DESIGN OF MANNED MISSIONS TO NEAR EARTH ASTEROIDS

Transcription

1 DESIGN OF MANNED MISSIONS TO NEAR EARTH ASTEROIDS 4TH INTERNATIONAL CONFERENCE ON ASTRODYNAMICS TOOLS AND TECHNIQUES (ICATT) 2010 EUROPEAN SPACE ASTRONOMY CENTRE (ESA/ESAC), MADRID, SPAIN 3 6 MAY 2010 Jesus Gil-Fernandez (1), Raul Cadenas-Gorgojo (1), Diego Escorial-Olmos (1) (1) GMV, Tres Cantos, 28760, Madrid (Spain), jesusgil@gmv.es ABSTRACT Manned missions to asteroids are gaining momentum due to scientific, exploration and outreach reasons. Human-related constraints pose stringent requirements on the mission design. A feasibility analysis is presented for missions launching between 2020 and The system constraints are derived from on-going ESA and NASA developments, leading to two different architectures. The man-rated constraints, including precursor robotic missions, are applied in an exhaustive trajectory search over the entire launch window for each mission architecture. The candidate missions are further analyzed considering (1) the terminal relative navigation, mainly target detectability, and (2) safety, regarding fast Earth-return emergency trajectories and backup optical navigation system performances. The results show how the candidate mission set is reduced until the best asteroid is finally selected, providing a large flexibility for the mission design. 1. INTRODUCTION Manned missions to Near Earth Asteroids (NEAs) are becoming popular in the space community. The risk that NEAs pose to the humankind has motivated numerous studies to design mitigation strategies. Specific missions for characterization shall be launched before a mitigation strategy can be selected. The benefits of crewed missions to NEA include at least science, outreach and exploration technology [1]. Manned missions provide significant scientific benefits compared to robotic missions, in terms of quantity and quality of in situ experiments and acquisition of samples. In addition, such mission will be a great inspiration for the general public since would be a deep space voyage beyond Earth bounds. Finally, a NEA mission is an intermediate step in a manned mission to Mars, where multiple technologies can be tested. The analysis of the manned missions to asteroids must consider not only typical system design criteria such as delta-v budget or payload mass, but the usually more demanding human-related constraints. Among the most important particular design limitation of a human mission to NEO are system constraints, such as total mission duration or the Earth entry velocity, crew safety issues, such as backup navigation system and abort trajectories, and asteroid mission specificities, such as target detection and identification. These aspects have been addressed during a GMV internal study to select the most promising missions in a realistic time frame. GMV has been involved in several studies on missions to NEO and on manned missions to the Moon. In NEO missions, the Guidance, Navigation and Control (GNC) during the terminal phase is one of the enabling technologies. Under different ESA contracts, GMV designed and validated the GNC for the terminal phase of kinematic impact missions [2], for the approach phase of robotic rendezvous missions [3], and for the proximity operations of sample return missions [4]. Regarding manned missions, GMV have analyzed different aspects of human lunar exploration such as crew safety constraints on the ascent trajectory design [5], return trajectories in case of contingency [6], and the backup optical navigation system [7]. The learned lessons and the tools developed during these studies on NEO and manned missions have been properly tailored to the feasibility analysis of manned missions to NEA. 2. TRAJECTORY ANALYSIS The first step is to check the set of asteroids that are reachable considering the system constraints of a human mission. The JPL asteroid database [8] has over asteroids. Manned-system constraints are applied to the trajectories computed in the 30-year launch window. Similar constraints have been applied in some NASA studies [1,9,10,11] but in this study the trajectory search is more systematic and exhaustive using DOIT tool [12], (powered by AstroTool toolbox [13]). The entire JPL asteroid database is processed to get all the trajectories in the launch window from 2020 to The trajectories are defined by the following sequence, (1) departure delta-v, (2) ballistic flight, (3) arrival delta-v to cancel the SC relative velocity with respect to the asteroid, (4) stay time in the asteroid, (5) asteroid departure delta-v, and (6) ballistic flight to arrive to the Earth.

2 2.1. SYSTEM ARCHITECTURE In order to constrain the mission delta-v to reasonable values, and analysis of the proposed transportation systems for human missions has been carried out. This analysis has taken into account the transportation capabilities of the systems proposed by NASA Constellation Program [14], and ESA Architectures for Exploration [15]. From NASA s architecture there have been selected the Crew Exploration Vehicle (Orion), the lunar lander Altair and the launch system AresV and AresI. From the ESA s architecture it has been selected the propulsion stage EDS24. With these elements three different mission architectures have been considered in order to constrain the different delta-vs. - Minimum mission, a train of 2 EDS24 and a CEV assembled in LEO. The two propulsion stages will provide the departure delta-v while the service module of the CEV will be responsible for the remaining manoeuvres; asteroid rendezvous and Earth return. - Intermediate mission, CEV will dock in LEO with an EDS24 launched by AresV. The Earth Departure Stage of AresV will provide the departure delta-v to the EDS24+CEV, which will be responsible for the remaining manoeuvres of the mission. - Full-scale mission, CEV and Altair will be used as in a human mission to the surface of the moon. AresV will provide the Earth departure delta-v while the descent module of Altair, ascent module and CEV service module will be responsible for the remaining manoeuvres of the mission. This full-scale mission could be used to test and validate some of the systems and procedures for a human mission to the Moon surface, but unfortunately the system offers low performance for a NEA mission (as it is sized now for a lunar mission). If an increase in the performances is achieved in the future (increase in the EDS excess velocity capabilities), this architecture could become interesting. Thus, the trajectories to all the asteroids within the prescribed launch dates are filtered considering the following constraints, - Total mission duration lower than 250 days - Hyperbolic excess velocity at Earth arrival lower than 4.5 km/s - Stay time at the asteroid longer than 5 days and shorter than 30 days - Maximum delta-v Configuration Depart. ΔV Post escape ΔV Minimum 3668 m/s 1560 m/s Intermediate 3800 m/s 3729 m/s 2.2. CANDIDATE MISSIONS For the minimum configuration only 4 asteroids were found fulfilling all previous manned system constraints, while for the intermediate configuration number of 26 asteroids are feasible targets. For each asteroid, there are multitude missions in the 30-year launch. Another requisite for an asteroid to be considered as a suitable target for a human visit is to allow for a precursor robotic mission [1,9,10]. This precursor mission will perform a detailed characterization the target asteroid. Given the risks of manned missions beyond Earth orbit, such information is critical even if Earth observation campaigns provide a lot of information about the asteroid. In particular, close observations will prove the existence of safe landing regions on the surface of the asteroid (one of the concerns during the design of ESA Marco Polo [16] sample return mission). Another important utility of the precursor robotic probe is as navigation aid at the arrival of the manned SC. For this application, the SC must be active at the arrival of the manned SC. Thus, the maximum time interval between the arrival of the robotic mission and the arrival of the manned mission is driven by the maximum operational time of the robotic mission around the asteroid. The spacecraft must be still operative at the arrival of the human mission, since it will be used for relative navigation during the approach phase of the manned mission. This time interval has a minimum limit in order to have enough time to characterize the minor body before the arrival of the manned mission. For the computation of the robotic missions, a systematic search on an expanded launch window ([ ]) is performed. Only the outbound trajectory is computed with the following constraints. - Maximum transfer time of 400 days - Launcher escape velocity of 4.5 km/s (taking into account that a delta-v Earth gravity assist transfer can be added) - Maximum delta-v budget for asteroid rendezvous of 1.5 km/s - Time interval between arrival of robotic mission and arrival of manned mission between 180 days and 2000 days. Fig. 1 presents the characteristics of the robotic (red) and manned (blue) trajectories to asteroid 2000 SG344 with the minimum configuration. Fig. 2 and Fig. 3 present the available launch opportunities for the asteroids that have feasible manned and precursor robotic missions with the minimum and intermediate configurations, respectively. The ranges below the horizontal grid line indicate the launch window for the robotic missions, while the upper lines indicate the launch window for the manned missions. Note that the launch windows are quite wide (months). The frequency of opportunities for direct transfer

3 depends on the synodic period of the asteroid and the Earth. Given the tight delta-v and time-of-flight constraints, only asteroids with orbit close to the Earth appear, which offer poor repeatability in Figure 1. Minimum configuration: robotic and manned mission opportunities to asteroid 2000 SG344. Figure 2. Minimum configuration: launch window opportunities for manned and robotic missions. 3. TARGET DETECTION ANALYSIS The approach phase starts when the target NEO is detected and relative navigation can be carried out. In the manned mission, even if the precursor robotic mission will be used as beacon for relative navigation, this phase is critical since the relative velocity is large (few km/s) and the braking shall be precise and take short time. The sooner and further the asteroid can be detected, the better. The only reasonable sensor to provide direct relative measurements of the NEA at long distances (up to 1e6 km) is a camera. The optical navigation system should be designed to be able to accomplish the nominal manned mission with no aid from the precursor robotic mission, as a contingency scenario (the other possibility is to abort the mission and go back to the Earth). The navigation during the approach and proximity operations can be analysed in detail with GMV s CHILON [3] and CHILON-2 [4] tools. The lessons learned during the design of the guidance, navigation and control (GNC) system are applied to the selection of the candidate missions. There are limitations on the maximum distance and the time to go (time to impact or closest distance) at which the asteroid can be detected by the on-board camera. Apart from the camera sensitivity, the trajectory and the asteroid brightness define the further distance for detection. The approach phase angle (Sun-NEO-SC) is a geometric parameter determined by the trajectory and has a significant impact on the detectability of the NEO by the navigation camera. The instrument magnitude can be approximated from the phase angle, the SC-NEO distance, the asteroid heliocentric distance, the asteroid absolute magnitude, and the slope parameter. The maximum distance and time to go for detection of the asteroid is computed assuming a sky magnitude of 22 mag/arcsec2. The results for all the manned missions are presented in Fig. 4 for the minimum configuration and in Fig. 5 for the intermediate configuration. In order to have enough margin to perform the approach safely, a minimum limit of 0.75 day is imposed. Only two asteroids present feasible missions with minimum configuration (2000 SG344 and 2001 FR85). For the intermediate configuration, two more NEO are available (2007 YF and 1999 AO10). Fig. 6 depicts the parallel coordinates plot (PCP) [17] of all the feasible missions to asteroid 2000 SG344. These figures are useful to summarize the thousands of missions for each asteroid, above all to identify the boundaries of the different parameters and the check correlations between variables. Figure 3. Intermediate configuration: launch window opportunities for manned and robotic missions.

4 4. EMERGENCY TRAJECTORY ANALYSIS In case of contingency during the outbound trajectory (from the Earth to the asteroid), the SC shall return to the Earth as soon as possible fulfilling the entry velocity constraint and executing a manoeuvre smaller than the delta-v allocated in the SC. The analysis of these return trajectories is performed on the reference missions. The reference trajectories are selected from the PCP and are presented in Table 1. As illustration example, the trajectory of the manned mission with minimum delta-v to 2000 SG344 is depicted in Fig. 7. The delta-v required to return to the Earth for different times of flight is computed at different points in the outbound trajectory. Figure 4. Minimum configuration: time to impact at asteroid detection. Figure 5. Intermediate configuration: time to impact at asteroid detection. Figure 7. Minimum configuration: minimum delta-v manned mission to 2000 SG344. Figure 6. Intermediate configuration: parallel coordinates plot of manned missions to 2000 SG344. Figure 8. Delta-V for Earth return for different transfer times (2000 SG344).

5 Asteroid Departure DVlaun TOFout DVarr DVdep TOFin Vinf TOF DVsc T2go [MJD2000] [km/s] [days] [km/s] [km/s] [days] [km/s] [days] [km/s] [days] Config 2001 FR & SG & SG YF AO Table 1. Reference manned missions for contingency analysis. The figures of merit of the emergency Earth return trajectories are the required delta-v and the hyperbolic excess velocity for a given time of flight. Thus, for each point in the nominal trajectory, the Earth return trajectories for different times of flight are computed. The results for the mission to 2000 SG344 are presented in Fig. 8 and Fig. 9. From these maps, the minimum transfer time for Earth return fulfilling the constraints on maximum delta-v and hyperbolic excess velocity at Earth can be computed. The results for the reference missions of Table 1 are depicted in Fig. 10. Should any limit on the emergency return time is be imposed, then some missions might be removed, or alternatively allow a higher SC delta-v budget or reentry velocity. From the missions presented in Table 1 and considering the results of Fig. 10, we select the asteroid 2000 SG344. The main reasons are that it allows missions with minimum and intermediate configuration and provides high flexibility to the mission design for intermediate configuration. For instance, permits the shortest mission, with high delta-v budget, or long missions with low delta-v. In addition, the launcher delta-v is the lowest. Figure 9. Hyperbolic excess velocity for Earth return for different transfer times (2000 SG344). Figure 10. Minimum Earth-return time for all the reference manned missions. 5. OPTICAL NAVIGATION ANALYSIS Manned missions are required to have a backup navigation system in case of problems in the communication system or in the ground segment [7]. Hence, the backup on-board navigation system must be autonomous. The most reliable and realistic solution is an optical navigation system, which in case of an asteroid mission is also needed for the approach phase. An analysis of the optical navigation performances assessment for the candidate trajectories is carried out with MATON (Mission Analysis Tool for Optical Navigation), that has been used for other studies [17]. The autonomous navigation system when far from the Earth can use distant celestial objects as beacons for triangulation. The celestial objects available as navigation beacons are computed considering the visibility by the camera including operational constraints as listed below. - Apparent magnitude of the beacon lower than the camera maximum magnitude (12 mag). - Angular separation with respect to the Earth (beacon-sc-earth angle) above a minimum to avoid dazzling (the minimum is the maximum of the Earth angular size to avoid occultation and a constant exclusion angle).

6 - Sun-SC-beacon angle above 40º to avoid Sun dazzling. - Sun-SC-beacon angle below 140º for thermal reasons. Fig. 11 presents in polar coordinates the available asteroids and Fig. 12 presents the available planets (the origin is the Ecliptic North Pole and the 180º circle is the Ecliptic South Pole). Fig. 13 present the total number of beacons during the trajectory, near the end of the trajectory there are few available beacons (only Earth, Moon, Mars and Hygiea). The navigation performances are assessed through a covariance analysis. Fig. 13 presents the worst case navigation performances considering the navigation point solution. No a priori in-formation or dynamics model are included in the covariance analysis. The main sources of navigation error are the asteroids ephemerides uncertainties and the camera (and image processing) accuracy. The asteroid ephemeris error is assumed 500 km (1-sigma). This value is selected because we consider only the brightest asteroids, which are the best known and observation campaigns can easily improve the knowledge. For the planets the assumed uncertainty is 25 km (1-sigma). The error in the line-of-sight (LOS) measurement depends on the camera accuracy and the image processing performances. The assumed total LOS error is 1 pixel (1-sigma). The optical camera during the emergency phase is a star tracker with a relatively large FOV of 10º. Thus, in this conservative approach we have the disadvantages of a large FOV (large pixel size) and we are not taking advantage of the sub-pixel accuracy reachable with defocused optics and oversampling. The results show that even in the period with only 4 beacons, the reference mission is flyable autonomously with the backup optical navigation system. It is worth mentioning that the navigation accuracy obtained with a real navigation filter including the dynamics (either sequential or batch) will be few times better. Figure 12. Visibility of asteroids throughout the reference manned mission to 2000 SG344. Figure 13. Number of visible objects throughout the reference manned mission to 2000 SG344. Figure 11. Visibility of asteroids throughout the reference manned mission to 2000 SG344. Figure 14. Worst performances of optical navigation along the reference manned mission to 2000 SG344.

7 6. CONCLUSIONS Manned missions to NEA have great interest for science and exploration because of the flexibility of the crew to perform complex tasks and to adapt to off-nominal situations or uncertainties in real time. The mission analysis of such manned missions cannot be based on conventional optimization of cost or delta-v budget. The feasibility analysis of candidate missions in a large time window (2020 to 2050) has been carried out considering human-related constraints. A preliminary mission architecture analysis leads to two different configurations, each with its own system constraints. A systematic search of missions to NEA considering the system constrains of the two architectures has been presented. One of the constraints is to have a precursor robotic mission before the arrival of the manned mission for characterization and for relative navigation. From these initial constraints a set of 22 asteroids was available for the intermediate configuration and only 4 for the minimum configuration. Then the detectability of the asteroid by the on-board camera was analyzed and only 4 asteroids present missions with at least 20 hours of visibility prior to the encounter. From the different options, 5 reference missions to the four NEA were selected. The emergency trajectories were also analysed. The asteroid 2000 SG344 was selected for its flexibility in the mission design. The preliminary optical navigation analysis shows that enough beacons are available for autonomous navigation in case of emergency. 7. REFERENCES 1. Abell, P.A. and the NEO Mission Study Team (2009). Exploration of Near-Earth Objects via the Orion Crew Exploration Vehicle: A Planetary Defence Rationale, 1st IAA Planetary Defence Conference, April 28, 2009, Granada, Spain. 2. Gil-Fernandez, J. et al. Impacting Small Near Earth Objects (NEOs) (2008). Advances in Space Research, Vol. 42, pp , doi: /j.asr Gil-Fernandez, J., Cadenas, R. et al. (2008). Autonomous GNC Algorithms for Rendezvous Missions to Near-Earth-Objects, AIAA/AAS Astrodynamics Specialist Conference and Exhibit, August, Honolulu, Hawaii, AIAA Gil-Fernandez, J. et al. (2010). Autonomous GNC for Descent and Landing on Small, Irregular Bodies, AAS , 20th AAS/AAIA Space Flight Mechanics Meeting, February 14 17, San Diego, USA. 5. Gil-Fernandez, J., Graziano, M. et al. (2009). Emergency Trajectories for the Crew Transfer Vehicle. Acta Astronautica, Vol. 65, 2009, pp , doi: /j.actaastro Corral, C. Cadenas-Gorgojo, R. Gil-Fernández, J., Graziano, M. (2010). On the Use of the Earth- Moon Lagrangian Point L1 for Supporting the Manned Lunar Exploration, Space Manifold Dynamics (Novel Spaceways for Science and Exploration), 1st edition, Springer, pp Corral, C., Prieto-Llanos, T., Gil-Fernandez, J., (2008). Optical Navigation for Lunar Transportation Systems Contingency Scenarios, AIAA/AAS Astrodynamics Specialist Conference and Exhibit, August, Honolulu, Hawaii, AIAA JPL s Near Earth Object Program [online database], URL: 9. Korsmeyer, D. J., Landis, R. R., and Abel, P. A. (2008). Into the beyond: A crewed mission to a near-earth object, Acta Astronautica, Volume 63, Issues 1-4, July-August 2008, Pages Landis, R. R., Korsmeyer, D. J., Abell, P. A., Adamo, D. (2008). A Piloted Orion Flight to a Near- Earth Object: A Feasibil-ity Study, AIAA , SpaceOps 2008 Conference, May Wagner, S., Wie, B., (2009) A Crewed 180-Day Mission to Asteroid Apophis in , IAC- 09.D Cadenas, R. et al, (2010). DOIT: Design of Interplanetary Trajectories, 4th International Conference on Astrodynamics Tools and Techniques, May Cadenas, R. et al, (2007) AstroTool Software User Manual, Aug Chris Culbert (2008), Constellation Programme Overview, Joint Annual Meeting for LEAG- ICEUM-SRR, Cape Canaveral, Florida, October 28-31, Architectures for Exploration, Final Review presentations, (2009) ESTEC, The Netherlands, January 19, Agnolon, D., Marco Polo Mission Requirements Document, Issue 4, SCI-PA/ /Marco- Polo, Arney, D., and Wilhite, A., Visualization of the Multidimensional Human Interplanetary Mission Design Space, Journal of Spacecraft and Rockets, Vol. 46, No. 6, November December 2009, doi: /