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GNC technologies for NEO missions
GNC technologies for NEO missions C. Cavel*, G. Jonniaux*, G. Flandin*, M-C. Perkinson** *EADS ASTRIUM SAS, 31 rue des Cosmonautes, 31402 Toulouse Cedex 4, France (Tel: +33 (0)5 62 19 98 10; e-mail: [email protected]). ** EADS ASTRIUM Ltd, Gunnels Wood Road, Stevenage, UK (Tel: +44 14 3877 8935; e-mail: [email protected]).
Abstract: Compared to planetary landing missions that have been already widely studied in the past years, the design of missions to small bodies is still an open field for innovation. This is especially true for the Guidance, Navigation and Control system (GNC) for which the particular environment of Near Earth Objects (NEO) both constrains the requirements and allows some flexibility in the design. Focusing on the Marco Polo mission, the autonomy and GNC requirements are derived from mission needs. A descent scenario is proposed and the guidance strategy is described. Candidate navigation solutions (for both relative and absolute navigation) are introduced, coming from past and on-going technology development projects, and a possible baseline is highlighted.
Keywords: Near Earth Objects, Guidance, Navigation & Control, autonomy, absolute navigation, visionbased navigation. 1. INTRODUCTION 1.1 The Marco Polo mission Near Earth Objects – asteroids and comets – are primitive fragments of rock, metal and ice remaining from the formation of our solar system. Their composition, structure and surface properties are a record of the history of the solar system, and much can be learnt from them. This is why the European Space Agency (ESA) is currently considering a sample return mission to an asteroid, the so-called Marco Polo mission, in the frame of the Cosmic Vision Programme. Astrium Satellites led during the past months one of the parallel contract Marco Polo Assessment Studies for ESA. During this 12 month study the entire mission design has been analysed in detail to determine an elegant baseline mission solution which fulfils the asteroid surface sample return objective. The Marco Polo mission comprises a single spacecraft design, using chemical propulsion for all transfers and proximity operations, launching in 2018 on a Soyuz Fregat 2-1B. It will reach the asteroid 1999 JU3 in 2021/2022, and after a mapping phase of some months, the spacecraft will descend to the surface to land for 10 minutes, up to three times. Sampling will take place on landing, using a fast corer sampling mechanism on a robotic arm to obtain a sample of around 30 grams, while the thrusters fire to hold the spacecraft on the surface. The spacecraft will then return to orbit, verify the acquisition of a meaningful sample, and perform the return transfer to Earth. The Earth Re-entry Capsule (ERC) would impact the ground in December 2024.
Fig. 1. Marco Polo mission outline 1.2 Specificities of GNC technologies for NEO missions The major difference between planetary landing and small body landing is the intensity and the regularity of the gravity field. For missions to Moon and Mars the control and propulsion system is usually designed to ensure a thrust to weight ratio sufficient to break the high entry velocity. The situation is opposite when landing on a small body: the gravity field is very low and the sphere of influence of the body is only a few kilometres wide. That means that the relative velocities are very weak and that the design of the descent trajectory is not driven by consumption issues: a large variety of descent scenarios can be imagined, from a
classical vertical descent to a horizontal flyby of the landing area or even hovering phase above the ground. On the other hand, the low gravity environment leads to some constraints on the GNC system. The classical use of high acceleration to get the 3D observability in pure vision-based navigation is no longer possible: a new solution must be imagined. Another particularity of missions to NEO is that the knowledge of the target is relatively poor. Whereas landing sites can be easily defined in advance for missions to Mars or the Moon, the shape and surface relief of the asteroid is not known, thus requiring a dedicated characterization campaign once in orbit. The GNC system shall therefore support the task of finding a safe landing site and ensure a very accurate landing (position accuracy requirement is a few meters which is far below what has been achieved in past missions). The mass, rotational state and gravity field are known to a certain level of accuracy, depending on the performances of the characterization phase. The GNC design shall therefore be robust to environment uncertainties greater than for planetary landing. Thanks to the low gravity environment a multiple landing attempts strategy is foreseen. Contrarily to a planetary landing (which is a “one shot” mission) a specific Failure Detection Isolation and Recovery (FDIR) strategy (including collision avoidance algorithms) shall be designed for proximity operations. The results of the Hayabusa mission (failure of the first landing attempt) show that the robustness of the GNC design is of prime importance for such missions.
The science behind the Marco Polo mission demands a precision landing on the surface of an asteroid which can only be safely achieved after global and local characterization of the object and potential landing sites. The GNC system for proximity operations must meet the requirements for all proximity phases. 2.2 Mission requirements The driving mission requirement for the descent and landing is that the spacecraft should be able to land at any location on the illuminated side of the asteroid. The descent strategy must therefore be designed for compatibility with landing at any latitude within half the rotation period of the asteroid (~3.7 hours). The GNC system and descent trajectory must be designed to provide a soft landing with an impact velocity below 30cm/s vertical and 5cm/s lateral, the relative angle between the spacecraft attitude and the local slope must also be less than 10 degrees. These requirements must be met to ensure the current simple landing gear design and hold down strategy can be used. The accuracy of the landing shall be better than 3.5m which has never been achieved in past missions (Hayabusa achieved about 30m). Such accuracy can only be reached by high-performance image based navigation.
2. AUTONOMY AND GNC REQUIREMENTS 2.1 Science requirements The Marco Polo mission has essentially only one key objective (ref. 1): to bring back to Earth a good, pristine asteroid sample from a primitive asteroid together with the necessary context about the asteroid and landing site from where the sample was collected. To ensure access to the most scientifically valuable areas on the asteroid the GNC system must be able to target a landing at any latitude, the only constraint placed on landing site location is that it should be on the illuminated side of the asteroid enabling the use of visual navigation. Because of the diversity of asteroid environments it is also necessary to collect the contextual information about the asteroid where the sample was collected. This information includes mapping of the mineralogy, morphology and environmental conditions. This information will be collected from a number of closed orbits and hovering positions within the sphere of influence. Measurements are needed at different altitudes to provide the required resolution for the global mapping and the detailed characterization of the potential landing sites.
Fig. 2. Stability geometry of lander showing critical design parameters
Table 1. GNC requirements at landing
The descent phase design also needs to take account of the contamination requirements. Contaminants or undesired
particles coming out of the thrusters must have a density lower than 1014 (1013 as a goal) molecules/cm2 on the NEO surface. 2.3 Autonomy requirements The low gravity environment around NEO modifies the classical task repartition between the ground and the board. The guidance task (designation of the landing site and generation of reference trajectories) can be performed by the ground (as done successfully for Hayabusa) using the results of the characterization phase and the landing rehearsals. The landing rehearsals can be considered as part of the GNC system as they provide valuable information for the guidance task (generation of reference trajectories). During the orbital phase (radio science and global characterization) the spacecraft can be operated from the ground using ground tracking (and possibly visual measurements using a posteriori data fusion) and uploading a pre-determined manoeuvre plan, as the orbit is relatively stable and safe. On the contrary an autonomous navigation & control function is required for the descent phase (including local characterization, landing rehearsals, landing and re-ascent) for which the communication time with ground is similar to the duration of the phase. This is also required to meet the accuracy specifications at landing.
variety of absolute navigation function (the pointing shall be close to the one used during characterization phases). Moreover only a partial knowledge of the asteroid is required for the absolute navigation (whereas a Hayabusa-like scenario relies on the full asteroid model to navigate). - Safety of the scenario and contamination minimization: any colliding trajectory shall be avoided when the relative velocity is high. Vertical control shall be inhibited in the last tens of meters to avoid landing site contamination.
The deltaV allocation for the descent sequence is weak and can be seen as a degree of freedom (a limit of 3m/s is proposed). This is an important tuning parameter to reduce the effect of perturbations and shorten the sequence. A preliminary baseline has been proposed to comply with these constraints. It is based on a transfer from a home position (which is the global characterization orbit at 2km altitude) to a gate position (500m above the selected landing site) followed by a controlled vertical descent. It is a good compromise between system requirements (among which the battery sizing is a critical issue) and GNC needs (this trajectory ensures optimal performances of the navigation function) while ensuring the safety of the spacecraft by appropriate tuning of deltaV. Simulation results on a typical scenario show a total deltaV of 1.85 m/s for the complete sequence. A complete trade-off with a Hayabusa-like scenario is still to be performed to refine the solution.
3. DESCENT STRATEGY & GUIDANCE 3.1 Descent strategy The design of the descent phase is subject to a lot of different constraints: - Illumination of the landing site at touchdown: the use of vision-based navigation requires some contrast in the scene, but in the meantime the shadows shall not be too large. A local solar time around 10.00 am at landing seems a good compromise. This constrains the duration of the landing sequence. - Power: the battery-powered phase shall not be too long to avoid over sizing of the battery and penalizing the mass budget. - Communication: images used for navigation shall be relayed to Earth if possible. - Robustness to the perturbations: a short scenario minimizes the effect of perturbations. This is also linked to the selection of the guidance law. - Performances at touchdown: vertical descent is required at some stage to cancel horizontal velocity induced by the rotation of the asteroid. - Compatibility with the navigation function: a vertical descent improves the observability of lateral velocity (which is the key parameter to be controlled) and allows a wide
Fig. 3. A candidate solution for the descent sequence 3.2 Guidance law Thanks to local characterization and landing rehearsals, a very accurate hazard mapping can be done in the landing site area. Using this information the ground is able to designate a safe landing site convenient for sampling operations. Since a very accurate landing is specified, no autonomous hazard avoidance function is required. In this configuration the complexity of reaching a safe landing site is transferred to the
ground (for landing site designation) and to the navigation function (for reaching this landing site). This simplifies a lot the GNC architecture and avoids unexpected failure of the HDA (Hazard Detection and Avoidance) function at proximity of the surface, as it was the case for the Hayabusa mission (the hazard detection sensor was deactivated for the second landing attempt).
Two different methods are available to build a reference trajectory in position and velocity:
4. NAVIGATION SOLUTIONS Two different kinds of navigation function are required to ensure an accurate landing: - An absolute navigation function, in charge of recognizing the landing site that has been pre-designated by the ground. This is the key element of the GNC as the relative navigation function will not be able to compensate for an initial error of absolute position. It is supposed to be used at some specific times and not periodically. This function is the critical one for achievement of the 3.5m position accuracy at landing.
- Define a thrust profile: in this case the guidance law is a succession of deltaV to be commanded along the descent. The command orders are directly sent to the actuators in a feed-forward scheme, and the control is only in charge to compensate for the small discrepancies during the descent. This technique is to be used when the environment (and especially the external forces) is well known as any uncertainty on the external forces will modify the trajectory.
- A relative navigation function, to control periodically the last part of the descent with respect to the landing site (once identified), and to control the spacecraft in hovering phases.
- Define a position / velocity profile: In this case the key element of the system is the control. The control system shall compute in closed-loop the thrusts that are required to follow the P/V profile. The P/V profile is only a geometrical profile to reach a point P2 from a point P1 in a given frame, without any consideration of external forces. Therefore this kind of guidance law is not sensitive to environment uncertainties, because the spacecraft observes in closed-loop its position and velocity and computes the required thrust to stick to the reference trajectory. This kind of guidance can lead to a wide range of deltaVs depending on the environment uncertainties (thus the thrusters configuration shall allow some flexibility) but has the great advantage to be robust to these uncertainties: the difference between the effective trajectory and the reference trajectory is only linked to the capacity of the spacecraft to observe and control its position / velocity at the required frequency.
3D models matching method (3D / 3D)
The first solution is well adapted to the first phase of the descent sequence (transfer from home to gate position) as the altitude is sufficient to consider that environment uncertainties are weak. Moreover the altitude of the gate position is high enough to be robust to some guidance errors, as the only requirement at gate position is to have the selected landing site in visibility to trigger the second phase, and this can be ensured by proper sizing of the field of view (FOV) of the navigation camera. For the vertical descent from gate position to landing the second solution shall be selected. Indeed in this case the environment uncertainties become more and more significant (this is especially true for the gravity field of the asteroid). Moreover this critical phase shall ensure the accuracy of the landing: therefore the spacecraft shall stick accurately to the vertical descent to minimize transverse velocities at landing and position errors.
Different vision-based solutions have been analysed for both absolute and relative navigation functions. 4.1 Absolute navigation for pinpoint landing
The first solution is an innovative technique, based on the matching of 3D shapes. Considering having the 3D model of the asteroid on-board, it is possible to use it for comparison: the asteroid portion visible along the final descent (or earlier) can be reconstructed on-board with a 3D dense reconstruction algorithm, using the motion of the camera to have stereo-like algorithm. Once the 3D portion is recovered, a 3D matching with the on-board 3D model allows estimating the pose of the lander.
Fig. 4. 3D matching principle: recognition of a DEM portion, from an elevation map estimated onboard during the descent
Model-Image matching method (3D / 2D) The 3D model can also be compared with the 2D descent images directly. The comparison can then be done by comparing the observed contour of the asteroid in the image, and the contour as simulated on-board with the 3D model of the asteroid and a 3D graphic engine. A more complex solution is to have a full 3D model of the asteroid, including
texture, and then compare an on-board simulated image and the actual images.
- Navigation filter covariance for {{position}{attitude}} = {{2500m , 200m , 1000m}{0.1°}}, 1 sigma - Actual error = {{1000m, 200m, 500m}{0.03°}} - Landmark 3D relative position error: 2m. The parameters that were changed to measure their impacts are: - The sun elevation and azimuth, between the orbital image and the in-flight one.
Fig. 5. Image-3D matching principle: recognition of the orientation of the asteroid from a descent image (left) and the on-board DEM (right).
- The pitch of the camera (while the altitude remains constant). - The number of feature points.
Landmarks extraction and matching method (2D / 2D, MAGELLAN) Another approach is to use database image patches of some landmarks, and to locate them on-board in descent images, to finally recover the pose of the camera. Astrium Satellites has already developed such algorithms in the MAGELLAN R&D study (MAtching of Ground ELements and Landmarks for Approach Navigation), which was designed for lunar environment and mission. Although it is not directly applicable on asteroids, this method is a good baseline for a new method dedicated to small bodies. The MAGELLAN algorithms are based on a succession of correlation methods: first, a rough correlation is applied on the images within a search box, whose size depend on the pose error at the output of the navigation filter. Secondly, a sub-pixel fine estimation is applied to precisely recover the Line of Sight (LOS) of the landmarks. Those steps use the 3D model of the asteroid to reshape the appearance of the images, in order to match actual and database point of views. Finally, after a rejection of badly matched landmarks, MAGELLAN can either deliver the LOS of the landmarks, or directly estimate the relative position and attitude of the camera from the 3D known position of the landmarks. Both of these measurements can be used in a Kalman-like filter to improve the estimation of the spacecraft’s pose, obtained from other sensors. Simulations have been performed on the lunar case to validate the algorithms and assess their performances. The image generation tool PANGU (Planet and Asteroid Natural Scene Generation Utility from University of Dundee) has been used to generate lunar images. The lunar scenario used for the robustness and performances tests is: - Altitude = 21 km (coast phase) - Camera pitch: from -90° (nadir) to -40° - Sun position in orbital image: Elevation and Azimuth = 30° - Sun position in the in-flight image: Elevation from 2° to 90°, Azimuth = 30°
Fig. 6. MAGELLAN simulation on the Moon: real position of the landmarks (left), a-priori position of the landmarks, from the navigation parameters (middle), and estimated position of the features, after matching (right). The results are fairly good for the targeted application (see Table 2). The pose estimation methods (DLT and Homography characterization) are only based on a single image, hence providing the performances at worst. Note that there is no error on the landmarks 3D positions, which can be considered as a bias on the final results. Position error
Attitude error
Average LOS error on the feature points
• Illumination conditions : similar • Pitch = -85 (quasi-nadir)
20 m
0.03
0.037
• Sun azimuth difference = 40 • Sun elevation difference = 40 • Pitch = -85
72 m
0.2
0.040
• Illumination conditions : similar • Pitch = -50
232 m
0.5
0.067
Table 2. Performances of absolute position and attitude estimation for a lunar scenario (coast phase) Good robustness with respect to Sun azimuth and elevation has been observed. Robustness to pitch is relatively important too. The simulations results indicate a quasiconstant performance while the LOS is not further than ~30° from nadir, and then a rapid degradation. Regarding the performance values themselves in the ideal cases (no important pitch, no important sun position difference between database and in-flight images), the average LOS precision is better than 1 pixel, so it proves that the sub-pixelic matching has sense. The rejection is an important step, as about 15% of the features are badly matched for each image.
Finally, the position restitution error remains below 100m (or 0.5% of the altitude) 3 sigma in nadir conditions, i.e. better than the initial position error. Further works should be done to study the improvements of introducing the features LOS restitution in a Kalman-like filter. A ratio of at least 2 can be expected. The ratio of 0.5% of the altitude may not be constant along the flight, so that more realistic test cases are needed to confirm this ratio. 4.2 Relative navigation for safe and accurate landing Astrium Satellites has developed for many years the NPAL system (Navigation for Planetary Approach and Landing, ref. 2), based on unknown features extraction and tracking in order to recover at high frequency the 3D relative state during the descent. This technique is a good candidate to provide the relative navigation function for Marco Polo, provided that a few adaptations to the asteroid case are brought to the system. In classical planetary landing the V/d ambiguity (zoom effect) that appears in the system is solved using acceleration sensing. This is not possible for Marco Polo since the gravity is very weak: another solution has to be found to solve this ambiguity and increase observability and robustness of the system. Different methods have been identified: - NPAL + hovering stations: The principle is to use a hovering station to calibrate the velocity of the spacecraft (the velocity is null when all features are motionless). - NPAL + altimeter: Here the idea is to use a direct range measurement to solve the V/d ambiguity. This is a good option with respect to health monitoring and FDIR as the altitude of the spacecraft is directly measured. - NPAL + target markers (TM): The idea is to release a few (2 or 3) target markers on the ground during the landing rehearsal, that can be easily detected by a flash camera. Then a picture of the landing area is processed by ground operators in order to determine the exact location of the TM. These TM are then used during the descent as known landmarks in order to solve the V/d ambiguity, as done for the Hayabusa mission. Robustness and FDIR strategy shall be considered as key elements to trade these different solutions and select a baseline. Thus the use of an altimeter appears an interesting option. The NPAL system has been adapted to the Marco Polo mission for a preliminary assessment of performances, on a 500 m vertical descent at 20 cm/s. No additional measurement has been hybridized with NPAL. Due to weak perturbations a pulse counting method coupled with a dynamics model has been used instead of an accelerometer. In this test case the NPAL system tracks 20 points at 0.1Hz, with a typical white noise standard deviation of 0.2 pixels. Initial estimation errors are set to 1m in position and 1 cm/s in velocity (all axes).
Fig. 7. Performances of NPAL without hybridization on a 500 m vertical descent at 20 cm/s
Table 3. Performances of NPAL without hybridization on a 500 m vertical descent at 20 cm/s The results show that it is possible to design a high accuracy relative navigation function on the base of NPAL hybridised with a relevant additional measurement. In particular the estimation of cross-track position and velocity is very good (no drift with respect to initial error is observed), which is the most critical point during the final vertical descent to ensure the landing accuracy. The degradation observed on the alongtrack axis (vertical axis) is due to the classical V/d ambiguity and is easily corrected by proper hybridization of NPAL with an additional measurement. 5. CONCLUSION The GNC system appears the most critical one for the challenging Marco Polo mission. A preliminary baseline has been defined on the base of past and on-going technology R&D activities led by Astrium Satellites, using advanced vision-based navigation techniques for both absolute and relative navigation. The detailed design of GNC algorithms and associated FDIR strategy is about to be started in the coming weeks. A complete validation in a real time environment and under representative dynamics will be performed to demonstrate performances and maturity of the GNC system. REFERENCES [1] ESA, “Marco Polo Mission Requirements Document”, issue 4 revision 1, May 2009. [2] Bodineau G., Boléat C., Flandin G., Frapard B., Polle B., Mancuso S. “Vision navigation for European landers and the NPAL project”. 17th IFAC Symposium on Automatic Control in Aerospace, June 2007. [3] J. Kawaguchi, S. Aida, H. Morita, “Hayabusa Detailed Guidance and Navigation Operations during Descents and Touchdowns”, AIAA 2006-6536.
Abstract: Compared to planetary landing missions that have been already widely studied in the past years, the design of missions to small bodies is still an open field for innovation. This is especially true for the Guidance, Navigation and Control system (GNC) for which the particular environment of Near Earth Objects (NEO) both constrains the requirements and allows some flexibility in the design. Focusing on the Marco Polo mission, the autonomy and GNC requirements are derived from mission needs. A descent scenario is proposed and the guidance strategy is described. Candidate navigation solutions (for both relative and absolute navigation) are introduced, coming from past and on-going technology development projects, and a possible baseline is highlighted.
Keywords: Near Earth Objects, Guidance, Navigation & Control, autonomy, absolute navigation, visionbased navigation. 1. INTRODUCTION 1.1 The Marco Polo mission Near Earth Objects – asteroids and comets – are primitive fragments of rock, metal and ice remaining from the formation of our solar system. Their composition, structure and surface properties are a record of the history of the solar system, and much can be learnt from them. This is why the European Space Agency (ESA) is currently considering a sample return mission to an asteroid, the so-called Marco Polo mission, in the frame of the Cosmic Vision Programme. Astrium Satellites led during the past months one of the parallel contract Marco Polo Assessment Studies for ESA. During this 12 month study the entire mission design has been analysed in detail to determine an elegant baseline mission solution which fulfils the asteroid surface sample return objective. The Marco Polo mission comprises a single spacecraft design, using chemical propulsion for all transfers and proximity operations, launching in 2018 on a Soyuz Fregat 2-1B. It will reach the asteroid 1999 JU3 in 2021/2022, and after a mapping phase of some months, the spacecraft will descend to the surface to land for 10 minutes, up to three times. Sampling will take place on landing, using a fast corer sampling mechanism on a robotic arm to obtain a sample of around 30 grams, while the thrusters fire to hold the spacecraft on the surface. The spacecraft will then return to orbit, verify the acquisition of a meaningful sample, and perform the return transfer to Earth. The Earth Re-entry Capsule (ERC) would impact the ground in December 2024.
Fig. 1. Marco Polo mission outline 1.2 Specificities of GNC technologies for NEO missions The major difference between planetary landing and small body landing is the intensity and the regularity of the gravity field. For missions to Moon and Mars the control and propulsion system is usually designed to ensure a thrust to weight ratio sufficient to break the high entry velocity. The situation is opposite when landing on a small body: the gravity field is very low and the sphere of influence of the body is only a few kilometres wide. That means that the relative velocities are very weak and that the design of the descent trajectory is not driven by consumption issues: a large variety of descent scenarios can be imagined, from a
classical vertical descent to a horizontal flyby of the landing area or even hovering phase above the ground. On the other hand, the low gravity environment leads to some constraints on the GNC system. The classical use of high acceleration to get the 3D observability in pure vision-based navigation is no longer possible: a new solution must be imagined. Another particularity of missions to NEO is that the knowledge of the target is relatively poor. Whereas landing sites can be easily defined in advance for missions to Mars or the Moon, the shape and surface relief of the asteroid is not known, thus requiring a dedicated characterization campaign once in orbit. The GNC system shall therefore support the task of finding a safe landing site and ensure a very accurate landing (position accuracy requirement is a few meters which is far below what has been achieved in past missions). The mass, rotational state and gravity field are known to a certain level of accuracy, depending on the performances of the characterization phase. The GNC design shall therefore be robust to environment uncertainties greater than for planetary landing. Thanks to the low gravity environment a multiple landing attempts strategy is foreseen. Contrarily to a planetary landing (which is a “one shot” mission) a specific Failure Detection Isolation and Recovery (FDIR) strategy (including collision avoidance algorithms) shall be designed for proximity operations. The results of the Hayabusa mission (failure of the first landing attempt) show that the robustness of the GNC design is of prime importance for such missions.
The science behind the Marco Polo mission demands a precision landing on the surface of an asteroid which can only be safely achieved after global and local characterization of the object and potential landing sites. The GNC system for proximity operations must meet the requirements for all proximity phases. 2.2 Mission requirements The driving mission requirement for the descent and landing is that the spacecraft should be able to land at any location on the illuminated side of the asteroid. The descent strategy must therefore be designed for compatibility with landing at any latitude within half the rotation period of the asteroid (~3.7 hours). The GNC system and descent trajectory must be designed to provide a soft landing with an impact velocity below 30cm/s vertical and 5cm/s lateral, the relative angle between the spacecraft attitude and the local slope must also be less than 10 degrees. These requirements must be met to ensure the current simple landing gear design and hold down strategy can be used. The accuracy of the landing shall be better than 3.5m which has never been achieved in past missions (Hayabusa achieved about 30m). Such accuracy can only be reached by high-performance image based navigation.
2. AUTONOMY AND GNC REQUIREMENTS 2.1 Science requirements The Marco Polo mission has essentially only one key objective (ref. 1): to bring back to Earth a good, pristine asteroid sample from a primitive asteroid together with the necessary context about the asteroid and landing site from where the sample was collected. To ensure access to the most scientifically valuable areas on the asteroid the GNC system must be able to target a landing at any latitude, the only constraint placed on landing site location is that it should be on the illuminated side of the asteroid enabling the use of visual navigation. Because of the diversity of asteroid environments it is also necessary to collect the contextual information about the asteroid where the sample was collected. This information includes mapping of the mineralogy, morphology and environmental conditions. This information will be collected from a number of closed orbits and hovering positions within the sphere of influence. Measurements are needed at different altitudes to provide the required resolution for the global mapping and the detailed characterization of the potential landing sites.
Fig. 2. Stability geometry of lander showing critical design parameters
Table 1. GNC requirements at landing
The descent phase design also needs to take account of the contamination requirements. Contaminants or undesired
particles coming out of the thrusters must have a density lower than 1014 (1013 as a goal) molecules/cm2 on the NEO surface. 2.3 Autonomy requirements The low gravity environment around NEO modifies the classical task repartition between the ground and the board. The guidance task (designation of the landing site and generation of reference trajectories) can be performed by the ground (as done successfully for Hayabusa) using the results of the characterization phase and the landing rehearsals. The landing rehearsals can be considered as part of the GNC system as they provide valuable information for the guidance task (generation of reference trajectories). During the orbital phase (radio science and global characterization) the spacecraft can be operated from the ground using ground tracking (and possibly visual measurements using a posteriori data fusion) and uploading a pre-determined manoeuvre plan, as the orbit is relatively stable and safe. On the contrary an autonomous navigation & control function is required for the descent phase (including local characterization, landing rehearsals, landing and re-ascent) for which the communication time with ground is similar to the duration of the phase. This is also required to meet the accuracy specifications at landing.
variety of absolute navigation function (the pointing shall be close to the one used during characterization phases). Moreover only a partial knowledge of the asteroid is required for the absolute navigation (whereas a Hayabusa-like scenario relies on the full asteroid model to navigate). - Safety of the scenario and contamination minimization: any colliding trajectory shall be avoided when the relative velocity is high. Vertical control shall be inhibited in the last tens of meters to avoid landing site contamination.
The deltaV allocation for the descent sequence is weak and can be seen as a degree of freedom (a limit of 3m/s is proposed). This is an important tuning parameter to reduce the effect of perturbations and shorten the sequence. A preliminary baseline has been proposed to comply with these constraints. It is based on a transfer from a home position (which is the global characterization orbit at 2km altitude) to a gate position (500m above the selected landing site) followed by a controlled vertical descent. It is a good compromise between system requirements (among which the battery sizing is a critical issue) and GNC needs (this trajectory ensures optimal performances of the navigation function) while ensuring the safety of the spacecraft by appropriate tuning of deltaV. Simulation results on a typical scenario show a total deltaV of 1.85 m/s for the complete sequence. A complete trade-off with a Hayabusa-like scenario is still to be performed to refine the solution.
3. DESCENT STRATEGY & GUIDANCE 3.1 Descent strategy The design of the descent phase is subject to a lot of different constraints: - Illumination of the landing site at touchdown: the use of vision-based navigation requires some contrast in the scene, but in the meantime the shadows shall not be too large. A local solar time around 10.00 am at landing seems a good compromise. This constrains the duration of the landing sequence. - Power: the battery-powered phase shall not be too long to avoid over sizing of the battery and penalizing the mass budget. - Communication: images used for navigation shall be relayed to Earth if possible. - Robustness to the perturbations: a short scenario minimizes the effect of perturbations. This is also linked to the selection of the guidance law. - Performances at touchdown: vertical descent is required at some stage to cancel horizontal velocity induced by the rotation of the asteroid. - Compatibility with the navigation function: a vertical descent improves the observability of lateral velocity (which is the key parameter to be controlled) and allows a wide
Fig. 3. A candidate solution for the descent sequence 3.2 Guidance law Thanks to local characterization and landing rehearsals, a very accurate hazard mapping can be done in the landing site area. Using this information the ground is able to designate a safe landing site convenient for sampling operations. Since a very accurate landing is specified, no autonomous hazard avoidance function is required. In this configuration the complexity of reaching a safe landing site is transferred to the
ground (for landing site designation) and to the navigation function (for reaching this landing site). This simplifies a lot the GNC architecture and avoids unexpected failure of the HDA (Hazard Detection and Avoidance) function at proximity of the surface, as it was the case for the Hayabusa mission (the hazard detection sensor was deactivated for the second landing attempt).
Two different methods are available to build a reference trajectory in position and velocity:
4. NAVIGATION SOLUTIONS Two different kinds of navigation function are required to ensure an accurate landing: - An absolute navigation function, in charge of recognizing the landing site that has been pre-designated by the ground. This is the key element of the GNC as the relative navigation function will not be able to compensate for an initial error of absolute position. It is supposed to be used at some specific times and not periodically. This function is the critical one for achievement of the 3.5m position accuracy at landing.
- Define a thrust profile: in this case the guidance law is a succession of deltaV to be commanded along the descent. The command orders are directly sent to the actuators in a feed-forward scheme, and the control is only in charge to compensate for the small discrepancies during the descent. This technique is to be used when the environment (and especially the external forces) is well known as any uncertainty on the external forces will modify the trajectory.
- A relative navigation function, to control periodically the last part of the descent with respect to the landing site (once identified), and to control the spacecraft in hovering phases.
- Define a position / velocity profile: In this case the key element of the system is the control. The control system shall compute in closed-loop the thrusts that are required to follow the P/V profile. The P/V profile is only a geometrical profile to reach a point P2 from a point P1 in a given frame, without any consideration of external forces. Therefore this kind of guidance law is not sensitive to environment uncertainties, because the spacecraft observes in closed-loop its position and velocity and computes the required thrust to stick to the reference trajectory. This kind of guidance can lead to a wide range of deltaVs depending on the environment uncertainties (thus the thrusters configuration shall allow some flexibility) but has the great advantage to be robust to these uncertainties: the difference between the effective trajectory and the reference trajectory is only linked to the capacity of the spacecraft to observe and control its position / velocity at the required frequency.
3D models matching method (3D / 3D)
The first solution is well adapted to the first phase of the descent sequence (transfer from home to gate position) as the altitude is sufficient to consider that environment uncertainties are weak. Moreover the altitude of the gate position is high enough to be robust to some guidance errors, as the only requirement at gate position is to have the selected landing site in visibility to trigger the second phase, and this can be ensured by proper sizing of the field of view (FOV) of the navigation camera. For the vertical descent from gate position to landing the second solution shall be selected. Indeed in this case the environment uncertainties become more and more significant (this is especially true for the gravity field of the asteroid). Moreover this critical phase shall ensure the accuracy of the landing: therefore the spacecraft shall stick accurately to the vertical descent to minimize transverse velocities at landing and position errors.
Different vision-based solutions have been analysed for both absolute and relative navigation functions. 4.1 Absolute navigation for pinpoint landing
The first solution is an innovative technique, based on the matching of 3D shapes. Considering having the 3D model of the asteroid on-board, it is possible to use it for comparison: the asteroid portion visible along the final descent (or earlier) can be reconstructed on-board with a 3D dense reconstruction algorithm, using the motion of the camera to have stereo-like algorithm. Once the 3D portion is recovered, a 3D matching with the on-board 3D model allows estimating the pose of the lander.
Fig. 4. 3D matching principle: recognition of a DEM portion, from an elevation map estimated onboard during the descent
Model-Image matching method (3D / 2D) The 3D model can also be compared with the 2D descent images directly. The comparison can then be done by comparing the observed contour of the asteroid in the image, and the contour as simulated on-board with the 3D model of the asteroid and a 3D graphic engine. A more complex solution is to have a full 3D model of the asteroid, including
texture, and then compare an on-board simulated image and the actual images.
- Navigation filter covariance for {{position}{attitude}} = {{2500m , 200m , 1000m}{0.1°}}, 1 sigma - Actual error = {{1000m, 200m, 500m}{0.03°}} - Landmark 3D relative position error: 2m. The parameters that were changed to measure their impacts are: - The sun elevation and azimuth, between the orbital image and the in-flight one.
Fig. 5. Image-3D matching principle: recognition of the orientation of the asteroid from a descent image (left) and the on-board DEM (right).
- The pitch of the camera (while the altitude remains constant). - The number of feature points.
Landmarks extraction and matching method (2D / 2D, MAGELLAN) Another approach is to use database image patches of some landmarks, and to locate them on-board in descent images, to finally recover the pose of the camera. Astrium Satellites has already developed such algorithms in the MAGELLAN R&D study (MAtching of Ground ELements and Landmarks for Approach Navigation), which was designed for lunar environment and mission. Although it is not directly applicable on asteroids, this method is a good baseline for a new method dedicated to small bodies. The MAGELLAN algorithms are based on a succession of correlation methods: first, a rough correlation is applied on the images within a search box, whose size depend on the pose error at the output of the navigation filter. Secondly, a sub-pixel fine estimation is applied to precisely recover the Line of Sight (LOS) of the landmarks. Those steps use the 3D model of the asteroid to reshape the appearance of the images, in order to match actual and database point of views. Finally, after a rejection of badly matched landmarks, MAGELLAN can either deliver the LOS of the landmarks, or directly estimate the relative position and attitude of the camera from the 3D known position of the landmarks. Both of these measurements can be used in a Kalman-like filter to improve the estimation of the spacecraft’s pose, obtained from other sensors. Simulations have been performed on the lunar case to validate the algorithms and assess their performances. The image generation tool PANGU (Planet and Asteroid Natural Scene Generation Utility from University of Dundee) has been used to generate lunar images. The lunar scenario used for the robustness and performances tests is: - Altitude = 21 km (coast phase) - Camera pitch: from -90° (nadir) to -40° - Sun position in orbital image: Elevation and Azimuth = 30° - Sun position in the in-flight image: Elevation from 2° to 90°, Azimuth = 30°
Fig. 6. MAGELLAN simulation on the Moon: real position of the landmarks (left), a-priori position of the landmarks, from the navigation parameters (middle), and estimated position of the features, after matching (right). The results are fairly good for the targeted application (see Table 2). The pose estimation methods (DLT and Homography characterization) are only based on a single image, hence providing the performances at worst. Note that there is no error on the landmarks 3D positions, which can be considered as a bias on the final results. Position error
Attitude error
Average LOS error on the feature points
• Illumination conditions : similar • Pitch = -85 (quasi-nadir)
20 m
0.03
0.037
• Sun azimuth difference = 40 • Sun elevation difference = 40 • Pitch = -85
72 m
0.2
0.040
• Illumination conditions : similar • Pitch = -50
232 m
0.5
0.067
Table 2. Performances of absolute position and attitude estimation for a lunar scenario (coast phase) Good robustness with respect to Sun azimuth and elevation has been observed. Robustness to pitch is relatively important too. The simulations results indicate a quasiconstant performance while the LOS is not further than ~30° from nadir, and then a rapid degradation. Regarding the performance values themselves in the ideal cases (no important pitch, no important sun position difference between database and in-flight images), the average LOS precision is better than 1 pixel, so it proves that the sub-pixelic matching has sense. The rejection is an important step, as about 15% of the features are badly matched for each image.
Finally, the position restitution error remains below 100m (or 0.5% of the altitude) 3 sigma in nadir conditions, i.e. better than the initial position error. Further works should be done to study the improvements of introducing the features LOS restitution in a Kalman-like filter. A ratio of at least 2 can be expected. The ratio of 0.5% of the altitude may not be constant along the flight, so that more realistic test cases are needed to confirm this ratio. 4.2 Relative navigation for safe and accurate landing Astrium Satellites has developed for many years the NPAL system (Navigation for Planetary Approach and Landing, ref. 2), based on unknown features extraction and tracking in order to recover at high frequency the 3D relative state during the descent. This technique is a good candidate to provide the relative navigation function for Marco Polo, provided that a few adaptations to the asteroid case are brought to the system. In classical planetary landing the V/d ambiguity (zoom effect) that appears in the system is solved using acceleration sensing. This is not possible for Marco Polo since the gravity is very weak: another solution has to be found to solve this ambiguity and increase observability and robustness of the system. Different methods have been identified: - NPAL + hovering stations: The principle is to use a hovering station to calibrate the velocity of the spacecraft (the velocity is null when all features are motionless). - NPAL + altimeter: Here the idea is to use a direct range measurement to solve the V/d ambiguity. This is a good option with respect to health monitoring and FDIR as the altitude of the spacecraft is directly measured. - NPAL + target markers (TM): The idea is to release a few (2 or 3) target markers on the ground during the landing rehearsal, that can be easily detected by a flash camera. Then a picture of the landing area is processed by ground operators in order to determine the exact location of the TM. These TM are then used during the descent as known landmarks in order to solve the V/d ambiguity, as done for the Hayabusa mission. Robustness and FDIR strategy shall be considered as key elements to trade these different solutions and select a baseline. Thus the use of an altimeter appears an interesting option. The NPAL system has been adapted to the Marco Polo mission for a preliminary assessment of performances, on a 500 m vertical descent at 20 cm/s. No additional measurement has been hybridized with NPAL. Due to weak perturbations a pulse counting method coupled with a dynamics model has been used instead of an accelerometer. In this test case the NPAL system tracks 20 points at 0.1Hz, with a typical white noise standard deviation of 0.2 pixels. Initial estimation errors are set to 1m in position and 1 cm/s in velocity (all axes).
Fig. 7. Performances of NPAL without hybridization on a 500 m vertical descent at 20 cm/s
Table 3. Performances of NPAL without hybridization on a 500 m vertical descent at 20 cm/s The results show that it is possible to design a high accuracy relative navigation function on the base of NPAL hybridised with a relevant additional measurement. In particular the estimation of cross-track position and velocity is very good (no drift with respect to initial error is observed), which is the most critical point during the final vertical descent to ensure the landing accuracy. The degradation observed on the alongtrack axis (vertical axis) is due to the classical V/d ambiguity and is easily corrected by proper hybridization of NPAL with an additional measurement. 5. CONCLUSION The GNC system appears the most critical one for the challenging Marco Polo mission. A preliminary baseline has been defined on the base of past and on-going technology R&D activities led by Astrium Satellites, using advanced vision-based navigation techniques for both absolute and relative navigation. The detailed design of GNC algorithms and associated FDIR strategy is about to be started in the coming weeks. A complete validation in a real time environment and under representative dynamics will be performed to demonstrate performances and maturity of the GNC system. REFERENCES [1] ESA, “Marco Polo Mission Requirements Document”, issue 4 revision 1, May 2009. [2] Bodineau G., Boléat C., Flandin G., Frapard B., Polle B., Mancuso S. “Vision navigation for European landers and the NPAL project”. 17th IFAC Symposium on Automatic Control in Aerospace, June 2007. [3] J. Kawaguchi, S. Aida, H. Morita, “Hayabusa Detailed Guidance and Navigation Operations during Descents and Touchdowns”, AIAA 2006-6536.