The Satellite Sweeper Approach for the Solution of the Space Debris Problem

SPACE PROPULSION 2016, MARRIOTT PARK HOTEL, ROME, ITALY / 2-6 MAY 2016 The Satellite Sweeper Approach for the Solution of the Space Debris Problem P. Pergola (1) and A. Pasini (2) (1) Sitael S.p.A., Via A. Gherardesca 5, 56121 Ospedaletto, Pisa, Italy, Email: pierpaolo.pergola@sitael.com (2) Civil and Industrial Engineering, University of Pisa, Via Girolamo Caruso, 8, 56121, Pisa, Italy, Email: angelo.pasini@unipi.it KEYWORDS: deorbiting, space debris ABSTRACT: Due to the increasing thread space debris pose for the future space exploration, an active space debris removal mission is nowadays considered as one of the main priorities for maintain the space usage sustainability. The active debris removal approach proposed in this study exploits a so called Space Garbage Truck (SGT) in charge of collect multiple dead satellites and move all of them in a single controlled re-entry trajectory. Debris are not collected directly by the SGT but a Waste Collector Robot (WCR) equipped with a standard docking interface and an autonomous, reusable and (if required) refillable propulsion system for orbiting maneuver and close proximity operations. One or more WCRs are lunched integrated in the SGT. Once the SGT reaches a nominal orbit in the proximity of the target debris to remove; the WCR undocks and reaches one target. The debris is then brought back to the SGT by means of the WCR propulsion system and thrown into a dedicated section of the SGT. After all the WCRs complete their mission, the SGT performs the final controlled re-entry maneuvers with all the debris collected. 1. THE SPACE DEBRIS PROBLEM Since the Sputnik-1 about 6000 satellites have been launched and of these less than 20% is still operational. About 6000 tons and 15000 objects are catalogues as debris, including results of fragmentations, in-orbit collisions and explosions, dead satellites, spent upper stages and uncontrolled objects [1,2]. These objects are prone to produce further debris colliding with other objects causing a cascade effect, the well known Kessler Syndrome [3]. In the worst scenario there might be in future entire orbital regions forbidden for further missions where the object density is at a critical point. Accordingly an Active Debris Removal (ADR) mission is a cornerstone mission to maintain the possibility to explore the near-earth space in future. Such a mission is especially required for past launches, since nowadays the United Nations Committee for the Peaceful Uses of Outer Space (COPUOS) [4] and the Inter-Agency Space Debris Coordination Committee (IADC) [5] pose some guidelines for modern satellites. In particular, the limit of deorbiting within 25 years is imposed for LEO missions, while GEO spacecraft are required to be re-collocated at the end of operational lifetime in a proper graveyard orbit. Many of the objects already in orbit, however, do not cope with these guidelines. This is particularly true for some specific regions, like sunsynchronous orbit, geostationay orbit, polar orbit and low Earth orbits where an abundance of objects might preclude the future exploitation. These objects are mainly dead satellites and launcher upper stages designed and launched without any deorbiting system. To remove such objects an active deorbiting mission is required and several approaches have been already proposed [6]. These spans from a deorbiting platform in charge of attaching/docking (e.g. by means of harpoons, nets, etc.) and actively bring one of these objects up to a controlled Earth re-entry; to methods where a device is attached to the target to reduce its natural lifetime by increasing the atmospheric drag action (e.g. sails, foam, etc.) thus leading to an uncontrolled atmosphere burn-up. In general ADR concepts can be classified as electromagnetic methods (i.e. electrodynamic tethers and magnetic sails), momentum exchange methods (i.e. solar sails, drag augmentation devices and foam-based methods), remote methods (i.e. lasers) and capture methods (i.e. nets). Each method may represent a valuable solution for space debris belonging to specific classes or types, or orbiting in particular space regions. The ADR mission concept here proposed integrates a single large platform in charge of the 1/8

final controlled re-entry and an in-orbit collector platform able to reach one or more target debris. The mission conceived is a multi target mission where the in-orbit operations and the final atmospheric re-entry are the two main functions divided between two different platforms. It has been proven [7] that impacts between intact satellites and rocket bodies cause much more debris than impacts involving small debris already in orbit. Accordingly the idea of this study is to consider large intact satellites as targets for the ADR mission. Objects of the IRIDIUM constellation are favourable targets since these are multiple objects rather close each other and the loss of control on one of these satellites might cause multiple collisions with the other constellation elements. Thus the ADR concept is applied to the IRIDIUM constellation to have a reference scenario to assess the overall mission performance. The constellation has been chosen also because it is composed by a large number of medium size objects populating one of the most appealing region for LEO telecommunication satellites. The Iridium-33/Cosmos-2251 collision in 2009 [8], indeed, is a clear example of the high level of criticality of such a region and of the control of the objects of this constellation. The IRIDIUM constellation is composed by 66 satellites orbiting in six different orbital planes, based on the LM700 series satellite operated by U.S.-based Iridium Satellite LLC. The satellites were manufactured by Motorola and Lockheed Martin and are designed to provide L-band mobile telephone and communications services [9]. The paper presents the idea in detail in Sec. 2 and presents the main mission performance in Sec. 3. Sec. 4 proposes a further concept extension replacing the chemical propulsion system with a low thrust device and Sec. 5 show the preliminary spacecraft sizing for both the SGT and the WCR. fulfilled by Waste Collector Robots (WCR) characterized by a lighter weight w.r.t. the SGT in such a way that the propellant consumption for catching the debris can be significantly reduced w.r.t. moving the SGT. Moreover, the SGT can be equipped with a high performance bipropellant propulsion system to perform the final controlled deorbiting mission. 3. PERFORMANCE ESTIMATION: THE IRIDIUM CONSTELLATION CASE To assess the main performance of the concept presented a preliminary mission analysis has been carried out. It is based on the assumption to remove a target debris list, where the order of the objects to deorbit is not given in advance. As reference case it is assumed to remove the whole IRIDUM constellation. 70 objects with an equal dry mass of 556 kg of the telecommunication constellation are considered. They orbit in almost circular orbits at about 777 km altitude and are split into six equally spaced orbital planes inclined around 86.4 degrees. Figure 1 presents the orbital parameters of the object considered. 2. THE SATELLITE SWEEPER APPROACH The idea behind the satellite sweeper approach is quite simple: as for common rubbish several garbage bins are emptied into a single garbage truck and only the garbage truck goes to the garbage dump, in the same way it is better to collet several debris into a single spacecraft and perform only a final deorbiting maneuver instead of providing each single debris of its own deorbiting system. This idea is feasible if the Space Garbage Truck (SGT) is located at the beginning of the mission in the vicinity of the target debris (i.e. more or less in the same orbit) and the collecting task is Figure 1: Orbital parameters of the 70 objects of the IRIDIUM constellation considered. From the left: orbital semi-major axis, eccentricity, inclination and RAAN The preliminary assessment of the deorbiting system performance is based on the study of the missions required to the Waste Collector Robots to catch all the objects and bring those back to the Space Garbage Truck. The SGT is assumed to be placed at the same orbit of the object of 2/8

constellation and it is not considered any change of its orbit. The WCR so is in charge of moving between the SGT and the most convenient debris of the list. To identify the specific target an iterative numerical scheme has been implemented based on the delta V requirement to reach the other objects. The delta V estimation is based on the Homann transfer with an additional burn for the inclination change at the apocenter of the transfer orbit. For the specific case under investigation, since the objects are all placed in rather similar orbits several missions are actually rephrasing missions. Two different approaches have been considered for the RAAN change. Changing the RAAN is a rather expensive maneuver, especially when it is required to move the WCR from a given set of IRIDIUM targets placed almost in the same orbital plane to another set; up to 126 deg of RAAN change might be required (clearly too high for a practical single multi-target mission). In LEO, however the J2 influence can be properly exploited to move the WCR RAAN. The RAAN variation due to the Earth oblateness is dependent from the orbital altitude and since for the specific constellation considered the orbits have a very similar altitude, multiple revolutions might be required to reach the target RAAN. This causes long waiting times (where the WCR is docked to the SGT) that can be however avoided by performing propulsive RAAN change maneuvers. Both scenarios are considered for this preliminary assessment. During the WCR missions also the atmospheric drag is taken into account. It acts against the propulsive thrust during the outward orbits (thus requiring an increased propellant mass), whereas it helps the transfer when the WCR moves toward the Earth. The cross section area of the WCR is assumed to be the one corresponding to its solar panels. 1 kw of on board power is considered (for the chemical case) with a power density of 70 W/m 2. Once docked with the debris, instead the cross sectional area is determined as the maximum between the one of the WCR and the one of the debris. For the IRIDIUM case, an average cross sectional area of 4.75 m 2 is considered, where it is assumed that the two IRIDIUM solar panels are not aligned against the velocity vector. One of the most critical aspects of the whole mission concept is related to the docking and undocking operations. The WCR, indeed, has to undock from the SGT, dock with the debris and redock with the SGT delivering the debris. The close proximity operations have to be autonomous and possibly also compliant with potential noncooperative objects. For a preliminary assessment a fix delta V of 20 m/s is assumed for the docking and undocking operations. In the baseline chemical scenario the WCR is equipped with a bipropellant propulsion system delivering up to 300 N thrust with 290 s Isp. The initial WCR mass is set at 1975 kg with a dry mass of 1600 kg (see Sec. 3 for details). A WCR mission is considered completed when the whole propellant mass has been depleted. At this point there are two possible options: or the SGT provides a WCR refueling or the WCR docks with the SGT after its last mission and the whole system performs the controlled re-entry. This affects the number of launches to be performed and affects the mass of the SGT, but does not change the in-orbit performance of the WCR. Under these assumptions, Figure 2 shows that 12 WCR missions are required to completely deorbit the IRIDUM constellation. Figure 2: Relevant figures (total semi-major axis change, inclination change, delta V, propellant mass, number of object deorbited and mission duration) for the 12 missions required to deorbit the whole IRIDIUM constellation with the ADR approach proposed (natural RAAN change) Beside the first mission that is able to bring 8 debris to the SGT and the last one that is required to catch only the last two objects (where a relevant inclination change of about 1.1 deg is required), a typical WCR mission is able to catch and bring 3/8

down to the SGT 6 debris. This would correspond to a SGT final payload of 3336 kg, which requires a SGT of about 6.8 ton to be placed in LEO (see Sec. 3). Since this mass value is compatible with medium class launchers (e.g. Soyuz, Falcon 9 etc.), it is assumed to conclude each deorbiting mission after 6 debris are collected into the SGT. As a consequence the refueling of the WCR is not required and after the last debris is brought back to the SGT the WCR re-docks and is deorbited together with the SGT and the 6 IRIDIUM satellites. A typical mission requires 301 kg of propellant, thus about 74 kg of propellant are left into the WCR for contingency and for additional maneuver requirements during the docking and undocking phases. Such a margin is a consequence of the value assumed for docking/undocking operations (20 m/s); changing this value, indeed, it is possible to have a better exploitation of this remaining propellant. With 5 m/s, for instance, 7 debris can be deorbited in a typical mission with 351 kg propellant usage (and only 24 kg left as unused margin). A typical mission delta V is around 429 m/s corresponding to about 71.5 m/s for a single IRIDUM object deorbiting mission. As shown in Fig. 3 the specific propellant mass per debris changes typically between 46 and 54 kg, although it reaches 102 kg for IRIDIUM 96 where a significant inclination change is required (see Fig. 1). Almost all missions require about 1.68 hours of transfer; however, in this scenario of natural RAAN change, up to 50.1 days might be required to obtain the proper RAAN alignment. Such a long idle time is mainly due to the close proximity of the orbits of the target objects. undocking operations. It can be noticed how the code autonomously leaves the most demanding missions (where significant inclination and/or RAAN changes are required) as last missions, when no more other missions, with lower delta V, can be performed. Figure 4: Time evolution of WCA semi-major axis, inclination, RAAN and propellant mass for each of the 12 missions to remove the IRIDIUM constellation with the ADR approach proposed (natural RAAN change) All in all, as presented in Fig. 5, the deorbiting of the whole IRIDIUM constellation requires approximately 2.89 years with a cumulative firing time of only 5 days. This long overall mission duration is again related to the idle times required for a proper RAAN alignment. Such a time is spent in the SGT orbit, the lowest one among the debris where also the J2 influence is higher. Figure 3: Relevant figures (transfer time, propellant mass, delta V and waiting time) to remove each IRIDIUM satellite with the ADR approach proposed Figure 4 shows the evolution of semi-major axis, inclination, RAAN and propellant mass for each WCR mission. The propellant mass discontinuities represent the mass required for the docking and Figure 5: Time evolution of WCA mass, semi-major axis, inclination and RAAN along the whole 12 missions to remove the IRIDIUM constellation with the ADR approach proposed (natural RAAN change) As already pointed out, the RAAN can be changed also by means of a propulsive maneuver. In this 4/8

case a higher propellant mass requirement per mission is expected (thus a larger number of total missions) but also a much shorter overall mission duration (actually the same of the firing time). Fig. 6 shows the mass, semi-major axis, inclination and RAAN change for the active RAAN change scenario. Indeed, the RAAN evolution is not continuous, differently from inclination and semimajor axis. SGT attitude control and orbit maintenance and for refueling the WCA. Figure 6: Time evolution of WCA mass, semi-major axis, inclination and RAAN along the whole 12 missions to remove the IRIDIUM constellation with the ADR approach proposed (powered RAAN change) In this scenario 34 missions are required and typically 2 IRIDIUM debris are brought to the SGT per mission. In this scenario the possibility to refueling the WCA would increase the overall mission return. Indeed the same SGT can be designed for both scenarios (powered and natural RAAN change) with the final aim to deorbit 6 debris, considering refueling capabilities only in case a short overall mission time is sought. Figure 7 shows that the code selects missions with increasing delta V, but in many cases the proposed WCA configuration (with 375 kg of propellant) is not able to deorbit neither a single debris. In particular the great majority of missions require a significant RAAN change maneuver [10] and if such a scenario is sought a modified WCA design is required. For the final burn for the controlled re-entry of the SGT filled with the collected debris a bipropellant thruster with 325 s Isp has been considered. Starting from an apogee at 7150 km, Tab. 1 shows the delta V and the propellant mass required to target a final perigee of 60 km. This altitude is considered for a direct re-entry trajectory and requires having a flight path angle at the 120 km below -1, assuring a narrow footprint before the final braking maneuver. Table presents several scenarios from a single up to six IRIDIUM debris deorbited with the SGT. The propellant required for the final re-entry is assumed to be stored in the same tanks of the bi-propellant thrusters used for Figure 7: Relevant figures (total semi-major axis change, inclination change, delta V, propellant mass, number of object deorbited and mission duration) for the 34 missions required to deorbit the whole IRIDIUM constellation with the ADR approach proposed (powered RAAN change) Nr of IRIDUM Initial Delta V Propellant satellite deorbited mass [kg] [m/s] mass [kg] 1 6480 198.2 390.7 2 7005 198.2 422.3 3 7530 198.2 453.9 4 8055 198.2 485.6 5 8580 198.2 517.3 6 9105 198.2 548.9 Table 1: initial mass, delta V and propellant mass requirement for the final controlled reentry of the SGT with a different number of IRIDUM satellites stored onboard As shown in the worst case (maximum payload of 6 IRIDIUM satellites) a fuel margin of 42 kg is considered in the SGT design. 5/8

4. LOW THRUST EXTENSION A further option for improve the WCA in-orbit performance is to replace the chemical propulsion system with an electric one. A 5 kw Hall effect thruster is considered as baseline, with a nominal thrust of 330 mn and 1700 s Isp. The numerical algorithm has been modified [10] by using the Edelbaum low thrust optimization algorithm; a simplified method for optimising continuous low thrust transfers between two circular inclined orbits [11]. Atmospheric drag, J2 effect and thruster switch off during eclipse periods have been included in the numerical scheme. In this case the actual firing time is around 5-6 days per debris, but it is compensated by a much smaller propellant mass consumption, typically below 10 kg (besides for IRIDIUM 96 where 1.1 deg inclination change are required), see Fig. 8. Figure 9: Relevant figures (total semi-major axis change, inclination change, delta V, propellant mass, number of object deorbited and mission duration) for the 3 missions required to deorbit the whole IRIDIUM constellation with the low thrust extension of the ADR approach proposed (natural RAAN change) Figure 8: Relevant figures (transfer time, propellant mass, delta V and waiting time) to remove each IRIDIUM satellite with the low thrust extension of the ADR approach proposed (natural RAAN change) In this case, since the transfer time for each single mission is much longer, the RAAN natural variation requires much longer times to match the target RAAN values. Indeed up to two years might be required for some specific objects, but an active RAAN changing maneuver can always be considered. In this electric propulsion scenario only three missions are sufficient (with the same WCA sizing considered for the chemical case) to completely deorbit the IRIDIUM constellation. Figure 9 presents the relevant figures for the three missions: up to 35 debris can be brought to the SGT per mission. Assuming however a maximum of 6 debris for the final controlled re-entry, the WCA can be sized differently making it significantly lighter/smaller and with less propellant to assure a maximum of 6 debris collected per mission without refuelling. The last mission is required to deorbit only the last constellation object (the out-of-plane IRIDIUM 96) lasting about 198 days with 14.5 days thrusting time and 24.7 kg propellant mass. Figure 10 shows the WCA mass evolution, semimajor axis, inclination and RAAN evolution during the three missions. Figure 10: Time evolution of WCA mass, semi-major axis, inclination and RAAN along the whole 3 missions to remove the IRIDIUM constellation with the low thrust extension of the ADR approach proposed (natural RAAN change) 6/8

5. PRELIMINARY SYSTEM SIZING A rough estimation of the mass budget of both the Space Garbage Truck and the Waste Collector Robot can be obtained using typical percentages of the different subsystem and assuming that the payload of the spacecraft is the debris to be deorbited. Typical value of the mass budget distribution has been obtained from open literature [12], choosing as reference case for the SGT the typical distribution of mass of the ATV cargo [13]. Table 2 summarizes the mass budget of the WCR used for the mission analysis performed in section 3 for the baseline chemical scenario. The overall mass of the WCR with the payload and the initial mass of propellant is roughly 2500 kg, however, its initial mass at the first undocking from the SGT is 1975 kg. Once the WCR has reached the target debris a grasping mechanism allows for capture it. A quite high value of mass budget for this system has been foreseen: 15% of the dry mass that in terms of mass is equal to 315 kg. This value represents only a first tentative assessment and future deeper analyses are necessary to better estimate the mass of this critical component. According to the mission analysis, the WCR can collect 6 IRIDIUM satellite without refuelling and the mass budget is fully compatible with a pressure regulated bipropellant propulsion system that confirms the feasibility of assuming the quite high value of the specific impulse (i.e. 290 s) used in the computations. A preliminary assessment of the mass budget of the SGT is reported in Table 3. This spacecraft is characterized by a quite high value of the payload percentage (37% of the dry mass) that is similar to the one of the ATV cargo [13]. It is worth noticing that the percentages of structure, TC, power, TT & C and ADCS are the same for the WCR and the SGT. At the end of the collecting mission, the mass of the SGT before the final deorbiting manoeuvre is roughly 9 ton with six IRIDIUM satellites as payload (see Table 3). The launch mass of the entire system comprises the sum of SGT overall mass (without the IRIDIUM payload), the WCR propulsion system and propellant mass and the grasping mechanism. Therefore, the overall launch mass is around 6800 kg. Even in this case, the mass budget is fully compatible with a pressure regulated bipropellant propulsion system with a thrust level higher than 300 N that confirms the feasibility of assuming the very high value of the specific impulse (i.e. 325 s) used in the computations. Table 2. WCR mass budget. WCR BUDGET WET [%] DRY [%] [kg] PAYLOAD (1xDEBRIS) 21 25 525 GRASPING MECHANISM 12,6 15 315 STRUCTURE 16,8 20 420 THERMAL CONTROL 2,52 3 63 POWER GENERATION SYSTEM TELEMETRY TRACKING AND COMMAND ATTITUDE DETERMINATION CONTROL SYSTEM 15,96 19 399 3,36 4 84 6,72 8 168 PROPULSION SYSTEM 5,04 6 126 PROPELLANT 15 N/A 375 SYSTEM LEVEL MARGIN 1 N/A 25 SGT BUDGET Table 3. SGT mass budget. WET [%] DRY [%] [kg] PAYLOAD (6xDEBRIS) 34,595 37 3150 STORING MECHANISM 6,545 7 596 STRUCTURE 18,7 20 1702 THERMAL CONTROL 2,805 3 255 POWER GENERATION SYSTEM TELEMETRY TRACKING AND COMMAND ATTITUDE DETERMINATION CONTROL SYSTEM 17,765 19 1618 3,74 4 341 7,48 8 681 PROPULSION SYSTEM 1,87 2 170 PROPELLANT 6,25 N/A 568 SYSTEM LEVEL MARGIN 0,25 N/A 23 6. CONCLUSIONS A preliminary analysis of feasibility of the Satellite Sweeper Approach for the solution of the space debris problem has been successfully performed. The application of this concept to a specific test case (the deorbiting of the whole IRIDIUM constellation) has confirmed the possibility of collecting several debris and deorbiting them in a single controlled re-entry maneuver performed with a high specific impulse storable bipropellant propulsion system (not feasible with an electric propulsion system). Moreover, the launch mass of the Space Garbage Truck obtained from the 7/8

mission analysis and the mass budget estimation is fully compatible with medium class launchers. The same approach can be used for different classes of debris providing them to be of similar shape or interface in order to simplify the design of both the WCR and the SGT. 7. ABBREVIATIONS AND ACRONYMS ADCS = Attitude Determination Control System SGT = Space Garbage Truck TC = Thermal Control TT & C = Telemetry Tracking and Command WCR = Waste Collector Robot 8. REFERENCES 1 Jenkin, A. B., Sorge, M. E., et al (2011). 100- Year Low Earth Orbit Debris Population Model. AAS 11-410, AAS/AIAA Astrodynamics Specialist Conference, Girdwood, Alaska. 2 Peterson, G. E. (2011). Effect of Future Space Debris on Mission Utility and Launch Accessibility. AAS 11-414, AAS/AIAA Astrodynamics Specialist Conference, Girdwood, Alaska. 9 Pratt S. R., Raines R.A., et al. (1999). An Operational and Performance Overview of the IRIDIUM Low Earth Orbit Satellite System. IEEE Communication. Surveys, vol. 2, no. 2. 10 Pergola P., Andrenucci M. (2015). Electric Propulsion Tug for Multi-Target Active Space Debris Removal Missions, 66th International Astronautical Congress, Jerusalem, Israel, IAC-15- C4.6.9. 11 Kechichian J. A. (1997). The Reformulation of Edelbaum s Low-thrust Transfer Problem using Optimal Control Theory. Journal of Guidance, Control, and Dynamics, vol. 20, no. 5. 12 Wertz, J., Everett, D., et al., (2011). Space Mission Engineering: The New SMAD. Space Technology Library, Microcosm Press. 13 ATV Edoardo Amaldi, ATV 3 fact sheet, (2012). ESA Erasmus Centre, Directorate of Human Spaceflight and Operations, ESA-HSO- COU-023 3 Kessler, D. J., Johnson, N. L. (2010). The kessler syndrome: implications to future space operations, 33rd Annual American Astronautical Society, Rocky Mountain Section, Guidance and Control Conference. 4 Committee on Peaceful Uses of Outer Space (COPOUS). (2006) Scientific and technical subcommittee, working group on Space debris: Progress report of the working group on space, U.N. Doc. A/AC.105/C.1/L.284. 5 Interagency Space Debris Coordination Committee: IADC space debris mitigation guidelines. (2007). rev 1. 6 Janovsky R., Kassebom M., et al. (2004). Endof-life de-orbiting strategies for satellites. Science and Technology Series 109. 7 Peterson, G. E. (2012). Target Identification and Delta-V Sizing for Active Debris Removal and Improved Tracking Campaigns. 23rd International Symposium on Spaceflight Dynamics, Pasadena, California, Paper No. ISSFD23-CRSD2-5. 8 Weeden, B. (2010). 2009 Iridium-Cosmos collision fact sheet. Secure World Foundation. 8/8