Clean-up your space: INTEGRAL low cost end-of-life disposal

Transcription

1 SpaceOps Conferences May 2016, Daejeon, Korea SpaceOps 2016 Conference / Clean-up your space: INTEGRAL low cost end-of-life disposal Jutta M. Hübner 1, Richard Southworth 2, Klaus Merz 3 ESA/ESOC, Darmstadt, Germany Dave Salt 4 TPZ VEGA, Darmstadt, Germany Claudia Dietze 5 CS, Darmstadt, Germany Alastair McDonald 6 CGI, Darmstadt, Germany Gerald Ziegler 7 SCISYS, Darmstadt, Germany Stefano De Padova 8 Serco, Darmstadt, Germany Space debris in orbit poses a serious threat: Even a tiny screw or a piece of paint travelling at very high velocity can cause major damage if it hits a satellite. Any single impact can cause a cascade of collisions and eventually lead to a domino effect: Each collision generating space debris increases the likelihood of further collisions creating even more space debris. Up to now the growth of the debris population has been dominated by object break-ups (intentional or unintentional). However, the Iridium- Cosmos collision illustrates the risks associated with debris population growth driven by the collisional domino effect. The numerous close approaches, near-misses and the increasing risk of collisions have prompted space agencies across the world to selfcommit to guidelines aiming to limit space debris population growth and the resulting collision threats. The Gamma-Ray observatory INTEGRAL with its 3.5 tons is ESA s heaviest science mission ever flown. Analyses have shown that once out of fuel, INTEGRAL would not re-enter during the next at least 200 years without intervention from ground. In this long period, it would initially pose a modest debris hazard in the protected sub 2000 km and the geostationary zone but could eventually become a major debris contributor in case of a collision. Therefore, despite the fact that INTEGRAL was launched before the ESA space debris guidelines were formulated, an investigation of potential disposal options has been performed several years after launch including graveyard orbit, disposal manoeuvre, break up and re-entry analyses. This paper summarises the different disposal options as well as the disposal concept and strategy ultimately chosen via parametric numerical propagations and break-up analyses. Furthermore, the in-depth planning of the disposal manoeuvres and their successful execution in 2015 are presented. Finally, the detailed assessment of the resulting consequences for the scientific mission and the routine mission operations are discussed including the re-entry characteristics and possible fine-tuning of the re-entry longitude, position and profile in the next decade. 1 Spacecraft Operations Engineer, HSO-OAI, jutta.huebner@esa.int 2 Head INTEGRAL Spacecraft Operations Unit, HSO-OAI, richard.southworth@esa.int 3 Space Debris Analyst, OPS-GR, klaus.merz@esa.int 4 Spacecraft Operations Engineer, OPS-OAI, dave.salt@esa.int 5 Flight Dynamics Engineer, OPS-GFE, claudia.dietze@esa.int 6 Flight Dynamics Engineer, OPS-GFE, alastair.mcdonald@esa.int 7 Flight Dynamics Engineer, OPS-GFE, gerald.ziegler@esa.int 8 Spacecraft Operations Engineer, HSO-OAI, stefano.de.padova@esa.int 1 Copyright 2016 by ESA. Published by the, Inc., with permission.

2 E I. Introduction SA s INTErnational Gamma-Ray Astrophysics Laboratory short INTEGRAL was launched into a highly elliptical orbit more than a decade ago in October 2002 and has been observing the high-energy sky simultaneously in gamma-rays, X-rays and visible light ever since with its unique payload the imager IBIS [1], the spectrometer SPI [2,3], as well as the complementary X-ray and optical cameras JEM-X [4] and OMC [5]. INTEGRAL is able to perform multispectral observations of the most violent and exotic objects in the universe, such as gamma ray bursts and supernova explosions [6,7]. With its 3.5 tons, INTEGRAL is ESA s heaviest science mission ever flown. Today, eight years beyond the design lifetime of 5 years (2.5 years nominal lifetime and the extended lifetime of another 2.5 years), it is still the most sensitive gamma ray observatory in space with a performance far exceeding design specifications and the interest from the scientific community in observing with INTEGRAL remains very high. On the basis of the originally planned operations, the lifetime-limiting fuel on board would have allowed for operations well into the next decade. Once out of fuel, INTEGRAL would not naturally re-enter for at least 200 years and could eventually become a major debris contributor in case of a collision. Beyond the 200 year period, the orbital evolution becomes difficult to predict due to accumulated uncertainties. This paper summarises the different disposal options, the planning and successful execution of the disposal manoeuvres implemented and illustrates the potential of extending INTEGRAL s mission well into the next decade - despite having spent extra fuel on disposal manoeuvres - including re-entry characteristics and the possible fine-tuning of the re-entry longitude, position and profile in the next decade. II. Review of the different disposal options In this section, different disposal options to avoid that INTEGRAL could eventually become a major debris contributor due to collision are summarised including graveyard orbit, disposal manoeuvre, break up and reentry analyses. Background: INTEGRAL long term orbital evolution INTEGRAL was launched in its initial orbit in Since then, only the longitude of perigee is maintained to control ground station coverage. The orbit has been allowed to evolve naturally which to a large extent was favourable for the mission operations. The INTEGRAL orbit is subject to considerable variation: the natural INTEGRAL long-term orbital evolution mainly caused by the natural orbital perturbations (Earth's oblateness and the luni-solar gravitational perturbations) is illustrated in Figure 1. The perigee altitude and orbital inclination varies enormously, occasionally leading to perigee altitudes as high as km and below 2000 km and to repeated crossings of the protected GEO and LEO regions. Analysis by the Space Debris Office in ESOC shows that this behaviour will continue for the next 200 years at least. This orbital evolution results in perigee altitudes below 6000 km, where INTEGRAL passes through the inner Van Allen radiation belts and therefore is exposed to trapped proton radiation at perigee. This poses a threat to the spacecraft component s performance. In addition, the orbital evolution results in visibility changes affecting ground station coverage [8]. Besides the effects of low perigee altitudes on INTEGRAL, INTEGRAL presents an occasional modest debris hazard in the protected sub 2000 km zone and the geostationary zone. Despite the fact that INTEGRAL is not covered by the space debris guidelines that came into force well after its launch in 2002 (a European code of conduct for space debris mitigation [9] was signed by ESA in 2006 and on this basis ESA developed its own requirements for space debris mitigations [10] that entered into force in 2008), investigations on disposal options for INTEGRAL were performed and analysed by ESA s Space Debris Office in 2013/

3 Figure 1: INTEGRAL long-term evolution of perigee altitude and orbital inclination in the period The perigee altitude varies enormously mainly caused by the Earth's oblateness and the lunisolar gravitational perturbations. When the perigee altitude drops below 6000 km, INTEGRAL passes through the inner proton belts. Constraints and requirements for the disposal The analysis of disposal options for INTEGRAL was based on projections of the available remaining propellant and the expected evolution of the orbit over time. Ideally, the total remaining potential science operations lifetime shall be maximised. Besides propellant consumption, further drivers for the disposal options were robustness as well as safety and risks. As a guideline for risk threshold calculation, the ESA Space Debris regulations were referenced, although not strictly applicable. Propellant budget and margin Due to the accurate insertion into its final orbit by the launcher as well as its efficient fuel usage and minimal number of safe modes (only 5 in 13 years), INTEGRAL still has a significant fuel margin. The INTEGRAL propulsion system uses hydrazine as a monopropellant, pressured by nitrogen gas in blow-down mode and separated from it by a flexible membrane. The remaining propellant within the spacecraft s four tanks can be estimated via three independent methods: Thruster usage book-keeping: By accounting for the duration of each thruster valve s actuation, ESOC Flight Dynamics (FD) calculated the exhausted mass flow via a model of the thruster s performance as a function of feed-line pressure and temperature. PVT method calculation by FD: estimation of the remaining propellant volume within each tank using derived values of each tank s pressure and temperature to calculate the nitrogen pressurant s volume via real gas law equations Independent PVT method calculation by industry (TAS-I) [11] All above mentioned estimates of the remaining propellant as per 1/5/2014 closely agreed: FD book-keeping: kg ± 10 kg FD PVT: 98.8 kg ± 18 kg Industry PVT: kg ± 20.3 kg These sets of residual propellant estimates are considered a sound basis to assess the possible future disposal strategies. In 2014, with an annual consumption in the order of 6 kg, this gave a remaining lifetime of another 15 years even allowing for some uncertainties. 3

4 INTEGRAL disposal scenarios Several disposal scenarios were analyzed in detail by the Space Debris Office in ESOC which focused on the long-term orbit evolution, orbit raising and de-orbit opportunities. A range of possible options was outlined and parametric numerical propagations were performed to identify an optimal solution [12] : Earth s re-entry trajectories: Direct or indirect force of an atmospheric re-entry by performing delta-v manoeuvres opposite to the direction of the velocity vector at perigee to lower the apogee or at apogee to lower the perigee Graveyard orbits: Long-term raise of the perigee to avoid the 2000 km protected region about LEO by performing delta-v manoeuvres at apogee in the direction of the velocity vector (also including moon escape options and resonance cases of the satellite s and the moon s orbital period) These simulations were accompanied by a large set of numerical propagations [13]. A methodology was developed aimed at reducing the amount of propellant needed. This mathematical multi-objective optimisation problem was solved by means of an evolutionary algorithm based on biological evolution. The computational effort was limited by employing a semi-analytical propagator. At the beginning of the analyses, it was shown, that no feasible option for a direct forced, i.e., controlled reentry has been found. Therefore, indirect, uncontrolled re-entry scenarios were further studied. Options have been identified leading to a long-term clearance of the LEO region for 200 years or to re-entry into the Earth s atmosphere within the next 200 years. Typically, both options require a large manoeuvre rather early, i.e. before Indications suggest that a disposal by re-entry may be preferred while maintaining enough fuel for science operations for at least the rest of the current decade. Re-entry considerations and characteristics Analysis by ESOC Space Debris Office showed that the propellant cost of executing a disposal manoeuvre increases with time over the next few years until mid 2017: up to mid 2017, manoeuvres with the currently available propellant are possible, leading to a re-entry in with some cases leaving fuel for further science mission extension. By late 2017, it becomes impossible with the predicted remaining propellant to re-enter Earth s atmosphere [12]. While a re-entry before 2029 is not possible with the remaining propellant, a re-entry significantly later than 2029 is possible but will lead to a higher casualty risk since the latitude of the re-entry is much less controllable and predictable so far into the future. Limitations of the simulated re-entry characteristics In the simulations and propagations, no burn attitude or ground station coverage constraints have been taken into account and the manoeuvres have been considered as instantaneous delta-v burns performed at either apogee or perigee, i.e., not split over several orbits or with a finite duration. The latter is assumed to be noncritical at this stage, but has an impact on the actual fuel consumption. The former has to be analysed in more detail once a particular strategy is envisaged, in particular if this leads to the situation that manoeuvres have to be performed at true anomalies significantly different from perigee or apogee. Break-up analysis A perigee decay rate analysis was performed to determine to which extent the orbit circularizes and fragments spread in latitude. For the scenarios involving re-entry, a break-up model using the tool SCARAB was developed by industry [14] allowing ESA to quantify the on ground risk and helping to determine the re-entry location provided the perigee altitude decay rate is large enough. Break-up analyses of INTEGRAL were performed for different re-entry scenarios providing information regarding the number and mass of the fragments reaching the Earth s surface. Simulations were carried out for a range of last perigee altitudes, i.e. first entry into the Earth s atmosphere, between 50 km and 117 km. In all re-entry cases, fragments of INTEGRAL reach the ground with a total mass of between 200 kg and 1000 kg. The affected surface area is with typically less than 25 m 2 comparably small. The most critical finding was that the steepness of the re-entry trajectory critically determines the survivable fragments reaching the ground: a shallower re-entry will result in an atmospheric contact over several consecutive revolutions and lead to fewer fragments surviving. However, the resulting effect will be the circularization of the orbit, which as a consequence will cause a larger spread of the impact latitude. Therefore, a trade-off analysis of the size of the debris footprint versus the surviving mass has to be performed [15]. 4

5 III. Selected disposal strategy In this chapter, the disposal concept and strategy ultimately chosen via numerical propagations and break-up analyses are illustrated. A more detailed assessment and refined analysis were required before an appropriate disposal strategy could be selected as a credible candidate: delta-v losses related to extended manoeuvre burn times were taken into account as well as attitude off-pointing requirements resulting from spacecraft operational constraints and nonoptimal manoeuvre location resulting from limited ground station availability [15,16]. The favoured option was to perform an apogee lowering manoeuvre at perigee in early 2015 to amplify natural orbital third body perturbation effects. This would lead to an atmospheric re-entry in 2029 (see Figure 2) [15]. This disposal option would guarantee an ESA Space Debris policy compliant re-entry while using less than half of the remaining propellant at that time and therefore maintaining sufficient propellant to ensure continued science operations for a reasonable number of years after a commitment to this strategy. Both the date and latitude of the re-entry are very predictable even more than 10 years before the actual re-entry. For all cases, the re-entry location of most fragments following the break-up will fortunately be in areas of very low population density over the Southern hemisphere: the latitude of most fragments will be in the range of -45 and -70 (see Figure 3). A casualty risk estimate is below 2.5e-05 for all cases completed, which is well compliant with new missions where the ESA policy risk is < 1.0e-04 [14]. Figure 2: INTEGRAL long-term evolution of the perigee altitude with (green) and without (red) a disposal manoeuvre in January Figure 3: Potential location of ground impact of most INTEGRAL fragments following re-entry. 5

6 IV. Detailed planning and execution of the disposal manoeuvres In this section, the feasibility of the chosen disposal option is assessed and the in-depth planning and successful execution of the disposal manoeuvres are presented. Defining the target orbit and the date for the disposal manoeuvre In 2014, the FD team at ESOC performed an analysis to determine the target orbit acceptable for future operations and designed a manoeuvre campaign to achieve this target orbit [16]. The propellant consumption of the disposal manoeuvre to achieve re-entry in 2029 critically depends on the disposal manoeuvre date: it increases with time - the earlier the apogee lowering the less propellant is needed (see Figure 4). For this reason, it was important to execute the disposal manoeuvre as early as possible, possibly already in early Once executed, INTEGRAL will re-enter in 2029 regardless of any future operations of the satellite. Date Delta-v [m/s] needed for re-entry Jul Oct Jan Apr Jul Oct Jan Figure 4: Evolution of re-entry delta-v requirement with time [15] : a manoeuvre in early 2015 achieves reentry with a delta-v of roughly 24 m/s, whereas by mid 2015 the required impulse would have been more than 26 m/s. By late 2017, it becomes impossible with the predicted remaining propellant to perform a reentry manoeuvre. Requirements on the target orbit: The requirements on the selected target orbit were as follows: The maximum apogee radius was provided by ESOC s Space Debris Office The resulting orbit shall be a repeat orbit for maintaining a recurring ground station coverage pattern to ease mission planning have an optimal phasing, i.e., tune the target longitude to maximise ground station coverage above the radiation belts The resulting science operation time during the INTEGRAL revolution post delta-v campaign shall be maximized. Operational constraints on the disposal manoeuvres to achieve the selected target orbit: The requirements on the disposal manoeuvre to ease the disposal operations at the same time were as follows: Seasonal constraints: Eclipse season is excluded for reasons of operational complexity and robustness. Ground station constraints: Ground station coverage is mandatory during the execution of the manoeuvres as well as before and afterwards for preparatory and post-burn activities. Burn strategy: Due to restrictions to the manoeuvre placement and attitude a long duration single manoeuvre would be quite inefficient. Therefore, to increase efficiency, the burn shall be split into two or more burns over two or more perigee passes. Nevertheless, splitting the manoeuvre further increases the planning complexity. Attitude constraints during the actual manoeuvres: The periods when these burns can be scheduled will also be subject to off-pointing constraints during perigee passage, which are driven by two basic requirements: To prevent the Earth or Moon passing through or nearby the star tracker field of view in order to avoid star tracker blinding. To ensure that the angle of the Sun with respect to the solar arrays, i.e. the solar aspect angle, does not exceed 5 in roll and 40 in pitch. Maximum science during the delta-v campaign: The disposal activities shall be planned with a minimum impact on science operations. The science return during the delta-v campaign shall be maximised. Therefore, as far as possible routine mission planning processes shall be used during the manoeuvre campaign to continue with routine science observations. 6

7 Choice of repeat orbit: Possible target orbits with different repeat cycles were analysed regarding their fuel efficiency and repeat cycle length (see Figure 5). The solution for the shortest repeat cycle of 3 revolutions in 8 sidereal days was selected, i.e., each orbit lasting 2 2/3 sidereal days, since it is closest to the minimum effective delta-v possible and therefore increasing the post-manoeuvre science lifetime. Furthermore, this shortest repeat cycle simplifies the mission planning process. Delta-v [m/s] Impulsive at perigee Repeat pattern [orbits / sidereal days] Apogee radius [km] 0 (original) 1/3 (original) 160,310 87, ,870 81, /8 147,064 81, /13 144,350 79, /18 14, ,139 Figure 5: Possible repeat orbits after the disposal manoeuvres [16]. Semi-major axis [km] Choice of phasing: The target longitude was selected such that the station visibility of the Kiruna ground station above the radiation belts is maximised over the three revolutions of the repeat orbit. For this repeat orbit, the longitude at perigee is shifted by 120 for each of the three revolutions of the repeat cycle. A visibility analysis by Flight Dynamics was performed and a longitude of 105 was targeted at every third perigee [16]. Re-confirmation of the re-entry: The chosen target orbit with a repeat cycle of 3 orbits in 8 days and a phasing of 105 was re-assessed by the ESOC Space Debris Office. It was confirmed, that for this target orbit, re-entry of INTEGRAL is achieved in Disposal manoeuvre characteristics Taking all operational constraints to achieve the selected target orbit as summarised in section IV.1 into account, the window for the disposal manoeuvre execution was restricted to January to March Additionally, the manoeuvre attitude constraints relative to Sun, Earth and Moon and the requirement on ground station coverage had to be respected and result in a slight loss in efficiency. Three burns plus touch-up burn strategy: Initial assessments suggested that delta-v manoeuvres performed in early 2015 at perigee in order to lower the apogee to ensure re-entry in 2029 should have a burn magnitude of 24 m/s (see Figure 4) assuming they are executed impulsively exactly at perigee in the exact opposite direction to the velocity vector of the satellite. However, the actually required delta-v respecting all the constraints is much higher than the magnitude indicated by the numerical analysis carried out. he necessary delta-v to achieve the target orbit has a duration of >45 min. This results in an additional loss in efficiency due to both, the manoeuvre not being executed at perigee and the velocity vector changing direction during the manoeuvre, since the manoeuvre inertial attitude is fixed. Therefore, the manoeuvre should be split into several burns. One of the main challenges was to perform a trade-off of manoeuvre efficiency vs. deviation from the perfect manoeuvre attitude as well as the time of the manoeuvre from perigee. In particular, this was the case for the first burn where there was no ground station coverage at perigee. Furthermore, the science operation time post delta-v campaign shall be maximised. To minimise the overall losses in efficiency while still respecting all constraints, a three burn plus touch-up burn strategy was chosen: Burn #1: Initial burn used to adjust the ground station coverage, i.e., favourable longitude over perigee Burn #2: Major part of the disposal manoeuvre Burn #3: Remaining part of the manoeuvre used to achieve a repeat orbit Burn #4 if needed: Touch-up burn to compensate for any misperformance of burn #3 After adjusting the ground station coverage with burn #1 that had to be performed before perigee under ground station coverage, the subsequent manoeuvres could be executed at perigee. Except for the first manoeuvre all burns therefore required an off-pointing from the optimal attitude (see Figure 6) to avoid attitude constraints, i.e., to postpone the Earth blinding after perigee. 7

8 Risk mitigation Delay of delta-v activities: For each delta-v manoeuvre, a back-up opportunity was available at one of the subsequent perigee passages in case they were missed due to ground station or operational issues. There was sufficient margin to allow two opportunities to be missed. This would have caused a delay of several days in the overall disposal orbit acquisition. Nevertheless, a delay of the complete disposal manoeuvres campaign past the early 2015 window would have implied significant fuel cost. Furthermore, it would then have been impossible to achieve the repeating 3 revolutions in 8 day orbit which was targeted for the science operations post disposal manoeuvre. Double ground station coverage: For robustness, double ground station coverage was scheduled for the period of the delta-v operations wherever possible. Misperformance of manoeuvres: Possible back-up strategies and contingency plans were prepared by Flight Dynamics to compensate for any impact due to manoeuvre non-performance or severe mis-performance. Compensating for any significant misperformance will also have an impact on the fuel budget. Disposal manoeuvre campaign Extensive mission and operational planning was required to allow for a successful disposal manoeuvre campaign. The mission planning process and resulting products had to be adapted to the special requirements of the disposal manoeuvres (see above). Test delta-v: Since the last INTEGRAL manoeuvres were executed during LEOP in late 2002, a full scenario test delta-v was performed. The objective was to train the team, re-gain experience of manoeuvres with a fully re-newed ground system, test the software as well as the manual delta-v operations including preparatory and post-burn activities. Furthermore, all manoeuvres performed during LEOP were executed at apogee, while the disposal manoeuvres were the first manoeuvres executed at perigee. Additionally, pre-launch formulated constraints by industry suggested specific issues related to star tracker use below km and slew performance below km [17]. The test manoeuvre was successfully performed in September Preparation of disposal manoeuvres: For every delta-v manoeuvre, the manoeuvre operations had to be performed manually and in real-time, since the delta-v manoeuvre attitudes became constrained shortly after the manoeuvre [16] and therefore cannot be scheduled via the routine mission planning process. Before every delta-v manoeuvre, preparatory activities had to be performed: slewing to the correct manoeuvre attitude as well as spacecraft configuration into thruster control mode and calibration of the sensors for the manoeuvres (IMU drift calibration). Execution of disposal manoeuvre: In early 2015, the unique opportunity was taken: INTEGRAL performed an apogee-lowering manoeuvre campaign that will lead to an ESA Space Debris Policy compliant re-entry of the spacecraft into the Earth s atmosphere in February 2029 with minimal propellant consumption. The manoeuvre campaign was executed successfully from January 12 th to February 12 th The disposal manoeuvres are summarised in Figure 6 including the planned characteristics, the results of the disposal delta-v campaign and the key operational and performance details. After the pre-burn configuration activities (see above), the manoeuvre operations were performed manually according to a pre-defined count-down sequence specified within the delta-v flight control procedure including monitoring of the manoeuvre. The actual execution of the delta-v was performed autonomously by the attitude control computer using its thrust control mode sequence. Due to the performance of the first three manoeuvres, a touch up delta-v was necessary. Post manoeuvre operation steps and activities: After every disposal manoeuvre, post-burn re-configuration activities had to be performed: AOCS had to be configured back for slews on reaction wheels. Furthermore, a slew back to an unconstrained attitude was performed. Post manoeuvre, the thruster firings were assessed by both the Flight Control Team and by Flight Dynamics and an orbit determination was performed. With the latter, the ground station predictions were updated if necessary. Finally, the resulting changes to the planning of the remaining manoeuvres were assessed and updated accordingly. 8

9 Execution Performance Orbit (post delta v) Disposal delta v # Revolution # Date (dd/mm/yyyy) 12/01/ /01/ /02/ /02/2015 Start time(z) 23:51:01 16:17:01 15:15:55 15:02:07 End time(z) 00:07:42 16:48:10 15:31:39 15:03:22 Duration actual (hh:mm:ss) 00:16:41 00:31:09 00:15:44 00:01:15 Expected nom. duration 00:15:21 00:28:32 00:14:09 00:01:20 Expected max. duration 00:19:18 00:35:07 00:17:52 00:02:50 Attitude off set (deg. in plane) Total thruster force mean (N) Delta V planned (m/s) Delta V estimated (m/s) Deviation (m/s) Propellant use predicted (kg) Propellant use measured (kg) Deviation (kg) Apogee (km) Perigee (km) Eccentricity Semi major axis (km) Period (hrs) 69:52:12 65:41:29 63:53:40 63:49:25 Inclination (deg) Right Ascension (J2000, deg) Declination (J2000, deg) Target longitude at perigee (deg) Achieved longitude at perigee (deg) (*) Figure 6: Summary of the results of the disposal delta-v campaign and key operational and performance details [15,16]. (*) The final target longitude of 105 was achieved by a combination of manoeuvre #4 and the subsequent routine reaction wheel biases. Manoeuvre performance: The spacecraft performed nominally during each of the four disposal manoeuvres, with all key telemetry parameters and their trends reflecting the expected behaviour throughout the whole set of operational activities. Analysis of the thruster on-times by Flight Dynamics indicated that the following propellant was consumed by the burn/during the whole manoeuvre: Mano #1: kg/ kg Mano #2: kg/ kg Mano #3: kg/ kg Mano #4: kg/0.318 kg The total usage was about 48 kg. Post disposal manoeuvre calibration campaign Following the successful delta-v campaign, the following AOCS calibrations were performed in the revolution after the last delta-v manoeuvre to ensure that INTEGRAL s AOCS performance is maintained after the delta-v manoeuvres resulting in significant changes in both propellant mass and tank pressure: a full thruster torque calibration an external torque calibration Additionally, monitoring and long-term planning activities had to be taken into account as well: After the manoeuvre campaign, ESOC s Space Debris Office performed a robustness analysis of the long-term evolution of INTEGRAL s orbit assuming an orbit control strategy to maintain a repeat pattern of 3 orbits in 8 days and the longitude at every third perigee in the band of 105 ± 5 [18]. A total of 290 Monte Carlo propagations were performed taking effects of potential variations in the orbit control and uncertainties in the perturbation modeling into account. All of the propagations lead to a re-entry close to the nominal time with differences less 9

10 than 2 days at the point when the altitude was low enough to make re-entry unavoidable, i.e., ignoring the final re-entry process. Therefore, it can be concluded that all trajectories will lead to a re-entry and break-up starting late February Science operations during the disposal manoeuvre campaign The disposal activities were planned with a minimum impact on science operations. In total, only about 68 hours science time were lost for the disposal manoeuvre campaign with the majority being consumed by postmanoeuvre calibrations. V. Consequences for the scientific mission and the routine mission operations The resulting consequences for the scientific mission and the routine mission operations were assessed in detail. The impact of the orbit changes on the future INTEGRAL mission operations and scientific programme execution was determined, any critical effects mitigated and the suitability of the mission planning system to continue the science mission verified. In this context, the following activities were performed: The scope for potential mission lifetime post disposal manoeuvres was assessed [19]. The science mission operations after the disposal manoeuvres were prepared. It was verified that the mission planning system can cope with the shorter post-disposal revolution length as well as the modified ground station coverage. VI. Collision avoidance close to re-entry The orbital evolution of INTEGRAL after the disposal manoeuvres will result in an increased duration of the presence in the sub 2000 km LEO region in 2020/2021 as well as 2028/2029 up to the re-entry in The Space Debris Office in ESOC assessed the potential risk associated with this and concluded that the accumulated collision risk is equivalent to less than that of two days of a typical polar LEO satellite on its operational orbit [20]. Consequently, no collision avoidance is considered necessary for this period. On the contrary, the strategy of a permanent disposal via re-entry of INTEGRAL offers a definitive solution with zero collision and break-up risk after the re-entry in 2029 and avoids all future crossing of the LEO and GEO protected region. VII. Re-entry in 2029 A trade-off analysis of the size of the debris footprint versus the surviving mass was performed. The main re-entry characteristics of the particular scenario chosen for implementation were simulated using the SCARAB tool and the results are summarised in the following [14] : 1 st contact with the atmosphere on 26/2/2029 Re-entry over 8 INTEGRAL orbits Predicted surviving (ground) mass: 266 kg Predicted number of fragments reaching the ground: 18 Predicted casualty risk: 3.0E-06 Predicted casualty area: 10.9 m 2 On-ground fragments: south of latitude -40 with parts from two fragments at -32 and -25 latitude The first contact of INTEGRAL with the atmosphere is predicted for February 26th, This represents the start of the re-entry. The on-ground casualty risk was computed with the tool Oriundo [21], taking into account the population density in 0.25 latitude bins around the re-entry location of the fragments. As in general for uncontrolled and eccentric re-entries, the longitude of the re-entry cannot be guaranteed but the latitude will remain fairly stable, the population density in the latitude bin is averaged out over the entire longitude circle. The overall risk is dominated by the two most Northern fragments, which contribute with < 3.0E

11 VIII. Fine-tuning of the re-entry in the next decade In this section, possible fine-tuning of the re-entry longitude, position and profile in the next decade is discussed. The disposal manoeuvre campaign in 2015 used approximately half of the propellant left at that time. After the campaign, a total of 47.8 kg of propellant was left, which at the rate of consumption at that time of around 16 g/day would support further INTEGRAL operations until October This simple linear projection assumes a margin of 10 kg of propellant to cover uncertainties in fuel bookkeeping and final propellant depletion. Other factors, like available power or financial constraints, may limit the spacecraft s useful lifetime to before that date. However, part of the remaining propellant could also be used for a later trim manoeuvre to better constrain the re-entry conditions and thus further reduce the on-ground casualty risk by reducing the re-entry latitude spread and by selecting latitude regions such that the satellite re-entry takes place with minimum risk to cause damage to populated areas of the Earth. Furthermore, fuel efficiency studies conducted by the INTEGRAL team in 2015 and 2016 have proven that consumption can be reduced by possibly two thirds if two measures are employed: reaction wheel bias with zero tranquilisation (already implemented in early 2016, current fuel saving of ~60%) and running the reaction wheels through zero speed also in stable pointing (to be implemented by June 2016, additional theoretical propellant saving of up to 65%) [22]. Capturing these fuel savings would potentially allow ESA to continue the INTEGRAL scientific operations phase until 2029 and possibly even allow for a controlled re-entry in 2029 with some residual propellant available for fine-tuning of the re-entry longitude, position and profile. IX. Potential extension of INTEGRAL s mission well into the next decade INTEGRAL with its unprecedented robustness of design and components has operated almost perfectly throughout the mission, with no significant unrecoverable failures, and its overall performance is still far above design specifications: all prime units are still in use maintaining full redundancy, no major failures have occurred and the degradation of spacecraft components is minimal [8]. Currently, the only life limiting elements for INTEGRAL are fuel and power. Due to its very good margin on consumables and all limited-lifetime components, there are no open issues for the continued successful operation of INTEGRAL. Until today, there is still significant interest in INTEGRAL s science data with a healthy over-subscription at each announcement of opportunity and an undiminished rate of target of opportunity requests. Given the continued interest by the scientific community, INTEGRAL has the potential to provide excellent scientific data well into the next decade [23]. X. Conclusion The nominal mission lifetime of 2.5 years and the extended lifetime of another 2.5 years have already been far exceeded for INTEGRAL. However, even now, more than 13 years after launch, INTEGRAL is still the most sensitive and accurate soft gamma ray observatory in space and the scientific return of INTEGRAL is excellent [6,7]. With the performed disposal manoeuvre campaign in early 2015 leading to a safe re-entry of the spacecraft in the Earth s atmosphere at the end of 2029, INTEGRAL is going to comply with ESA s space debris mitigation regulations, even though it was not designed for re-entry. This prevents that ESA s heaviest science mission ever flown eventually becomes a major debris contributor in space in case of a collision. Approximately half of the propellant available at that time was used for this manoeuvre with the residual propellant being sufficient to operate the satellite for at least another five and possibly even eight or more years. With further low cost fuel saving options implemented, it may even allow for a controlled re-entry in This would be the first controlled re-entry ever performed by a mission in such a highly elliptical orbit as well as by an astronomy satellite. 11

12 Acknowledgments The authors would like to thank the combined INTEGRAL/XMM Flight Control Team, Flight Dynamics, Space Debris Office, ISOC and ISDC as well as the instrument PIs for their dedicated work and invaluable effort throughout the mission. INTEGRAL is an ESA Science Mission in cooperation with Russia and the United States with instruments and contributions directly funded by ESA Member States. References [1] Ubertini, P., Lebrun, F., et al., IBIS: The Imager on-board INTEGRAL, 2003, A&A 411, L131-L139 [2] Roques, J. P., et al., SPI/INTEGRAL in-flight performance, 2003, A&A 411, L91-L100 [3] Vedrenne, G., Roques, J. P., et al., SPI: The spectrometer aboard INTEGRAL, 2003, A&A 411, L63- L70 [4] Lund, N., et al., JEM-X: The X-ray monitor aboard INTEGRAL, 2003, A&A 411, L231-L238 [5] Mas-Hesse, J. M., et al., OMC: An Optical Monitoring Camera for INTEGRAL. Instrument description and performance, 2003, A&A 411, L261-L268 [6] Winkler, C., et al., The INTEGRAL mission, 2003, A&A 411, L1-L6 [7] Winkler, C., et al., Space Science Reviews, Volume 161, Issue 1-4, pp [8] Huebner, J.M., et al., INTEGRAL Revisits Earth - Low Perigee Effects on Spacecraft Components, AIAA SpaceOps2012 Conference, Proceedings, Stockholm, Sweden [9] Anselmo, et al., European Code of Conduct for Space Debris Mitigation, Issue 1.0, June 28, 2004 [10] ESA, Requirements on Space Debris Mitigation for ESA Projects, ESA/ADMIN/IPOL(2008)2, Annexes 1, Paris, 1 April 2008 [11] TAS, Residual Propellant Analysis, Technical Note (INT-TN-AI-0296), 2014 [12] Merz, K., Disposal orbit options for Integral, Technical Note (INT-REN-TN HSO-GR) [13] Armellin, R., et al., End-of-life disposal of high elliptical orbit missions: The case of INTEGRAL, Advances in Space Research 56 (2015) [14] Merz, K., et al., Re-entry options for INTEGRAL - synopsis of SCARAB Analyses, Technical Note (INT-REN-TN HSO-GR), 2014 [15] Salt, D., Southworth, R., INTEGRAL Disposal Strategy & Scope for Future Science Operations, Technical Note (INT-OPS-TN-1012-HSO-OS), 2015 [16] Dietze, C. et al., INTEGRAL End-Of-Life Disposal Manoeuvre Campaign, Proceedings of 25th International Symposium on Space Flight Dynamics ISSFD, Munich, 2015 [17] Smargiassi, M., et al., INTEGRAL User s Manual, INT-MA-AI-0001 (section ), 2002 [18] Merz, K., et al., Integral Disposal - Post-Manoeuvre Assessment, Technical Note (INT-REN-MEM HSO-GR) [19] Southworth, R., Integral Mission Lifetime Scope, Technical Note (INT-MIS-TN-1005-HSO-OAI), 2014 [20] Merz, K., et al., Re-entry option for INTEGRAL - collision risk, Technical Note (INT-REN-MEM HSO-GR), 2014 [21] Merz, K., et al., Re-entry options for INTEGRAL - summary of SCARAB run by SDO, Technical Note (INT-REN-MEM HSO-GR), 2014 [22] Huebner, J. M., et al., Run, INTEGRAL, run! Low wheel speed operations for fuel savings, AIAA SpaceOps2016 Conference, Proceedings, Daejeon, Korea [23] Huebner, J. M., et al, INTEGRAL operations beyond the design lifetime - Challenges of running an 11 years old mission, AIAA SpaceOps2014 Conference, Proceedings, Pasadena, California 12