PLANNED YET UNCONTROLLED RE-ENTRIES OF THE CLUSTER-II SPACECRAFT Stijn Lemmens(1), Klaus Merz(1), Quirin Funke(1) , Benoit Bonvoisin(2), Stefan Löhle(3), Henrik Simon(1) (1) European Space Agency, Space Debris Office, Robert-Bosch-Straße 5, 64293 Darmstadt, Germany, Email:[email protected] (2) European Space Agency, Materials & Processes Section, Keplerlaan 1, 2201 AZ Noordwijk, Netherlands (3) Universität Stuttgart, Institut für Raumfahrtsysteme, Pfaffenwaldring 29, 70569 Stuttgart, Germany ABSTRACT investigate the physical connection between the Sun and Earth. Flying in a tetrahedral formation, the four After an in-depth mission analysis review the European spacecraft collect detailed data on small-scale changes Space Agency’s (ESA) four Cluster II spacecraft in near-Earth space and the interaction between the performed manoeuvres during 2015 aimed at ensuring a charged particles of the solar wind and Earth's re-entry for all of them between 2024 and 2027. This atmosphere. In order to explore the magnetosphere was done to contain any debris from the re-entry event Cluster II spacecraft occupy HEOs with initial near- to southern latitudes and hence minimise the risk for polar with orbital period of 57 hours at a perigee altitude people on ground, which was enabled by the relative of 19 000 km and apogee altitude of 119 000 km. The stability of the orbit under third body perturbations. four spacecraft have a cylindrical shape completed by Small differences in the highly eccentric orbits of the four long flagpole antennas. The diameter of the four spacecraft will lead to various different spacecraft is 2.9 m with a height of 1.3 m. They are spin atmospheric entry conditions and are thus predicted to stabilised at 15 rpm. The mass amounts to 1200 kg of show significantly different behaviour. Given the rare which 650 kg are propellants, which are expected to be opportunity of a repeatable, predictable, yet depleted by the end of the mission. After in-depth uncontrolled re-entry event, and of the object’s mission analysis, manoeuvres executed during 2015 will relatively simple geometries, the definition of multiple ensure a re-entry of all four spacecraft between 2024 in-situ observation campaigns and expected science and 2026 [1]. The re-entry epoch can already now be return in terms of break-up model and risk validation predicted with a high temporal accuracy, of a few days, algorithms are explored. due to the third body perturbations driving the re-entry, which is in stark contrast to re-entries driven by 1 INTRODUCTION atmospheric perturbations where a ±20% on the remaining orbital lifetime is considered the standard International space debris mitigation standards request a accuracy [2]. This effectively limits the final re-entry of permanent clearance of the Low Earth Orbit (LEO) and all four spacecraft to Southern latitudes, minimising the Geostationary Orbit (GEO) protected regions. risk for people on ground. Furthermore, the risk on-ground, following a potential atmospheric re-entry, shall be constrained by clear During the past decades space agencies, and more recent safety limits. Corresponding disposal options are well space industry and academia as well, have developed established for missions in GEO and LEO, and consist software models to evaluate the break-up and demise of of near circular graveyard orbits or atmospheric decay. spacecraft during re-entries. Due to very limited Science missions, however, sometimes operate on observational data for validation of these models and the highly-eccentric orbits (HEO) to achieve their mission operational needs of the agencies, these models are goals, such as astronomical observations or focussed on re-entries from nearly circular orbits or measurements of the Earth's environment. HEOs controlled re-entries. This in turn implies that their describe a group of orbits with perigees in or close to usability for atmospheric re-entry from highly eccentric the LEO region and eccentricities above those of orbit is based on extrapolation, in some cases ignoring Geostationary Transfer Orbits (GTOs) (approximately the physical differences. This work reviews the limited 0.73). The dominant perturbation forces on these orbits available observational data on uncontrolled spacecraft are typically caused by the gravity fields of Sun and re-entries with high flight path angle. It describes the Moon. simulated break-up of the Cluster-II spacecraft during re-entry and the associated uncertainties which heavily The European Space Agency’s (ESA) Cluster II depend on the exact re-entry conditions. Due to the spacecraft are part of an international collaboration to relatively simple design of the spacecraft, the possibility Proc. 7th European Conference on Space Debris, Darmstadt, Germany, 18–21 April 2017, published by the ESA Space Debris Office Ed. T. Flohrer & F. Schmitz, (http://spacedebris2017.sdo.esoc.esa.int, June 2017) Figure 1 Evolution of perigee and apogee height for the four Cluster spacecraft without manoeuvre. Table 1. Naming conventions for the Cluster-II spacecraft. Name Rumba Salsa Samba Tango Numbered Name Cluster-1 Cluster-2 Cluster-3 Cluster-4 Flight Model FM5 FM6 FM7 FM8 COSPAR ID 2000-045A 2000-041B 2000-041A 2000-045B US catalogue ID 26463 26411 26410 26464 Predicted Re-entry 2025-11-04 2024-09-07 2026-08-21 2026-08-21 of observing individual component releases, and hence In the frame of the mission extensions granted in 2012, categorising the overall break-up mechanism, is it was decided to study the orbital evolution and addressed. manoeuvre options in more detail and develop models to assess the on-ground casualty risk in case of re-entry. 2 RE-ENTRY PREDICTABILITY It was obvious that advancing the re-entry of Cluster-1 (Rumba) from 2038 to the mid-2020s would be In order to have a clear naming convention, we will use beneficial to lower the on-ground risk since then the spacecraft names as given in Table 1 and use the term perigee would be located over Southern instead of “Cluster” for the overall mission (i.e. dropping the “II”), Northern latitudes. Since shortening the on-orbit time by in particular Cluster-1 and Cluster-2 will refer to the more than a decade also reduces the risk of an on-orbit specific spacecraft and not to the original or recovered collision or break-up, by mid-2014 a baseline strategy overall mission. For Cluster the investigations of re- for Cluster-1 (Rumba) was selected involving a entry trajectories have a long history and was discussed manoeuvre during the first quarter of 2015 leading to re- from 2007, when mission extensions beyond 2010 wre entry in 2025 which was subsequently implemented. under debate but the orbit of Cluster-2 (Salsa) had to be adjusted as part of the constellation change manoeuvres The dominating orbit perturbation of these HEOs are the to avoid the otherwise expected re-entry in spring 2010. Sun’s and Moon’s gravity leading to strong variations in The overall constellation change manoeuvres were the eccentricity, and therefore altitude of perigee since executed successfully in mid-2010. This led to natural the semi major axis is essentially constant. Secular rates long-term evolutions of the perigee altitude implying re- of the orbital elements per orbit averaged over an orbital entries in the years 2024 – 2026 for three of the revolution of the third body are given in [3] and indicate spacecraft and one in 2038 for the last one (Cluster-1, in particular no secular change in semi major axis and a Rumba). Figure 1 shows the orbital predictions at the change of the eccentricity resp. perigee altitude time of the mission extension review in 2012. depending on the location of the perigee. Figure 2 Perigee altitude [km] and orbital period [hours] for the last perigee passages before re-entry for Cluster-1 (top left), Cluster-2 (top right), Cluster-3 (bottom left) and Cluster-4 (bottom right) based on a full numerical prediction extending an operational orbit of 2016 For the argument of perigee (measured with respect to on whether the long-periodic and higher order effects the third body’s orbital plane) being in the 2nd or 4th nearly cancel each other or are amplified by quadrant the eccentricity decreases and the perigee superposition in which case the perigee altitude of the altitude increases, whereas eccentricity increases and last orbits may reach several 10s of km difference from perigee altitude decreases for the argument of perigee one orbit to the next. In case the pattern is steep enough, being in the 1st or 3rd quadrant. The secular effect due the object will enter the deeper layers of the atmosphere to the Moon is about double the effect due to the Sun. without first circularizing its orbit, as is usually the case Long-periodic effects are described in [3] as well, with a for the more typical uncontrolled re-entries from lower period of half the orbital period of the third body, i.e. orbits. half a year for the Sun and approximately two weeks for Since then the breakup occurs near the location of the the Moon. For the long-periodic oscillations the effect perigee, this would allow predicting and therefore of the Sun is about six times larger than that of the potentially controlling the latitude band of the break-up Moon in terms of amplitude of eccentricity or perigee process and therefore limiting the distribution of altitude and can reach about 200 km due to the Sun fragments on ground in terms of latitude. Even more, if perturbation and some 10s of km due to the Moon. the orbital period could be controlled precisely at EOL These analytical considerations are useful to understand and is not perturbed by air drag until the very last the overall evolution, however for a detailed analysis a revolution, one may even predict and control the time of higher order semi-analytical or full numerical the re-entry and therefore the sub-satellite longitude of propagator is required.