spacecraft, for mono and bi-propellant options respectively [Macdonald et al., 2013]. If controlled re-entry is not required, i.e. the object does not pose a threat to human life upon uncontrolled re-entry, then a hybrid-controlled deorbiting approach would be much more mass efficient, with fuel savings of up to 40-50% with respect to the aforementioned values.

Electric Propulsion. Is known for its high specific impulse, which leads to substantial propel- lant mass savings compared to chemical propulsion. The considerable power consumption, however, constitutes a serious usage constraint. Electric propulsion systems include arcjets, resistojets, ion thrusters, pulsed plasma thrusters (PPT), field emission electric propulsion (FEEP), and station- ary plasma/hall current thrusters (SPT/HCT) [Burkhardt et al., 2002]. These systems can offer low thrust, high impulse propulsion transfers. Conversely, for this option to be used, the space- craft needs to be operational over an extended period of time in order to control the direction of thrust of the continuous deorbit burns. For controlled deorbiting if the lifetime chosen is close to 25 years, then this option becomes unattractive due to both operational costs and the lifetime of most electric thrusters. For hybrid-controlled deorbiting, however, the deorbiting period can be greatly reduced. If the low-thrust propulsion system is already on-board, this option is likely a very attractive option, as the mass of the required fuel can be below 1-2% of the total mass of the spacecraft for a wide range of spacecraft in LEO [Macdonald et al., 2013].

Electrodynamic Tethers. They are conductive wires that link the main spacecraft with the deployed end-mass. As the usual kilometer size length tether passes through the ionospheric plasma it picks up electrons that travel to the cathode held at the end-mass, where the electron emitter would expel them back into the ionosphere. The current along the tether interacts with Earth’s magnetic field to produce a Lorentz force against the orbital motion direction of the spacecraft as

shown in Figure 2.22 (left). The orbital kinetic energy is then dissipated as resistive heating of the tether and the surrounding plasma. This passive method is preferred for a deorbiting option over that where current is constantly provided from the main craft to the tether to induce drag.

Mechanical tethers are a reasonable mature technology, but the deployment and post-deployment electromagnetic effects of electrodynamic tethers are not so straightforward. Entanglement with an uncontrollable tumbling host, current and ohmic heat management, or the dynamic coupling between the gravity-gradient and Lorentz forces on a deployed tether, are some of the inherent risks of this technology. The company Tethers Unlimited Inc. has developed an 27 kg commercial deorbit system known as the Terminator Tether [Hoyt and Forward, 2000]. It uses a feedback control system in the deployment unit that stays attached to the spacecraft to manage the unspooling process and subsequently manage the tether. Hence, since electrodynamic tethers are not an inert system, they require continuous active control to remain operational. This poses some restriction in the deorbiting mission lifetime considered for them, which is normally sized to last for less than a year. In [Hoyt and Forward, 2000], the length of the tether required to derive enough power for the tether control unit from the tether current, thus enabling an autonomous system, was found to be in the order of 5-10 km. It is also shown that only above the 650-700 km range altitude, the area-time product for hosts that use a Terminator Tether system would be smaller than that of aerodynamic drag acting alone on the spacecraft. The area-time product can be used to compare different deorbiting options, as it is proportional to the risk of the host and deorbiting system colliding with another body during the descent. In addition, it is stated that deorbiting efficiency of electrodynamic tethers considerably drops for orbit inclinations higher than about 75◦. This is because the orbital path and the Earth’s magnetic flux lines move into near-alignment, thus reducing the size of the Lorentz force that can be generated. It is worth mentioning that in [Voronka and al., 2005], a low-cost Nano-Terminator tape system with a mass of just 100 g was proposed for cubesat-type nano-satellites. However, for the larger size tether systems envisioned, the mass of the deorbiting system would be of about 1-5 % that of the total spacecraft mass.

Figure 2.22: Tethers Unlimited’s Terminator TetherT M (left) [Hoyt and Forward, 2000]; and AeroAstro’s SPORTT M Aerobrake configuration (right) [Gloyer et al., 2001].

Solar Sailing to gain Atmospheric Drag Augmentation. Solar sailing is assumed to be inapplicable for orbits below 750 km altitude due to the dominance of atmospheric drag in this region [McIness, 2004b]. However, solar sails can be utilised in a hybrid-controlled way if passive stabilisation is achieved by design as in [Lappas et al., 2011]. Reflective balloons have also been studied to exploit the enhanced solar radiation pressure on the gossamer surface in [Lucking et al., 2011, Lucking and Colombo, 2012]. Above that region they can utilize solar radiation pressure to actively manoeuvre to an altitude where atmospheric drag can then be harnessed with the same gossamer structure. Therefore, the host would only need to be operational for the first phase of the deorbiting period, which should not take longer than 10% of the mission lifetime if this method is to be effective. However, in [Macdonald et al., 2013] it is shown that the sizes of the solar sail required to comply with the aforementioned time constraint can be significant, if either, the mass of the spacecraft is large, or its initial orbit altitude is high. For the latter, re-orbiting to a higher graveyard orbit above 2000km using solar sailing, was found to be of limited value for the reference cases studied. As shown in [McIness, 2004b], for equatorial orbits, a simple switching law can be used during the solar sailing controlled deorbiting phase, that requires a slow slew of 90◦ twice per orbit. For polar orbits, the sail attitude can be fixed relative to the Sun, but the sail must yaw 360deg per orbit to align the sail thrust vector opposite to the velocity vector.

Atmospheric Drag Augmentation. These systems rely on decreasing the ballistic coefficient of the host by deploying a gossamer structures, that reduces significantly the mass-to-area ratio of the spacecraft. The systems that are used for drag-assisted deorbiting can be classified in two big types according to their directionality: spherical-shaped systems like balloons, that are omnidirectional. This means that they always offer the same amount of surface area to the residual atmosphere, regardless of their orientation, and thus they do not strictly require stabilisation; and drag sails, which are directional, and require stabilisation in order the maximize the exposed area to the stream of free-molecular flow particles. Drag sails can be completely flat, similar to common solar sails designs, or have a pyramidal shuttlecock-like shape as shown in Figure 2.22 (right). The result of many studies [Kumar, 1996, Roberts, 2004, Harkness, 2006, Roberts and Harkness, 2007], can be summed up by stating that flatter designs will generate greater drag forces, but smaller restoring torques than their sharper counterparts.

Nonetheless, as explained in [Macdonald et al., 2013], the simplicity and robustness of a spherical envelope comes at the cost of a mass penalty when compared to a shaped gossamer structure. Considering alone the material required, the surface area of a sphere goes as 4πr2, where r is the characteristic length, e.g. radius, while the surface area of a flat disc or square scales as πr2

or r2, respectively. However, as proposed in [Nock et al., 2013], one should consider the lower risk of debris-generating impacts that the majority of the deployable area offers when calculating the Area-Time-Product (ATP) that is commonly used to compare the collision risk of different deorbiting concepts.

The main advantage of drag augmentation devices is that they are the only technology that does not have to rely on the host still being operational or manoeuvrable, either at the time of activa- tion or throughout the deorbiting phase. Therefore, they can truly provide uncontrolled-deorbiting capability to the host. They can also act as a “fail safe” system, that if the spacecraft suffers a

catastrophic failure would activate. The latter requires a completely independent deorbit device such as the one proposed in [Fernandez et al., 2014a]. As explained in [Harkness, 2006], such a device could be contained in a small bolt-on package which would be attached onto new space- craft at launch. It could conceivably remain dormant thereafter, only awakening when instructed from the ground after the host has reached EOL and requires a deorbit manoeuvre, or after it becomes unresponsive following the activation signal of a “watch-dog” counter. However, according to [Harkness, 2006, Macdonald et al., 2013], the use of an atmospheric drag augmentation system is applicable only to low and medium mass spececraft (>1000 kg), or spacecraft that are unlikely to survive atmospheric re-entry, hence minimizing risk to human life. Also, they are generally only effectively applicable in the low and medium LEO regions below 800 km.