65th International Astronautical Congress, Toronto, Canada. Copyright ©2014 by Scott Fisher. Published by the IAF, with permission and released to the IAF to publish in all forms IAC-14-A6.4.7 DEBRIS CREATION IN GEOSTATIONARY TRANSFER ORBITS: A REVIEW OF LAUNCH PRACTICES 2004-2012 Scott Fisher Space Generation Advisory Council (SGAC)*, United Kingdom, [email protected] Emmanuelle David German Aerospace Center (DLR), Germany, [email protected] * This research was completed whilst the author was under the employ of the DLR. Objects left in Geostationary Transfer Orbit (GTO) pose a threat to operational satellites in Low Earth Orbit (LEO) and Geostationary Earth Orbit (GEO). In addition to regularly crossing both of these belts at high relative velocities, objects in these orbits are often large in size and mass (i.e. upper stages, fuel tanks and payload adaptors). Each successful satellite launch to GEO can generate 1-4 large pieces of debris (dependent upon the launch vehicle), a significant portion of which is left in transfer orbits that do not decay. This paper analyses the past practices of all launch vehicles that placed satellites in GEO from 2004- 2012. The orbits of the 294 pieces of debris that were identified are discussed further in depth with reference to their threat to the Low and Geostationary Earth Orbits and the likelihood of their decay in 25 years as per international guidelines. Results are grouped for each of the 17 distinct launch vehicles used, and rankings of launch vehicles with regards to the amount of mission related debris are developed. Furthermore, the physical mechanisms with regards to orbital decay are investigated in order to ascertain the impact of various orbital parameters on re-entry time. The European Space Agency’s (ESA) Debris Risk Assessment and Mitigation Analysis (DRAMA) tool specifically is used to conduct Monte-Carlo analyses on nominal upper stages in GTO. I. INTRODUCTION The following sections of this paper will analyse From 2004-2012, 210 launch vehicles operated by the mitigation measures of launch vehicle upper five independent nations and two international stages from 2004-2012. Furthermore, the physical organisations placed satellites in Geostationary Earth mechanisms with regards to orbital decay will be Orbit (GEO; altitude 35,786km). In almost all cases, investigated to determine the impact of various orbital each successful launch left one or more pieces of parameters on the rate of orbital decay. debris (e.g. upper stages, payload carriers or auxiliary The survey of GTO launches extends on work fuel tanks) in highly elliptical Geostationary Transfer previously conducted by Johnson [2] where a similar Orbits (GTO). GTOs typically have perigees in Low analysis was conducted from 2000-2003. Earth Orbit (altitude < 2000km) and apogees near- Data for this study has been obtained from Space GEO. Inclinations for GTO vary depending upon Track (https://www.space-track.org), and orbit launch site, and can be as high as 52°. Debris in GTO propagation and re-entry calculations utilise the is particularly concerning for operational satellites as European Space Agency’s (ESA) Debris Risk it frequently crosses both GEO and the crowded LEO. Assessment and Mitigation Analysis (DRAMA) tool. The Inter-Agency Debris Coordination Committee (IADC) has defined two “protected regions with regard to the generation of space debris” [1]: Region A, a “spherical region that extends up from Earth to an altitude of 2000km” (i.e. LEO) and Region B, a segment of a spherical shell near GEO (shown in Figure 1) [1]. The IADC Guidelines strongly recommend disposing objects that pass through these protected regions at mission end-of-life, either by placing them in ‘parking orbits’ (IADC Measure 5.3.1) or by placing them in orbits that will decay with the debris re-entering within 25 years (IADC Measure 5.3.2). Fig. 1: IADC Protection Zones [1] IAC-14-A6.4.7 Page 1 of 10 65th International Astronautical Congress, Toronto, Canada. Copyright ©2014 by Scott Fisher. Published by the IAF, with permission and released to the IAF to publish in all forms II. BACKGROUND II.I.II J2 Effect As mentioned in Section I, debris in GTO is a The Earth is not a perfect sphere (it is closer to an cause for concern as it crosses both the congested oblate ellipsoid), and hence it generates a non- LEO and GEO orbits. In addition, the orbits of uniform gravitational field. The gravitational pull of objects in GTO are difficult to propagate and predict the Earth on the object is hence not directly towards due to intricacies in their orbital perturbations. This the centre of the Earth, but is offset slightly towards section will provide background information on both the equator. Any orbiting object is pulled orbital perturbations and allowable debris mitigation preferentially towards the equator, creating a torque orbits for objects in GTO. on the orbit. This torque does not affect inclination, and instead causes gyroscopic precession of both the II. I Orbital Perturbations in GTO RAAN and the AoP. Objects in GTO are very susceptible to perturbations that change their orbital parameters. II.I.III Third Body Effects This can lead to interesting and counter-intuitive Gravitational perturbations due to the Sun and the effects. The primary perturbations arise from Moon directly affect the RAAN, Argument of atmospheric drag, the oblateness of the Earth (i.e. the Perigee, inclination and eccentricity of an orbit. They ‘J2 effect’) and third body effects. do not however affect the semi-major axis [3]. The A brief outline of the effects of these orbital Sun and the Moon apply an external torque to the parameters is presented here. For additional details orbits and cause the angular momentum vector to please refer to [3], [4] and [5]. rotate. The effect on RAAN and argument of perigee is small compared to the J2 effect. II.I.I Atmospheric Drag For objects in GTO, the effect of third body forces The density of the Earth’s atmosphere decreases on eccentricity is important as a small change in exponentially with increasing altitude. It hence has eccentricity can result in a large change in perigee. the largest impact at low altitudes, i.e. near perigee. The effect of the sun is much larger than the effect of Atmospheric drag works to decrease the object’s the moon [4]. orbital velocity at perigee, which results in a lower Depending on the relative position of the sun, the apogee during subsequent orbits. It has negligible eccentricity of the orbit will either increase or effect on the Argument of Perigee (AoP), Right decrease. As the semi-major axis is unaffected by Angle of Ascending Node (RAAN) and inclination of third body effects, an increase in eccentricity will lead an orbit, instead steadily decreasing the semi-major to a decrease in perigee. This is illustrated in Figure 2. axis and eccentricity. The magnitude of drag is difficult to predict in II.I.IV Sun Synchronous Resonance advance as the atmospheric density is highly For certain combinations of semi-major axis, dependent upon solar activity. There are a variety of eccentricity and inclination, the J2 effect can result in methods used in its estimation, as further outlined in a solar angle (itself dependent upon the RAAN and [6]. AoP; shown in Figure 2) remaining approximately constant over an extended period of time. Third body effects can then result in significant increases or decreases in the orbit’s eccentricity, and therefore its perigee, over time. This effect can be exploited to Fig. 2: Dependence of perigee increase/decrease on Fig 3: Example of sun-synchronous resonance [4] location of the Sun [3] IAC-14-A6.4.7 Page 2 of 10 65th International Astronautical Congress, Toronto, Canada. Copyright ©2014 by Scott Fisher. Published by the IAF, with permission and released to the IAF to publish in all forms Stages left Other mission-related GTO Perigee GTO Apogee Number of Launch System in GTO per debris left in GTO per Range (km) Range (km) launches launch launch Ariane 5 1 0-1 222-658 33,476-38,467 43 Atlas 2AS 1 0 207-236 35,262-35,534 2 Atlas 3Aa 1 0 191 35,080 1 Atlas V 1 0 110-6,231 34,644-85,173 6 Long March 3A 1 0 134-405 35,794-42,062 8 Long March 3B 1 0 156-197 30,472-49,260 4 Long March 3B/E 1 0 105-268 14,886-49,902 11 Long March 3C 1 0 161-295 34,272-42,402 9 Delta 4M+(4,2) 1 0 6,595-6,604 35,127-35,173 3 GSLV 1 0 172-192 31,659-35,486 2 H-IIA 1 0 180-263 34,048-34,899 4 Proton-K/DM 0-1 0-2 240-361 35,699-36,180 9 Proton-M/Briz-M 0-1 1 310-14,676 12,092-42,470 54 PSLV-XLa 1 0 282 21,342 1 Soyuz-FG Fregata 1 0 311 66,081 1 Zenit-3Fa, b 0 0 N/A N/A 1 Zenit-3SLc 1 0 132-11,250 35,710-39,378 26 TOTAL 185 Table 1:2004-2012 launch systems examined aOnly one mission was flown during 2004-2012 bUpper stage was left in GEO; specific GTO unknown cIncludes both Zenit-3SL and Zenit-3SLB expedite re-entry of debris in GTO. Conversely, if it GEO will ensure that the debris does not occurs unexpectedly (e.g. due to changing solar enter either protected orbits. Specifically: conditions over several years), it can result in a much a.