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GNC 2021 ABSTRACT BOOK

Contents GNC Posters ...... 7 Poster 01: A Software Defined Radio and GPS SW receiver for real-time on-board Navigation for space missions ...... 7 Poster 02: JUICE Navigation camera design ...... 9 Poster 03: PRESENTATION AND PERFORMANCES OF MULTI-CONSTELLATION GNSS ORBITAL NAVIGATION LIBRARY BOLERO ...... 10 Poster 05: EROSS Project - GNC architecture design for autonomous robotic On-Orbit Servicing ...... 12 Poster 06: Performance assessment of a multispectral sensor for relative navigation ...... 14 Poster 07: Validation of Astrix 1090A IMU for interplanetary and landing missions ...... 16 Poster 08: High Performance Control System Architecture with an Output Regulation Theory-based Controller and Two-Stage Optimal Observer for the Fine Pointing of Large Scientific ...... 18 Poster 09: Development of High-Precision GPSR Applicable to GEO and GTO-to-GEO Transfer ...... 20 Poster 10: P4COM: ESA Pointing Error Engineering For Telecommunication Missions ...... 22 Poster 11: Qualification Results of Honeywell’s 3-Axis HG4934 Space Rate Sensor ...... 23 Poster 12: REVLANGNC flight demonstrator: Development and maturation of GNC technologies for Descent and Landing ...... 25 Poster 13: Autonomous rendezvous GNC system development - overview and recent advances ...... 27 Poster 14: BATCH ORBIT DETERMINATION COMPARISON FOR GEOSTATİONARY SATELLITES ...... 29 Poster 15: Uncertainty propagation in a Guidance, Navigation & Control system for landing operations using Differential Algebra ...... 30 Poster 16: Image simulation for space applications with the SurRender software ...... 32 Poster 17: Real-time thermal infrared image generation for testing future vision-guided autonomous navigation systems ...... 34 Poster 18: Study of an Orbital Transition Scenario to NRHO Considered Rendezvous of Two Spacecrafts . 36 Poster 19: Highly efficient damping system for structural resonances of flywheel masses in reaction wheels ...... 38 Poster 20: SPACE RIDER: TAEM GNC, focusing on TAEM Hybrid Navigation of the Future European Reusable Space Transportation System ...... 40 Poster 21: Towards a Scalable Technology for Albedo Insensitive Sun Sensors in Silicon Carbide ...... 42 Poster 22: Guidance, Navigation and Control for Orbit Station-Keeping with In-Situ Gravity Estimation ...... 44 Poster 23: DOCKING VIA FUSION OF MONOCULAR VISION-BASED SYSTEMS ...... 46 Poster 24: ALINA GNC – Architecture & Design ...... 48 Poster 25: Application of Microthrusters for Space Observatory Precision ...... 50 Poster 26: Design of the Spacecraft Guidance, Navigation and Control Software Functions for Impulsive Manoeuvre Execution on the Coronagraph Spacecraft of the Proba-3 Mission ...... 51 Poster 27: Spacecraft Flexible Attitude Dynamics Modeling for Accurate Control Design ...... 53 GNC for Small Body Missions ...... 55 Centralized Relative Navigation of Multiple Spacecraft with reduced visibility: validation campaign of the vision-based approach ...... 55 Impact of Complex-shaped Small Bodies on a Quadcopter-based Controller for Near-surface Operations ...... 57 LOCALIZATION AND MAPPING MERGING SILHOUETTES INFORMATION AND FEATURE TRACKING FOR SMALL BODY APPLICATIONS ...... 59 Mission and GNC system design of the Juventas CubeSat on-board the Hera mission ...... 61 HERA GNC Subsystem – Preliminary Design ...... 63 Current Space Missions ...... 65 GNC NEO: a full electric platform for telecommunication satelittes, safe mode on star tracker and electrical orbit raising ...... 65 The AOCS and FGS verification tasks at System level ...... 66 AOCS INNOVATIONS FOR NEO FULL-ELECTRIC PLATFORM ...... 67 GNC for Future Space Transportations Systems ...... 68 Robust Control for Reusable Rockets via Structured H-infinity Synthesis ...... 68 Design and testing of the GNC for the HERACLES Lunar Ascent Element ...... 70 Dynamic Modelling and Control of an Aerodynamically Controlled Capturing Device for "In-Air-Capturing" of a Reusable ...... 71 Trends in AI for GNC Systems ...... 73 Pose estimation of a non-cooperative target based on silhouette imagery using convolutional neural networks ...... 73 Using Convolutional Neural Networks for Relative Pose Estimation of a Non-Cooperative Spacecraft with Thermal Infrared Imagery ...... 75 An unsupervised deep learning approach to on-board AOCS FDI(R), first results and conclusions ...... 77 Advances in Control ...... 79 Sloshing AOCS/Fluidic coupled analysis ...... 79 Development of a robust GNC architecture for a flexible spinning spacecraft with long wire booms ...... 81 Thrusters Off-modulation for Attitude Control during Orbital Manoeuvers ...... 83 EVALUATION OF NONLINEAR MODEL PREDICTIVE CONTROL FOR CUBESAT ATTITUDE CONTROL; A HARDWARE-IN-THE-LOOP SIMULATION ...... 84 NMPC Based Guidance and Control for Observation Missions ...... 86 Convex guidance for Close Rendezvous operations in cislunar Near Rectilinear Halo Orbits ...... 88 Sentinel-2: In-orbit Benchmark of Precision Star Tracker / Gyro Attitude Estimation and Smoothing ...... 90 High-Order Sliding Mode Controller for the test mass stabilization of the LISA Mission: Preliminary Results ...... 92 ASCENT FLIGHT CONTROL SYSTEM FOR REUSABLE LAUNCH VEHICLES: FULL ORDER AND STRUCTURED H∞ DESIGNS...... 94 Legs on Aerial Robotic Drone: Control Design Challenges of Agile Thruster-Assisted Legged Locomotion . 96 ROM-Based Feedback Design for Thruster-Assisted Legged Locomotion ...... 97 Advances in Sensors and Actuators ...... 99 Flash LiDAR for and Docking Missions ...... 99 RVS®3000-3D LIDAR – Gateway Rendezvous and Lunar Landing ...... 101 FaintStar - An Intelligent Single-Chip Sensor Head for Star Trackers - Evaluation Results ...... 102 The Terma T1 Star Tracker Qualification Test Results and New Developments ...... 104 NEWTON CMG PACKAGE : Making CMG control as easy as wheels ...... 105 Reaction Wheels with Internal Speed Control ...... 107 Magnetorquers design for improved demisability in uncontrolled re-entry LEO platforms ...... 108 AQUILA: development of a European high-accuracy accelerometer for space applications ...... 109 A new iXBlue Fiber Optic Gyroscope offering accuracy and reliability at a fraction of their historical cost ...... 110 ARIETIS and ARIETIS-NS, new highly competitive gyro solutions for space applications ...... 111 Evolving SiREUS ...... 112 IBIS and MAUS, the Jalapeno’s under the Sunsensors...... 113 Challenges of LEO missions ...... 114 Utilizing for a Space Interferometer Laboratory by Three Satellites ...... 114 GNC design solution for the deployment of BIOMASS large deployable reflector ...... 116 Lightweight algorithms for collision avoidance applications ...... 118 AstroBus product AOCS architecture for LEO missions ...... 120 Microsatellite AOCS design for the agile mission MICROCARB...... 121 Passive rate damping of non-operational satellites in Low Earth Orbit to enable Active Debris Removal 123 High Performance Pointing Systems ...... 125 High accuracy image stabilization system for GEO High Resolution Missions ...... 125 Reduction of flexible modes excitation during slews...... 127 Characterization of SADM induced disturbances and their effects on spacecraft pointing errors ...... 128 On-board Real-time Calibration of Non-synchronous Attitude Sensors and Gyros ...... 129 Spacecraft Line-of-Sight Jitter Mitigation and Management Lessons Learned and Engineering Best Practices ...... 131 In-Orbit Experiences and Demonstrators ...... 133 In-flight Experience on GNC Design Challenges for BepiColombo ...... 133 AOCS in flight return of experience ...... 135 Visual Sensor Suite – Flight Experience on MEV ...... 136 Flight of a New Miniaturised and Integrated Spaceborne GNSS Receiver ...... 137 Auriga CP and Auriga SA flight heritage ...... 139 The imaging LiDAR of the in-orbit demonstration RemoveDebris ...... 140 Aeolus AOCS In-flight Performance ...... 141 In-Orbit Experiences with the Spin Stabilized Attitude Control System of Eu:CROPIS ...... 142 AOCS support for optical communication experiments with the small BIROS ...... 144 Sensors Data Fusion and Autonomous Navigation ...... 146 GENEVIS: Generic Vision-Based Navigation for Descent & Landing ...... 146 Computational Guidance & Navigation for Bearings-Only Rendezvous – methods and outcomes of GUIBEAR ...... 148 Distributed Cooperative Visual Odometry for Planetary Exploration Rovers ...... 150 A deterministic and high performance parallel data processing approach to increase guidance navigation and control robustness...... 152 GPS Constellation Modernization Impact on GPSR at GEO ...... 153 Current Space Mission Validation & Verification ...... 155 SmallGEO Product Line FlatSat - A powerful tool for AOCS verification ...... 155 Independent V&V of the ExoMars 2020 Cruise GNC, through analytical techniques, simulation-based analyses and tests ...... 156 Automatic code generation of the GNC software for Spacebus Neo ...... 158 A LISA mission simulation environment: TAS NUMES simulator framework a powerful tool for future missions ...... 159 Future Space Missions ...... 161 GNC design for Rendezvous Autonomous CubeSats Experiment (RACE) Mission ...... 161 Sample Return – Test campaign for Near Range Image Processing on European Proximity Operations Simulator ...... 163 LICIA GNC baseline for DART-Didymoon impact tracking ...... 165 The LISA DFACS: overview of the design activities for the drag-free mode ...... 167 Investigation of Multi-Body/Multi-Actuator Modeling Techniques for Applicability to Future Space Observation Missions ...... 169 Closed-Loop Guidance for Low-Thrust Interplanetary Trajectories Using Convex Programming ...... 170 Autonomy, Fault Tolerant Control and Operations ...... 171 Autonomous Guidance for Electrical Orbit Raising ...... 171 Measuring Resilience of Autonomous Controllers to Spacecraft Missed Thrust Events ...... 172 GNC for In Orbit Robotic Operations ...... 173 Modelling and attitude control design for autonomous in-orbit assembly ...... 173 EROSS Project – Coordinated control architecture of a space robot for capture and servicing operations ...... 175 Results of the COMRADE project: Combined control for and manipulator in servicing missions: Active Debris Removal and Re-fuelling ...... 177 Critical GNC Aspects for ADR missions ...... 179 Real-time combined control for active debris removal for a satellite with a robot arm implemented on a SBC connected to a detailed multi- and VR simulation ...... 181 GNC for Planetary Exploration ...... 182 Development, tests and results of onboard image processing for JUICE ...... 182 Application of Advanced Navigation Techniques for Lunar and Mars Pinpoint Landing ...... 183 Autonomous Guidance, Navigation and Control Systems Development and Validation ... 185 Guidance, Navigation, and Control System for NASA’s Mission ...... 187 Control Challenges for NASA’s Mars Helicopter ...... 189

GNC Posters

Poster 01: A Software Defined Radio Galileo and GPS SW receiver for real-time on-board Navigation for space missions Bidaux-Sokolowski A1, Kobyłkiewicz A1, Wildowicz S1, Paśnikowski M1, Sarnadas R2, Giordano P2 1GMV Innovating Solutions Sp. z o. o., 2ESA/ESTEC The use of GNSS receivers in space has become quite common for the Low Earth Orbits navigation offering the advantage of continuous position velocity and time knowledge on-board, possibility of precise orbit determination (few centimetres accuracy in post processing), as well as some scientific instruments as reflectometry or radio-occultation. Up to now, most of the receivers for space applications are based on HW design – ASIC or FPGA and the orbital navigation as part of the GNC is external to the receiver or duplicates the inner navigation of the receiver. The novelty of the approach presented in the current article lays in implementing all the receiver functionalities, including the signal processing, in SW together with the navigation. This approach gives a large flexibility in modifying or tuning the SW including the signal processing for different objectives, in particular non-standard ones like higher sensitivity, launcher dynamics, thrusting or relative GNSS among others. An additional advantage, when one looks at navigation performance and measurements noise reduction is to implement consistent coupling techniques or aiding schemes for more accurate tracking as for example using directly the in phase and quadrature measurements from the tracking loops in the navigation filter and/or using an IMU aiding scheme for rapid change in accelerations. The receiver is thought to be adaptable to different platforms targeted and has been tested on: PowerPC, Zynq-based & GR740. The adaptability to multiple platforms has twofold objectives: to integrate the receiver on an on-board computer of the client with limited integration effort and to upgrade the receiver when more powerful computers/processors becomes avialable. The receiver, which was designed for single CPU at first, has been adapted in depth to allow for parallel execution of tasks with the recently updated version of RTEMS for Symmetric MultiProcessing (SMP) on the relatively new GR740 platform (4 cores). This parallelization approach makes the receiver scalable to multi core processors with minor SW updates, and also allows to host various applications SW on the same computer: the GNSS receiver being possibly run on one to four cores, giving the possibility to host for example a guidance, navigation and control SW on one core and the receiver on the remaining three. The evolution of the real-time operating system: RTEMS used for more than two decades for space application for SMP allows the system to assign automatically tasks to free cores of the microprocessor – not currently busy at executing a task and/or let the user assigned manually task to the defined core – so called affinity. The GNSS SW receiver consists of the five following tasks: • A Main task which is responsible for parsing TM/TC, getting data from the Radio Frequency Front End, navigation and controlling RTEMS tasks, • one acquisition (ACQ) asynchronous task for searching of the available satellites by mean of FFT based search, • 3 tracking (TCK) tasks for the tracking channels and the generation of raw measurements. Each tracking task being composed of a set of tracking channels. Each task is synchronised by RTEMS event manager which is a high-performance method for inter-task communication with a minimum delay impact. The high level schematic architecture of the GNSSW receiver is presented in the block diagram attached to this abstract. It depicts the functional interface between the different functions and data flow. The GNSS core of the receiver lays in the main 2 signal processing functions: the acquisition which is in charge of detecting the GNSS satellites in view and the tracking of the signal for all the channels maintaining a continuous estimate and control of the GNSS code delay, Doppler and carrier phase in the tracking loops. While the acquisition mainly relies on FFT/IFFT techniques, the tracking uses highly optimised time-based correlation techniques. The FFT acquisition permits decreasing the search space in cold mode to a two degrees of freedom: GNSS satellite and Doppler search space while a time based search would require an additional dimension to the search space on the code delay. The main trade-off for the acquisition machine is between the sensitivity and the time to acquire a satellite, the sensitivity can be increased at the cost of the execution time by performing FFT on larger portion of the signal incoming. This parameter is left to the user tuning depending on the application and required sensitivity. In LEO, where high signal to noise is available for a zenith pointing antenna, the time to first fix in cold start of the receiver is quite fast with only three minutes mean and 6 minutes worst case obtained from Monte Carlo analyses. The tracking instead is performed continuously and in real time on the full incoming signal from the FE, different methods has been implemented to increase the stability and performance of the tracking including carrier aiding, pilot signal tracking for Galileo E1, long coherent integration time and tuneable order and bandwidth of the tracking loops. The article will depict the raw measurements and navigation results of the receiver for various applications including Low Earth Orbiting satellites, micro launcher, formation flying with carrier phase based dual difference and increased sensitivity of the receiver for higher orbital regimes. The tests are performed using the target receiver HW including the integrated processing board with the RF FE, using the latest generation of Spirent GNSS signal simulator available at the Warsaw Military University of Technology and ESA/ESTEC Facility in Noordwijk. The results attached to the abstract were obtained for a 400 km sun-synchronous orbit . For this test Ionospheric errors were disabled and ephemerides errors were enabled. The navigation filter results were compared with the Spirent outputs. Overall performance is quite satisfactory, with 3D RMS of the position error around 1 m and velocity 3D RMS error around 1 cm/s. The receiver developed is planned to be flown on two missions in the short future: on GOMX-5 with an accurate dual frequency version on a GR740 processing platform and on the MIURA-1 micro launcher on a Zynq 7030 based computer.

Poster 02: JUICE Navigation camera design Gorog F1, Charavel R1, Arnolfo M1, Dervaux S2, Gherardi D3 1Sodern, 2Airbus Defense and Space, 3ESA-ESTEC The Jupiter Icy Moon Explorer (JUICE) is an ESA mission whose aim is to study the Jovian system : Jupiter itself, its icy (Europa, Ganymede and Callisto) and the magnetosphere. For this challenging mission, Sodern was selected to provide the Navigation Camera (NavCam).

NavCam is a key instrument for the spacecraft navigation. During its lifetime JUICE does numerous fly-bys, however the V is limited by the available propellant in the spacecraft. To achieve such precise navigation in Jovian environment, ground based radiometric measurements are not sufficient, therefore it is paramount to measure in-flight, with high accuracy, the position of the Jupiter moons.

To achieve this measurement, the NavCam assesses the position of the moon and stars in its field of view in an inertial reference frame. The combination of these two measurements gives the accurate position of each of Jupiter icy moons. The stars processing is similar to that of a star tracker and is performed inside the camera, whereas the moons detection is done on-ground from images captured by the camera. In addition to the above primary mission, the navigation camera will also be used by the spacecraft attitude control loop during close fly-bys in order to improve the spacecraft absolute pointing performance. Jovian environment is particularly severe. The mission profile imposes considerable radiations constraints on the spacecraft, which are uncommon in typical space programs. Jupiter magnetic moment is the largest of the system (over 10,000 times higher than the Earth’s). It induces a high total dose exposure at electronic parts level (100krad), high number of Single Event Effects and internal charging effects. Furthermore, to achieve the best performances, the NavCam is mounted on the optical bench near sensitive instruments; so it must comply with stringent thermal control and magnetic cleanliness to avoid disturbances to the other instruments. Finally, the measurements of moons and stars in the same field of view induce complex optical design to mitigate stray light. Star trackers and instrument experience paired with radiation and optical engineering know-how have put Sodern in the position to solve all these challenges. JUICE NavCam passed the CDR stage; the paper describes the final features and the ongoing model activities.

Poster 03: PRESENTATION AND PERFORMANCES OF MULTI- CONSTELLATION GNSS ORBITAL NAVIGATION LIBRARY BOLERO Delong N1, Guérin P2 1CNES, 2CS-SI 1. INTRODUCTION

Compared to a simple GNSS receiver, the interest of an orbital navigation filter is to improve the quality of the position and clock estimation of the receiver, taking benefit from a dynamical and a measurement model. It can be used to extrapolate position and time, for mission purpose, to support the receiver for signal acquisition, or eventually to provide input data to an Autonomous Orbit Control module, improving therefore the autonomy of the spacecraft. BOLERO, a navigation library developed by CNES since 1998, was recently adapted to multi-constellation purpose. This article presents the functionalities of the library and the validation and performances analysis, showing the advantages of multi-constellation GNSS navigation.

2. BOLERO LIBRARY PRESENTATION

BOLERO means Bibliothèque d'Objets Logiciels Embarqués de Restitution d'Orbite, that is « on board orbit determination objects library ». BOLERO was developed in C language and is compliant with embedded code constraints. The first version was limited to mono constellation GNSS measurements and was integrated on several receivers and used in different space missions. In 2019, it was decided, due to emergence of multi-constellation GNSS receivers, to adapt the library to the possibility to handle any GNSS constellation measurements together. The library is adapted for any type of orbit, from Low Earth Orbit (LEO) to Geosynchronous Earth Orbit (GEO), due to its very complete dynamic model. BOLERO is based on a Kalman filter. The state vector represents the parameters that can be estimated by the filter: they are divided into the dynamical parameters, used to compute the movement of the spacecraft and into the measurements parameters, such as the receiver clock bias model, the carrier phase ambiguity and inter-constellation bias for example. BOLERO uses as input various GNSS measurements: pseudo-range and carrier phase, single or dual frequency, for single or multi-constellations.

3. VALIDATION AND PERFORMANCE

The aim of the validation and performance study is to confirm the proper functioning of the estimation filter, and to assess the precision on the estimated parameters. In particular, validation tests allow to verify the correct convergence of the Kalman filter and to evaluate its convergence duration. They check the rejection of wrong measurements. For performance tests, we are interested in particular in evaluating the precision of position and clock model of the receiver, in real time as well as after an extrapolation period. For that reason, the validity of the estimated dynamic model is to be evaluated. This concerns for example the solar pressure and the drag coefficients, the maneuvers as well as the instantaneous pole of Earth rotation. The validity of the measurements parameters is also checked: inter-constellation bias, carrier-phase ambiguities. A particular attention is payed to compare the interest of multi-constellation with mono-constellation. 3.1. Methods and principles These tests were done first based on simulated measurements, then using real flight measurements:

The interest of simulation in these validation tests is to have a total control of dynamic and measurement models. The comparison between the estimated parameters and the reference gives directly the error made by the filter. For this study, a 3 days’ ephemeris of a LEO satellite is generated with a full precision dynamic model, using the CNES precise orbit determination and extrapolation software named ZOOM. For Galileo and GPS constellation data, we use a real precise International GNSS Service (IGS) orbits and clocks solution. The receiver clock is produced on ground by a DORIS like ultra-stable oscillator. Bi-frequency GPS and Galileo measurements are then simulated, including perturbations like carrier phase ambiguity, inter-constellation bias, windup, emitter and receiver clock biases.

For the second part of the tests, we use real flight measurements from a bi-frequency and bi-constellation GPS-GAL receiver, on a LEO orbit. In this case, the reference orbit was produced on ground during launch and early orbit phase operations, using a full model precision for dynamic and measurements and using IGS solution for GNSS orbits and clocks in differed time.

For both series of tests, bi-frequency measurements are treated using the ionosphere-free combination, to remove the major part of this perturbation. Pseudo-Range and Carrier-phase measurements are used. The navigation function gives the possibility to feed the navigator with broadcast navigation data, but also IGS precise solution. The tests are done with different configurations: GPS only, GAL only and both constellations together, with the estimation of the inter-constellation bias.

3.2. Results and comments The extended analysis of the tests is presented in the whole article. In summary, we observe a very good behavior of BOLERO for convergence, wrong measurement rejection and estimation.

For both simulated and flight measurements, the accuracy is lower than 50 cm 3D RMS on position when using bi-constellation measurements with broadcast ephemeris. The bi-constellation estimation appears to be more accurate than the single constellation one, due to the increase of the minimum number of GNSS satellites in visibility. Furthermore, the greater number of measurements allows to smooth the error of GNSS broadcast ephemeris error. We notice that GAL-only solution is often better than GPS-only one, since it benefits from the good quality of its broadcast navigation data.

4. CONCLUSION BOLERO confirms with these validation tests that it has migrated to a high performance multi-constellation orbital filter. Is accuracy make it a good candidate for integration in modern navigators and for use in an autonomous orbit control system. It has already been chosen for some future missions to be launched in the coming years. It will be a good opportunity to test it in real conditions with multi-constellation configuration.

Poster 05: EROSS Project - GNC architecture design for autonomous robotic On-Orbit Servicing Dubanchet V1, Casu D1, Andiappane S1, Béjar Romero J2, Torralbo Dezainde S2 1Thales Alenia Space, 2GMV Current robotic developments for space application are making On-Orbit Servicing (OOS) missions closer to reality, while also paving the way for future challenges like autonomous rendezvous with celestial bodies or Active Debris Removal (ADR). Such technologies are being developed and integrated towards an experimental ground demonstration in the scope of the H2020 project called EROSS (European Robotic Orbital Support Services). The current paper presents the proposed Guidance, Navigation and Control (GNC) architecture along with the Autonomy layer which are developed to further assess the reachable performances and robustness in the scope of a collaborative rendezvous between two spacecrafts. The paper focuses on the overall architecture proposed to handle the GNC and autonomy requirements, with their implementation through the Mission and Vehicle Management (MVM) that triggers and schedules the equipments and processing algorithms.

The work presented in this paper is led in the frame of the H2020 project “European Robotic Orbital Support Services” (EROSS), which received funding from the European Union’s Horizon 2020 research and innovation programme under grant agreement No 821904. Therefore, the EROSS GNC and autonomy layers are built upon the past Operational Grants (OG) outputs led in the scope of the Strategic Research Cluster (SRC) on Space Robotics [3-4-5]. A particular focus is given on the ERGO autonomy framework of OG2, on the INFUSE data processing framework of OG3, and on the I3DS sensors developed and integrated within OG4.

The EROSS mission of application considers the servicing of a Client satellite designed to be serviced and whose behaviour is actively controlled to ease the rendezvous with the Servicer spacecraft. This Client is then said to be “prepared” and “collaborative”. From the Servicer GNC point of view, it means that the Client is equipped with dedicated features to ease the relative navigation or the servicing tasks during the rendezvous manoeuvers, while its own control is made compatible with the mission steps. EROSS scenario mainly considers the last phases of a rendezvous mission.

A traditional rendezvous mission covers the following key steps [1,2]: A. Launch & Orbit Injection B. Orbit Phasing C. Far-Range Rendezvous (or Homing), D. Close & Proximity Operations (or Closing), E. Mating Operations (including forced motion and capture), F. Servicing Operations, G. Departure Operations. The EROSS project mainly focuses on the last steps of the rendezvous with the “E - Mating Operations” (MOP) and the “F - Servicing Operations” (SOP).

The GNC control loop of the Servicer follows a classic automation scheme with a mix of physical and numerical elements. This split with the functions embedded in the “On-Board Software” (OBSW) block, segregated from the physical units within the “Dynamics, Kinematics and Environment” (DKE) block. This latter will be partially implemented through physical units for the actuators and sensors related to the rendezvous functions, while numerical models will represent the rest of the platform hardware to mimic their behaviour during the experiments. On top of this traditional control loop, the autonomy layer is implemented within the generic “Mission and Vehicle Management” (MVM) block that triggers the GNC modes and manages the related configuration of the GNC functions and of the scheduling of the equipments and their proper activation. This blocks is also interfaced with the external environment (i.e., ground segment or other spacecraft if any) to exchange meaningful information to support the definition of the Servicer level of autonomy (e.g., E1, E2, E3, E4 according to [7]) depending on the high or low level of the commands received and on the needs of processing on-board.

The “Mission and Vehicle Management” (MVM) block is developed within the ERGO framework coming from the OG2 project, allowing to reach the required autonomy level for the EROSS mission (i.e., mainly E3 level will be demonstrated). The MVM function is made up by three components: Ground Control Interface, Mission Planner and Command Dispatcher. These MVM components are in charge of sending commands or splitting the goals of the different mission phases into subsystem commands (e.g. platform GNC commands, robotic arm GNC commands, ...) in order to fulfil the mission objectives. It thus triggers the GNC modes at system level, and then set up the different G-N-C sub-modes for the Servicer platform hardware and for its robotic arm. The Guidance functions and MVM framework are handled through the OG2-ERGO project and extended by GMV [3], the Image Processing Navigation ones are built upon the OG3-INFUSE project and applied on EROSS by Space Applications Systems [4], while the GNC architecture with the logic behind the MVM, the Navigation Data Fusion and the Control functions are based on expertise. On the hardware side, most of the sensors used within EROSS are cameras coming from the OG4-I3DS project [5-6]. One of the most challenging equipment being integrated is the robotic arm with its related GNC functions for a coordinated control of the Servicer platform and arm during the capture, based on the expertise of the National Technical University of Athens (NTUA), of the Polish company PIAP-Space, and on the support of the Canadian company MacDonald, Dettwiler and Associates (MDA). This paper will detail the GNC & MVM architecture being designed for the EROSS project with the different trade-offs made at each G-N-C level and how the functions are orchestrated at the MVM level.

[1] Fehse W., “Automated Rendezvous and Docking of Spacecraft”, 2003. [2] Flores-Abad A. et al., “A review of space robotics technologies for on-orbit servicing”, 2014. [3] Ocón J. et al., “ERGO: A Framework for the Development of Autonomous Robots”, 2017. [4] Dominguez R. et al., “A common data fusion framework for space robotics : architecture and data fusion methods”, 2018. [5] Dubanchet V., Andiappane S., “Development of I3DS: An integrated sensors suite for orbital rendezvous and planetary exploration”, 2018. [6] Dubanchet V., et al., "Experimental assessment of I3DS performances: a suite of sensors for on-orbit rendezvous", 2019. [7] ECSS-E-70-11C, “Space Engineering - Space segment operability”, 2008.

Poster 06: Performance assessment of a multispectral sensor for relative navigation Dubanchet V1, Dubanchet V1, Weigel T2, Fiengo A3, Bravo Pérez-Villar J3, Rosa P4, Sanchez Gestido M5 1Thales Alenia Space, 2Thales Alenia Space, 3DEIMOS Space S.L.U., 4DEIMOS Engenharia , 5ESA / ESTEC Multispectral imaging techniques have been widely used in space at instrument and payload level for the observation of the Earth or of any celestial body. These techniques often result in complex instruments able to handle the different spectral bands of interest, because different detectors and an accurate temperature control system are required to stabilize their performances during the mission. Here, an alternative approach of the multispectral technology is presented to tackle the issue of the relative navigation of a spacecraft around either another spacecraft or a celestial body. Advanced image processing and filtering techniques are applied to the images acquired by a dual sensor with two detectors in the visible and thermal infrared bands in order to estimate the relative pose of the targeted object. This paper focuses on the performance assessment of the overall navigation solution made up of a multispectral sensor combined with the related processing using a high-fidelity simulator with synthetic images.

The work presented has been performed in the frame of the “Multispectral Sensing for Relative Navigation” (MSRN) study funded by ESA. It is organized as follows: a first chapter is dedicated to the selection of the spectral bands of interest considering different rendezvous missions of applications, then the resulting design of the multispectral sensor is presented, before describing shortly the image processing and filtering technique. Eventually, the obtained performance results are commented on a set of validation scenarios.

The design of such a sensor has to adapt to the difficulties of the multi-mission paradigm where high performance are expected in a compact equipment fulfilling a variety of mission conditions. The initial trade- off focuses then on the identification of the most interesting spectral bands for each of the scenarios considered initially for the MSRN study. It includes the derivation of a common set of mission requirements to drive the design of the sensor architecture and of the navigation solution.

For rendezvous missions, the most challenging scenario identified is the non-cooperative rendezvous, since no data are exchanged between the target and chaser spacecraft while a high level of autonomy is required. For this scenario, the conditions of the e.Deorbit mission have been selected, focusing on the navigation around , with the Earth possibly appearing in the background of the images taken by a camera on the chaser spacecraft. A multispectral camera is helpful in this case, especially for the segmentation of the target vehicle prior to the image processing and filtering. It allows increasing the robustness of the navigation solution, both at close and long ranges. The activity has also shown the drawbacks of the hyper-spectral bands such as the low robustness to orbit environment, the need for longer exposure times (i.e., not suited to fast moving objects), on top of the additional complexity brought at hardware/software levels.

The second type of scenario considered is the approach and inspection of a binary asteroid. In the frame of this work, the definition of the former Asteroid Impact Mission (AIM), now Hera, to explore Didymos has been selected. For this type of target, several spectral bands of interest have been identified to prepare for the relative navigation with pose estimation, as well as for the asteroid 3D reconstruction. Spectral bands in the ultraviolet or infrared are used to extract spatial features exploiting the different compositions of the ground, while their combination can greatly increase the robustness of the features extraction. The infrared band is of particular interest to maintain the ground features detection above a given threshold when illumination conditions are not favourable.

Based on the set of requirements and on the key spectral bands identified for both missions, a hardware design of the multispectral sensor is proposed at the optical and detector levels. The preliminary design favored a Si-based CMOS detector for the ultraviolet to near infrared bands, while an uncooled µ-Bolometers one is preferred to handle the thermal infrared range. The biggest trade-off in this design phase is to promote the detector technology complying with the expected chosen bands while not requiring complex and costly cooling systems that would dramatically impact the mass and power budgets of the final sensor. In this work, the spectral resolution in visible bands is obtained by a rotating filter wheel placed in the optical path before the visible detector. On the optical side, a WALRUS configuration is used to keep the equipment’s compactness and to allow a single path for both the visible and thermal infrared detectors. This choice is discussed in the paper as it optimizes compactness at the expense of a lower optical performance on the thermal infrared detector.

The image processing and the filtering techniques used for the relative navigation have already been thoroughly presented in [1]. They are based on the combination of machine learning – by using a codebook onto which the relative pose is projected, for each spectral band – with an Extended Kalman Filter (EKF) for fusing the resulting relative measurements with inertial sensors. The codebooks are tuned and trained prior to the mission on a given set of synthetic images, which must be as representative as possible of the optical and electronic design of the multispectral sensor.

Eventually, the performance assessment is conducted on a set of scenarios derived around the two missions of interest mentioned above, with the rendezvous with a spacecraft around the Earth and the inspection of an asteroid in space. The validation of the proposed design has been performed through extensive analyses considering a large panel of simulation scenarii. It has shown the performances of the whole navigation solution, and validates the camera design along with the navigation solution on the different missions considered. The work presented clearly confirms and quantifies the advantage of a multispectral camera on the pose estimation problem during space rendezvous phases with challenging illumination conditions.

References [1] Fiengo A., Bravo J.I., Rosa P., Guercio N., Christy J., and Dubanchet V., “Multispectral Sensing for Relative Navigation”, in Proceedings of 8th European Conference for Aeronautics and Space Sciences (EUCASS), 2019.

Poster 07: Validation of Astrix 1090A IMU for interplanetary and landing missions Masson S1, Loubières P1, Cros G1, Dupuis A1, Pouzargues O1, Bonnefois J2, Shariati S2, Ollivier C2 1Airbus Defence and Space, 2Ixblue In the early 2000s, SAS developed, in collaboration with Ixblue company, and with CNES and ESA support, a family of inertial reference units (IRU) for a wide range of space applications. This family of products, called "Astrix™", is based on solid-state Fibre Optic Gyro (FOG) technology, with the Astrix 200 and Astrix 120 providing the highest performance solutions for mission such as . In 2014, Airbus Defence and Space and Ixblue have developed and qualified Astrix 1090, a 3- axis gyro in a compact and single box configuration, dedicated to mid-level performance applications in high reliability and quality standards, such as Telecom platforms. More recently, in the frame of Exomars 2020, an enhanced version of the Astrix 1090 was developed in order to fulfil the challenging needs of planetary exploration missions such as interplanetary cruise, and Entry, Descent and Landing, which require high dynamic measurements and fine resolutions. Astrix 1090-A configuration was modified in order to implement 3 acceleration sensing axes in addition to the 3 gyro axes, leading to a compact IMU, based on European technology. The Astrix 1090-A was submitted to a specific validation campaign, representative of the mission constraints and has been successfully qualified with respect to Exomars requirements.

This paper will deal with: - The main modifications applied to the Astrix 1090 gyro function to match the planetary exploration requirements - The overall architecture of the added accelerometer function - The main inertial performances and the characterization. - The overall validation results

Technical content: The nominal Astrix 1090 optical design has been upgraded in order to improve the gyro scale factor budget error. During qualification, we proved that the gyro is able to measure rotation rates up to 400 °/s, with angular accelerations up to 2200 °/s² and angular jerks up to 33000 °/s3. In those extreme dynamics, the scale factor error over the temperature range [-10 °C +50 °C] (considering 4th order thermal model) was lower than 30 ppm. The scale factor linearity error has the same order of magnitude for rates up to 20 °/s. This improvement kept the original low noise value, which is mainly driven by Angular Random Walk (below 3.5 10-3 °/√h Beginning Of Life). Three acceleration measurement channels have been specifically developed. Each acceleration channel is an accelerometer unit QA3000 from Honeywell and an analog signal processing. The 3 QA3000 devices are located on the same mechanical frame as gyro sensing head, in order to guarantee the lowest alignment variation between gyro and acceleration sensing axes. The accelerometer signal is integrated by an analog electronics and then digitalized, in order to provide delta-velocity increments with a very low noise. As for gyro channels, a post-processing unit provides thermal compensation for bias, scale factor and alignment. The accelerometer function was designed to measure acceleration up to 20g with a quantization noise smaller than 820 µg @100Hz. The bias stability (over 1h) is smaller than 2 µg, as shown on plot below and the bias error due thermal sensitivity in the range [-10 50 °C] (considering 4th order thermal model) is was measured below 30 µg. The accelerometer scale factor error over the temperature range (considering 4th order thermal model) was measured below 40 ppm, and the linearity error is below 300 µg/g in the range +/- 1g. Astrix 1090A unit has been submitted to an extensive qualification test campaign, which has validated the ability of the equipment to withstand the mission environment within the GNC chain. The main environment requirements were: - Sine environment Frequency All directions (0-peak or max shaker amplitude) Qualification sweep rate 5 Hz 1 g 2 oct/min 5-100Hz Sweep up and down 30 Hz 25 g 100 Hz 25 g

- Random vibration environment Random Vibration Environment Qualification Axis Frequency (Hz) Duration 2 min

Perpendicular and parallel to mounting plane All axes 20 – 40 +6 dB/oct 40 - 450 0.16 g²/Hz 450 - 2000 -6 dB/oct Overall 11.00 grms

- Shock environment Axis Frequency (Hz) Acceleration (g)

Perpendicular & Parallel to mounting plane (3 axes, 3 shocks by axis, total : 9 shocks) 100 25 1800 950 10000 950

- Thermal vacuum cycling in the range [-35 °C , 60 °C], 8 cycles. - EMC environment (RE, RS, CE, CS)

Poster 08: High Performance Control System Architecture with an Output Regulation Theory-based Controller and Two-Stage Optimal Observer for the Fine Pointing of Large Scientific Satellites Fogliano V1, Zanzi M1 1University Of Bologna The work presented in this paper is based on the design and analysis of the Attitude Control Subsystem (ACS) architecture embedded in the Guidance, Navigation and Control (GNC) Unit responsible for the repointing manoeuvre of a satellite. The novelty of the paper contribution is the presentation of a realistic space application of an output- regulation theory based robust controller coupled with a two-stage filter for the state and disturbance estimation capable to improved performances in term of pointing accuracy and stability. In particular, as a real case scenario, the so called Fine Pointing Mode of the EUCLID satellite has been considered as the reference case study. Euclid is a medium class astronomy and astrophysics space mission planned to be launched for 2022 by ESA, aimed at understanding the expansion of the Universe. Differently from the actual implementation of the spacecraft ACS designed by the Euclid Consortium team for this particular flight condition, in the presented case study the satellite has been enabled to use 4 reaction wheels coupled with a micro propulsion system as actuators (already implemented in the current Service Module configuration). The overall control system architecture has been thought in order to minimize the computational burden of the on-board computer but at the same time maximizing the performances in terms of pointing error accuracy. The first part of the work provides a realistic representation of the kinematics and the dynamics of the satellite by means of a software simulator within the Matlab/Simulink development environment. This have been achieved thanks to an approximation of the sloshing motion of the propellant inside the tanks and an accurate modelization of the service module actuators and sensors. The second part of the work consists in the design of the ACS of the GNC Unit. The ACS architecture can be split in three different parts: an optimal state observer, a disturbance observer and a regulator. The design of the state estimator has been based on the well-known Multiplicative Extended Kalman Filter (MEKF) model. It has been designed starting from an idea of Markley and Crassidis (Fundamentals of Spacecraft Attitude Determination and Control, Springer, 2014) and uses an angular position error vector, a bias error vector and a scale factor error vector as state vector. The main idea is to use the angular position error vector for the quaternion correction and the other two error vectors for the angular velocity correction. Successively, the state estimation algorithm has been modified in order to correct for the measurement processing delay of the star tracker by means of a redefinition of the sensitivity matrix based on the state transition matrix. The estimated state measured by the star tracker has been propagated till the current instant of time. This operation has been accomplished by means of a multiplication of only two matrices, the sensitivity matrix and a new correction matrix. A novel idea, based on the propagation of the correction matrix, has been illustrated in order to strongly reduce the computational effort and the memory burden for the generation of the above mentioned correction matrix. As it has been reported, the norm of the error affecting the angular position is of the order of 10^(-3) degrees, providing a very good accuracy for the purpose of the mission. In cascade to this filter, a second observer has been implemented. The task of this estimator has a different nature. In particular, while the MEKF internally implements the angular position error dynamics based on the dynamics of the bias and scale factors, this second filter exploits the dynamics of the angular velocity of the spacecraft as a function of the torques given by the actuators and by external/internal disturbances. The disturbance observer has been implemented with the aim of getting an estimate of the process noise, i.e. the component of disturbance torque acting on the spacecraft. This has been possible by means of the implementation of a reduced observer model which is exploiting the information on the state coming from the MEKF and the input generated by the spacecraft attitude controller. Eventually, the controller design has been based on the output regulation theory. In particular, thanks to the observers, the regulator can make use of the full knowledge of the angular position vector, angular velocity vector, disturbance torques to be rejected and reference trajectory to be tracked. Therefore, this controller consists of a two parameter architecture: a gain K, responsible for the stability of the overall system, and a second gain L, responsible for the zeroing of the tracking steady error. The former gain matrix has been designed by means of the optimal control theory. The cost index design has been based not only on the transient response of the spacecraft but even on the minimization of the quantity of propellant used and of the torque given to the reaction wheels. The computation of the latter gain implements a state-to-output matrix as identity matrix, exploiting therefore the full knowledge of the angular position and velocity vectors bringing to a simple solution of the regulator Sylvester equations. This regulator has been proved capable of tracking not only step reference trajectories but also sinusoidal time-varying signals. The reported plots of simulated manoeuvres highlight the high performances of the algorithm, capable of fully satisfying a stringent pointing error requirement of the order of 10^(-3) degrees on the angular position. Hence, the results obtained by the output regulator showed a remarkable ability in zeroing the tracking errors not only for asymptotically stable reference trajectories but also for periodic unstable ones.

Poster 09: Development of High-Precision GPSR Applicable to GEO and GTO-to-GEO Transfer Harada R1, Kumagai S1, Kawakami S1, Kasahara M1, Yamamoto T2, Nakajima Y2 1NEC Space Technologies Ltd., 2Japan Aerospace Exploration Agency (JAXA) The Global Positioning System Receivers (GPSR) applicable to geostationary orbit (GEO) satellite have released in recent years. The high-precision GPSR can promote autonomy of satellite operation and contribute to reduction of operation costs. It can also facilitate signal acquisition of inter-satellite optical communication, and can reduce the time required for it. The GPSR application to all electric propulsion satellites has also noticed. Frequent orbit determination operations using Range and Range Rate (R&RR) measurements during electric orbit raising from GTO to GEO require high operation costs. The high- precision GPSR can eliminate the need for the operations and dedicated ground stations for the purpose.

In order to achieve high-precise GPS navigation at GEO, it is necessary to solve the following problems. Because received GPS signal power at GEO is lower about 20 dB than that at low earth orbit (LEO), GPSR for GEO is required higher sensitivity to the received power than GPSR for GEO. Furthermore, dilution of precision (DOP) of GPSR for GEO is worse than that of GPSR for LEO because GPSs exist only in the earth neighborhood from the angle of GEO satellites. On the other hand, Doppler frequency and received power dynamically vary depending on the orbit stages of GTO-to-GEO transfer. Furthermore, GPS navigation has outage time because the number of visible GPS satellites are not enough for GPS navigation depending on orbit and attitude conditions of the user satellites at GTO-to-GEO transfer.

This paper shows the development of the high-precision GPSR (hereinafter called GEO GPSR) applicable to the satellites at GEO and even for GTO-to-GEO transfer. The development of GEO GPSR is based on the on- orbit heritages of GPSR for LEO (hereinafter called LEO GPSR) which NEC Space Technologies, Ltd. has developed in 2013. A functional diagram of the developed GEO GPSR is shown in Image 1. GEO GPSR consists of a L1 antenna (GPSA) for GEO (hereinafter called GEO GPSA), two L1 antennas for GTO-to- GEO transfer (hereinafter called GTO GPSA), three low noise amplifiers (hereinafter called GPSL), and a processor unit (hereinafter called GPSP). The GPSP processes GPS L1 coarse acquisition (C/A) signals and outputs navigation results which fulfill the precision shown in Image 2. The detailed problems and the countermeasures of the satellites at GEO and GTO-to-GEO transfer in developing the GEO GPSR are described in order below.

At GEO, received power before the GPSL was required above -153 dBm, which was obtained by the visibility analysis of GPSR at GEO, to get enough GPSs for executing GPS navigation. In order to improve the receiver sensitivity at GEO, The GEO GPSA was designed as a patch array antenna which had narrow beam width. The integration time of correlation power at the GPSP signal processing was also adjusted according to the evaluated received power. Furthermore, a tracking algorism using frequency-locked loop (FLL) and Kalman Filter was introduced. In order to improve the bit error rate (BER) of decoded Ephemeris messages, majority decision was performed between the repeated Ephemeris messages. By improving the receiver sensitivity, self-correlation and cross-correlation power of received signals may cause false locks against wrong code-phases and wrong GPS satellites. In order to avoid the false lock of wrong code-phases, signal acquisition was designed to record the self-correlation power of all code-phases temporally, and to detect the peak power as a correct signal in the all self-correlation power. In order to avoid the false lock of wrong GPS satellites, signal tracking was designed to monitor the self-correlation power of tracking signals periodically, and to unlock the false locked signals if those power were incorrect.

At GTO-to-GEO transfer, wide dynamic range of Doppler frequency and received power are problems at GTO-to-GEO transfer in addition to the problems at GEO. For applying on GTO-to-GEO transfer, the GTO GPSA was designed as patch antenna which had wide beam width as well as the antennas for the LEO GPSR. The integration time of correlation power and Doppler range were adjusted according to the evaluated received power and velocity of the user satellite. In the case of GTO-to-GEO transfer, acquisition order of GPS satellites and selection algorism of user antennas are also important because of orbit and attitude varieties. In order to improve the visibility, GPS satellites which faced to the user satellite was selected preferentially. Appropriate user antenna for each GPS signal was selected by the vector and coverage information of each user antenna. As well as the LEO GPSR, navigation algorism using the Kalman Filter with the pseudo-range and carrier-phase was adopted. However, navigation outage time exists at GTO-to- GEO transfer when the orbit and attitude of the satellites is unsuitable for GPSA. Therefore, high-precision prediction models for the Kalman Filter were introduced, for example of earth gravity model EGM96 to degree 20, solar and lunar perturbation models, oven controlled oscillator (OCXO) of long term stability.

According to the GEO GPSR design described above, GPS navigation simulation was performed assuming the conditions of GEO and GTO-to-GEO transfer. In the case of GTO-to-GEO transfer, typical orbit examples were selected. In the simulation, positions of GPS satellites were varied to evaluate the statistical performance of the designed GEO GPSR. Simulation results confirmed that the designed GEO GPSR met the performance shown in Image 2 at GEO and GTO-to-GEO transfer. Finally, proto-type model (PM) of the GEO GPSR was produced and put in the qualification test (QT). Product pictures are indicated on Image 3. The QT results confirmed that the developed GEO GPSR achieved its designed function and performance at the flight conditions of GEO and GTO-to-GEO transfer.

The new GEO GPSR was developed applicable to GEO and GTO satellites. The functions and performance of the developed GEO GPSR for GEO satellites will be evaluated on the Japanese Data Relay System satellite (JDRS). The full functions and performance of the GEO GPSR will be put to practical use at GEO and GTO-to-GEO transfer on the JAXA Engineering Test Satellite-9 (ETS-9). Poster 10: P4COM: ESA Pointing Error Engineering For Telecommunication Missions Hirth M1, Ott T2, Girouart B3, Deslaef N3, Vogel T4, Closs M5, Campuzano San Miguel J6, Guercio N7, Mosson N8, Neumann N9 1Astos Solutions GmbH, 2Airbus Defence and Space, 3ESA/ESTEC, 4ESA/ECSAT, 5Airbus Defence and Space, 6Airbus Defence and Space, 7Thales Alenia Space, 8Thales Alenia Space, 9OHB Systems AG In 2011 ESA published the ESA Pointing Error Engineering Handbook. In the following years, the Pointing Error Engineering Tool (PEET) has been prototyped and further developed under ESA contracts jointly with the European space industry to support users in applying the elements in the handbook and in the ECSS control performance standard. The handbook – and with it also the PEET software - has been used in several ESA space mission studies and projects in the last years and became a well-known and broadly accepted reference in the European space community. So far, the application was mainly focused to the field of Earth Observation and Science mission where high-accuracy pointing is generally crucial. A key benefit of the tool in this respect is its capability to assess the budgets with advanced statistical methods, i.e. probability density estimation, frequency domain characteristics and cross-correlation information of error sources. No commercial tool is available to assess error budgets with such advanced methods. Moreover, “classical” budget assessments via spreadsheet processing do not and cannot implement such methods sufficiently accurate. In recent years, interest in a standardized pointing error engineering approach has also risen in the telecommunication sector. First, such an approach is expected to improve and simplify the design and development process. Second, also in this field, specific missions have to cope with increasingly stringent pointing requirements – e.g. for hosted payload concepts or communication via inter-satellite links. As response, ESA initiated a development study in the Advanced Research in Telecommunications Systems ARTES. This study – with all European telecommunication primes directly involved as consultants - is called Pointing Error Engineering for Telecommunication Missions, P4COM, and started in 2018 with foreseen end in mid-2020. This paper presents the following results of the study: • features of the new PEET software release, • challenges and conclusions of applying the ESA Handbook analysis methods to the telecommunication reference case missions, • proposed evolutions for the ESA Pointing Error Engineering Handbook based on the lessons learned over the last years and the special needs of telecommunication missions. While the last bullet is covered in detail in a separate paper, this article focuses on the PEET software aspects. It further presents and discusses its application to telecommunication mission study cases. The first part of the paper provides an overview of the updated PEET software and its new features. The identification of the actual needs was based on results of a survey conducted among the existing PEET user community at the beginning of the study to ensure the development of an industrial reliable tool in terms of stability, user-friendliness, modelling and reporting functionalities. Further - as the application of the ESA handbook and tool was new in this sector – the tool was complemented with focus on the feedback received from the European telecommunication primes, especially with respect to specific analysis features. The second part of the paper introduces an overview of the setup and analysis of the telecommunication study cases in P4COM, namely SmallGEO (OHB Systems), E3000 Broadcast Mission (Airbus), SPACEBUS NEO (Thales Alenia Space) and EDRS Global (Airbus). This selection of study cases aims to reflect missions with a high interest to ESA and industry in terms of pointing requirements, pointing challenges and pointing error engineering process as well as to cover the specific interests of all involved primes. The definition, setup and analysis of these reference cases was carried out in close iterative co-engineering between core study team and telecommunication consultants with the ultimate goal to consolidate PEET for the application in this field. Finally, the lessons learned and benefits of applying the ESA handbook methodology and the PEET software to telecommunication missions are discussed from the perspective of the industry consultants involved in the study.

Poster 11: Qualification Results of Honeywell’s 3-Axis HG4934 Space Rate Sensor Horkheimer D1, Shick J2 1Honeywell Aerospace, 2Honeywell Aerospace Honeywell unveiled the design of the HG4934SRS, a Micro Electro-Mechanical Systems (MEMS) based Inertial Reference Unit (IRU) in 2019. The HG4934SRS was developed to address the needs of the growing small satellite market with its advantageous capabilities in terms of Size, Weight, Power and Cost (SWAP- C). The Space Rate Sensor development effort maintained Honeywell’s market-leading MEMS gyro performance while also establishing a design robust to the launch environment and natural space radiation effects while incorporating improvements in materials, processes, and part selection.

This paper presents details on the product’s 6-year 1100 km Polar Low Earth Orbit (LEO) reference design mission, its radiation environment, and how it influenced test strategy, part selection and qualification. The reference mission’s 1100 km altitude exposes the product to significant fluxes of trapped protons and electrons in the Van Allen Belts, and when traversing over the poles the product encounters solar particles and galactic cosmic rays including high-energy heavy ions. The HG4934SRS product also seeks to support a wider variety of mission applications beyond LEO.

The HG4934SRS reference mission is not a typical low-altitude low-inclination LEO mission with minor radiation exposure. It requires a more comprehensive approach to space Radiation Hardness Assurance (RHA) than the minimalist radiation qualification methods often applied for low-cost LEO applications operating in mild radiation environments. One method historically proposed uses exposures to apply Total Ionizing Dose (TID), Total Non-Ionizing Dose (TNID), and Single Event Effects (SEE) simultaneously in a single test. Recent literature and Honeywell’s own experience during development of the HG4934SRS show that a proton only test approach, with protons low effective Linear Energy Transfer (LET) can fail to reveal a component’s Destructive Single Event Effects (DSEE) sensitivity and thereby a satellite mission to significant residual risk.

Another method takes a commercial product with no radiation hardening applied, performs some radiation test exposures, then simply sets the product’s radiation rating at the capability level observed in test. While these methods may be appropriate in certain applications, for long-life satellite missions and more severe radiation environments, both methods tend to fail to address a variety of radiation effects concerns, including heavy-ion Single Event Latchup (SEL) and other destructive SEE, Enhanced Low Dose Rate Sensitivity (ELDRS), and radiation response variability from lot to lot, from part to part, and from fabrication process drift over time.

The RHA approach developed for the HG4934SRS attempts to address TID, TNID, SEE, and spacecraft charging effects in accordance with standards, guidelines, and processes developed for traditional space applications. However, Honeywell has adapted those standards and guidelines for small satellite and low-cost space applications in a variety of ways, as described in this paper. A creative approach to hardness assurance allows the HG4934SRS to meet the intent of many demanding radiation requirements specifications while maintaining the cost advantage of Commercial Off the Shelf (COTS) or automotive-grade components and maintains alignment with a high-volume production factory.

Results of system level TID and Single Even Effects SEE analysis and qualification are presented. As part of the development program a detailed NOVICE radiation shielding model of the HG4934SRS was created to facilitate system level TID analysis and to allow Honeywell to generate accurate mission dose predictions for different customer orbits and satellite installation configurations as a service to customers. TID predictions for a GEO mission with different shielding configurations are shared as an example benefit of the NOVICE model.

The HG4934SRS’s Environmental qualification testing is now complete and the high-level test flow and test results are presented for the design. Testing included both operational and non-operational mechanical shock, random vibration, thermal shock, and Thermal Vacuum testing in addition to other tests. Testing confirmed the structural and thermal analysis predictions of the design. Testing also established that the HG4934SRS design is also insensitive to magnetic fields and exposure to a Helium environment commonly associated with some launch vehicles. Electromagnetic Interference (EMI) qualification testing was also performed per tailored MIL-STD-461F conducted and radiated environments. With every Flight unit going through a thorough suite of acceptance tests the overview of the test flow and summary test results are highlighted.

Lastly, the results of system level Reliability analysis for the HG4934SRS are presented. The HG4934SRS Probability of Success against the reference mission were assessed using MIL-HDBK-217F Notice 2/VITA51.1 methodology. The analysis includes the effects of typical component stresses (e.g. voltage, temperature, etc.) and in addition where appropriate including the effects of radiation in unit level Probability of Success (Ps) calculations. Effort also generated a detailed component level Failure Modes, Effects and Criticality Analysis (FMECA) as well as a MIL-STD-822 Safety Analysis.

With the initial Production build and delivery of early Flight units to customers now complete and the next build scheduled the paper concludes with a development road map for future upgrades and improvements to the HG4934SRS. Future development includes a 28V MEMS Inertial Reference Unit (MIRU) to offer compatibility with a common satellite primary bus voltage.

Poster 12: REVLANGNC flight demonstrator: Development and maturation of GNC technologies for Descent and Landing Garcia L2, Cojocaru A1, Nechita S1, Grozea I1, De Zaiacomo G2, Rosa P3, Huertas I4, Yabar Valles C4, Martinez Barrio A4, Ionita A1 1Deimos Space S.R.L., 2DEIMOS Space S.L.U, 3DEIMOS Engenharia S.A, 4ESA – ESTEC Nowadays there is great interest for reducing the cost of space transportation systems, which led to a considerable focus on developing different technologies that will allow reusability. So far, the reusability concept has manifested in two solutions: reusable re-entry vehicles and recovery of system components, such as the launch vehicle’s fairing, or stages/boosters. Autonomously operating parafoil systems are considered as suitable means that can precisely and safely descend and land different spacecraft or spacecraft components to a wide range of on ground areas. Moreover, a guided parafoil system can perform an autonomous descent and landing either to land on ground with high accuracy or to perform a Mid-Air Retrieval (MAR) with high accuracy. In this context, DEIMOS Space has the objective of further developing and maturing GNC technologies for Descent and Landing (D&L) that could be applied to future European space applications. Building on REVLANSYS, a previous DEIMOS-ESA activity, in the REVLANGNC project the design of innovative GNC solutions for parafoils is consolidated taking into account reference scenarios fully representative of future space applications. In REVLANGNC, the G&N solution preliminary identified in REVLANSYS shall be revisited and consolidated with respect to consolidated requirements, and the GNC solution shall be completed with the implementation of the Control function, targeting a fully autonomous and robust parafoil GNC solution. In order to cover the different landing solutions, two strategies are taken into account: - Pinpoint landing, where the GNC shall steer the parafoil toward a predefined, fixed, landing site, and high accuracy is required for the Guidance. - Mid-air retrieval, where the GNC shall steer the parafoil toward a predefined location and maintain a given heading in order to allow a recovery vehicle to approach and perform the recovery manoeuvre. A more relaxed accuracy requirement can be accepted. The Descent System is composed of a 2-stage system: Pilot/Subsonic Drogue and Parafoil. The system can be deployed in a feasible combination of altitude and velocity, which defines the energy that the vehicle possesses at the end of the TAEM phase. The Descent System will allow the vehicle to manage the last part of the re-entry mission until landing. The Pilot + Drogue will slow down the vehicle to a speed compatible with the opening of the Parafoil. The Parafoil shall provide control authority in the last phases of the descent, allowing the system to achieve a soft and precise landing. The proposed GNC solution will cope with the challenges imposed by autonomous parafoils. Unlike other powered vehicles, parafoils do not have the capability to ascend. This means that only one attempt can be made at landing. Furthermore, the GNC must be able to compensate for dispersions at the start of the descent, but also for any disturbances encountered during the descent. The parafoil guidance problem consists of generating a trajectory from a given initial configuration in terms of position and heading at an altitude h_0 to a given terminal configuration, considering the final position, heading, and altitude. The parafoil Guidance will be composed of a Trajectory Generator, responsible for generating the reference trajectory profile, and a Trajectory Tracker, which tracks the reference trajectory and computes the required control commands. Moreover, the parafoil system flies at velocities that are often comparable to the wind velocity, thus making the system very susceptible to air influences. Online wind estimation is as a key improvement in order to achieve the required autonomy of the system during descent and landing. The control function considers feedback compensation, therefore allowing the attitude of the parafoil to be controlled. In case a pinpoint landing on a prepared terrain has to be performed, the guidance will change to a landing flare mode once the altitude decreases below a certain threshold. In this phase, the GNC focuses on reducing as much as possible the position error with respect to the target point, and also performs a heading alignment such that the landing is executed headwind in order to reduce the horizontal velocity. Moreover, a final flare manoeuvre shall be commanded in order to perform the touchdown respecting strictly the vertical velocity requirements. This paper will discuss the developments and innovations implemented for the descent and landing phase of a re-entry vehicle and MAR capture of a launch vehicle’s fairing implemented in the framework of the REVLANGNC project. The parafoil GNC solutions designed will be described in detail, and the results of the GNC verification carried out for the GNC design review will be reported to demonstrate their applicability to the considered space application scenarios. These results will be based on a GNC performance assessment carried out in MIL making use of a Functional Engineering Simulator. Particular attention will be given to key aspects related to the parafoil GNC design and verification, such as the modelling of the parafoil, ensuring the dynamics for control and for simulation are properly grasped, by using LFT and nonlinear models, respectively. Also, the Guidance problem will be discussed in detail, focussing in particular on the performance improvements provided by advanced algorithms able to optimise online the trajectory to be flown, in comparison to classic methods that have to follow a pre-defined trajectory with limited update capabilities (e.g. X-38 like). In addition, different wind estimation techniques will be traded off based on several criteria, in order to improve the accuracy of the Navigation. In terms of Control design, H/mixed- approaches have been adopted in REVLANGNC, and the preliminary results obtained will be provided in this paper.

Poster 13: Autonomous rendezvous GNC system development - overview and recent advances Klionovska K1, Rems F1, Risse E1, Burri M1 1German Aerospace Center (DLR) Spaceflight and space technologies have a high impact on our daily life (e.g. communication, navigation, Earth observation). Since the first satellite in 1957, several thousands of satellites have been launched into orbit. Facing growing population of operational spacecraft and objects, technologies for space debris removal and lifetime extension of satellites have to be developed. It is important to remove large retired satellites to protect operative satellites in nearby orbits, to avoid collision and additional, new debris. Further, the lifetime of satellites should be increased, such that satellites need not be replaced as early as nowadays. The rendezvous phase, i.e. the approach of a target spacecraft by a chaser spacecraft, is an important but also challenging phase of an on-orbit servicing mission. If there is no permanent contact between control center and servicing satellite like in Low Earth orbits, the rendezvous has to be done on-board in an autonomous way. The paper describes the recent GNC system developments at the German Aerospace Center (DLR) in the field of autonomous rendezvous and visualizes the main components and tasks which are needed to prepare the rendezvous phase of an on-orbit servicing mission: (i) rendezvous sensors, (ii) on-board GNC system, (iii) advanced on-board computers, (iv) ground segment and (v) simulation and test. Sensors for rendezvous such as cameras or LiDARs provide the input for the GNC system. In the worst case, the target spacecraft is completely passive and has no operative attitude and orbit control system. Since technologies like (relative) GNSS cannot be used in this case, optical sensors can be applied for relative navigation during the approach to the target. Different sensor types have different strengths and weaknesses: For example, with a single monocular camera the distance can only be measured if an accurate 3D model of the target is available. Dependent on the resolution and the field of view, distance measurements can be very in-accurate and noisy. LiDARs and other active sensors like a photonic mixing device camera (PMD) directly deliver good distance measurements; however other components of the pose like translational position components could be measured with 2D cameras often more accurately. Thus a combination of sensors could be used to obtain a robust and stable navigation result. The on-board GNC system lets the chaser approach the target spacecraft in a safe, controlled way using the rendezvous sensor data as input. The guidance sub-system generates trajectories for straight-line approach, fly-around, hold-points and attitude changes. The system receives tele-commands from ground and generates guidance trajectories based on the current position and attitude and the commanded guidance values. The navigation system determines an estimation of the current state of the target spacecraft. It consists of a navigation filter and sensor-dependent algorithms for pose estimation (based on 2D image processing or 3D data processing). The navigation system is also commandable, for example the different sensors can be activated, the pose estimation can be started and stopped and it can be selected via tele-command which sensors should be used by the navigation filter. The filter is an Extended Kalman filter, which can also handle delayed measurements and can perform sensor fusion. Sensor data is fused in the filter correction step. Different gaining weights depending on the sensor type can be used to weight the single components of the pose measurements individually. The control sub-system compares the guidance trajectory for the servicer with the actual pose and computes force commands for the actuators such that the service satellite moves as desired. During the GNC system developments, it is important to keep in mind that the final software has to be executed on-board of a satellite. Therefore, at all development stages the GNC system has to be tested also on real on-board computer hardware. On the one hand, on-board computers have to be improved such that their performance is sufficiently high to allow executing applications with image processing elements. On the other hand, the algorithms chosen for pose estimation have to be optimized for and adapted to on-board computer software frameworks. The counterpart to the on-board GNC system is the ground-system with its communication system, mission operations system and special elements for on-orbit servicing. Telemetry of the service spacecraft is received on ground during contact phases with the satellite. Tele-commands are sent by an operator at the rendezvous console. Also so-called "science telemetry" like compressed camera images are sent to ground during contact phases for visual monitoring and checks. Last but not least, the complete end-to-end chain, from operations in the control room at the German Space Operations Center (GSOC) to the rendezvous sensor payloads has to be continuously tested and verified. This is done by using DLR’s end-to-end simulation framework which has been established in the last years and is continuously improved and extended. The rendezvous sensor payload, the on-board computer and an exemplary target satellite mockup are integrated at the EPOS 2.0 (European Proximity Operation Simulator) 2.0 facility. The satellites’ orbit and attitude dynamics are simulated in software, as well as the communication path between space and ground. At ground, the entire operations infrastructure has been established, such that simulations and tests can be done in a very realistic environment. Especially the interactions between space and ground, realistic contact durations, as well as the performance of the autonomous on-board system can be tested and proved.

Poster 14: BATCH ORBIT DETERMINATION COMPARISON FOR GEOSTATİONARY SATELLITES Koker A1, Aydın S2, Dag E3, Yılmaz Ü4, Şakacı C5, Tekinalp O6 1TUBITAK UZAY, 2TUBITAK UZAY, 3TUBITAK UZAY, 4TURKSAT, 5TURKSAT, 6Middle East Technical University The orbit determination process is one of the vital areas for orbit control of geostationary satellites. Efficient station-keeping maneuver planning with lowest fuel consumption, accurate orbit determination is critical. There are two approaches for orbit determination: determination from batch data and sequential orbit determination. The batch approach is generally used for off-line activities and it uses the whole measurement set at once in order to estimate initial conditions. The sequential determination is generally used for on-line application when orbit information is required continuously. Standard orbit determination procedure requires initial orbit information, measurements and dynamic and static parameters, such as satellite mass, surface area, solar radiation pressure coefficient etc. During actual operations, initial orbit information is calculated by propagating previous orbit determination results. Raw ranging data are generated by measuring time delay of signal from the station to the spacecraft (up-link) and back (down-link) using the on-board transponder. Raw pointing data can be obtained from down-link auto track ground antenna with narrow beam width. Pre-processing algorithm rejects undesired data from raw data by comparing it with the standard deviation of total data. Pre-processed measurements and initial conditions are used to initialize nonlinear batch filter for orbit determination. In nominal operations, 48 hour measurements with a frequency of once per hour are used to estimate initial position, velocity, solar radiation coefficients, and measurement bias, provided that a maneuver is not executed. When an E/W (East/West) or N/S (North/South) maneuver is carried out , 96-hour measurement sets, with once per hour measurement data are used to estimate similar parameters, as before with the addition of maneuver velocity increments. These 96 hour duration measurements are taken from ground station antennas for two days before and two days after the maneuver. In this work, batch orbit determination of GEO satellites using of the real ground station measurements (azimuth, elevation and range) is presented. Various factors such as different type of measurement configurations, incorporating measurements from multiple ground stations, effect of orbital maneuvers are investigated. For this purpose the developed software is compared with the reference software. The final manuscript will present the orbit determination algorithm. A parametric study will also be presented to show the effects of various parameters, such as data collection frequency, the data amount, maneuvers, dynamic parameters, etc. on the orbit determination accuracy. Orbit determination algorithm is developed for orbit control software used at TUBITAK UZAY ground station. Orbit control software is named as “UYK”. Estimated results from UYK software are also compared with reference orbit determination software used at TURKSAT ground station during real station keeping operations. TURKSAT is one of the world's leading companies providing all sorts of satellite communications through the satellites of TURKSAT as well as the other satellites. TURKSAT Company is located at Ankara, Turkey. All measurements used in orbit determination analysis are taken from ground stations antennas used for real-time T4A satellite operations at TURKSAT.

Poster 15: Uncertainty propagation in a Guidance, Navigation & Control system for landing operations using Differential Algebra Konstantinidis K1, Izzo D1 1Advanced Concepts Team, ESA-ESTEC Future landing and close-approach missions (e.g. asteroid landing for mining, space debris removal, pinpoint planetary landing) will require increasingly complex Guidance, Navigation & Control (GN&C) systems. These missions share the need for certain high-level GN&C functions, namely navigation, hazard detection and avoidance, and guidance. Due to signal delay and limited knowledge of the operation environments, these functions must be performed with a high degree of autonomy.

It is important that the design and analysis of these non-linear and complex GN&C systems can be performed accurately and efficiently. Uncertainty propagation in models of GN&C systems has been traditionally performed either by linearizing them at the cost of accuracy or by performing accurate but computationally expensive Monte Carlo simulations.

A compromise between the two is offered by Differential Algebra (DA), a numerical technique to automatically compute high order Taylor expansions of a given function. In the DA framework, the Taylor expansion of a system model function is calculated once at relatively high computational expense. This expansion however can then be used to evaluate any input value around the nominal at negligible computational cost. DA thus offers the possibility to reduce the computational effort significantly while maintaining adequate accuracy.

In this paper, we describe the application of the DA framework for non-linear uncertainty propagation in a GN&C model for landing operations. We then investigate the computational efficiency and accuracy of this novel approach compared to the standard one.

A demanding asteroid-landing scenario is used as reference, where a lander must land accurately and safely on the surface of an asteroid, near one of a set of preselected landing sites. The position of the preselected landing sites is considered imperfectly known, as is that of the landing hazards at the general landing area.

Three high-level functions are necessary: a navigation function must navigate the spacecraft relative to the landing sites and the local terrain, as well as track interesting landmarks. A hazard detection and avoidance (HDA) function must sense the local terrain for hazards and command a retargeting to a new landing site if the original is deemed unsafe. A guidance function must calculate a feasible fuel optimal trajectory to the chosen landing site.

We created a nominal model of this GN&C system based on the usual floating point algebra.

For navigation, an Extended Kalman Filter-based Simultaneous Localization and Mapping (EKF-SLAM) approach is used. The navigational state of the lander is updated and corrected along with that of observed landmarks, based on interoceptive (e.g. IMU) and exteroceptive (e.g. camera, lidar) sensor measurements.

For HDA, hazard information (terrain slope and roughness, obstacles) by different sensors is merged to a single hazard map, which is in turn merged with other pertinent information, such as the scientific interest of the candidate landing sites, their reachability etc. A final landing score map is produced from which the best new landing site is chosen. Various methods have been used in the past for this fusion under uncertainty into a single landing goodness map, such as fuzzy, probabilistic, and evidential reasoning. These methods are investigated and one is selected for implementation.

For guidance, we use an adapted Sims-Flanagan direct method where the trajectory is broken down to segments of constant thrust. The constraints of the problem are the end state (landing target point), the thrust magnitude, and the maximum flight time. The optimization controls are the thrust per segment and the total time of flight. The problem is then formulated as NLP and is solved in SNOPT.

As previously mentioned, DA operates in the algebra of truncated Taylor polynomials. For a given function to be compatible with DA, all mathematical operations within it must be able to be performed in this algebra, which is not trivial. Accordingly, some work must be done to modify the three main GN&C functions mentioned above in order for them to be compatible with DA. The necessary modifications are investigated and applied. For guidance, an Artificial Neural Network (ANN) is trained, in order to map any initial state around the nominal to a fixed final state. For HDA, a DA-based information fusion method is investigated and implemented. For navigation, a DA-based implementation of EKF-SLAM is investigated and possibly implemented.

Monte Carlo (MC) simulations using the nominal model and the DA-based model are performed for the reference asteroid landing scenario described above. Landing accuracy and reliability are calculated with both models. Results from the DA-based model are compared to those of the nominal one for increase in computational efficiency and loss of computational accuracy, for a varying number of MC iterations and for varying orders of expansion for the Taylor polynomial of a given system model function. It is thus investigated at which level of required simulation accuracy the DA-based model becomes more efficient than the standard modelling approach. In a wider sense, the results of this work investigate the feasibility and the efficiency of this new and promising DA-based approach for the design and analysis of future complex landing GN&C systems.

Poster 16: Image simulation for space applications with the SurRender software Lebreton J1, Brochard R1, Baudry M1, Berjaoui A1, Jonniaux G1, Kanani K1, Masson A1, Panicucci P1, Robin C1 1Airbus Defence And Space Summary: Vision-Based Navigation (VBN) solutions require reliable image simulation tools. In this paper we provide a comprehensive explanation of why traditional rendering engines may present limitations that are potentially critical for space applications. We introduce Airbus SurRender software v6.1 and illustrate how it uniquely models specific physical effects that are crucial for image processing algorithms. We show how our simulation capabilities are at the heart of the development processes of our computer vision solutions.

The simulation of space scenes presents specific challenges, which are typically not handled by general purpose image simulators. Vision-based navigation solutions require training and validation datasets that are as close as possible to real images. Our team and partners develop computer vision algorithms for space exploration (Mars, Jupiter, , the Moon), and for in-orbit operations (rendezvous, robotic arms, space debris removal). Of course “real images” are rarely available before the mission. Ground-based test facilities such as robotic test benches embarking mock-ups or experiences with scaled mission analogues (mars terrain analogue, drones flights, etc.) are useful, yet they are limited. For example it is very difficult to capture the scale of space scenes in a room-sized facility (such as a small objects illuminated by an extended light source). Also limited numbers of images are available from previous missions or from lab experiments, when thousands are needed to represent the variety of possible configurations that the algorithms will encounter. Another decisive of simulations is that their ground-truth is perfectly known, whereas real-life experiments are prone to errors and biases, which are hard to estimate or lack accuracy. Some of the effects visible in space images are not of particular importance for traditional image simulators. For example, for far-range rendezvous, very low SNR targets (SNR ~ 1) must be simulated with high radiometric fidelity. Space-qualified cameras often have unusual optical distortions and achromatism and the geometrical performance relies on properly modelling them. The Point Spread Function and associated effects (resolution, blooming) are fundamental parameters for image quality and they need to be simulated physically. Defocus is often encountered and shall be well simulated. Although 3D OpenGL-based renderers can implement some physical effects, it is well-known that the raytracing technique is needed for the rendering of perfectly reliable images in physical units. With the increase of computing performance, general purpose rendering engines are starting to implement raytracing techniques. However they are not optimized for space applications: the sparsity of scenes and large scale ranges call for specialized methods for the rays to target the scene objects, in order to bind the required computing resources. SurRender intrinsic use of PSF models for ray sampling (as opposed to post-processing) guarantees radiometric and geometric accuracy at the subpixel level. SurRender also has specialized routines for the rendering of stars which is systematically overlooked by other engines. Furthermore in some cases it is necessary to render each independent pixel of pixel line sequentially in which case the simulation needs to be done in the time domain. For example when implementing rolling shutter, push-broom or LiDAR sensors, one must take into account the photons optical path and the camera relative motion during target acquisition. Finally it is essential for a simulator to be flexible in the sense that software architecture allows the input of new models painlessly. In this paper, we provide a quick listing of available rendering engines and discuss their limitations for space computer vision applications. We show how Airbus SurRender software, already introduced by Brochard et al. 2018 (IAC-18,A3,2A,x43828), attempts to go beyond these limitations. Specifically we focus on illustrating key physical effects that SurRender naturally simulates and that could easily have been missed by other engines (secondary illumination, small targets, achromatism, BRDF, etc.). We highlight our past and ongoing efforts for a formal validation of the software geometrical and radiometrical performances. We introduce the latest features available in SurRender v6.1 (LiDAR, variable PSF, distributed computing, open- source interface). Last we present new use cases of SurRender for machine learning applications - image augmentation for dataset synthesis - that we are currently evaluating for Earth observation applications. Currently the SurRender software is at the heart of the development process for many image processing solutions at Airbus mostly in VBN. We use it from early prototyping, to extensive performance test campaigns and hardware-in-the loop experiments. Various API (Python, Matlab, Simulink, C++, etc.) are available to interface SurRender with different simulations environments (GNC environment simulator, optical stimulators, etc.). In conclusion, in the context of a growing need for autonomy and artificial intelligence in space applications, our team is pursuing a constant effort in the development of the SurRender software and its diffusion.

Poster 17: Real-time thermal infrared image generation for testing future vision-guided autonomous navigation systems Martin I1, Dunstan M1, Sanchez Gestido M2 1University Of Dundee, 2ESTEC, ESA Vision systems are becoming well-established in mission studies to aid to spacecraft navigation and hazard detection to support safe and precise landings [1]. Future missions may also use thermal cameras for navigation and guidance, sensing radiation in the infrared, to obtain different or additional information than vision-based systems. There are existing tools that generate realistic thermal simulations on relatively low- resolution models but not in real-time for large, high-resolution terrain models [2]. To test and validate thermal image processing algorithms being developed for future navigation and guidance, a system to generate realistic thermal images in real-time for closed-loop simulations is proposed. Extensive thermal data of the Moon has been provided by the Lunar Reconnaissance Orbiter (LRO) Diviner instrument [3] and thermal models for the lunar surface have been derived from that data [4][5]. This abstract summarizes the adaptation and extension of the thermal models developed from Diviner data to implement a thermal renderer running on a high-speed Graphical Processing Unit (GPU) to generate representative thermal images of asteroids and the lunar surface in real-time. The results will be evaluated by comparing the simulated images with equivalent real thermal image data.

Generating simulated images to test navigation instruments at high speed with model resolution ranging from kilometres to centimetres is a challenging task due the large model sizes, the image resolution range and the requirement to support closed-loop testing in real or near-real time (i.e. 10 Hz). Generating simulated thermal images presents additional problems due to the more complex physical process of heat absorption and emission than for reflected visible light. The approach taken in this study is to develop a thermal renderer based on Diviner data [3] and models derived from that data [4][5], then incorporate it into the established Planet and Asteroid Natural scene Generation Utility (PANGU) tool which generates simulated visual and LiDAR images of planetary surfaces and small bodies. PANGU incorporates a sophisticated surface modelling system [6] with a high-speed, GPU-based visual renderer which includes a physics-based camera model [7].

The NASA Diviner Lunar Radiometer Experiment (DLRE) from the Lunar Reconnaissance Orbiter has produced a large about of thermal data of the lunar surface [3]. Vasavada et. al used Diviner data to estimate the thermal properties of the equatorial regolith of the Moon and devised a model to estimate surface temperature [4]. They used a one-dimensional thermal model to predict temperatures with a range of parameters including albedo, emissivity, latitude, heat flux, bulk density, heat capacity and thermal conductivity. Their model represents the regolith as two layers with a highly insulating low-density upper layer abruptly increasing in density to replicate the sharp drop in temperature following sunset followed by a slower cooling throughout the night. The temperature measurements of the Moon during a day-night period follow a distinctive pattern for the non-polar regions as a function of time of day in hours from noon with seasonal variations. The maximum temperature is at local noon falling off towards sunset. During the night, temperatures fall almost linearly with a smooth transition at sunset and sunrise. The temperature rises rapidly during the morning until the next noon.

To implement this model in PANGU, the diurnal and seasonal temperature responses of the Moon or target body are encoded into Look-up Tables (LUTs) which empirically determine the surface temperature at each pixel based on the local slope, illumination conditions, and material properties such as emissivity and thermal inertia. The physics-based camera model then computes the emitted radiance in each detector channel based on the spectral emissivity and temperature of the surface. This is added to the reflected radiance in each channel, defined by a Bidirectional Reflectance Distribution Function (BRDF). Regolith and rock have different thermal inertia, so in shadowed areas, rocks will be brighter in the infrared than the surrounding soil. Differences in emissivity and thermal inertia mean that a target landing site may have similar boulders in visual images but distinctive boulder types in the infrared.

The thermal behaviour of small bodies is less well known at present although the -2 and OSIRIS- REx missions will provide additional processed thermal images in the next few years. However, since asteroids are mostly rocky airless bodies their temperatures are expected to follow a similar day/night profile as the Moon. Their faster rotation periods mean that the surface temperatures may not reach the same maximum values as on the Moon and the night-time temperatures are likely to be different. Small bodies are unlikely to have their own residual heat from their core, but the faster rotation means that the daily and yearly temperature variations won’t penetrate as deep into the surface as on the Moon.

The full paper will present results that include thermal material property variations and a more detailed description and justification of the approach taken. Example mission scenarios for the Moon and an asteroid will be described and evaluated through image comparisons of simulated and thermal data. PANGU was developed by the University of Dundee for ESA and is being used on many European activities aimed at producing precise, robust planetary lander and rover guidance systems.

References [1] M. Dunstan and K. Hornbostel, “Image processing chip for relative navigation for lunar landing”, in 9th International ESA Conference on Guidance, Navigation, and Control Systems (GNC 2014), 2014. [2] ESATAN, https://www.esatan-tms.com/. [3] Diviner data, https://www.diviner.ucla.edu/single-post/2018/04/20/New-Diviner-Level-4-Data-Products. [4] A. R. Vasavada et al. “Lunar equatorial surface temperatures and regolith properties from the Diviner Lunar Radiometer Experiment. Journal of Geophysical Research E: Planets”, 2012. [5] J. P. Williams et al., “The global surface temperatures of the Moon as measured by the Diviner Lunar Radiometer Experiment”, Vol 283, pp. 300–325, Feb 2017. [6] I. Martin, S. Parkes, M. Dunstan, “Modeling Planetary Surfaces with Real and Synthetic Terrain”, IEEE Trans. Aerospace and Electronic Systems, vol. 50, no. 4, 2014. [7] I. Martin, S. Parkes, M. Dunstan, M. Sanchez Gestido, G. Ortega, “Simulating planetary approach and landing to test and verify autonomous navigation and guidance systems”, ESA GNC 2017, Salzburg, May 29th–June 2nd, 2017.

Poster 18: Study of an Orbital Transition Scenario to NRHO Considered Rendezvous of Two Spacecrafts Matsumoto Y1, Yamamoto T1, Ueda S2, Ikenaga T2, Goto D2 1Japan Aerospace Exploration Agency, 2Japan Aerospace Exploration Agency At present, study that a new lunar space station in orbit around the moon as a base for manned and Mars will be built is underway around the world. Large rockets with high launch capabilities are being developed mainly in the United States, to transport spacecraft to the lunar space station. If two rockets are used and the rockets’ capabilities are used well, there is a possibility that the spacecraft can be transited to the lunar space station without using a large rocket. Image1 shows an example scenario. First, a spacecraft (hereinafter referred to as supply spacecraft) which is the main body to be transited to the moon, stays in a certain orbit around the earth with the first rocket. One month later, the second rocket is launched with an empty payload and put into a parking orbit (orbit altitude: 300 km x 6000 km, orbit inclination: 30 deg). After that, the supply spacecraft rendezvous with the second stage of the second rocket (hereinafter called the second-stage spacecraft). Furthermore, after docking, the second-stage rocket engine is re-ignited in orbit around the earth, and the supply spacecraft is shifted to the orbit where the new lunar space station stays. In the conventional problem of spacecraft rendezvous with the International Space Station (ISS), a spacecraft (chaser) is launched after the ISS (target) in accordance with the ISS orbit. In orbital rendezvous flight operation, the point is that the supply spacecraft which is a chaser is launched first aiming at the injection trajectory of the second-stage spacecraft (target) that is launched later, and then the second- stage spacecraft is launched. There are several design guidelines for transition orbits from near Earth to the Near Rectilinear Halo Orbit (NRHO), which is a candidate for orbit where the new lunar space station stays. Among them, the low-energy ballisitc lunar transfer orbit using solar tidal force is considered to be a promising orbit for unmanned vehicles for transportation because it takes a long time but can keep ΔV small. The low-energy ballisitc lunar transfer orbit from the orbit around the earth to NRHO has been studied so far, but the problem considering the rendezvous of two spacecrafts near the earth has not been studied yet. In this paper, we propose a trajectory design solution for scenario that the supply spacecraft launched by the first rocket and the second-stage spacecraft launched by the second rocket were rendezvous-docked in orbit around the earth, and then put into orbit around the moon from orbit around the earth. Furthermore, the setting of the second rocket is discussed in terms of ΔV. The following is a brief result. First, if both the first and second launch vehicles are launched as scheduled, it takes about 11m/s for the rendezvous and it takes about 32m/s to transit from a near Earth orbit to NRHO. Therefore, the total ΔV required for the supply spacecraft is about 43m/s. Next, the results for case that the launch of the second rocket is delayed are described. If the launch date of the second rocket is delayed, an additional ΔV is required mainly due to two factors. The first is necessary for rendezvous adjustment, and the second is necessary for returning to the NRHO nominal transition orbit from TLI, which is Earth escape maneuver, to NRHO. In the former case, the phase of the second-stage spacecraft that has been put into the parking orbit is different from that of the nominal, therefore the supply spacecraft must adjust the phase in order to rendezvous with the second-stage spacecraft. As the phase adjustment period becomes longer, the right ascension of the ascending node, the perigee argument and the orbit inclination change due to the perturbation caused by the Earth's gravitational field. An additional ΔV is required to adjust these orbital elements. In the latter case, after rendezvous and docking of the supply spacecraft and the second-stage spacecraft, the orbit propagates until the TLI time, but if the orbit propagates to the TLI time with the phase shifted from the nominal orbit, the maneuver cannot be performed at the original TLI time, and in the worst case, the TLI time will be delayed (or advanced) by a half period of the parking orbit. In order to correct the TLI time lag, an additional modified maneuver is performed in the middle of the low- energy ballisitc lunar transfer orbit to return to the nominal trajectory. The additional ΔV of the rendezvous and the additional ΔV for the deviation of the TLI are shown in Image 2 and 3, respectively. As can be seen from Image 2, even if the launch of the second rocket is delayed, the ΔV required for rendezvous is suppressed to about 23m/s. Considering that the nominal case requires about 11m/s, the additional required ΔV is 12m/s. Also, as can be seen from Image 3, the ΔV required to correct the TLI deviation is about 20m/s at the maximum, and considering the use of about 16m/s for the DSM maneuver in a nominal scenario, the additional required ΔV is at most 4m/s. In this study scenario, it was found that even if the launch of the second rocket was delayed, it was possible to transition to NRHO with an additional ΔV of about 16m/s on any day within a week, regardless of the launch. Therefore, depending on the design of the spacecrafts, it is possible to secure launch opportunities every day. From the above, the proposed scenario is a practical and useful design solution to the problem of transitioning a spacecraft to NRHO using two rockets.

Poster 19: Highly efficient damping system for structural resonances of flywheel masses in reaction wheels Mueller I1, Ehinger M1 1Rockwell Collins Deutschland GmbH Scope: Reaction and momentum wheels are used for stabilization and altitude control of satellites. The market demand for large reaction wheels with an angular momentum of 100Nms and more to maneuver heavy satellites for telecommunications up to 7000 kg has risen sharply in recent years. At the same time the requirements of high vibration loads are increasing, especially during launch or separation of satellite and rocket. To satisfy this demand, Rockwell Collins Deutschland GmbH has developed a heavy reaction wheel with a highly efficient damping system for structural resonance of flywheel masses.

Special requirements: Reaction wheels usually consist of a rotating flywheel mounted on a ball bearing unit, a drive motor, housing and - in case of an internal wheel drive electronic - an electronic assembly. The ball bearing technology has prevailed successfully for many years. The limiting factors of the vibration loads are the resulting forces on the ball bearings. Therefore, the bearing loads are calculated to ensure that the required bearing life-time of over 15 years will be reached. FEM analysis shows that the first axial resonance of the flywheel mass has a predominant impact to the bearing load. The damping of this resonance is characterized by a quality (Q) factor. Therefore, it is necessary to define an efficient damping concept for flywheel masses.

Constraints: The damping concept of a heritage TELDIX™ flywheel - such as a RSI (reaction wheel with integrated wheel drive electronics) 68 wheel type - relies on friction which extracts vibration energy from the system through rubbing between two bodies with relative movement. With this flywheel type a Q-factor of 15 is typically achieved in a wheel assembly. Without this damping system, the Q-factor would be significantly higher and the bearing would be overstressed. Alternatively, a much bigger bearing unit would have to be used with negative impact on important performance parameters as friction torque, size, weight and cost.

Solution: A special evaluation method for non-linear FEM simulation was used to investigate the optimized flywheel design. For evaluation of the simulations, the relative movement from damping system and flywheel mass were considered. The criteria for the best design are the expected total weight, the practicability for manufacturing, the minimization of unbalance, the relative movement between damping system and flywheel mass based on friction, the displacement of the mass rim and the acceleration amplitude of the mass rim.

During the development of larger sized flywheels, as required by market, initial FEM simulations of the expected bearing loads showed that in case of resonance of the flywheel mass an overstrain of the ball bearing occurs. In order to further use the well-proven TELDIX™ bearing unit with comparable high launch loads as for traditional Rockwell Collins TELDIX™ reaction wheels, an improved damping concept had to be developed. Here, the existing friction based damping concept has been extended by new damping concept based on vibration absorber.

Vibration absorbers are weakly damped oscillators which are tuned to the main resonance of the system. By out-of-phase oscillation the vibration absorber suppresses those parts of the vibration whose frequency coincides with the resonance frequency of the vibration dampers. According to the special, patented way the spring-mass system is added to the main structure. As a consequence, the resonance of the system breaks down into two new, neighboring peaks adjacent to the suppressed target frequency of the main structure. By friction, these two peaks are coupled to one peak and the Q-factor is reduced.

To validate the results of the FEM analysis several hardware measurements are carried out. Beginning with a plain flywheel mass without damping system, followed by a flywheel mass assembly with mounted damping system and finally the whole reaction wheel assembly are vibrated. The plain flywheel mass is axially vibrated in a hard mounted configuration. The deviation to the FEM model is less than 2 percent. Afterwards, the main eigenfrequency of the damping system itself must be tuned to the flywheel mass by FEM simulation which takes place by weight variation of the mass dampers and the rigidity of the damping system. The frequency deviation of the flywheel mass assembly with mounted damping system between simulation and test is approximately 11 percent. A Q-factor decrease of about a factor of 4 is achieved by adding the damping system. Finally, by integrating the damped flywheel mass into the reaction wheel assembly the Q-factor is additionally reduced about a factor of 1.5.

Summary: In order to avoid overstrain of the ball bearing unit in case of resonance of the flywheel mass, large and heavy flywheel masses of reaction wheels require a damping mechanism. In this abstract, an existing damping concept based on friction is extended by a second damping concept based on vibration absorbers. Due to the excellent damping performance of the new flywheel design, the loads on the ball bearing unit can be reduced in comparison to the initial design approach. This improvement makes it possible to sustain higher vibration levels, relative to a flywheel without mass damper. The new damping system allows the usage of the well- known reliable TELDIX™ bearing unit. In addition, a resizing of the ball bearing unit or an external damping system to protect the ball bearing unit is not necessary. Future flywheel masses can be extended to an angular momentum of 150Nms, without any ball bearing unit modification using the developed damping mechanism.

Poster 20: SPACE RIDER: TAEM GNC, focusing on TAEM Hybrid Navigation of the Future European Reusable Space Transportation System Recupero C1 1Deimos Space S.l.u. Abstract

This paper presents the Space Rider (SR) TAEM (Terminal Area Energy Management) GNC, focusing on TAEM Hybrid Navigation, at the closure of phase B2/C of the Space Rider ESA program. Space Rider is an ESA development program that aims to provide with an affordable, independent, reusable end-to-end integrated space transportation system for routine access and return from low orbit. It will be used to transport payloads for an array of applications, orbit altitudes and inclinations. The paper presents an overview of the TAEM GNC, especially focusing on the Hybrid Navigation solution, as developed in phase B2/C of the Space Rider program, designed to take the vehicle from TAEM Entry Point (TEP) (Mach 2.5, 30.5km) to subsonic conditions at the parachute release (Mach 0.73, 16km). The paper provides information on the TAEM GNC concept, especially focusing on the Hybrid Navigation design, definition and justification. The promising simulation results obtained demonstrate a high level of performance, as well as robustness against conservative (enlarged) model uncertainties and mission parameters dispersion, proving the effectiveness of DEIMOS Hybrid Navigation solution for a demanding flight phase such as the TAEM.

Introduction

On February, 11th, 2015, the successful flight of the Intermediate eXperimental Vehicle (IXV), demonstrated the European independent capability to return from space. IXV was a vehicle with two movable flaps for aerodynamic control that performed a suborbital mission, allowing in-flight demonstration of critical technologies for hypersonic flight conditions and successive re-entry from LEO. After being injected by in a 400km altitude orbit, the IXV performed a successful entry targeting the desired parachutes triggering conditions, to start the final descent for a safe splashdown in the Pacific Ocean. Leveraging on IXV’s development, qualification and mission success, intended as an European “intermediate” step toward multiple future space applications, the (ESA) initiated an effort to develop a sustainable reusable European space transportation system in Space Rider, integrated with the VEGA C launcher, currently under development, to enable routine launch and return space missions.

The Space Rider Re-Entry Module (SR RM) will have a multi-purpose cargo bay able to integrate a number of modular payloads to fulfill multiple mission objectives and to perform experimentation of payloads for multiple space applications. SR is designed to be an operational demonstrator able to perform 6 missions. It will have to support orbital operations in multiple orbital scenarios, from SSO to equatorial, deorbit and flight back to Earth with high maneuverability and controllability throughout all flight regimes (i.e. hypersonic, supersonic, transonic, subsonic) to perform a safe and precise soft-landing on ground under parafoil. The vehicle is therefore required to have the flexibility to ensure that environmental and operational unexpected events are mitigated and to guarantee the accomplishment of the mission objectives in compliance to stringent safety constraints in case of failure. The activity has been performed under the SR programme in phase B2/C, funded by the European Space Agency. In SR programme, Thales Alenia Space Italia (TASI) and AVIO are Prime contractors of Space Rider, with TASI as responsible for the SR Re-Entry Module (RM), SENER is the GNC subsystem responsible, and DEIMOS Space is the responsible for the Entry GNC and the TAEM GNC for SR, as part of the overall GNC subsystem activities and as a natural continuation of a role that was successfully executed in IXV.

The heritage from IXV for re-entry phase GNC is expected to be highly applicable to the Entry phase of SR RM, which is an hypersonic phase with the aim of getting the vehicle to the desired conditions (position, velocity, attitude) at the beginning of the TAEM phase. The Entry GNC will be further developed in the following phases of the programme. On the other side, the TAEM phase, goes from low hypersonic to high subsonic, incluiding the delicate transonic pass. The TAEM GNC, new wrt IXV, has the aim of steering the vehicle to the desired conditions (position, velocity, attitude) at the DS triggering. It has been undergoing a detailed design process, being important and critical, as the transonic regime was not flown by IXV, that used a supersonic drogue chute. Moreover, SR is an operational vehicle, that will support a range of mission types. Hence, SR RM presents new challenges during the re-entry phase and requires the development of European capabilities to fly the lifting vehicle aeroshape through the transonic regime to a subsonic drogue chute, which places in particular constraints and challenges on the GNC and the vehicle’s flying qualities during the lower Mach regime from Mach 1.2 to Mach 0.7. For that reason, the paper will mainly focus on the TAEM GNC activities carried on until CDR, and especially on the TAEM Hybrid Navigation developed for the phase C of the SR programme.

TAEM Navigation overview

The TAEM Navigation function is in charge of computing the navigation solution, i.e., inertial position, velocity and attitude vectors and a set of navigation derived parameters, to be used by the TAEM Guidance and TAEM Control functions to accomplish SR mission, based on the input data provided by the navigation sensors, an IMU and a GNSS receiver. The core of the TAEM Navigation function is the inertial navigation based on the IMU measurements. The computed navigation solution, consisting of position, velocity and attitude vectors, is improved by using the GNSS receiver PVT measurements when they are available. The attitude of the vehicle is inertially computed using IMU measurements and then improved through en EKF-based filter, every time GNSS receiver PVT measurements are available. With the estimate of the position and velocity vector, it is possible to derive a set of parameters needed by the GNC Flight Management and TAEM Guidance and Control functions.

Functional Architecture of the TAEM navigation is composed of four different sub-functions: • Inertial navigation system; • GPS filter; • Product generation; • Navigation FDIR.

The description, discussion and trades on the TAEM Navigation concept and results, as per phase C of the project, will be presented in the full version of the paper.

Poster 21: Towards a Scalable Technology for Albedo Insensitive Sun Sensors in Silicon Carbide Romijn J1 1Delft University Of Technology Sun sensors are vital parts in the attitude control of spacecraft, which orientates the spacecraft and its instruments in the correct direction for its mission. This entails large, heavy and expensive satellites as well as the emerging small, lightweight and lower budget Cube and NanoSats [1]. For both types of satellites, as well as the missions they embark on, it is important that the sun sensors are developed accordingly. Current sensor implementations face challenges like high assembly costs and radiation damage due to the use of predominantly silicon detectors, which require additional shielding. Most importantly though, all state-of-the- art sun sensors suffer from albedo sensitivity to some degree. As a result, the outstanding performance that many of these devices have, cannot be guaranteed in situations where the albedo is dominating the field-of- view (such as at sun rise and sun set). This work focuses on enabling a technology that is inherently insensitive to the albedo, due to the use of the wide-bandgap material silicon carbide [2]. Since technology migration typically goes hand-in-hand with cost increase, efforts are made to make this step more cost- effective. Hence, this work also focuses on the complete monolithic integration of the sensor optics in silicon carbide on a wafer scale.

The solar spectrum ranges from UV to IR and thus allows sun sensors to operate in a wide variety of light spectral ranges. However, the Earth albedo absorbs certain bands in the spectrum due to specific molecules that are present in the atmosphere. As a result, there are several opportunities for sun sensors to operate in ranges where the atmosphere is absorbing and thus be intrinsically albedo insensitive. Since silicon-based devices are sensitive in a similar spectral range as the sun (which is a good match for photography), one has to include spectral filters to only allow those spectral ranges where the albedo is fully absorbing incident light. Since silicon carbide photodiodes are only sensitive to UV light below wavelengths of 300 nm, which lies outside the albedo spectrum [6-7], they allow for an inherently albedo insensitive device, without requiring the incorporation of spectral filters. Due to the wide-bandgap characteristics, silicon carbide-based sensors are furthermore not prone to radiation sensitivity. Other approaches that target the IR range exist [8- 9], but these require spectral filters and still suffer from radiation sensitivity as they do not incorporate a wide-bandgap substrate such as silicon carbide.

Modern day sun sensors are implemented using standard silicon IC technology with some light masking layer aligned on top of it. The light sensitive elements are photodiodes, which are typically controlled through integrated CMOS electronics. Although the principle of a sun sensor is simple [10] and can be implemented using just four photodiodes, state-of-the-art implementations incorporate large imaging sensor-like arrays with complex readout electronics [11]. The use of standard silicon IC technology is relatively cheap, as the infrastructure for it is completely developed and the production quantity is high. However, the light masking layer is typically implemented on die-level or in the packaging. Due to the alignment of IC and light masking layer that each device requires, this part of the fabrication is relatively costly and would benefit from fabrication techniques compatible with wafer-scale production.

Another advantage of implementing a technology for sun sensors that incorporates silicon carbide and complete monolithic integration is that it allows for miniaturization of the devices, which is in line with the miniaturization trend of satellites. Since silicon electronics are affected by the harsh environment of space, they require shielding to ensure reliable operation over extended periods of time [10]. This adds volume and weight to the sensor, which limits the potential of miniaturization as well as translates to added launch costs.

In conclusion, we propose a concept for a novel, scalable technology for sun sensor devices that are inherently albedo insensitive. Moreover, a scalable and wafer scale approach enables a cost-effective migration from the existing silicon-based technologies to the silicon carbide-based technology. To profit from existing approaches, it is essential that silicon carbide photodiodes are integrated with on-chip readout electronics for signal conditioning as well as a scalable fabrication method for the light masking layer that facilitates sensor design freedom. At present, a silicon-based platform is under development to investigate approaches for monolithic integration of the light masking layer on wafer scale. In parallel, silicon carbide- based CMOS circuitry and imager designs are fabricated to deduce the level of integration that is currently feasible. By later combining these two research branches, a sun sensor implemented through the monolithic integration in silicon carbide technology can be obtained.

Poster 22: Guidance, Navigation and Control for Asteroid Orbit Station-Keeping with In-Situ Gravity Estimation Sanchez J1, Biggs J2, Bernelli-Zazzera F2, Vazquez R1 1Universidad de Sevilla, 2Politecnico di Milano The exploration of asteroids, and other small bodies is of great importance to understand early solar system history and planetary processes, see [1]. Asteroid sample return missions such as OSIRIS-REX, see [2], and Hayabusa 2, see [3], are currently on-going. Moreover, future planned missions for the next decade include visiting the Jupiter , see [4], and the rare metallic object of the , see [5]. Designing missions to non-visited small bodies is challenging because little is known about the target object except its orbit, spin-rate and pole orientation. A long in-situ characterization campaign (~ 1 year) is usually carried out to properly determine the small body shape and gravity field, see [6]. The non- homogeneous gravity field is considered the major disturbance when performing low-orbit operations, see [7], and must be considered for efficient guidance, navigation and control (GNC) design.

Close rendezvous station-keeping and a stationary attitude are required when estimating in-situ the unknown asteroid gravity field. To this end, GNC systems have to be designed accordingly to meet these challenging requirements. For this mission phase, the navigation part is of paramount importance since not only the spacecraft state has to be determined but also the gravity parameters (directly or indirectly). An indirect approach is undertaken in [8] where an extended-state observer estimates the gravity disturbance and a least- squares method provides the gravity parameters. On the other hand, a direct approach is used in [9] where the gravity properties are obtained simultaneously with the state using an unscented Kalman filter (UKF).

Integrated guidance and control strategies using model predictive control (MPC) can be employed if the model parameters can be accurately determined using a navigation filter. MPC usually relies on linearization and discretization procedures such that it can be employed on-board and autonomously, see [10] where MPC is employed for low Earth orbit satellites control. In addition, computational burden should also be limited in the navigation strategy by estimating a reduced number of gravity parameters.

In this work, a GNC strategy is developed for station-keeping in the vicinity of an asteroid while maintaining a stationary attitude, as well as performing in-situ estimation of the gravity parameters. The navigation filters rely on the UKF whereas the guidance and control logic is MPC-based. The orbiting probe is assumed to be equipped with optical sensors such as a camera and a laser imaging detection and ranging (LIDAR), which combined provide relative position with respect to landmarks on the asteroid surface, see [11]. The attitude sensors are star-trackers providing inertial orientation and gyroscopes for angular velocity estimation.

The translational motion is modelled by the Gauss variational equations for modified equinoctial elements (MEE). This way, the singularities only arise for equatorial retrograde orbits. The rotational motion is characterized by the modified Rodrigues parameters. The complete GNC scheme is shown in the paper where it can be seen that attitude and orbit filters work independently at different frequencies but exchange their gravity parameters estimation. The orbit reference is generated allowing a "natural" evolution of the non- controlled parameters assuming the controlled ones (semi-major axis and eccentricity) are constant. The orbit MPC tracks the controlled parameters by solving a quadratic-programming optimization.

Some results for a 32 km circular controlled orbit around Eros 433 are shown in the paper. The effectiveness of nullifying the out-of-plane control to prevent unnecessary controls, see [10], is demonstrated. The control performance and gravity parameters estimation are assessed for several initial orbit inclinations and a comparison is carried out for the case where the MPC lacks information about the gravity parameters.

Finally, potential improvements, such as the consideration of a satellite constellation sharing its gravity estimation to obtain a joint estimate, is identified.

[1] Castillo-Rogez, J. C., Pavone, M., Hoffman, J. A., and Nesnas, I. A. D., “Expected science return of spatially extended in-situ exploration at small Solar system bodies,” 2012 IEEE Aerospace Conference, 2012, pp. 1–15. doi:10.1109/AERO.2012.6187034. [2] Berry, K., Sutter, B., May, A., Williams, K., Barbee, B., Beckman, M., and Williams, B., “Osiris-rex touch-and-go (tag) mission design and analysis,” Advances in the Astronautical Sciences, Vol. 149, 2013, pp. 667–678. [3] Watanabe, S.-i., Tsuda, Y., Yoshikawa, M., Tanaka, S., Saiki, T., and Nakazawa, S., “ Mission Overview,” Space Science Reviews, Vol. 208, 2017, pp. 1–14. doi:10.1007/s11214-017-0377-1. [4] Stanbridge, D., Williams, K., Williams, B., Jackman, C., Weaver, H., Berry, K., Sutter, B., and Englander, J., “Lucy: Navigating a Tour,” AAS/AIAA Astrodynamics Specialist Conference, Stevenson, United States of America, 2017. [5] Williams, D., Elkins-Tanton, L., Bell, J., Lawrence, D., Weiss, B., Russell, C., Wenkert, D., and Amiri, N., “The NASA Psyche Mission: a Journey to a Metal World,” GSA Annual Meeting, Indiana, United States of America, 2018. doi:10.1130/abs/2018AM-316784. [6] Miller, J., Konopliv, A., Antreasian, P., Bordi, J., Chesley, S., Helfrich, C., Owen, W., Wang, T., Williams, B., Yeomans, D., and Scheeres, D., “Determination of Shape, Gravity, and Rotational State of Asteroid 433 Eros,” Icarus, Vol. 155, 2002, pp. 3–17. doi:10.1006/icar.2001.6753. [7] Ceccaroni, M., and Biggs, J., “Analytic perturbative theories in highly inhomogenous gravitational fields,” Icarus, Vol. 224, 2013, pp. 74–85. doi:https://doi.org/10.1016/j.icarus.2013.01.007. [8] Biggs, J. D., and Ciccarelli, E., “In-situ tracking of a solar sail’s characteristic acceleration using a robust active disturbance estimator,” 5th International Symposium on Solar Sailing, Aachen, Germany, 2019. [9] Stacey, N., and D’Amico, S., “Autonomous Swarming for Simultaneous Navigation and Asteroid Characterization,” AAS/AIAA Astrodynamics Specialist Conference, Snowbird, United States of America, 2018. [10] Tavakoli, M., and Assadian, N., “Model predictive orbit control of a Low Earth Orbit satellite using Gauss variational equations,” Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Vol. 228, No. 13, 2014, pp. 2385–2398. [11] Vetrisano, M., and Vasile, M., “Autonomous Navigation of a Spacecraft Formation in the Proximity of an Asteroid,” Advances in Space Research, Vol. 57, 2015. doi:10.1016/j.asr.2015.07.024. Poster 23: DOCKING VIA FUSION OF MONOCULAR VISION- BASED SYSTEMS Sharif H1,2, Suppa M1 1Universität Bremen, 2DLR Space exploration is crucial for satisfying our to learn about the solar system and challenging our technological limits. Navigation is still a critical factor in preparing any exploration mission, as remote operation limits the objectives of the mission while onboard sensors alone assume high risks in unmanned navigation of the spacecraft. In this paper, multimodal monocular imaging sensors are employed in conjunction with acceleration sensors to offer autonomous navigation for rendezvous and docking with a non- cooperative spacecraft in orbit. The technique is also applicable for capturing decommissioned spacecraft for space debris removal missions and for providing on-orbit servicing or refuelling of one.

In this work, visual odometry is drawn from long-wavelength infrared and monochromatic visible images to evaluate the position and orientation of the subject relative to the cameras for tracking the subject. FlirOne dual imaging cameras with 240x320 in infrared and 480x640 in visible pixel resolution are employed to acquire images at 1Hz frequency. Dual modalities are used as imaging in each wavelength compensates for the other’s shortcomings. Visible imagery often offers more details of the subject which is ideal for accuracy of feature matching. However infrared imaging can still view the contours of the subject in extreme lighting conditions, when visible imaging is blinded.

This study employs a novel camera calibration technique where visible and infrared cameras are calibrated using a single passive checkerboard for both modalities of the camera. The calibration of both cameras in parallel ensures high precision in extraction of geometric transformations of the subject.

In order to obtain a robust vision-based tracking of the motion, images are filtered by a non-local means- based technique to reduce noise and further smooth out the pigmentations. Then SURF (Speeded-Up Robust Features) feature detector extracts features of the target. This feature detector was selected as it offers robustness to illumination and orientation variations, ideal for rendezvous and docking of two spacecraft in most lighting conditions. The outliers are eliminated by applying a RANdom SAmple Consensus (RANSAC) to the matched features. Despite application of RANSAC, not all outliers are accounted for, which tremendously impacts the system’s accuracy.

Inertial Measurement Unit (IMU)’s accelerations is also collected at 200Hz. Due to high drift, it is unreliable as a sole measurement instrument. However, pose estimations (relative translation and rotation vectors) from both images are fused with the IMU measurements using an Extended Kalman Filter (EKF) to further improve the accuracy in detection. Accelerations are defined as the control as there are more updates available, while the EKF is updated with the visual odometry of either or both visible and infrared images whenever a measurement update is available.

The rendezvous and docking was simulated at the state-of-the-art Testbed for Robotic Optical Navigation (TRON) facility at DLR. TRON lab enabled a controlled test environment where the data acquisition instruments were mounted on a KUKA robotic arm as it approached a fixed reference target. The custom pre- programmed robotic arm allowed for emulating the same rendezvous and docking path and speed multiple times during the study.

An illustration of the data acquisition setup with the model is given in the paper. Moreover, the target, a 75 : 1 scale model of a capsule, was 3D printed via a gray Poly Lactic Acid (PLA) to mimic similar texture as to the spacecraft’s Multi-Layer Insulation (MLI) covered external surface. The model was also coated in select regions with aluminium foil to emulate contrasting thermal emissivities of certain regions.

Preliminary results demonstrate the detected motion for visual odometry in each modality, ground truth measurements, as well the EKF outcome. It can be observed that the EKF struggles when thermal imaging is unavailable. However, the fusion of sensors improved the performance to accuracies below 1m in range. As the camera becomes too close to the target, there are very few features in field of view, which significantly weakens the visual odometry’s performance. At the same time, IMU’s drift is exponentially increasing, which explains the high range of variance for EKF towards the end. It is believed that the overall performance of the EKF can be further improved by reducing the noise in the instruments and eliminating outliers in the measurements.

In summary, the results illustrated that although each sensor alone suffers from inaccuracies, the sensor fusion presents a robust method towards autonomous guidance and navigation of a spacecraft. The proposed vision-based navigation technique is appealing for space applications, as it employs limited onboard memory and computational power. Additionally, the docking of two spacecraft can be achieved without the need for extensive knowledge of the spacecraft nor former customizations made to it so that the instruments can detect the reflectors to dock with. Poster 24: ALINA Moon Lander GNC – Architecture & Design Winter M1, Krüger H1, Redondo Gutierrez J1, Farì S1, Razgus B1, Magalhaes Oliveira W1, Wenzel A1, Solari M1, Bremer S1, Woicke S1, Theil S1, Seelbinder D1, Maass B1, Schlotterer M1, Heidecker A1, Schwarz R1, Hervieu C2, van Camp A2, Eichhorn H2 1German Aerospace Center (DLR), 2Planetary Transportation Systems GmbH (PTS) With the Commercial Lunar Payload Services program (CLPS) by NASA and the recent ESA decision to start the robotic exploration of the lunar surface, the development of unmanned Moon landers becomes an important engineering challenge for the early 2020s. This paper describes the architecture and design of the AOCS/GNC system developed by DLR and Planetary Transportation Systems GmbH (PTS) for the Autonomous Landing and Navigation Module (ALINA). ALINA is the combined transfer stage and Moon lander currently under development by the NewSpace company PTS [1]. Originating from the , of which the project received a total of $750,000 of prize money, PTS plans to perform the first European landing on the Moon in the early 2020s with the target of science experiments (e.g. in-situ resource utilization), technology demonstrations and commercial payloads.

Driving requirements for the system are the need for complete autonomy during the landing phase and delta-v maneuvers, the required absolute landing accuracy of 500m along-track and 750m cross-track and the touch down conditions (i.e. horizontal speed < 2m/s, vertical speed < 4m/s, yaw/pitch angle < 10°). Moreover, the nature of a non-abortable landing maneuver sets strong demands on the robustness and fault-tolerance of hardware and software, especially as safe mode occurrences or computer reboots are not an option during this critical phase. Finally, as the mission baseline requires extensive delta-v, nearly 70% of ALINA’s initial mass is propellant. This leads to a spacecraft with strongly changing mass, inertia and center of mass over the mission duration and the need to consider sloshing effects in the design.

The paper starts with an overview about the overall mission plan (see attached document). It then describes the main engine configuration, which is based on a cluster of 15 engines, the attitude control thruster configuration and the sensors configuration, which includes redundant unit sphere coverage by sun sensors, star trackers, Combined Radar Altimeter and Doppler Velocity meters, IMUs and navigation cameras for localization relative to the lunar surface. It expands on the redundant computer architecture utilizing dedicated GNC computers for running the guidance, controllers, actuator management algorithms, navigation filters as well as the Mission Vehicle Management (MVM) software. Additionally a crater-based lunar navigation system called “CNAV” [2] is run on two hot-redundant VPUs (based on Zynq UltraScale+ EV MPSoC).

The AOCS architecture assumes two phases: The Cruise and Orbital Phase (COP) and the Descent and Landing Phase (DLP). The AOCS modes are defined and the switching logic is presented as well as the objectives, the initial, target and exit conditions and the actuator/sensor configuration for each mode. The paper expands on the 3 Degree of Freedom GNC system used to control the attitude/position during the Cruise and Orbital Phase and the 6 Degree of Freedom GNC system controlling the position/velocity (outer- loop) and attitude/rate (inner loop) during the Descent and Landing Phase.

The criticality of the landing process and the lunar capture sets strong demands on the robustness and fault- tolerance of hardware and software. The paper gives an overview about the redundancy concepts and general FDIR approach on AOCS level used to achieve the required reliability in the face of a not well characterized landing environment and the usage of COTS components with a mentionable probability of radiation event occurrence.

The paper gives a high-level overview about the approaches/technologies used in all major algorithmic areas. For the guidance, it will concentrate on the off-line trajectory optimization approach for the powered descent phase from the 15 km perigee of the descent orbit to the lunar surface as well as the online selection of the currently active trajectory point (running variable) during execution. For control, a look is taken at automated controller tuning processes during the design stage, using optimal and robust control techniques to deal with the strongly changing mass, inertia and center of mass of the spacecraft as well as sloshing effects. For the actuator management, an implementation of the SIMPLEX algorithm is shown, which finds the optimal thruster actuation for the cluster of 15 main engines plus 8 dedicated attitude thrusters to realize the demanded torques and forces. The implementation also considers limitations to the number of thrusters switched on per 200ms time interval mandated by the pressurization system and the electrical system. Moreover it shows a graceful degradation in performance in case of thruster failures. On the navigation side, a look will be taken at CNAV, a crater-based localization system which allows navigation relative to the lunar surface [2]. Moreover, an automated generation process for the (extended) Kalman filters is presented.

Finally, Model in the Loop test results from the ALINA GNC simulator are presented. The simulator is implemented in Matlab/SIMULINK and supports the realistic simulation of each mission phase from launcher separation, delta-v maneuvers, the Moon transfer and Moon capture to the descent orbit and powered descent. Results of a Monte-Carlo Campaign show compliance with the accuracy requirements defined for the landing position and the touch-down conditions.

1. Karsten Becker, Kate Arkless Gray (2017). Back to with PTScientists' Mission-1. Beijing, Global Space Exploration Conference, GLEX-2017,3,2A,8, x39094. 2. Guilherme Fragoso Trigo, Bolko Maass, Hans Krüger, Stephan Theil (2018). Hybrid optical navigation by crater detection for lunar pin-point landing: trajectories from helicopter flight tests. CEAS Space Journal. Springer. DOI: 10.1007/s12567-017-0188-y ISSN 1868-2502.

Poster 25: Application of Microthrusters for Space Observatory Precision Attitude Control Wolf A1, Dennehy C2 1Jet Propulsion Laboratory, 2NASA Engineering & Safety Center This paper describes the results of a NASA investigation into the benefits of microthrusters compared to reaction wheels on future observatory-class missions with tight pointing stability requirements. Pointing repeatability and stability (i.e., jitter) requirements are key for space telescope missions of the future. For example, managing jitter is essential to being able to “image” planets orbiting distant stars. Jitter requirements for missions in this class are difficult to meet with current reaction wheel-based architectures. The reaction wheels are typically the largest pointing disturbance on the spacecraft. Disturbances from reaction wheels can be mitigated, typically by mechanically isolating the wheels, which imposes system complexity and cost. Thrusters capable of thrust forces in the micronewton (μN) range (referred to as micro- thrusters or micronewton thrusters) have been developed to support the Laser Interferometer Space Antenna (LISA) mission, which requires drag-free control to place a test mass in near-perfect free-fall Beyond the drag-free control application, micro-thrusters could be used as a substitute for reaction wheels or as a supplement to wheels for fine pointing control. Used in this fashion, micro-thrusters have potential for reducing the cost and technical risks of achieving demanding pointing stability performance on observatory- class missions. Poster 26: Design of the Spacecraft Guidance, Navigation and Control Software Functions for Impulsive Manoeuvre Execution on the Coronagraph Spacecraft of the Proba-3 Mission Woo P1, Sobiesiak L1, Hamel J1, Galano D2, Grzymisch J2 1NGC Aerospace Ltd., 2European Space Agency Proba-3 is a high-precision formation flight mission, composed of two spacecraft—the Coronagraph spacecraft (CSC) and the Occulter spacecraft (OSC)—flying together in formation as a “large rigid structure”. The OSC is a large disk that blocks the Sun, while the CSC hosts the coronagraph instrument that points directly to the OSC in order to study the Sun’s . As part of nominal mission activities, the CSC executes impulsive manoeuvres using Monopropellant Thrusters (MPT).

This paper describes the design of the Guidance, Navigation and Control (GNC) software loop for impulsive manoeuvre execution by the CSC. In the context of the Proba-3 mission, this GNC loop is a part of the Spacecraft Services (SC-SRV) function that computes the actuator commands in order to realise the commanded delta-v (expected to range from 4 mm/s to 200 mm/s), while maintaining inertial attitude control. The main contribution of this paper is the design of an impulsive manoeuvre management function for a small satellite that addresses the design constraints stemming from the spacecraft configuration and the thruster characteristics.

The CSC has eight MPTs arranged in a way such that the preferred delta-v execution direction is in the body- fixed −Z-axis. As no acceleration measurements are available, the impulsive manoeuvres are executed in “open-loop”, based on on-board determination of the applied thruster force. In order to maintain attitude control, the thruster disturbance torque is determined from the on-board knowledge of the applied thruster force, the propulsion system geometry and the estimated spacecraft centre of mass location. There is an uncertainty in the centre of mass, in addition to its migration over the lifetime of the spacecraft, as propellant is depleted. Although eight MPTs are available, only six can be commanded at a time, due to the design of the unit that interfaces with the hardware. This results in a pairing constraint, where the two thrusters of the same pair cannot be actuated at the same time. The CSC also has three active reaction wheels available for attitude control. The reaction wheels have limited torque and momentum capacity when compared to the torque that can be generated by the thrusters. Thus, impulsive manoeuvres must be executed while minimising spurious torques. Furthermore, during delta-v manoeuvre executions, the thruster disturbance torques are too large to be absorbed by the reaction wheels, thus the thrusters must assume attitude control during the manoeuvres.

The MPTs operate in “Pulse Mode Firing”, where thrust is generated from a sequence of pulses, each of specific duration. The impulse generated from each pulse depends on several variables, such as the inlet pressure, the opening time, the pulse position within the pulse train, etc. Furthermore, the first pulses of a pulse train exhibit degradation from their nominal value. These MPT characteristics must be considered in the on-board estimation of the applied thruster force.

In order to execute the desired impulsive manoeuvre while accommodating for the various design constraints, the SC-SRV contains the following functions: - Navigation for the determination of the attitude states based on star tracker measurements, and for the estimation of the spacecraft’s mass, centre of mass, and inertia properties; - Guidance for the computation of the attitude control errors; - Control for the computation of the commanded force for delta-v execution and of the commanded torque for attitude control; - Thruster management for the computation of the thruster on-time commands.

More specifically, the thruster management function includes the following considerations: - The duty cycle command is computed based on a predefined actuation pattern that maps unitary force and torque commands to duty cycle commands. This actuation pattern minimises the residual spurious torques when a force is commanded and minimises the spurious forces when a torque is commanded. - The actuation period is divided into two half-cycles. A subset of up to six thrusters are commanded in the first half-cycle, and the remaining thrusters are commanded in the second half-cycle. The thruster subsets are selected according to the pairing constraint. - The duty cycle command is converted into on-time commands. Commands that are smaller than the thruster minimum on-time can either be held-over to be added to the next actuation cycle, or rounded down or rounded up in the current actuation cycle, depending on user selection. - The thruster applied force is estimated based on the on-board model of the thruster performance. Each thruster’s nominal force is computed based on the inlet pressure measurement. For the first pulses, the force degradation is determined based on a known mapping of the degradation factor with respect to the pulse number. - The thruster torque is estimated considering the centre of mass location as estimated by the Navigation function.

The proposed design is validated in-the-loop using a high-fidelity simulator, implemented in MATLAB/Simulink, which includes orbit and attitude propagation, as well as sensor and actuator dynamics. The performance of the impulsive manoeuvre execution is based on the following metrics: - delta-v magnitude realisation error; - delta-v direction realisation error.

A series of Monte Carlo campaigns are performed. The first series of tests consider two types of impulsive manoeuvres along the preferred direction (−Z-axis of the body-fixed frame): small manoeuvres with delta-v ranging from 3 mm/s to 20 mm/s, where first-pulse dynamics dominate the thruster performance, and large manoeuvres with delta-v ranging from 50 mm/s to 200 mm/s, where first-pulse dynamics are not significant.

The second series tests consider small delta-v manoeuvres along the non-preferred directions: +X, +Y, +Z- axes of the body-fixed frame.

The third series of tests evaluate the impact on the delta-v realisation performance due to different design parameters (such as the actuation period, the logic on handling on-time commands that are below the thruster minimum on-time).

Results show that the proposed design meets the impulsive manoeuvre performance requirements. In general, larger realisation errors are observed for the small delta-v manoeuvres. These results demonstrate the importance of an accurate on-board modeling of the thrusters.

Poster 27: Spacecraft Flexible Attitude Dynamics Modeling for Accurate Control Design Angelone C1, Paolini E2, Lavagna M1 1Politecnico di Milano Aerospace Science and Technology Department, 2OHB Italy Spa The fast development of space technologies represents, often, a fundamental source of possibilities and ideas: each time a new objective is achieved companies can push their limits a step forward, to face harder challenges. However, every time the level of complexity and ambition of a mission increases, several criticalities arise as a direct consequence: the impact of flexibility is one of these. The growing interest in the topic of the dynamic coupling between attitude dynamics and flexibility is due to several factors involved in the design and operations of a space mission. The increasing performances of payloads and antennas imply more severe pointing requirements and reflect in more power demanding devices. Thus, since the using of batteries is limited, larger and so more flexible panels became unavoidable. In some cases, depending on the mission, the combination of these payloads needs and of the elastic response of the panels may become critical from the AOCS point of view, affecting the pointing accuracy. In addition, it must be remarked the extremely limited volume available inside launcher, whose requirements impose the presence of folded and unfolded configurations. These opposite operating conditions require the introduction of deployable elements that keep the configuration of the satellite more compact during the launch, but open after the separation. The presence of deployable elements is, often, a source of vibrations because of the presence of joints, whose flexibility may be not negligible. In this context, Finite Elements Software provide very accurate results modeling the structural dynamics of the panels. These tools are typically used by Structural Engineers and required a lot of experience. However, the typical working environment of GNC and AOCS Engineers is MatLab-Simulink to design the closed-loop control system considering orbital disturbances and real loads of space environment. The interface with FEM tools reveals not to be straightforward because of the complexity of the structural software. Thus, the aim of the research is to provide a robust analytical model of the coupling between vibrational phenomena and attitude dynamics to be translated in an MatLab-Simulink tool and easily used by AOCS Engineers. The analytical base of the model has its foundations on the Ritz-Galerkin Approximation of flexibility: where flexible coordinates are substituted by a set of trial functions times a vector of generalized coordinates. The associated Equations of Motions are formulated according to the Lagrangian Formalism in Quasi-Coordinates. The output of the process is a set of coupled Ordinary Differential Equations, whose size depends on the desired accuracy of flexibility approximation. In general, the higher the level of accuracy, the larger is the number of generalized coordinates needed. For this reason, at this stage of the formulation, the number of Degrees of Freedom involved in the problem is not reasonable. Therefore, the entire dynamics is projected onto the Modal Space, by solving the eigenvalues problem associated to system’s dynamics. This last passage strongly reduces the size of the problem up to the number of modal shapes of interest. In this final form, the problem can be handled and integrated in time rapidly by the CPU thanks to a reduced computational effort. According to the aforementioned techniques, the numerical tool is able to provide the most critical quantities of the system, i.e. the rotational natural frequencies of the system that can be excited by attitude maneuvers. In addition, the Simulink part of the tool integrates in time the Augmented Euler Equations describing the coupled dynamics. Downstream to this introduction, it is worth to notice that, the important advantage of the tool relies in its ability to hide the complexity of the problem. Exploiting this tool, an AOCS Engineer can assess the most important information and results affecting the closed-loop and the de-pointing effects due to vibrations, without a real deep knowledge of the modeling techniques of the flexible dynamics problem. However, a clear understanding of the structure of the problem and of its outputs is required, so that, it is possible to reconstruct matrices in the Space State. Once these matrices arising from flexible dynamics are available, a first assessment to the open loop performances and the final synthesis of the closed loop control can be performed. By means of the research, it is possible to clearly state the mathematical fundations of the model and its validation. The validation strategy included three different steps: in the first part, the results of the attitude response computed by the model is compared with some simple but representative maneuvers, whose results can be easly predicted. Then, the structural results in terms of natural frequencies of vibration are compared with the results of Nastran, and finally, a comparison with literature results of closed-loop dynamics is presented. By means of this robust procedure, the structural and attitude predictions of the model are considered to be validated. Therefore, the tool is ready to be used in closed-loop by AOCS Engineers whenever fine pointing maneuver are required in presence of flexible elements.

GNC for Small Body Missions

Centralized Relative Navigation of Multiple Spacecraft with reduced visibility: validation campaign of the vision-based approach Piccinin M1, Silvestrini S1, Capannolo A1, Lavagna M1, Gil Fernandez J2 1Politecnico Di Milano, 2European Space Agency, ESTEC Introduction

Space agencies are planning missions (e.g. Dart, Hera), that involve the release of CubeSats in close proximity to the asteroid target. Such strategy allows to perform riskier operations and to gather higher-value scientific data, still granting the safety of the main spacecraft [1]. Centralized autonomous navigation for these spacecraft has been studied as a potential navigation strategy, even though it poses several challenges (low-observability sensors measurements, large dynamics uncertainties, measurements non-linearities) [2]. This paper presents numerical tests of an on-board, centralized autonomous navigation algorithm able to reconstruct the trajectories of a fleet of CubeSats relative to an asteroid binary system. In particular, the paper studies the possibility of a centralized estimation algorithm on-board Hera spacecraft, to estimate the trajectory of Apex and Juventas CubeSats around the Didymos system. The algorithm is executed on-board the main spacecraft, which takes optical relative measurements. The centralized navigation algorithm here studied is the result of a trade-off presented in [2], where the low-visibility of the CubeSats has shown to pose challenges to the navigation. In this paper, a thorough numerical testing campaign is conducted to assess the navigation approach robustness; the algorithm testing approach is presented in detail, showing results in nominal and non-nominal scenarios.

Centralized Navigation Architecture

This section briefly presents the architecture for the CubeSats centralized navigation; for a more complete description of the baseline algorithm please see [2]. The core of the navigation architecture is made by an Image Processing algorithm and a Filter. The image processing algorithm detects and tracks the CubeSats, deriving line-of-sight measurements. The IP algorithm pipeline is presented in the paper. It takes in input two subsequent images and its output is the CubeSats position in image coordinates or a not-found message. As additional relative measurement, inter satellite ranging is considered. In order to obtain a fully-observable solution from the low observability of the sensor measurements in the presence of large dynamics uncertainties and measurements non-linearities the filter design rely on the non- linear extension of the Kalman filter [3]. In particular, the Schmidt-EKF is selected as baseline in this paper. The ERTBP, expressed in the inertial frame centered in Didymain and assuming circular motion of the secondary, is chosen for implementation. In addition, in this paper the usefulness of adding a neural network for improving the filter performance is evaluated.

Numerical testing approach

Testing simulator architecture The architecture for testing is shown in the paper: the core of the simulator is the navigation architecture, that comprises the Image Processing algorithm and the Filter. The mission scenario is modeled thanks to high-fidelity trajectories computation and generation of the synthetic images acquired by Hera spacecraft. Finally, the navigation performance is assessed analyzing its outputs in a post-processing tool. Please note that the simulator is open-loop, so each block processes the whole batch of samples constituting the scenario under study.

High fidelity trajectory propagator High fidelity trajectories are generated for the reference mission scenario. In particular, the reference scenario is the Detailed Characterization Phase of Hera mission [4]. The Hera's trajectory has been selected from the planned hyperbolic passages sequence and is defined in the Keplerian model with SRP, as shown in the paper. The assumed Apex orbit is a Short Period Orbit around L4, with a period of 12 hours and 20 minutes. The CRTBP orbit is corrected to the SERTBP [5], plus the effect of SRP. The final trajectory has a bounded, non- periodic motion. Juventas CubeSat is located in a Self-Stabilizing Terminator Orbit, with a radius of 3.3 km and a period of 56 hours. Juventas's trajectory has been corrected in the SERTBP.

Image generation tool Representative images are synthetically generated using PANGU. For an IP numerical tests, if synthetic images are used, the assumptions made for modeling the scene objects and the electro-optical sensor are extremely important [6], [7]. Details on the modeling approach of will be reported in the paper, including Didymain, Didymoon, Apex, Juventas, stars background and camera model.

Numerical tests of the algorithm Numerical tests are performed with increasing level of complexity, as follows: • Filter validation test consist of the implementation of the filter coupled with the high-fidelity orbital simulator to validate the correct implementation of the measurement model and the filter itself. Moreover, the proper tailoring of the simulator to the reference scenario is validated. • IP validation test is performed on the reference scenario and trajectories. The algorithm performance is assessed by computing: false positives, false negatives and accuracy.The objective is to test the functional implementation of the IP and the derivation of measurements models. • An Open Loop validation tests the integration of the full navigation architecture on the reference scenario; the accuracy of the navigation is measured. • A Monte Carlo campaign evaluates the robustness of the proposed architecture.

Numerical Results

In the reference orbital scenario the Cubesats have a low visibility (Apex is visible only the 13\% of the time and Juventas the 5\%), due to geometric constraints. In fact, they are visible only when falling inside the camera FOV and when not in front or behind the asteroids [2]. From the Open Loop test, it emerges that the major challenge is to cope with the lack of visibility.

Results show that centralized autonomous navigation is able to deliver an estimation error in the order of 10^2 m for both CubeSats. When LOS measurements are not available, the ranging-only navigation degrades the estimation of Juventas trajectory, whereas it is beneficial for Apex. Results are fully commented in the paper, where tests additional to the nominal scenario are reported, focused on the filter optimal tuning by means of a statistical approach. Further analyses will be presented in the paper, that identify operational constraints such as maximum blind-time. As a result, the benefit to the central navigation is the estimation of the CubeSats orbit with a lower RMS error.

Acknowledgements The presented work was carried out under ESA funding ESA RFP/3-16012/19/NL/CRS/hh, item no.19.3EC.04.

Impact of Complex-shaped Small Solar System Bodies on a Quadcopter-based Controller for Near-surface Operations Martin M1, Belien F1, Falke A1 1Airbus Defence And Space GmbH A novel mobility concept for enhanced surface mobility on small solar system bodies (SSSBs) under low- gravity environments is presented. The idea is to utilize a quadcopter-based spacecraft design with thrusters instead of rotors to potentially hover in low altitude over the whole surface of the SSSB. The particular challenges for this design in terms of control requirements are investigated. The uncertainties in the low- gravity environment, the thruster performance and allocation are explicitly taken into account in the control model, and their effects on the controller performance are inspected. The analysis shows that these uncertainties have to be taken into account in the further development of the GNC design.

Exploration missions to small solar system bodies (SSSBs) are an important key to the understanding of the formation and evolution of the Solar System. Parts of these objects are believed to have remained largely unaltered after their creation [1]. This is why in-situ measurements are vital to enhance our knowledge on the solar system. So far, the mobility concepts have been limited to only a few locations and limited regions on a SSSB. The most recent examples are the OSIRIS-REx and Hayabusa-2 missions, which use the touch-and-go principle to obtain samples from the surface and return them to Earth [2], [3]. Another mobility concept is the hopping principle from MASCOT and the MINERVA rovers. A swing arm produces torque that allows the spacecraft to a generally uncontrolled ballistic movement to another location [4], [5]. A classical landing approach, where the lander is released and stays at one location, has been attempted with the lander from . All these approaches have in common that only a limited region of the SSSB can be explored at close range.

To circumvent these limitations, a novel mobility concept is proposed. The idea is to autonomously perform relocation manoeuvres between safe landing positions by hovering with a lander at a relatively low altitude (3 to 10 m) over the surface of the target body. The goal is to significantly increase the scientific return of SSSB missions by ideally covering the whole surface. The feasibility and details of the underlying GNC concept is investigated and developed in part by Airbus Defence and Space GmbH under the DLR-funded collaborative technology demonstrator project Astrone - Increasing the Mobility of Small Body Probes. It examines a reference scenario consisting of several relocations at the well-characterized 67P/Churyumov- Gerasimenko (67P/C-G). The comet is chosen based on the available detailed characteristics and its challenging gravitational environment and shape.

One of the key elements for the implementation of the mobility concept is to employ a quadcopter-based architecture using four electrical thrusters instead of rotors. The advantage of using only four thrusters is a compact and lightweight design. The latter is particularly important, because additional complexity of the propulsion system significantly increases the system mass. Electrical thrusters in the mN thrust range required for the given reference scenario are available (e.g. [6]) and increase the efficiency due to the high specific impulse. Due to this efficiency increase, the complete surface of larger SSSBs can be covered, possibly with multiple stopovers for recharging the batteries. Unlike the hopping approach, the relocations are conducted in an autonomous and controlled manner with slow movement, described using relative dynamics with respect to the surface.

This concept imposes new challenges on the GNC design of such a surface probe. The proposed architecture for the quadcopter-based Astrone system is shown. Flash LIDAR and inertial measurement unit (IMU) measurements are used as inputs for autonomous functions for the guidance and navigation, such as map generation, localization, landing site detection and path planning algorithms. These functions are computationally expensive, but need to run on space-grade CPUs. Therefore, the map update and path planning are executed at lower frequency than the control loop. This can be justified by the relatively slow movement.

This paper focuses on challenges for the control design. The quadcopter concept is an under-actuated system with a strong link between attitude, gravitational acceleration and thrust force. Compared to quadcopter controller designs on Earth, there are several differences to be considered. (i): Firstly, the spacecraft is supposed to cover the whole or at least large areas of the SSSB. Irregular bodies, such as the comet 67P/C-G, have a highly inhomogeneous gravity field close to the surface. This slowly changes the plant dynamics during a relocation flight. When considering the complete surface, the controller must be able to cope with a wide range in gravitational field strength. (ii): Secondly, the gravity field is uncertain because it can only be estimated based on pre-processed models or in-flight measurements. A quadcopter design on Earth assumes that the gravity vector is pointing in the “downward” direction. This might not be the case for the hovering scenario on the SSSB, so that the controller has to cope with this model uncertainty. (iii): Lastly, to increase efficiency, an electrical actuator has been chosen. These actuators have a slower reaction time compared to cold gas or chemical propulsion systems [6]. Additionally, they have limited thrust and a fixed orientation, impacting the controllability.

These aspects indicate that the robustness of the control concept has to be analysed. The uncertainties are included in the control model by a linear fractional representation (LFR). The time-varying plant (i) is addressed by linearization of the nonlinear dynamics for several surface locations. An LFR model is derived to express the gravity field (ii) and actuator (iii) uncertainties. To illustrate the approach, the effects of the uncertainties on a given controller in terms of the prescribed controller requirements are shown throughout the surface. The sensitivity of the system with respect to the uncertainties (i)-(iii) is compared to conventional uncertainties in a GNC system such as centre of mass shifts. Simulations of the relocation manoeuvre are used to validate the analysis. The analysis shows that the uncertainties in the given reference scenario have a considerable influence, which has to be taken into account in the further development of the GNC design.

LOCALIZATION AND MAPPING MERGING SILHOUETTES INFORMATION AND FEATURE TRACKING FOR SMALL BODY APPLICATIONS Panicucci P1, Bochard R3, Lebreton J3, Lefez R3, Zenou E1, Delpech M2 1ISAE-SUPAERO, 2CNES, 3Airbus Defence And Space The limited knowledge of small bodies properties yields numerous challenges in mission design and spacecraft operation. In particular, the time required for communication and the uncertainty in small bodies' parameters estimation require the development and improvement of autonomous navigation and decision making. During close approach, i.e. when the small body is several pixels in the mapping camera field of view, the probe slowly moves radially with respect to the observed small body which is rotating around its rotation axis. At this point of the mission, it is of crucial interest to reconstruct the small body shape and estimate the rotation pole orientation in the inertial space. The former is an important milestone to investigate the gravity field under the assumption of constant density and to enable the use of model-based tracking algorithms. The latter is important to understand the relative dynamics between the small body, the Sun and the observing camera. Currently small body missions approach relies on optical data, range and range rate measurements to solve for the relative geometry between the Earth, the Sun, the small body and the spacecraft. In an iterative procedure, shape reconstruction is performed by processing images on-ground. Current techniques rely on shape from shadowing, i.e. stereophotoclinometry, or shape from motion, i.e. stereophotogrammetry. The full pipeline requires not only precise orbit determination and radiometric tracking, but also not negligible downlinked optical data to be processed on ground. This approach has been used since the beginning of deep space exploration and, as a consequence, it has a high degree of reliability. Unfortunately, these algorithms cannot be used on board because of the extreme computational burden and the need of human-in- the-loop to control the convergence of the output solution. In the perspective of designing autonomous concepts to enable localization and mapping during the approach phase of small body missions, new solutions must be developed by limiting the needed data to the information available on board and by considering algorithms that could be run without delayed communication with Earth. Airbus Defence & Space (ADS) in the last decade has invested effort to implement, validate and test precise vision-based navigation (VBN) solutions to improve on-board GNC and increase pointing accuracy. Proposed algorithms have allowed to gain expertise on precise VBN and algorithms implementation in navigation filters on space-certified processors. Feature tracking and model-based tracking have been implemented, validated and tested in different environments. A high Technology Readiness Level (up to 5/6) has been demonstrated. The Advanced Studies teams (IP & GNC departments) are constantly investigating new methods to be tested and implemented. The present paper outlines recent research that has been carried out at ADS about shape reconstruction in collaboration with CNES and ISAE-SUPAERO. In our recent work, shape from polygonal silhouette has shown its applicability under the hypothesis of known camera poses and under different illumination conditions. A summary of the results is shown where the shapes of asteroids Itokawa and Bennu are reconstructed under. The main drawback of that technique relies in the fact that poses must be known and that no information about points inside the silhouette can be inferred. To overcome these problems, an algorithm is presented here that merges information from feature tracking, building block from ADS portfolio and based on the Themis library, and silhouettes to localize the camera with respect to the small body and to estimate small body shape. On the one hand, tracked features give information about concavities perpendicular to the image plane, they allow localization of the camera with respect to the small body and they improve the detection of concavities. On the other hand, silhouettes can constrain the tracked points to lay inside small body limbs, they provide information about points belonging to the limbs and they allow the gathering of small body convex approximation. First, the feature tracking is used to solve for the relative dynamics between the asteroid and the camera. This allows the gathering of the camera poses with respect to the small body and find the relative geometry between the Sun, the small body and the spacecraft. Second, the silhouette are extracted by considering the projection of the Sun in the image plane. This enable to recover from the silhouettes only the part of the body that is surely exposed to the Sun light. Third, the 3D points are extracted from the silhouettes and merged with the one deduced from the feature tracking. Finally, to show the applicability of the proposed algorithm, numerical results are presented and commented. In the simulations, a spacecraft, like Hayabusa or OSIRIS-REx, is slowly approaching an asteroid and takes images with the mapping camera, like AMICA or OCAMS. The trajectory is simulated and images are generated with the SurRender software, a high-fidelity rendering engine optimized for space applications developed by Airbus Defence & Space, to have a ground truth to compare the estimated poses and the reconstructed shape. Mission and GNC system design of the Juventas CubeSat on-board the Hera mission Moreno Villa V1, Palomino Aguado A1, Villadoniga Prendes J1, Balaceanu M2, Prioroc C5, Cabral F3, Goldberg H4, Garcia Gutierrez B5 1GMV, 2GMV, 3GMV, 4GomSpace, 5ESA-ESTEC Juventas is a deep space 6U CubeSat with the objective of performing a scientific and technological demonstration mission, aiming for the characterization of the binary asteroid, Didymos. The asteroid system is composed of two bodies, a 780-meter asteroid called Didymain and a smaller 160-meter asteroid called Didymoon orbiting around it. Juventas takes part in the Hera mission as one of the two CubeSats that will be released in the asteroid proximities. Hera is an ESA mission that will explore and characterize the Didymos asteroid system after the impact of DART, a NASA impactor with the aim of assessing planetary defence techniques by deviating the secondary component of the asteroid. GMV, as part of a consortium led by GomSpace, is in charge of Mission Analysis and the development of the GNC system of Juventas. Juventas will be inserted into a special set of orbits, called Self-Stabilized Terminator Orbits (SSTO). SSTOs, also known as photo-gravitational orbits, have been studied for a large number of small-body missions. The existence of this special type of orbits arises in environments where the solar radiation pressure (SRP) perturbation is comparable to the gravitational pull of the central body. SSTOs are quasi polar low- eccentricity orbits perpendicular to the sun direction. The SRP perturbation displaces the orbital plane up to an equilibrium offset, where the gravitational pull in the Sun direction and the SRP force cancel each other out, whilst the low eccentricity compensates for the plane rotation and the associated non-inertial forces. As detailed by this paper, it is observed that the selection of the SSTO as the baseline orbits for the observation phases of Juventas is an optimum solution that eases the operations of the CubeSat. Their self- stability reduces the necessity of station-keeping manoeuvres and the associated delta-v budget, thus passively improving the safety of the mission. Another significant aspect of these orbits lies in the scientific return obtained when orbiting them, partially due to the reduced number of operations creates the perfect scenario for highly precise navigation. In these orbits, the main payload of Juventas, a low-frequency radar, is accountable for mapping the interior of both asteroid bodies. Juventas equips an inter-satellite link (ISL) system for both communications and radiometric (range and range-rate) capabilities. This information may be processed after operations for radio science, to improve the reconstruction of the trajectory. This leads to an entire characterization of the asteroids dynamical properties as their masses and gravitational coefficients. For GNC sensors, Juventas is equipped with a navigation camera and a laser altimeter. The combined usage of these two types of measurements improves accuracy in the state estimation. Camera images provide information on the line-of-sight (LOS), while the laser altimeter ranging capabilities complete the information, fully characterizing the state of the CubeSat. This paper justifies the necessity of a semi-autonomous attitude guidance (SAG) GNC mode to ensure sufficiently accurate pointing in order to guarantee the operational safety of the mission. Attitude profiles uplinked by ground are worsened by both the distance to the target at which Juventas operates and the ground turn-around times (time between data downlink and command uplink), characteristic of deep space missions. The high latency of the information causes greater uncertainties in the predicted CubeSat state, while smaller distances to the target translate these errors into larger pointing errors. Hence, to prevent missing from the camera field of view the asteroid, on-board attitude corrections must be commanded by the GNC system to achieve continuous imaging of the asteroid. The SAG mode is accountable for this command which is generated making use of on-board image processing techniques to estimate the LOS. Nevertheless, ground operations are proven to be sufficient to perform required station-keeping manoeuvres, which are scheduled to reduce the operational cost of the mission while assuring the safety of the CubeSat. As an extended part of the mission, Juventas attempts to land on the surface of the secondary body, Didymoon. The landing process comprises a controlled touchdown and a series of uncontrolled bouncings until Juventas rests on the asteroid surface. During the bounces, Juventas uses an inertial measurement unit (IMU) to collect information about the physical properties of the Didymoon surface. Once the CubeSat is stabilised on the surface, a gravimeter payload will give further insight into the Didymoon dynamical properties. This paper depicts two main strategies for the landing. The first strategy is the most straightforward where the CubeSat follows a ballistic trajectory from the SSTO to the asteroid. Thus, the transfer is simplified and the operational complexity is significantly reduced. The second approach takes advantage of third-body dynamics, carrying Juventas through the second Lagrangian point before reaching the asteroid surface. After each touchdown, the bouncing is dependent on the arrival velocity and the physical characteristics of both, the surfaces of the asteroid and the CubeSat at the contact point. These characteristics can be gathered in a parameter called Coefficient of Restitution (CoR). The bouncing velocity could be high enough to make Juventas escape the gravitational influence of Didymoon. The low mass of the target body and the assumed dynamics (CR3BP) implies that an escape is possible for velocities below down to 4 cm/s in some considered scenarios. Therefore, the outbound velocity shall be minimized. At the same time, the inbound velocity must be high enough to allow bouncing, as previously mentioned for, but below a minimum threshold that could jeopardize the landing, since the CoR is hardly controllable. This imposes compelling constraints on the mission design, highlighting the impact and necessity of an autonomous GNC system. The highly perturbed dynamics together with the small size of the target makes the landing the most critical situation that Juventas will encounter. The latency of information and the manoeuvre execution errors drive the performance of the landing. Thus, novel techniques are conceived to maximize the probability of a successful landing.

HERA GNC Subsystem – Preliminary Design Pellacani A1, Kicman P1, Cabral F1, Birlan A1, Bodin P2, Larsson R2, Gil J4, Carnelli I4, Gerth I3, Fittock M3 1GMV Aerospace And Defence, 2OHB Sweden, 3OHB system, 4ESA AIDA is the international NASA- and ESA-supported collaboration that will combine the data obtained from NASA’s DART mission (which includes ASI’s LICIACube) and ESA’s Hera mission to produce the most accurate knowledge possible from the first demonstration of an asteroid deflection technology. AIDA is not a formal joint project of NASA and ESA, but instead an agency-supported collaboration among planetary defense and asteroid science researchers, most of whom are involved in the DART or HERA missions, to share information and contribute to the planning and execution of both missions so as to enhance their synergy. HERA is an ESA mission that will rendezvous with and explore the Didymos binary asteroid system approximately two years after the impact of DART, studying both Didymain and Didymoon in detail and examining the visible after-effects of the impact. HERA has been approved at the ministerial council Space 19+ and it will be one of the mission of the new Space Safety Program (S2P). GMV leads an international consortium composed by , Romania, Portugal and Poland that, together with our partners like OHB Sweden, is designing what can be considered “the brain” that drives the HERA spacecraft. The HERA GNC subsystem is based on a robust solution that will allow the spacecraft to be operated manually during the interplanetary phase and during the first part of the close proximity operations with the main purpose of safely study the low gravity environment around the binary asteroid. During this phase, the autonomous GNC will be tested in flight and after validation it will have authority over the spacecraft. In order to guarantee a high level of autonomy, the first step is to give to the Spacecraft enough on-board information to estimate its position with respect to the asteroids and, based on that estimation, to act accordingly following a preloaded sequence of commands or objectives. To achieve this requirement, a vision based GNC has been designed, which includes image processing algorithms and a navigation filter capable of processing the visual information and to get the desired estimation. The HERA GNC subsystem is capable of using not only a visual camera but also several other payloads, involving data fusion with a thermal infrared camera and a laser altimeter (PALT). These additional instruments will be able to increase the robustness of the strategy: the thermal infrared camera will be able to see where the AFC is limited by shadows and phase angle constraints; PALT will be able to give observability on the radial direction, which is poorly estimated in a monocular configuration of a vision based GNC (especially if the low gravity environment does not allow stable orbits and hyperbolic arcs are preferred for safety reasons). HERA close proximity operations have been investigated in the previous phase and the design has been updated in the frame of the current phase B2. A specific distance of 9 km is identified from a GNC driving factor. If the Spacecraft will be outside the 9 km sphere centered in the main asteroid, the AFC will give a full visibility of the body and centroiding techniques can be used for the state estimation. For closer fly-bys the entire camera FoV will be covered by the asteroid and a feature tracking technique is baselined to maintain the estimate of the relative state. Image Processing and navigation functions have been developed taking into account this factor and a detailed design showed its feasibility taking into account ground segment operations and pushing the validation of the on-board algorithms up to TRL 5/6 with Hardware-In-the-Loop tests. In order to use the navigation algorithm based on feature tracking image processing a precise initialization is needed and a range measurement (from PALT) is required. Considering that the altimeter is a payload and not a GNC unit/sensor, it is important to investigate alternatives in case PALT could not be used on-board HERA. Currently the HERA GNC team is investigating three different technologies/strategies in order to get closer than 9 km without altimeter. The paper will describe in detail these alternatives: the MOSAIC technique that artificially enlarge the camera FoV and allow to keep using the centroiding for smaller distances than 9 km; the Enhanced Relative Navigation (ERN) that is capable of matching a previously generated landmarks database in order to improve the initialization of the relative navigation algorithm; Didymoon based navigation using centroiding on the moon of the binary system when it is clearly visible and separated from the primary body in the camera plane. This paper will include the consolidated strategies of the vision based GNC designed for the HERA mission, together with the test campaign results. The focus will be directed towards the autonomous close proximity operations and the different technologies/strategies that will allow to map the crater with a target resolution of 10 cm per pixel. The final selection of the baseline depends from development and validation considerations, together with system constraints that will not be tackled in the paper, but the results reported will be used as inputs from mission prime (OHB System) and ESA in order to take the decision Current Space Missions

GNC Spacebus NEO: a full electric platform for telecommunication satelittes, safe mode on star tracker and electrical orbit raising Dandré P1, Boulange D1, Chevallier M1, Goffinet G1, Goulamhoussen J1, Sumelzo Martinez I1 1Thales Alenia Space Spacebus Neo is the new Thales Alenia Space telecom satellite product line, developed to meet the market needs and launcher landscape. It replaces the Spacebus 4000 product line and enables to embark all types of telecom mission payload with the usage of xenon plasma propulsions for orbit raising and geostationary orbit control.

The GNC safe mode (SHM) on the Space bus Neo product line is based on multi-head start tracker (STR). In the context of the coverage of a large range on the product line, the classical Coarse Sun Sensor (CSS) solution of the Spacebus 4000 platform is not retained. For robustness demonstration purposes, a safe mode solution based on CSS (Ultimate Safe Mode, USM) is developed for the first satellite of the product line only (KONNECT).

The orbit raising strategy on Spacebus Neo is purely based on xenon plasma propulsion. It includes the management of the first perigee crossings in Guided Waiting Phase (GWP during which S/C attitude follows a guidance law provided by the ground, based on star tracker measurement, but no thrust is performed. The electric propulsion (XPS) is used in Guided raising phase (GRP) to perform the electrical orbit raising while controlling (closed loop control) the angular momentum stored on the reaction wheels using the command on a pair of diagonal thruster orientation mechanisms (Two-Axes Thruster Mechanisms, TATM).

The GNC software on the product line is based on three applications for attitude and orbit control (AOCS), solar arrays management (SADM) and electrical thruster orientation management (TATM). All the applications are developed in a Matlab Simulink development environment and the GNC software is designed following the automatic code generation process described in [1].

The paper will present the developed concepts and the strategy used for the validation of the safe mode and orbit raising strategy on SpaceBus Neo, and will illustrate them by giving some results of simulations done during the development phases.

References 1. Chevallier M., Dandré D., Grossinho T., Lopez Negro P., Automatic code generation of GNC software for Spacebus Neo

The EUCLID AOCS and FGS verification tasks at System level Saponara M1, Bosco A1, Llorente S2, Rosso C3, Huzel D4 1Thales Alenia Space Italia, 2SENER Aeroespacial, 3ESTEC, 4Rockwell Collins Deutschland Thales Alenia Space in Italy (TAS-I) is prime contractor of the Euclid Medium Class mission that, belonging to the ESA 2015-2025 plan, is currently in the C/D phase. Euclid will be launched in 2022 and will operate for more than 6 years in large amplitude orbit around L2. The objective of Euclid is to elucidate the geometry and the nature of the dark energy and dark matter components of the universe, with unprecedented accuracy, in a large survey of the extragalactic sky by two main techniques, weak lensing and galaxy clustering. Both techniques require the ability to survey a large fraction of the extra-galactic sky over the mission lifetime, with very high system stability (telescope, focal plane, spacecraft pointing).

The extremely accurate pointing performance is achieved through the use of the Fine Guidance Sensor (FGS), which provide very precise attitude measurement to the Attitude and Orbit Control System (AOCS) control loop during the scientific mode. The FGS is accommodated inside the Payload Module (PLM) in order to limit the contribution of deformation between FGS and instruments field of view.

TAS-I has also driven the design of the Hybrid solution used in the scientific mode. It consists of the management of the AOCS actuators (reaction wheels (RWL) and cold-gas micro-propulsion subsystem(MPS)) in order to cope with spacecraft agility and fine torque commanding for high precision pointing. In this approach, the RWL are used only for the slews, bringing them to rest before each observation commences, where the MPS-based control loop is restored. This solution has required the execution of a RWL qualification test campaign for the start-stop approach, as well as refined modelling of the Micro-Propulsion System (MPS). In particular, different speed profiles have been exercised on an Engineering Qualification Model (EQM) simulating the expected usage of the RWL during the mission in different phases, like execution of slews or compensation of PLM induced disturbances: each of them have been repeated 168000 times, covering a so called “cycle”, and 116 cycles have been executed, arriving then to about two millions start-stop cycles. The main results was that no degradation on RWL friction performance has been identified at the end of this so stressing campaign, confirming the feasibility and expected performance of the Hybrid solution. The Euclid program is organized allocating all Sub-Systems to different European industries, and in particular, Sener and ADSNL are in charge of the AOCS design, implementation and verification. Furthermore, the FGS is developed by Leonardo Firenze, while the star catalogue is in charge of Astronomic Observatory of Turin (OATO).

Nevertheless, significant verification tasks at both AOCS and FGS level are performed by TAS-I. In particular, TAS-I has arranged an Avionic Validation Module (AVM), which provides a test bed to validate by test the electrical design of the Spacecraft, its operational and functional interfaces and the system checkout including software and database: in that sense, AOCS closed loop simulation campaign with real sensors/actuators hardware can be performed.

Furthermore, a FGS test campaign at system level has been specified for the functional and performance requirements verification.

Finally, TAS-I has developed together with Leonardo a tool for the management of the FGS star catalogue, in particular for what concerns its partitioning for on-board upload.

The paper will present the main results of the RWL start-stop qualification campaign, carried out by Rockwell Collins. Then, it will concentrate on the Euclid verification aspects, and in particular on the description of the AVM (with test results), on the FGS system tests validation campaign and on the FGS star catalogue management concept and tool.

AOCS INNOVATIONS FOR EUROSTAR NEO FULL-ELECTRIC PLATFORM Reuilh A1, ROUSSEL S1, BEROUD J1, ROSSO C2 1Airbus Defence And Space, 2ESA Airbus Defence and Space is developing the Eurostar Neo telecom satellite product line, supported by ESA in the frame of the ARTES 14 program. Eurostar Neo product is highly innovative with respect to E3000 telecom product, starting from the hardware related to AOCS: new high momentum reaction wheels, new solar arrays with semi-rigid panels, large range for platform size, new OBC. Eurostar Neo AOCS is based on a full-electric propulsion design, covering both the electric orbit raising (EOR) transfer phase and the electric station-keeping during on-station operational lifetime. From the AOCS perspective, the EOR is a demanding phase since the electric thrust direction has to be optimized at all times to reduce fuel consumption and EOR phase duration, which can last from 3 to 6 months. One of the main challenges is also the management of high disturbing torques, especially gravity gradient and aerodynamic torques: several days can be spent on low perigee orbits (below 400kms) during the early transfer phase, so as to be compatible with the largest range of existing & future launchers. A simple and innovative solution based on a wheel controlled cruise mode has been developed for that purpose. This wheel cruise mode can be used nominally before the EOR starts, and as a robust backup mode during the EOR. It relies on a specific attitude guidance law coupled with a dedicated Solar Array guidance, which aim at minimizing aerodynamic and gravity gradient torques close to the perigee. A dedicated on- board orbit propagator (OBOP) has also been designed to compute the orientation of the Earth-linked AOCS mission reference frame, in order to reach the required precision for SADM & attitude commanding, which is a key contributor to the mode robustness in the early EOR phase. In addition, Eurostar Neo AOCS has driven the introduction of Airbus Defence and Space auto-coding development process. This process is generic for all Airbus auto-coded products and allows reducing the development time and cost, as well as providing more agility in the AOCS development. It relies in particular on a unique functional architecture which groups the AOCS processing according to their processing type. This paved the way to the concurrent development of AOCS software for both Eurostar Neo full-electric platform and Quantum chemical platform at the same time. Eurostar Neo and Quantum AOCS design is indeed based on common architecture, functions and mode management. AOCS software is being qualified and soon in-orbit for both applications.

Last but not least, the spacecraft autonomy has been increased through simplified FDIR and recovery operations. The principle of AOCS FDIR is a hierarchical failure recovery adapted to the failure detection. The global principle of this hierarchical FDIR is to minimize the number of HW reconfiguration, to maintain telecommunication mission as long as possible, i.e. satellite Earth pointing, and to ensure recovery in a mode where all used equipment units are either monitored or have been reconfigured. The recovery from on station safe mode to nominal mode has been made fully automatic, without mission interruption in most anomaly cases, leading to an easy-to-use platform.

GNC for Future Space Transportations Systems

Robust Control for Reusable Rockets via Structured H-infinity Synthesis Sagliano M1, Tsukamoto T3, Seelbinder D2, Heidecker A2, Macés Hernandéz J2, Farí S2, Woicke S2, Schlotterer M2, Ishimoto S3, Dumont E2 1German Aerospace Center / Japan Aerospace Exploration Agency, 2German Aerospace Center, 3Japan Aerospace Exploration Agency The second decade of the new millennium has led to a complete disruption of the space sector, which has been shaken by the astonishing successes of SpaceX. The company led by Elon Musk has demonstrated that reusability is no longer a chimera pursued since the beginning of the era, but a logical and technological step which is now at our hand. This revolution could lead to an astronaut back to the Moon by 2024 and, even more ambitiously, to manned missions to Mars during the next 10 years. To speed-up the pace of space-missions cost sustainability governmental agencies are now moving with decision towards the reusability paradigm. With this long-term vision in mind and the aim to develop strategic technologies in the frame of a wider reusability-focused program the German Aerospace Center (DLR), the Japan Aerospace Exploration Agency (JAXA), and the French National Centre for Space Studies (CNES) joined in a trilateral agreement to develop and demonstrate the technologies that will be needed for future reusable launch vehicles. In the joint project CALLISTO (Cooperative Action Leading to Launcher Innovation in Stage Toss back Operations) a demonstrator for a reusable vertical take-off, vertical landing rocket, acting as first stage, is developed and built. As long-term objective this project aims at paving the way to develop a rocket that can be reused, and the joint efforts of the three agencies will culminate in a demonstrator that will perform its first flights from the Space Center (KSC), in . Within the trilateral agreement two lines of development of Guidance and Control (G&C) subsystems take place in parallel for CALLISTO. Specifically, DLR and JAXA decided to strengthen their synergy and proceed with the development of a unique, fully integrated G&C subsystem. The missions consist of multiple flight phases, which correspond to different aerodynamic configurations of the vehicle. Specifically, four main phases of flight can be defined to better frame the problem: the ascent phase, the boostback maneuver, the aerodynamic phase, and the powered descent and landing phase. Therefore, a plethora of modern methods is needed to successfully and autonomously complete such an ambitious mission. More specifically, one of the critical aspects to realize an ambitious program focused on reusability is the capability of the system to counteract disturbances and uncertainties acting on the vehicle while satisfying the strict accuracy requirements needed to realize pinpoint landing. This paper focuses on the feedback control strategy conceived for the aerodynamic phase of a reusable rocket, while its twin paper emphasizes the actions taken to control the rocket during the powered descent and landing phase. Each phase has peculiar aspects and different means of actuation. For what regards the aerodynamic phase the fins are used to actively track the desired attitude, which will result in the desired aerodynamic forces required to track the trajectory computed by the guidance system. The controller needs to be extremely robust while satisfying the tight requirements of the mission. Since tracking requirements, measurement noise and disturbances act at different frequencies loop-shaping techniques represent a valid set of tools for this problem. More specifically, a set of controllers is designed with the help of structured H techniques. The benefits of such choice are twofold: on one side the analysis power of frequency-domain optimization techniques coming from the well-established H framework gives important insights about the behavior of the controller. On the other hand, structured H synthesis allows for a a-priori definition of the controller to be tuned, having to the designer flexibility in terms of control structure complexity and performance. The synthesized controller will exhibit, in the limits of the possibilities of the chosen control structure, features like the ones showed by a full-order H technique. The paper is divided into an analysis and a synthesis part. In the former the baseline PD (proportional- derivative) controller is analyzed according to the classical stability metrics: gain margin, phase margin, and delay margin. Moreover, the parametric control law is analyzed both in frequency and time domain, to completely characterize its behavior. In the attempt to establish a parallel for descent scenarios to the literature available for launchers we propose a reconstruction of the baseline controller using the Hinfstruct framework. The recovery of the legacy controller through the framework allows the follow-up synthesis based on the same framework. The second part of the paper focuses on the synthesis of a robust controller for aerodynamically controlled reusable rockets: specifications to perform mixed sensitivity analysis are discussed, their connection to classical stability margin indicators established, and a full H controller is synthesized. A special care is given to the plant representing the attitude dynamics, where pairs of poles on the imaginary axis require special treatment. The synthesis is then moved towards the structured form, where we consider the trade-off between the degradation of performance due to the lack of degree of freedom of the controller, and the desirable simpler structure chosen for the structured synthesis. The achieved trade-off is described and motivated, together with the validation of the results for both the linear and the nonlinear plant. Fig. 1 shows the linearized system in 10 different points along the nominal trajectory. These points are used to approximate the error dynamics of the system as a set of different Linear Time Invariant (LTI) systems. As examples in the paper show the frequency domain performance for the baseline controller in terms of complementary sensitivity and sensitivity transfer function, respectively and the corresponding step response. The baseline is compared with Hinf techniques and results are compared in both frequency and time domain. The reconstruction of the baseline controller through Hinf techniques is given in frequency and time domain. While the baseline controller strategy is stable, robust stability and performance are not satisfying. The improvement of performance and stability in the frame of robust techniques is the focus of the paper. Design and testing of the GNC for the HERACLES Lunar Ascent Element Peters T1, Briz J1, Arroz P1, Duarte P1, Cometto F2, Berga M2, Perez J3, Cuffolo A3, Cropp A4 1Gmv, 2TAS-I, 3TAS-F, 4ESA-ESTEC This paper describes the design and testing of the GNC for the HERACLES Lunar Ascent Element (LAE). The LAE needs to carry samples from the lunar surface to the Lunar Orbital Platform – Gateway (LOP-G) from where the samples are eventually transported back to Earth. The GNC design integrates the GNC for the launch and ascent, the orbit transfer manoeuvres and rendezvous in a near-rectilinear halo orbit (NRHO) into a single simulator. The main focus of this paper is on the consolidated Monte Carlo simulation test campaign results. A discussion of the results and an analysis of the dependence of the results on the assumptions made to define the scenario is performed. The results show that the current design for the GNC is feasible and that the HERACLES LAE mission to the LOP-G can be successfully performed. Dynamic Modelling and Control of an Aerodynamically Controlled Capturing Device for "In-Air-Capturing" of a Reusable Launch Vehicle Singh S1, Stappert S1, Buckingham S3, Lopes S3, Kucukosman Y3, Simioana M2, Pripasu M2, Wiegand A2, Sippel M1, Planquart P3 1German Aerospace Center (DLR), 2Astos Solutions GmbH, 3Von Karman Institute for Fluid Dynamics VKI The 21st century has witnessed the development of multiple reusable launch systems. While the most successful launch vehicles by SpaceX are designed to Return To Launch Site and to perform Down-Range Landing, they require a significant quantity of fuel to perform landing. Further, winged stages require an additional propulsion system to perform horizontal landing, which also adds to the stage mass. An innovative approach patented as “In-air-capturing (IAC)” [1], provides the possibility of returning a winged stage back to the launch site without the need for a propulsion system for this phase. After a standard lift-off, the winged booster stage separates at Main Engine Cut-Off and re-enters the atmosphere in a ballistic trajectory. At an altitude of about 20 km, it decelerates to subsonic velocity and follows a gliding path. Somewhere between 8000 m to 2000 m, the stage is awaited by a large capturing aircraft, which captures the gliding stage and tows it back to the landing site [2].

For the capturing process, an Aerodynamically Controlled Capturing Device (ACCD) was found to be the most promising option [3,4]. The device is released when the two main crafts are descending close to each other in a parallel formation. The device does not have a propulsion system of its own but its position and orientation are strongly correlated to the aircraft and the towing rope connecting the two bodies. With a small capture window of about two minutes [5], a robust yet efficient controller would be required to actively control the ACCD. To achieve this, the complex dynamics associated with the interactions between the aerodynamics, flexibility of towing rope and control must be studied.

For the current study, ACCD consists of a main cylindrical body with a boat tail and four flaps to control its orientation and position behind the towing aircraft. The body roughly measures 2 m x 1.5 m (with symmetry in X-Z plane), and houses a capturing mechanism to enable safe docking with the returning stage [6]. Paper shows the aerodynamic pressure distribution of the ACCD at 0° angle of attack (AoA), and the gradient of pitch moment coefficient to AoA. The Center of Gravity position was estimated to be 0.75 m from the nose (subjected to change with further analysis). In this configuration, the ACCD was found to be stable while still allowing for sufficient controllability around all axes.

The current research is conducted under a Horizon 2020 project named FALCon (Formation flight for in-Air Launcher 1st stage Capturing demonstration). The project aims at accelerating the development of IAC technology with a combined effort from multiple partners across Europe. A realistic full-scale simulation of IAC will require detailed modelling of a number of aspects. Paper shows a preliminary architecture for the simulation with the role of each partner indicated by a unique color. It can be seen that the model consists of three main bodies, namely, the towing aircraft, the ACCD and the RLV, along with a flight controller. A more detailed overview is explained below: 1. The towing aircraft consists of a guidance trajectory, which influences the motion of the ACCD. Another important aspect that cannot be overlooked is the 3D flow field originating from the wake of the towing aircraft. The turbulence originating from the wake can cause significant perturbations in the rope dynamics, which impacts the attitude of the ACCD. A mean velocity and turbulence intensity map is used to model these effects, which are treated as external disturbances to the ACCD model. 2. The ACCD consists of detailed modelling of a number of parameters affecting its attitude. Apart from the environment, kinematics and dynamics associated with the body, external perturbations originating from the flexible rope are included using a multibody approach. The rope is modelled as a flexible body with up to 20 rigid segments connected by spring-damper elements. The associated aerodynamics with the control flaps are also accounted for in this model. 3. The RLV, which is currently assumed to follow a predefined path, consists of visual sensors (or cameras) to navigate the position of the ACCD. During the capture window, a set of two sensors will be able to provide the position and orientation of the ACCD with respect to the RLV. 4. The controller block will compare the absolute position of the ACCD against the relative position w.r.t the RLV to determine the necessary flap deflections. A basic PID controller and a self-tuning fuzzy PID controller will both be used to study the controllability of the system. Self-tuning fuzzy PID controller can show superior performances in the presence of uncertain dynamics and external disturbances, for applications like pitch control and trajectory tracking [7].

Hence, the final draft of the paper will demonstrate the dynamics and control characteristics of the ACCD performing “In-Air-Capturing” of a returning stage in the presence of multiple external perturbations.

[1] Patentschrift (patent specification) DE 101 47 144 C1, Verfahren zum Bergen einer Stufe eines mehrstufigen Raumtransportsystems, released 2003. [2] Stappert, S.; Wilken, J.; Bussler, L; Sippel, M.: A Systematic Comparison of Reusable First Stage Return Options, 8th EUROPEAN CONFERENCE FOR AERONAUTICS AND SPACE SCIENCES (EUCASS), 2019. [3] Sippel, M., Klevanski, J.: Progresses in Simulating the Advanced In-Air-Capturing Method, 5th International Conference on Launcher Technology, Missions, Control and Avionics, S15.2, Madrid, November 2003 [4] Sippel, M.; Klevanski, J.: Simulation of Dynamic Control Environments of the In-Air-Capturing Mechanism, 6th International Symposium on Launcher Technology 2005 [5] Sippel, M.; Bussler, L; Krause, S.; Cain, S.; Stappert, S.: Bringing Highly Efficient RLV-Return Mode “In-Air-Capturing” to Reality, HiSST 2018-1580867, 1st HiSST: International Conference on HighSpeed Vehicle Science Technology, Moscow, November 2018 [6] Stappert, S. : Aerodynamics of the ACCD for the In-Air-Capturing of the SpaceLiner Booster, SART TN-005/2006, December 2018. [7] Wahid, Nurbaiti, and Nurhaffizah Hassan. "Self-tuning fuzzy PID controller design for aircraft pitch control." 2012 Third International Conference on Intelligent Systems Modelling and Simulation. IEEE, 2012.

Trends in AI for GNC Systems

Pose estimation of a non-cooperative target based on silhouette imagery using convolutional neural networks Bettens A1,2, Comellini A1,3, Zenou E1, Dubanchet V3 1ISAE SUPAERO, 2The University of Sydney, 3Thales Alenia Space The current article proposes the use of Convolution Neural Network (CNN) to determine pose estimation of a non-cooperative spacecraft from silhouette images. In the scope of space rendezvous and close proximity operations with non-cooperative targets, the chaser (or servicer) can rely on vision-based navigation to obtain the measurement of the targets (or clients) relative pose (i.e. attitude and position). The pose acquisition (or initialization) problem in vision-based navigation consists of determining the 6DOF pose of an object without any prior knowledge of its position and orientation. The pose acquisition phase is needed to initialize the tracking algorithm and the navigation filter used by the chaser to estimate the target state. Tracking algorithms, on the other hand, usually rely on a local search around the estimated pose at the previous instants to compute the updated pose and on frame-to-frame features tracking.

In the case of a non-cooperative and non-prepared spacecraft, such as space debris, the acquisition algorithms cannot rely on visual aids such as markers or fiducials to help the pose initialization. The only a-priori information on which the acquisition and tracking algorithms can rely is the 3D geometrical model of the target. Model based acquisition algorithms, such as template matching, compare a currently observed estimate with a library of model templates to determine the current pose. However, the large number of templates required tends to make these algorithms slow, even when aided by a hierarchical pose tree to cluster object pose estimates and reduce the template resolutions. Alternatively, the use of LIDAR or cameras can provide a point cloud which can be used and processed to generate a model to initialize the pose tracking. Nonetheless, this approach is very computationally intensive as it requires multiple iterations on large datasets of Cartesian points. For this reason, the research industry has recently turned its attention towards deep learning techniques such as CNN’s as a means of model-based pose estimation based on monocular images: once trained, a CNN has a relatively low computational cost and can rapidly generate a pose estimate, at the expense of a specific training on ground to tune it with respect to the considered target. As long as a ground link is available between the chaser and the control center, this approach can even allow refining the pose estimation during the target inspection using the real images captured in space to re-train the CNN on ground before updating on board.

Deep learning neural networks have the ability to extract relevant features and a model can be trained to recognize patterns, rather than being reliant on a structure. A CNN performs mathematical image recognition operations using local receptive fields to create a feature map in the hidden layers. CNN's have shared weights and bias, meaning that each neuron in a given layer is trying to detect the same thing in an image, just in a different region of that image. This makes CNN's tolerant to translation and hence, good for pose estimation.

In-space visual sensing presents challenges, such as dealing with varying illumination or poor illumination, low Signal to Noise Ratio (SNR) and high contrast environments. Data fusion from multiple sensors provides a mean of coping with these problems. In space environments, multispectral imagery can provide robust image segmentation and extraction of the target silhouette. The fusion of data from the visible, the thermal infrared, and the near ultraviolet bands allows the proper extraction of the satellite silhouette regardless of the illumination conditions or of the presence of Earth and clouds in the background.

An original data set of silhouetted satellite images is generated to provide an accurate training data set, using Thales Alenia Space high fidelity image generator SPICAM. These silhouette images are then used to refine a regression-based CNN model. Image pre-processing in the form of sensor noise, translation and resizing expanded this data set to provide an adequate number of training images to prevent overfitting of the CNN. The knowledge of the spacecraft geometrical model without the need to render textures allows for fast generation of multi-observation synthetic data sets of a spacecraft to train the CNN. Furthermore, this provides a more robust method to prevent overfitting as the model is not learning textures present within an image.

Training a CNN on silhouette images has the advantage of providing a model invariant to differences in feature appearance and textures, and therefore to illumination conditions. Extracting a silhouette of the target satellite for use in a CNN provides a novel approach to pose estimation of non-cooperative targets and remains compatible with the high constraints on the space processor capabilities. Using a regression-based approach to a CNN framework allows for numerical attribution of estimates, where the network is capable of providing an output based on a 2D image. Regression analysis in a CNN makes it possible to solve a 6DOF pose estimation problem, without limiting the number of categorical discretizations. Furthermore, image pre- processing led to the generation of an original data set, large enough for training a CNN to perform accurate machine learning.

The current paper demonstrates that a CNN is capable of extracting a pose estimate of a spacecraft from silhouette imagery, offering a novel approach to pose estimation of a spacecraft for in-orbit-servicing and active debris removal applications. The trained CNN model has the potential to use silhouette images that are robust against illuminations, for onboard pose estimation of non-cooperative targets compatible with hardware and equipment constraints.

Using Convolutional Neural Networks for Relative Pose Estimation of a Non-Cooperative Spacecraft with Thermal Infrared Imagery Hogan M1, Ronao D2, Aouf N1, Dubois-Matra O3 1City, University Of London, 2Centre for Electronic Warfare, Information and Cyber, Cranfield University, 3European Space Agency, ESTEC I. INTRODUCTION Autonomous rendezvous and docking missions such as, ETS-VII and Orbital Express, have had the advantage of two cooperative spacecraft which are predetermined to be compatible. Markers and additional sensors on the target spacecraft, and a physical interface to grapple were used to ensures a safe rendezvous. However, recent interests have shifted to technologies for use in situations where the target is unresponsive such as Active Debris Removal missions. The lack of aid from the target puts higher responsibility on the chaser’s navigation, guidance, and control systems. To the author’s knowledge such a manoeuvre has only been achieved once with the capture of and relaunch of the IntelSat VI satellite which required human intervention for success. Passive sensors have an advantage in terms of smaller size and lower power consumption over active sensors such as LIDAR. However, visible light cameras struggle to match the reliability of continuous measurements of active sensors due to the poor illumination conditions in eclipse and oversaturation from reflective surfaces or direct light from the sun. The solution in this paper takes advantage of thermal infrared images which are less affected by ambient light conditions in order to provide a good alternative for space vehicle relative navigation. A deep Convolutional Neural Network (CNN) is developed to provide an initial pose estimate of the non-cooperative target for space rendezvous missions.

II. METHODOLOGY a. Background Recent studies suggest that Long Wave infrared based cameras offer a promising alternative to vision based or active technologies for non-cooperative rendezvous missions. However, the lack of relevant infrared images for training causes an immediate barrier to the development of IR-CNNs. Therefore, this study will examine the usability of a CNN that has been trained using predominantly visible light images to classify Infrared images. Any image processing technique that utilises model-based pose estimation must consider the possibility that the non-cooperative target may be damaged, may contain movable sections or that the training data might contain incorrect dimensions or be missing features. Therefore, it is very important that the CNN is good at generalizing features of the model. b. Our Approach The CNN adopted in this study is an adapted version of ResNet. This network has demonstrated good generalization from winning many competitions in image recognition and classifications. It uses skip connections between layers in order to avoid the issue of vanishing gradients which occurs when training networks using gradient descent algorithms. Current experiments have used the 18-layer version of Resnet as it has been found to converge faster than other 18-layer nets. Deeper networks are yet to be investigated for higher accuracy. In order to assist in training, Resnet utilizes batch normalization right after a convolution layer and before an activation layer. This helps the gradient to remain consistent and not get so large as it will slow down the network or prevent it from training. Two dropout layers are currently being employed in the fully connected layer. These turn off parts of the network to allow other nodes to train and to stops the nodes becoming co- dependent which will cause overfitting. Synthetic visible images of the non-cooperative target are used to create the training set. This dataset can be expanded further with augmented images with techniques such as colour jitter to randomise the brightness, contrast, saturation and hue of the images. This method will also improve the CNN’s ability to generalise the features of the target. After the synthesis of the images for the training set, the images need to be categorised into classes that best represents its relative coarse pose in relation to the chaser. Once trained the net should be able to correctly classify an image it had been supplied from a test set. This test set will initially be made up of synthetic IR images. The difference of modality between the training data and the test data is a difficult challenge that we use in this study to evaluate the performance and the robustness of our deep coarse relative pose estimation solution.

III. PRELIMINARY FINDINGS Initial testing using synthetic images of Envisat which were constructed by the Astos Camera Simulator show promising results. Using only generated visible images in the training data, the trained network was able to correctly identify 85% of the images in the test set which was composed entirely of generated LWIR images.

IV. Conclusion This paper presents a novel counter to the responsibilities placed on chaser spacecraft during approach of a non-cooperative target. Based on a novel multimodal Visible/Infrared deep network algorithm. It has the potential to provide continuous, fast pose estimates which could be used as a coarse pose system initialiser. Thus, eliminating the dependence on using an estimate from the inertial navigation system resulting in a truly autonomous system.

An unsupervised deep learning approach to on-board AOCS FDI(R), first results and conclusions Scharf A1, Murray C1, Hervas Garcia C1, Thomas D2 1Airbus Defence And Space Ltd, 2Airbus Defence And Space SAS Fault detection and diagnosis is an important problem in spacecraft operations and a critical aspect of on- board software with respect to safety, performance and reliability. In particular, the on-board Failure Detection Isolation and Recovery (FDIR) function is a key element within the Functional Avionics chain of any modern spacecraft. Ideally, the on-board FDIR shall monitor the spacecraft, identify when a fault has occurred, determine its type and its location (e.g. pinpointing the failed equipment) and trigger the necessary recovery action(s). Current on-board FDIR practices focus on the functional surveillances of spacecraft vital state (e.g. power, attitude / rates, temperature) and are limited with regard to the failure cause localisation. This limitation can result in undesirable and unnecessary interruption of the spacecraft mission/science, costly and overly pessimistic HW redundancy schemes or increased operational costs for the spacecraft/fleet. Besides, as highly ambitious new space missions are proposed, more demanding objectives of performance accuracy, availability and/or autonomy requirements arise, leading to increasingly complex spacecraft designs. Similarly, with the arrival of large spacecraft fleet, the exponential growth in number of spacecraft to operate translates into prohibitive volumes of telemetry data to process for on-ground diagnosis. Furthermore, interplanetary spacecraft where important telemetry time latency and/or periods with no communication from/back to Earth apply make it difficult for the operators to respond to any failure or recover from a Safe Mode in a short time. The on-board FDIR function is an enabler for all these cases and thus its functionality and performance will need to improve drastically. In this context, new techniques for anomaly/fault detection, isolation and recovery together with an increase in the computational power available on-board the spacecraft make it viable to transfer many of the FDIR functionality from the ground to the spacecraft itself to reduce the need for intervention from operators and thus the overall operational cost: increased autonomy and availability. In particular, the advent of Artificial Intelligence and the latest developments on its Machine Learning branch have proven outstanding performance on anomaly detection for a wide range of problems (e.g. finance). This paper presents the main results of “SMART-FDIR”, an innovation project at Airbus Defence and Space that has prototyped a generic and reusable deep learning approach for both anomaly detection and isolation for an AOCS-FDIR use case and benchmarked the solution against “classical” FDIR showing promising results. Through a novel "unsupervised adversarial learning” approach named MODISAN (MOdification DIScrimination Adversarial Networks), deep neural nets are trained on nominal telemetry for different equipment/channels using high-fidelity simulated data extracted during performance validation test campaigns. Thanks to the unsupervised nature of the technique no prior real or simulated failure data is needed to train the deep learning model. Moreover, no prior assumptions on the type of failures or failure modes of the equipment are needed. Thus, in the MODISAN approach the discriminator(s) are simply trained on a sufficiently representative dataset of nominal telemetry covering the envelope of nominal and acceptable degraded operations. Through the training process the discriminators are capable of learning an efficient representation of the envelope of nominal behaviour. The paper is structured as follows: A brief introduction to previous attempts to enhance the FDI(R) and health monitoring function, both on ground and on-board, through AI techniques, with a special focus on recent Machine Learning and, in particular, deep learning techniques. After a discussion of the pros and cons of different approaches, the paper introduces MODISAN and its main advantages: • Able to cope efficiently with all types of anomaly (e.g. point, contextual or collective) • Avoids further engineering of triggering logic • Provides means for isolation • Generic approach with little or no feature engineering involved

The reader is introduced to the main use cases where MODISAN has been prototyped and benchmarked against classical AOCS-FDIR techniques: FSS-GYR and FSS-GYR-CPS-ACC in the frame of Solar Orbiter. This includes a brief introduction to classical FDIR design and its main challenges and is followed by a definition of clear benchmarking criteria, including the definition of the failure set for the benchmark. Then “classical” versus “SMART-FDIR” benchmark results are presented, providing insight on the behaviour of key metrics: detectability, reactivity, isolation capability, false positive rate and, when possible, additionally derived system performance (e.g. safety, availability or propellant saved).

Some words will be dedicated to SMART-FDIR through the Functional Avionics chain and its proposed auto-coding process for neural nets: from Tensorflow to final OBSW validated code.

Finally, the paper provides conclusions of the work done so far and outlines the next steps to follow in order to see this solution flying soon.

Advances in Control

Sloshing AOCS/Fluidic coupled analysis Manuel Juanpere X1, Laurens P1, Regnier P1, Dalmon A1, Bvestrello H1, Levenhagen J2 1Airbus Defence and Space, 2Airbus Defence and Space Current complex missions’ requirements demand to equip spacecraft with huge tanks of propellant. Whatever the type of mission (science, telecommunications, observation…) the propellant sloshing is a major attitude disturbance which can lead to a degradation of the mission performances.

When liquid sloshing disturbance presents very low frequency content, the disturbance is observed and controlled by the AOCS in the limits of its actuation capacity: with a good design, one can expect that they are not an issue for the mission. When they present high-frequency content (e.g. transients during attitude or orbit control manoeuvres), the disturbance can be significantly above the control bandwidth: in this case the disturbance is not controlled thus the pointing performances can be degraded. The latter is simply covered by linear analyses or in simulations by feeding the model with the CFD in open-loop: such strategy is quite typical and has been used in several projects, as for instance in MetOp Second Generation (MOS).

Between both extreme points, there is a range where the liquid sloshing and the control can interact since the frequency content of the first mode and the control bandwidth are closer. In this case, control robustness is no longer ensured even if linear stability analysis can be performed to check at the first order that the controller is robust.

Traditional methods to model liquid sloshing in such analysis are based on mechanical models: pendulum or spring-mass. Pendulum permits to capture the stationary behaviour of the liquid when the spacecraft is submitted to acceleration; spring-mass is rather used in micro-gravity conditions, but no one of both is representative of the true behaviour of the liquid. Even if they permit to cover at the first order the robustness approach, their lack of representativeness affects the temporal analysis thus the complete AOCS performance evaluation.

To tackle this problematic, Airbus Defence and Space has prepared a roadmap to improve the mastery of fluids dynamics and its impacts on AOCS. After the successful ISS experience called Fluidics (CNES and Airbus Defence and Space, performed in 2017 by the French Astronaut Thomas Pesquet), CFD models have been improved and liquid dynamics better understood. With that in hand, a co-engineering work between AOCS and Fluidics teams has permitted to propose a flexible solution to extend AOCS simulation capacities to include liquid sloshing. One of the first results of this roadmap is a simple Model-Based generic interface which can be included in the AOCS simulator to model liquid sloshing disturbance using in-the-loop CFD software results (as illustrated by the block-diagram, please see the attached image).

To achieve this functionality, several difficulties have been overcome. At the end of the model setting up, the physics of the coupling has been cautiously validated through several reference cases. Even if liquids physics is never exactly caught by CFD models, these successful tests permit today to trust on the results of coupled AOCS/CFD simulations.

The first industrial application of this tool has been applied to ESA JUICE mission. The objective was to validate the AOCS strategy for the critical trajectory manoeuvres (Jupiter and Ganymede Orbit Insertion, Ganymede orbit circularization). During this phase, the Main Engine Boost Manoeuvre (MEBM) generates such forces that the solar panels, which are deployed at this stage of the mission, have a significant deflection. In case of FDIR triggering, the MEBM is immediately stopped; because of the criticality of this phase, it is required to resume the MEBM as fast as possible in order to ensure the JOI. Nevertheless, without MEBM thrust, the solar panels start to exchange all their deformation energy with the Satellite central body. Resonance frequencies are close to sloshing frequencies and close to AOCS bandwidth. A strategy to actively damp as fast as possible the solar arrays Out-Of-Plane flexible mode has been setup; even if the theoretical analysis based on an energetic approach proved the robustness of the strategy, the amount of liquids inside the tanks is significant and its interaction with the solar panels flexibilities was uncertain in terms of potential coupling. Fully representative temporal simulations at various critical tank filling ratios were necessary to finally confirm that the strategy would ensure the success of the critical manoeuvres in case of MEBM failure.

For that, the fluidic model has been introduced in the high representative simulator dedicated to JUICE. The operational sequence following the MEBM failure has been run coupled to the CFD software. The results have shown that the proposed design is effectively robust. The efficiency of the co-engineering task realized by the Fluidics and AOCS teams has proved the full-industrialisation of the approach.

The first application has confirmed that the tool is mature enough to be used in further applications and several projects are already interested on it.

This success is an intermediate step paving the way on the Airbus Defence and Space roadmap. Co- engineering is still ongoing in order to generalize simple inclusion of representative liquid models in AOCS analysis. For that, next steps target to find a generic and representative model permitting to drastically reduce the necessary amount of computation thus the necessary time to run each simulation. The Sloshing roadmap targets to in the TRL of such models with the industrialisation of a light model permitting to take into account liquid sloshing in any AOCS temporal analysis in a generic way.

Development of a robust GNC architecture for a flexible spinning spacecraft with long wire booms Passarin F1, Hervas Garcia C1, Cantiello I1, Whittle L2, Falcoz A2, Hyslop A3, Preda V4, Girouart B4 1Airbus Defence And Space Ltd., 2Airbus Defence and Space Sas, 3Vitrociset – A Leonardo Company (for ESA), 4European Space Agency – ESTEC Experiments involving extremely large flexible wire booms attached to a spinning body are of great interest to the science community. This is the case for ESA’s Cluster mission which uses a formation of spinning spacecraft equipped with wire booms in order to study the Earth’s magnetosphere. Similarly, NASA’s Magnetospheric Multiscale Mission (MMS), the Van Allen Probes, Polar, FAST and Themis employ the same kind of spacecraft. From an AOCS standpoint, the resulting dynamic system to control is challenging and, so far, all solutions have relied on the passive damping of wire boom oscillations. However, flexible spinner missions up to now have never required slow spin rates to perform their science. Low spin rate translates into lower flexible boom frequencies, which in turn lead to slower passive damping of the flexible oscillations. The latter implies that oscillation transients have a longer duration, putting the science availability of the wire booms at stake. The challenge thus is to be able to actively damp these oscillations and therefore make low-rate spinners a more effective option for the scientific community. In this context, this paper presents the dynamic modelling, GNC design and verification process followed for the development of a robust GNC architecture for a flexible spinning spacecraft in the framework of the ESA R&D TRP study “Robust Attitude Guidance and Control for Flexible Spacecraft”. The use case under investigation is a based on ESA’s mission, which is spinning at 12 deg/s with four 50 m long radial (spin-plane) wire booms and two radial booms. In order to account for the significant uncertainties and satisfy the specified requirements in the presence of disturbances while guaranteeing fast convergence, the use of advanced robust control and guidance techniques is paramount to the successful design of a high-performing GNC architecture. As a first step, the full nonlinear model of the spacecraft consisting of hub, four wire booms, two magnetometer booms and fuel tanks has been derived from the 6 degrees of freedom Newton-Euler equations. Sensors consist of a gyroscope and a three-headed star tracker, while the actuators used are a typical reaction control system (RCS). The resulting model has been validated with simulations against another model generated using the in-house Airbus dynamics engine DYCEMO, demonstrating strong correlation for a wide range of input commands. A symbolic linearisation has been performed and validated against a numerical linearisation of the dynamics engine both in the frequency and time domain, and against the full nonlinear simulator. Compared to the results obtained so far in literature for similar spacecraft, the derived linear model has the added benefit of not relying on a-priori decoupling of in-plane/out-of-plane dynamics and also includes centre of mass offsets and inertia cross products, which are crucial elements for the problem in question. An extensive observability and controllability analysis for the modes of oscillation of the wire booms has been performed. The low observability of the wire boom modes demanded a control architecture that includes wire boom deflection sensors and central-hub actuation in force and torque to dampen wire boom oscillations to the required level. In order to guarantee closed-loop stability and performance in the presence of system uncertainty, an uncertain system model in Linear Fractional Transformation (LFT) form is derived and validated. The LFT is used both during the controller synthesis phase as well as for stability and worst-case analysis. This ensures that the resultant design is robust to the large, inherent uncertainties within the system, whilst also meeting stringent and demanding performance constraints. To achieve this, the design requirements of the mission have been translated into corresponding frequency domain requirements on the closed-loop sensitivity functions. Taking into account the fundamental limits of the feedback system, the sensitivity goals are included as frequency domain weights in the robust synthesis problem. The control design problem is subsequently transformed into a worst-case gain optimisation problem and minimised using structured H infinity tools. Two independent robust control designs have been optimised and compared with a standard proportional rate controller, demonstrating the advantages of the proposed architecture with respect to a more classical solution: the standard rate controller is not able to meet all requirements specified for the mission, while the proposed settling architectures are able to dampen all oscillations to the specified level within the required time. Advanced guidance techniques have also been developed and implemented with the aim of reducing excitation of uncontrollable modes and the amplitude of oscillations resulting from a manoeuver such that the total manoeuver plus tranquilisation time and the total fuel usage are minimised. In particular, both closed loop and open loop guidance techniques have been developed and implemented in two different solutions. This paper also presents the verification results based on worst-case analysis (both non-linear and linear mu- analysis). The end goal of these procedures is to establish bounds on the robust stability and robust performance margins or to find parameter combinations that lead to worst-case performance. By performing both verification techniques, it is possible to demonstrate with high confidence that the GNC architectures can satisfy performance and stability requirements. Thrusters Off-modulation for Attitude Control during Orbital Manoeuvers Cederna L1, Paolini E2, Celiberti C3, Biggs J1 1Politecnico Di Milano, 2OHB Italia, 3OHB Italia LEO satellites are often required to perform frequent manoeuvres for orbit maintenance. Depending on the altitude, the duration of these manoeuvres can last up to 300 seconds. During the firing, the offset of the center of mass displacement from the center of geometry, uncertainties in the thrusters’ accuracy and misalignments will generate parasitic torques. In addition, high precision requirements on pointing do not allow the employment of magneto torquers, therefore the attitude is generally undertaken with reaction wheels. Consequently the reaction wheels must provide additional torque to compensate for these uncertain disturbances and, considering the size constraints together with the aforementioned firing times, they may reach saturation. However, by employing off-modulation, it is possible to guarantee longer firing times with the same reaction wheel configuration while avoiding saturation. This approach consists of controlling the spacecraft‘s attitude during the firing by switching off alternately one (or more) thrusters. In this way the satellite does not require any longer to divide one single orbital manoeuver into multiple shorter ones and, therefore, it acquires higher availability (more time can be dedicated to the mission).

The aim of this research is to develop a control algorithm capable of manoeuvring the satellite through an off- modulation of the thrusters. In addition this paper will demonstrate its compliance with respect to high precision requirements. It is shown that is possible to employ the off-modulation to efficiently compensate for the uncertain parasitic torque. This strategy improves manoeuver duration, and thus, from an operational point of view, maximizes the available time envelope for operations. The analyzed mission is characterized by 4 hydrazine thrusters arranged in a cross configuration with thrust axes normal to the mounting surface. With a 4 tilted thruster configuration it is possible to modulate the firing to desaturate the wheels. The off- modulation presented here is able to guarantee the same outcome, with a non-tilted thrusters configuration, avoiding wheel saturation, and can also operate when using only one couple of thrusters.

Thrusters’ technological restrictions on duty cycles do not allow the switch-off time duration and frequency to be chosen arbitrarily. They are permitted to work only at a pre-defined frequency and with a restricted set of possible choices of the off time (which should be the lowest possible in order to limit the effect on the delivered ΔV) and, as a consequence, they can generate only limited values of momentum. Moreover, since it is not possible to predict in advance the off-modulation duration, the new input for orbital manoeuvers is no longer the firing time but the desired ΔV. The evaluation of the delivered thrust is computed on board and the manoeuver is considered complete when this quantity reaches the required value. Off-modulation can be regarded as a hybrid control, where reaction wheels and thrusters work simultaneously to provide the required torque. In particular, thrusters generate discrete quantities of momentum reducing the load on reaction wheels, whereas the reaction wheels, capable of finer control, are exploited to reach the correct precision. The orbit control procedure is divided into two sub-states: the first one is performed to achieve the correct orientation both before the firing starts and after the firing stops (using only reaction wheels). The second one, instead, takes place simultaneously to the thruster firing and exploits the off-modulation to control the satellite.

This paper investigates in details the parasitic torques affecting the thrusters, identifying which disturbances are the most influential in reducing the occurrence of reaction wheels saturation. Subsequently the off- modulation algorithm is described in detail, including its validation and verification for both the nominal case and during state transition. The results of the simulations are presented together with an analysis of the excitation of deployable solar panels during the off-modulated firing. In the last section a Montecarlo analysis is performed in order to provide a rigorous investigation of the algorithm behaviour and to verify the controller robustness.

EVALUATION OF NONLINEAR MODEL PREDICTIVE CONTROL FOR CUBESAT ATTITUDE CONTROL; A HARDWARE-IN-THE-LOOP SIMULATION Bromose L1, Petersen J2, Kaas K1, Nielsen J2 1Space Inventor, 2Department of Electronic Systems, Aalborg University ABSTRACT Over the past years improvement to imaging technology has increased and likewise for the interest in using nanosatellites (e.g. CubeSats) for earth observations. This drives the need for star trackers. Using a star tracker poses constraints on the alignment of the boresight of the star tracker and the sun given the sensitivity to light exposure.

Model Predictive Control (MPC) allows for constraints, however, computational time has long been the issue for applications employing relatively small On-Board Computers (OBCs); especially considering optimal control problems employing nonlinear dynamics. With both developments in tooling and availability of processing power on small OBCs, MPC is enabled for smaller applications.

In this paper we evaluate the feasibility of MPC, employing a model with nonlinear dynamics and a cone- constraint, which can be used to ensure that a star tracker is not damaged by exposure to the sun. Using state- of-the-art tooling, a Hardware-in-the-Loop (HIL) simulation is developed, with a simulation environment feeding virtual sensor readings into an OBC, running a compiled MPC application. The evaluation shows that it is feasible to run MPC with nonlinear dynamics and cone-constraints on modern low power devices. Worst case computation times on a ARM Cortex-A53 (BCM2837) was found to be in the range of 250 ms during a sun avoidance maneuver. In the non-constrained region the computation time was in the range of 25 ms. The memory footprint of the MPC is 14.2 kB of ram.

1. INTRODUCTION The topic of slew maneuver control is treated in [WK05], where an optimal torque distribution controller is designed, based on a energy shaping method. In the papers treated in [EPK+17], it is concluded that the argument of computational overhead against using nonlinear MPC (NMPC) for small-scale aerospace applications no longer holds. This is also the conclusion in [LGK+17], where NMPC for constrained attitude maneuvering is designed and simulated, showing the potential for implementation on embedded platforms.

2. METHODS We propose a NMPC, using a quaternion-based model as described in [Yan12], employing soft-constraints to ensure that the spacecraft does not align the boresight of the star tracker with the sun, with a constrained margin of $\pm\SI{30}{\degree}$. The MPC computes control input (torques) that minimize tracking errors over a prediction horizon of 5 seconds. Constraints on rotational velocities and input torques are also enforced.

We employ CasADi, as presented in [AGH+19], for the purpose of describing the dynamics, posing and solving a multiple-shooting optimal control problem. CasADi enables code-generation, and the developed MPC has been compiled for a OBC (ARM Cortex-A53 (BCM2837), Raspberry Pi 3).

3. SIMULATION SETUP A space environment simulator, implemented in MATLAB / Simulink, is used for the HIL simulations. This simulator is based on the work presented in [JV05]. The simulator is built with the assumption of a 1 Hz controller.

The simulator sends virtual sensor readings over an TCP/IP connection to the OBC. The OBC then invokes the MPC algorithm, and the control inputs are returned the over same connection.

4. SIMULATION RESULTS A simulation was performed with an 10:30 Sun-Synchronous Orbit in order to investigate the behavior of the controller during an avoidance maneuver. The controller was setup to follow the Local Vertical Local Horizontal (LVLH) frame.

The tracking error is shown along with the angle from the star tracker to the sun. φ, θ, ψ denotes the roll, pitch and yaw axis, respectly. While the spacecraft is outside the constraint, it is tracking the LVLH frame very well. As the sun-angle is closing in the on the 30⁰ constraint, the controller is tracking the LVLH frame with best-effort, given that is does not violate the sun-angle constraint.

As the focus of this paper is to consider the feasibility of running an MPC on an OBC,the execution time and memory footprint of the controller is of interest. A correlation between the execution time and the sun-angle was found. This is expected as the optimization problem in the MPC formulation becomes harder when the sun-angle constraint is active. We find it likely that tuning solver setttings and investigating the formulation of key nonlinear dynamics could minimize the time jump, compared to the non-constrained execution times. The memory footprint of the MPC is 14.2 kB, the memory usage are guaranteed, as no dynamic allocation are performed in the generated code [AGH+19].

As a preliminary result, it is clear that the MPC implementation is feasible, as the maximum observed execution time in the range of 250 ms -- this is during the avoidance maneuver, with active sun-angle constraint. For all other cases, the execution is in the order of 25 ms.

The MPC is equally minimizing the axis-wise error angles, but it could easily be adapted to minmize the the boresigth error instead. This is of interest as many payloads are more sensitive to the boresigth error than the axis-wise error angles.

We have proven the feasibility of implementing a MPC with nonlinear dynamics and cone-constraints on an OBC. However, future work involves posing and analyzing stability and robustness, in order to guarantee performance.

NMPC Based Guidance and Control for Earth Observation Missions Pagone M1, Boggio M1, Novara C1, Massotti L2, Vidano S1 1Politecnico di Torino, 2ESA/ESTEC Autonomous mission planning and, in particular, autonomous guidance and control are at present important open problems in the context of advanced space missions. The development of strategies to deal with these issues would be fundamental in a wide range of space applications, such as low-thrust orbital transfer, orbit phasing, deep space maneuvers and rendezvous. Traditionally, delta-V guidance strategies are obtained by means of classical astrodynamics open-loop methods (e.g. Lambert’s problem solution) based on the concept of impulsive and instantaneous thrust action. However, these methods are not always feasible in practice, due to the technical limitations of real propulsion systems: the delta-V budget cannot be concentrated in a single impulse and then gravity and misalignment losses are introduced with a consequent increase in propellant consumption (key factor in space missions as it strongly affects their duration). A significant advance in this context could be represented by autonomous guidance approaches. From this point of view, Nonlinear Model Predictive Control (NMPC) has a great potential for the future of aerospace control and guidance systems. Indeed, it can both plan its trajectories autonomously and implement an unified optimal guidance and control strategy, ensuring the minimization of the fuel consumption to complete the mission. Furthermore, it is also able to systematically handle linear and nonlinear constraints and complex Multi-Input Multi-Output (MIMO) systems, reducing the effort of the mission plan design. In this paper, a novel NMPC framework for autonomous guidance and control with high-thrust quasi-impulsive maneuvers is presented. A key feature of the proposed NMPC framework is the use of different kinds of orbital motion models as possible internal prediction models in the NMPC algorithm. These models are based on: Classical Keplerian Elements, Cartesian Coordinates (i.e. function of position and velocity) and Modified Equinoctial Orbital Elements (MEOEs).

In the context of Earth Observation missions, many space programmes (i.e. GOCE, GRACE-FO, Sentinels, etc.) have been focused on remote sensing the Earth, in order to observe events and phenomena which cannot be studied in-situ. In the near future, satellite operations might require a spacecraft which, from a parking orbit, is able to quickly overfly, autonomously and with a narrow tolerance, any point on the Earth surface, after suitable alarms triggered on board or by the ground stations. These kinds of missions expect one or more consecutive orbital plane changes, without any modification of semi-major axis or eccentricity. They are suitable test benches for quasi-impulsive high-thrust guidance and control laws. The ESA Sentinel-2 mission is the baseline for the case study. The NMPC behaviour is studied using the different internal prediction models in order to look for the one that provides the best performance in terms of propellant consumption, reference tracking and elapsed time for visiting a given point. Furthermore, in order to analyse the benefits of using the NMPC as autonomous guidance and control system, the developed framework is compared to the ideal impulsive strategy and to the realistic open-loop maneuver. The ideal strategy for orbital plane change consists in the application of instantaneous impulsive thrust actions at the orbital node, while in the realistic open-loop one the spacecraft fires with a maximum constant thrust in a direction that, instant by instant, is perpendicular (in the orbital plane) to the satellite tangential velocity, for a proper burning time interval.

Modified Equinoctial Orbital Elements are effective to describe the orbital spacecraft dynamics in simulation. In the case of Cartesian Coordinates, the differential equations are strongly non-linear; therefore, despite the use of high-order integration algorithm, tight tolerances end up in a fairly high number of simulation steps per orbit. Instead, if the satellite motion is described in terms of MEOEs, the variation of the six orbit elements is much smaller than that of the Cartesian coordinates. In particular, the first five parameters are constant (if no perturbing forces are acting), while the true longitude increase linearly with time. All existing high-order integration methods have error bounds which depend on Taylor expansions of the state trajectory. Then, if the MEOEs are used as state variables, instead of the Cartesian vectors radius and velocity, the trajectory state variables will be smoother, and therefore the integration algorithm will be able to estimate them with a higher relative precision using much larger time steps. As for the Classical Keplerian Elements, they suffer from two main singularities: when the orbit is circular, i.e., when the eccentricity is zero and/or when it is equatorial, i.e., when the inclination is zero. The use of MEOEs allows us to overcome these singularities. Indeed, they suffer of mathematical singularities only in the rare case of retrograde equatorial orbits. The benefits of using the Modified Equinoctial Orbital Elements have also been seen with respect to classical methods. Indeed, the obtained results have proved that the NMPC configuration is able to provide better solutions with respect to the quasi-impulsive realistic maneuver, achieving not far performance indices with respect to the ideal impulsive case. So, if on one side the implemented configuration introduce a slight worsening in comparison with the ideal case, due to the engines action delivery in a finite small amount of time, on the other hand it leads to a significantly higher level of autonomy, less fuel consumption and more flexibility and adaptation capability with respect to traditional approaches.

Convex guidance for Close Rendezvous operations in cislunar Near Rectilinear Halo Orbits Blazquez E1, Lizy-Destrez S1, Ankersen F2, Capolupo F3 1ISAE-SUPAERO, 2European Space Research and Technology Center, 3Airbus Defence and Space The future of exploration points towards complex missions designed around the Earth- Moon Lagrangian points. The , assembled in a Near Rectilinear Halo Orbit (NRHO) about the second Earth-Moon Lagrangian point (EML-2), will provide easier access to the Moon surface and facilitate further exploration of the solar system. Rendezvous and Docking (RVD) operations are critical for the maintenance and assembly of such a complex infrastructure. There is extensive experience with RVD in the two-body problem in Low Earth and Lunar Orbits to various space stations, based on the Apollo missions or the ATV deliveries to the ISS. Despite that, the problem of RVD in non-Keplerian dynamics is a quite recent topic and no operational rendezvous has yet been performed in the vicinity of the Lagrangian points. The need for fully autonomous rendezvous is a driving force for modern Guidance, Navigation and Control (GNC) design and a first step towards fulfilling the requirements of future missions.

The scope of this research is to apply successive convexification and convex optimization methods to design on-board autonomous Close Rendezvous guidance about NRHOs. Such trajectories reflect real mission constraints and follow the international RVD interoperability standards. This paper highlights how to express and convexify the main dynamics and constraints involved in a cislunar RVD scenario, and makes use of the dynamical properties of non-Keplerian orbits to design passively safe trajectories. The method presented outputs a 6-degree-of-freedom (DoF) guidance profile for the chaser, in translation and attitude motion.

The research focuses on Close Rendezvous between two vehicles: a non-maneuvering target and a chaser. The target is located on a Near Rectilinear Halo Orbit about EML-2, with a reference perilune radius of 3270 km, and has its docking axis pointing towards the Sun. The chaser is modeled after the European ATV module and is Within a 100 km sphere about the target. The rendezvous problem is formulated as a finite-time optimal control problem, where the chaser’s translation and attitude dynamics are modeled by first-order differential equations and the cost function is the mass of fuel burned during RVD operations. The chaser’s trajectory must satisfy a certain number of path constraints. Those include line of sight translational and attitude constraints with the docking axis of the target, to allow camera and Lidar signal acquisition. Within the Rendezvous Sphere (10 km radius sphere around the target), the chaser stops at predetermined relative Hold Points to allow for navigation switching and other miscellaneous operations. In addition, at any point before entering the Approach Sphere (1 km radius sphere around the target), the chaser must remain in a safe area around the target for 24 hours following an abort. Finally, close RVD shall be performed within 6 hours.

The 6-DoF dynamical system modeling the chaser’s motion uses the full non-linear Circular Restricted Three Body Problem (CR3BP) equations of relative motion for translation, and Euler’s equations for attitude motion. Such a system is highly non-linear and involves continuous-time dynamics. Moreover, the problem’s constraints are for most of them non-convex. This results in a continuous-time non-convex optimal control problem. In order to achieve fast numerical solution and allow for autonomous on-board guidance, this paper uses a method called successive convexification. The full non-linear continuous system is approximated locally, about a reference solution, as a convex sub- problem. This step requires linearization and discretization of the dynamics to obtain a second-order cone programming problem with a finite number of decision variables. Path constraints are similarly convexified. The newly obtained convex sub-problem is then solved and reformulated by an iterative procedure that converges towards an optimal solution of the initial non-convex problem. In this paper, the translation dynamics of the chaser are linearized with respect to the target’s state. We linearize the attitude of the chaser with respect to a reference attitude profile, chosen as a free rotation with zero control. This choice increments slightly the number of iterations required for the convergence of the algorithm but in turn significantly enhances its robustness. The linearized dynamics are then discretized using a zeroth order hold by keeping the control input constant between the samples. Line-of-sight constraints are convexified according to the procedure described by Virgili-Llop and al. Terminal constraints are enforced by adding penalties directly to the cost function in order to improve convergence. The paper enforces passive drift safety constraints exploiting the dynamical properties of NRHOs, in particular the presence of unstable and central invariant manifolds.

The guidance algorithm is implemented in Matlab using the CVX tool for the convex problem transcription and MOSEK as the second-order cone-programming solver. The initial state of the chaser is randomly chosen within a 100 km sphere around the target, with random initial orientation of its body frame. Early results show very promising performance and convergence properties. The algorithm converges for all initial conditions within the search space, and takes less than ten iterations for more than 90% of the cases. The accuracy of the solution, integrated with the full non-linear dynamics, is within 100 m in position before terminal approach, and increased up to the centimeter for the final operations. These encouraging results suggest that the method is able to generate real-time highly constrained rendezvous trajectories in a challenging environment, while making use of the dynamical properties of libration point orbits to ensure passive safety. Future works involve firstly adding a model-switching functionality to the procedure, in order to further simplify the linearization and discretization of the dynamics. Secondly, we consider implementing a compiled language version of the algorithm in order to demonstrate on-board autonomy with a rendezvous simulator.

Sentinel-2: In-orbit Benchmark of Precision Star Tracker / Gyro Attitude Estimation and Smoothing Winkler S1, Fischer D1, Gockel W1, Gratadour J2 1Airbus Defence And Space, 2ESA / ESTEC Introduction

The ESA Copernicus Sentinel-2 mission comprises a constellation of two polar-orbiting satellites launched in 2015 and 2017 and placed in the same sun-synchronous orbit (half an orbit spacing). It aims at monitoring variability in land surface conditions using its wide-swath, high-resolution MultiSpectral Instrument (MSI). Airbus Defence and Space is the satellite prime contractor and also in charge of the attitude and orbit control system driven by precision attitude knowledge requirements.

Onboard attitude estimation accuracy is one of the main contributors to geolocation performance. Based on its high precision gyroscope unit and star trackers, Sentinel-2 serves as a benchmark of todays technology.

On board the Sentinel-2 satellites, star trackers and a gyroscope unit are fused within a so-called “gyro-stellar estimator” (GSE) for precision attitude knowledge determination. With the approximation of a time-invariant estimator (or filter) model, motivated by approximately time-invariant body-fixed angular rates, the original Kalman filter was simplified to a constant-gain filter. However, the filter gains were generated using the originally developed Kalman filter by covariance-tuning in a simulation campaign to achieve the required attitude knowledge accuracy. Basis for the tuning were the documented (!) performances for both gyroscope unit (Astrix 200) and star trackers (Astro-APS) and their nominal alignment.

It shall be explicitely pointed out that both the Astrix 200 gyroscope unit and the Astro APS star trackers on board Sentinel 2 are one of the most accurate attitude sensors available today. But it has still not been answered whether their true (i.e. in-orbit) measurement accuracy is fully used for attitude knowledge determination, i.e. for finally achieving the best possible geolocation performance.

This question leads to the two major contributions of this paper - analysis of the optimality of the current attitude knowledge accuracy on-board Sentinel 2 - improvement potential of attitude knowledge accuracy by on-ground smoothing

Application of optimality theory for stochastic filters

In general, each estimation/stochastic filter is derived based on assumptions made for the states that shall be estimated and the errors in the available measurements. Since these assumptions never exactly reflect the real world in which the filter shall be applied, the model used in the filter is just an erroneous approximation of the real world. Specifically, the system and error/noise model used in the filter never exactly reflect the reality, hence, filter tuning is used in practice to obtain the required estimation performance.

In a real world application like Sentinel 2 the question rises whether the tuned filter does really perform with optimal estimation performance (i.e. minimum estimation error covariance). On ground, where the true value (spacecraft attitude) is known, tests can simply be done using simulations computing the estimation error. But in orbit the true value is unknown. Therefore the estimation error can not be computed. Indeed, the filter, such as the GSE Kalman filter used here, provides an estimation error covariance. But this is only correct if the model used in the filter and the reality are identical (or the filter is perfectly tuned). In reality, this will never (initially) be the case.

Methods to test the optimality (i.e. consistency) of the tuned filter during a mission when the spacecraft is already in orbit are mainly based on (1) normalized error square (NES) (2) autocorrelation (3) normalized mean error of the measurement residual (innovation). In literature, a sufficient practical analysis for spacecraft-specific applications is missing. But Airbus has specifically treated that subject already in the frame of the HOREOS study financed by DLR in 2011. And this know-how is now applied to the Sentinel 2 satellite in orbit. As most robust innovation-based criteria for optimal estimation performance were identified (1) The innovations should be zero-mean (unbiased) and correctly reflected by the filter-computed innovation covariance matrix (2) The innovations should be white (i.e. not time-correlated). These criteria are applied to the Sentinel 2 gyro-stellar estimator (GSE Kalman filter).

Application of smoother theory

The focus is on estimation (filtering) in post-processing, thus batch processing. In order to achieve the best estimation accuracy, a forward-time filter is combined with a backward-time filter. Such filters are knows as smoothers.

Different smoother types are distinguished, usually fixed-interval, fixed-point and fixed-lag smoothers. The fixed-interval smoother uses the entire measurement batch (interval) to estimate the states at all points of time within this interval. Since the measurements of the entire batch are used to produce an estimate, this smoother provides the best possible estimate over that interval. Therefore, the smoother applied to Sentinel 2 is a fixed- interval smoother.

The overall attitude knowledge estimation algorithm consists of the following three parts: (1) Forward filter (like the on-board GSE) (2) Backward filter (like the on-board GSE but running backward in time) (3) Smoothing of forward and backward filter results For the Sentinel 2 analysis to be presented in the paper both forward and backward filter are Extended Kalman filters. At each sensor measurement date, the smoother provides a point solutions for minimum estimation error variance using the current a posteriory state estimate of the forward filter and the a priory state estimate of the backward filter.

One of the most convenient forms of such estimators is the RTS fixed-interval smoother (RTS: Rauch-Tung, Striebel). It combines backward filter and smoother into one single backward recursion. An RTS smoother is used as final solution for Sentinel 2 and will be further detailed in the paper.

Exemplary results in form of a star tracker measurement filtering are provided.

Final remark

This is not a theoretical paper. This is a paper about making sophisticated theoretical concepts work for real life problems. All results presented in the final paper will be based on measurements from both gyroscope unit and star trackers on board Sentinel 2. They will provide a benchmark of the attitude estimation accuracy achievable today. Such a benchmark is missing in literature.

High-Order Sliding Mode Controller for the test mass stabilization of the LISA Mission: Preliminary Results Bloise N1, Capello E1, Punta E2, Grzymisch J3 1Politecnico Di Torino, 2CNR-IEIIT, 3ESA The LISA mission is the ESA future observatory, consisting of three spacecraft separated by 2.5Mkm in a triangular geometry, with two free-flying test masses within each spacecraft. Gravitational waves are detected by measuring the change in relative distances between test-masses on different spacecraft, at pico-meter level. The main objective of this paper is the design of the test mass controller, able to robustly deal with large initial deviations of the release mechanism. A second-order sliding mode control (SMC) is proposed for this critical phase, which is able to handle uncertainties and noise introduced by the sensors system. In addition, limitations both of the actuation system (with saturations and delay) and of the update frequency of the controller are considered. Space gravitational observatories, such as LISA, will be complementary to the existing terrestrial laboratories to detect gravitational low frequency signals, not measurable from Earth. In the space, the observatory is not subject to any perturbations and disturbances, which are usually caused by seismic, thermal or geophysical noises, affecting the observatories on Earth’s surface. The LISA observatory, which is, as said before, one of the space gravitational observatories, is on near- circular Earth heliocentric orbit and each spacecraft contains two optical assemblies and two test masses, included in the laser interferometer (in a Michelson configuration). The main goal of the mission is to protect the two test masses of each spacecraft against all external disturbances, to have free fall on their own geodesic. The relative distance of test mass, continually measured, changes when a gravitation wave passes through. Therefore, the spacecraft are equipped with a drag free attitude control system, which compensates for the non-gravitational disturbances and guarantee free fall conditions. In this paper, the focus is on the control of the test mass release. Initially, the test mass is kept in the center of the case, grabbed by the two plungers until they are retracted. When the masses are released, they are free to move inside the case. So, the main aim of the control system is to steer and keep the test mass at the centre of the case. Thanks to LISA Pathfinder’s pioneering mission, the initial state conditions of our problem are obtained from estimations of the conditions at the release. In detail, LISA Pathfinder, an ESA mission launched on December 3 2015, was dedicated to carry out experimental demonstrations in views of the subsequent LISA mission, expected in 2034. The controller of the test masses must be design to deal with large initial offsets and velocities for the position and attitude states. A great challenge of this control system is given by the limited actuation authority of the electrostatic suspensions, which allow the application of electrostatic forces and torques to the test masses. These electrodes measure even all six degrees of freedom of each test mass by means of capability sensing techniques. Two different operation modes are provided by this actuation system: (1) one is related to the Wide Range (WR) mode of electrostatic suspensions, and (2) High Resolution (HR) mode is used to improve control accuracy. The first mode is activated few seconds after the TM release, since it allows higher forces but introduces more noises. The orbital simulator includes the following subsystems: (i) plant dynamics of spacecraft and test masses, (2) sensor model to obtain the measured values of the positions and attitude angles and (3) a state estimator for the observation of the velocities. These outputs are provided to the controller, given the error dynamics with the desired reference values, required to stabilize the TM system. The strict limits of the actuation system are taken into account for the evaluation of the effective forces and torques. Since the spacecraft and test mass dynamics is a nonlinear system, affected by uncertainties and disturbances, we propose a second order SMC algorithm, named Super-Twisting (STW), for both the position and attitude control channels. STW designs a continuous control law, which is able to steer to zero in a finite time not only the sliding output, but also its first time derivative. Another advantage regards the chattering phenomenon, which is attenuated improving the control accuracy. The STW SMC strategy does not require the availability of any time derivative of the sliding variable. However, the design of the sliding surface relies on the knowledge of signals, such as linear and angular velocities. Exploiting the measured signals (affected by noise) of positions and attitude, in the performed simulations, the estimated velocities have been provided by using an approximation of the derivatives of each signal, smoothed by a low pass filter with desired bandwidth. Moreover, in this work, the action of the controller is limited due to the small force, that can be exerted by the electrostatic actuators and due to the constraint on the switching frequency, imposed by the actuation system, of 10 Hz. These two limitations reduce the accuracy of the controller itself, in terms of position and velocity error. Montecarlo simulations with different initial conditions and parametric uncertainties are performed to compare the proposed controller with a first order controller, as shown and flown in LISA Pathfinder. Performance metrics, such as control performance, computational efficiency, and error accuracy, are considered. The proposed control strategy provides satisfactory results for position and attitude control of the test masses, even in presence of external disturbances, actuators and sensors noises. The simulation results are compliant with the strict requirements in terms of positions and attitude with their velocities both in WR mode and in HR mode.

ASCENT FLIGHT CONTROL SYSTEM FOR REUSABLE LAUNCH VEHICLES: FULL ORDER AND STRUCTURED H∞ DESIGNS. Macés-Hernández J1, Sagliano M1, Heidecker A1, Seelbinder D1, Schlotterer M1, Fari S1, Theil S1, Woicke S1, Dumont E1 1Deutsche Zentrum für Luft- und Raumfahrt In recent years, companies and space agencies have focused their efforts on bringing reusability technologies to the space domain, in an attempt to substantially decrease mission costs and development times. The usage of Vertical Take-off and Vertical Landing (VTVL) Reusable Launch Vehicles (RLVs) is on the way to becoming the new industrial standard; private companies like SpaceX and Blue Origin have already matured this technology and successfully demonstrated its capabilities with several flights. Governmental agencies, are now applying themselves to extending the application envelope of these technologies. The German Aerospace Center (DLR), in collaboration with the Japan Aerospace Exploration Agency (JAXA) and the French National Centre for Space Studies (CNES) have joined efforts to build and operate a reusable first- stage demonstrator, named CALLISTO. This paper gives an overview of DLR’s work on the robust ascent flight control system for this vehicle.

First we introduce the launcher dynamics in the form of a time variant, non-linear model; physical parameters of the vehicle like moment of inertia and center of gravity are strongly connected to the propellant and oxidizer consumption and, due to the vehicle’s velocity during its ascent flight, the induced aerodynamic forces change rapidly as the physical properties of the surrounding air mass also changes with the altitude. During ascent flight CALLISTO uses a throttleable engine for longitudinal control, thrust vector control (TVC) for the pitch/yaw motion and a reaction control system (RCS) for roll stabilization. We elaborate the process and the reasoning used for reducing this complex non-linear model to a set of linear equations of motion that include the dynamics of the vehicle, actuators and sensors at different points along the trajectory; this is the basis for our control designs. For robustness purposes we extend the linear models by including the envelope of parametric and model uncertainties in the form of Linear Fractional Transformations (LFTs).

It is well known that for both launch vehicles and RVLs the interaction of the vehicle with the lateral wind and gust perturbations is the most challenging phenomena to handle for control engineers. It is therefore important to limit aerodynamic loads induced on the vehicle by controlling the angle of attack; further, of vital importance is to design controllers accordingly. They must be able to maintain loads bounded while achieving demanding and tight tracking requirements. As second step, the problem is formulated in the H∞ form; loop-shaping functions that include trajectory tracking requirements, wind-rejection objectives, actuation performance and worst-case navigation noises were utilized to penalize the frequency response. Different control objectives were defined according to the flight instant, i.e. if the vehicle is near its maximal point of dynamic pressure, the load relief capability of the vehicle becomes more relevant, while at the beginning and the end of the ascent phase it is more important to have low attitude and position tracking errors, respectively. We propose a fully integrated attitude and position tracking controller that enables focusing in different control objectives and at the same time preserves the control architecture for every design point.

Synthesis models were used for the calculation of two sets of controllers for the ascent flight domain: The first set consists of full-order H∞ controllers. The scheduling of these controllers is performed on the controller output, i.e. all controllers run in parallel and the control command is obtained by interpolation of the output and a reference variable. The second is a set of controllers obtained by using structured H∞ synthesis; a predefined controller structure is imposed to the optimizer. This results into a set of static gains for each operating point. Scheduling is then based on linear interpolation of these gains on the non- gravitational velocity.

Both controller sets are tested in nonlinear, numerical simulation under nominal and dispersed conditions. Simulation results are presented and discussed accordingly. Our stabilizing controllers take advantage of the decreased dynamic pressure at the end of the ascent phase for improved terminal position tracking in the attempt of minimizing error propagations to the next phases, and hence maximizing chances of a successful landing at the end of the mission. Care has also been taken in the correct reduction of the aerodynamic loads, rejecting wind-induced perturbations and guaranteeing adequate values for the rigid body stability margins under nominal and perturbed conditions.

Legs on Aerial Robotic Drone: Control Design Challenges of Agile Thruster-Assisted Legged Locomotion Ramezani A1, Ramezani A1 1Northeastern University Despite many accomplishments by legged robot designers, state-of-the-art bipedal robots are prone to falling over, cannot negotiate extremely rough terrains and cannot directly regulate unilateral contact forces. Our objective is to integrate merits of legged and aerial robots in a single platform. We will show that the thrusters in a bipedal legged robot called Harpy can be leveraged to stabilize the robot’s frontal dynamics and permit jumping over large obstacles which is an unusual capability not reported before. In addition, we will capitalize on the thrusters action in Harpy and will show that one can avoid using costly optimization-based schemes by directly regulating contact forces using an Reference Governor (RGs). We will resolve gait parameters and re-plan them during gait cycles by only assuming well-tuned supervisory controllers. Then, we will focus on RG-based fine-tuning of the joints desired trajectories to satisfy unilateral contact force constraints. ROM-Based Feedback Design for Thruster-Assisted Legged Locomotion Dangol P1, Ramezani A1 1Northeastern University The main objective of this work is to take preliminary steps in the design and control of multi-modal robots capable of negotiating unstructured environments through a mixture of legged and aerial locomotion. These robots could be used in a broad number of applications including search and rescue operation, structural inspection, automated package delivery in residential spaces to name a few. We will assume for well-tuned supervisory controllers and will focus on fine-tuning the joints desired trajectories to satisfy the performance being sought. In doing this, we will devise an intermediary filter based on reference governors that guarantees the satisfaction of performance-related constraints. Since these modifications and impact events lead to deviations from the desired periodic orbits, we will guarantee hybrid invariance in a robust way by applying predictive schemes within a short time envelope during the gait cycle. To achieve hybrid invariance, we will leverage the unique features in our model, i.e., the thrusters.

Introduction: Raibert's hopping robots and Boston Dynamic's BigDog are amongst the most successful examples of legged robots, as they can hop or trot robustly even in the presence of significant unplanned disturbances. A large number of humanoid robots have also been introduced. Honda's ASIMO and Samsung's Mahru III are capable of walking, running, going up and down stairs. Despite these accomplishments, all of these systems are prone to falling over. Even humans whose performance easily outperform that of today's bipedal robot cannot recover from severe pushes or slippage on icy surfaces. Our goal is to enhance the robustness of these systems through a distributed array of thrusters.

From a control design standpoint, studying these systems can help extend principles for robot multi-modality and mobility in unstructured environments by building reduced-order models (ROM) for terrestrial and aerial locomotion of non-trivial robot morphologies. For instance, earlier legged locomotion works have shown us that reduced-order systems can be invaluable in uncovering basic dynamical structures. ROMs contain the smallest number of variables and parameters that exhibits a behavior of interest can be hypothesized as an attracting invariant submanifold on which the restricted dynamics take a form prescribed by a supervisory controller. In dynamical systems terminology, this is a collapse of dimension in state space, which would follow from the existence of a center or inertial manifold with a strong stable foliation. While mathematical models of legged robots of varying size and complexity are relatively well developed, models of multi-modal locomotion are not well established.

In this paper, we report our efforts in designing ROM-based closed-loop feedback for the thruster-assisted walking of legged systems, currently being developed at Northeastern University. These bipeds are equipped with a total of six actuators, and two pairs of coaxial thrusters fixed to its torso.

ROM-based feedback design for thruster assisted legged locomotion: The robot is modeled in the sagittal plane of motion as an equivalent three-link biped. Using Lagrangian mechanics, the single support (SS) phase equation of motion (EOM) are generated. The thrusters are considered to be inactive during SS phase to permit under-actuation and therefore the existence of zero dynamics. The gaits are then designed using the Hybrid Zero Dynamics (HZD) approach and we assume for a well-tuned controller that ensures finite time convergence to the desired gait. We focus on fine-tuning the desired trajectories to satisfy the performance constraints on the system. This is achieved through an optimization-free intermediary filter based on the emerging idea of reference governors. In this formulation, a Lyapunov argument is taken to manipulate reference trajectory according to the required constraint by setting an upper bound on a Lyapunov function such that limits are always satisfied.

Since these modifications and ground contact between gaits lead to deviations from the desired periodic orbits, we exploit the double support (DS) phase within a very short time envelope and leverage the thrusters to achieve hybrid invariance. The DS phase is modeled similar to SS phase with both feet fixed to the ground and the thrusters now active. Due to the over-actuation imposed by kinematic constraint and the need to maintain ground reaction forces (GRF), a model predictive scheme is opted for. An objective function is minimized such that the states follow a trajectory generated between the post impact and initial SS phase states to ensure impact invariance. The physical and constraints brought on by the need to maintain ground contact through regulating GRF are included as constraints in the optimizer. Through posture adjustment post impact, this scheme ensures hybrid invariance by guiding states back to the zero dynamics manifold while handling actuator redundancies.

Results: Under the proposed control scheme a three link equivalent model of the biped was simulated with each DS phase lasting 20ms. The nominal limit cycles obtained through implementing the described approach are shown.

The feasibility conditions are satisfied during the robustification process. The tangential to normal load ratio for each feet is less than or equal to the friction constant value at all times and the normal forces are always positive, which indicates that the feet were stuck to the ground throughout the DS phase. We note that that the normal forces spike at the same time as the thruster action, providing additional force required for the biped to maintain ground contact which would not be possible with control action on the joints alone.

Conclusion: By relying on the thrusters available on our robot, stable walking gaits were achieved in a robust fashion. The merit of our approach is that, unlike existing methods, satisfying performance-related constraints during the SS phase does not rely on expensive optimization approaches and hybrid invariance is achieved through exploiting the DS phase. In addition, this approach allows for increasing performance and robustness enhancing capabilities during specific parts of the gait cycle. These preliminary results obtained in the design and control of multi-modal robots will hopefully aid in the development of robots that are able to traverse through unknown terrains.

Advances in Sensors and Actuators

Flash LiDAR for Space Rendezvous and Docking Missions Shimizu S1, Katayama Y1, Okada N1, Yamamoto T1, Mizuno T2, Ikeda H2, Kondoh Y1, Ito N1, Imaoku T3, Kase T3, Kawahara A3 1Japan Aerospace Exploration Agency, 2Japan Aerospace Exploration Agency, 3NEC Corporation The International Space Station (ISS) Multilateral Coordination Board (MCB) met on August 6, 2019, and its members acknowledged the need for a human outpost in the lunar vicinity “Gateway” as a critical next step. Rendezvous and docking missions are required to build and resupply the Gateway, and accomplishing such missions entails measuring the relative position and attitude between a spacecraft (a chaser) and the Gateway (a target). The Japan Aerospace Exploration Agency (JAXA) is developing a relative navigation sensor called “Flash LiDAR” for such space rendezvous and docking missions. Relative navigation sensors rely on the use of radio waves, visible light images, infrared images, and a laser system. LiDAR (Light Detection and Ranging) is a sensor device that uses a laser system. The mechanism of LiDAR irradiates a target with pulsed laser light, detects the reflected light, and then measures the target’s light intensity, direction, and distance as a time of flight. The requirement specifications for a relative navigation sensor used in a rendezvous or docking mission, such as measurement distance range, field of view, and reflection characteristics of a target, vary from mission to mission. The light intensity, directivity, and wavelength of LiDAR can be flexibly designed with respect to the requirement specifications, thus making LiDAR a suitable type of relative navigation sensor. For these reasons, LiDAR has been adopted by technical test satellites “Orihime” and “Hikoboshi”, the H-II transfer vehicle (HTV) “Kounotori”, and the spacecraft “Hayabusa” and “Hayabusa 2”. The scanning and flash types of LiDAR can take 3D images. The scanning type projects a narrow divergence and high energy density beam, and scans a field of view by using moving mirrors. This type has a high S/N ratio and offers the advantage of a long measurement range compared to the flash type. However, the scanning type also has the disadvantages of having moving parts that could potentially cause hardware failure, and requiring a long scanning time. In contrast, the flash type projects wide divergence beam with low energy density and takes a 3D image instantaneously. This type offers the advantages of high reliability by having no moving parts, and taking a 3D image with a wide field of view in one shot. However, the flash type has the disadvantage of a short measurement range caused by a low S/N ratio compared to the scanning type. JAXA conducted a feasibility study on Flash LiDAR that overcomes the disadvantage of the flash type by employing a newly developed 3D image sensor chip called the Multi Pixel Photon Counter - Read Out IC (MPPC-ROIC). The MPPC-ROIC consists of the MPPC and ROIC that are bonded vertically. The MPPC is a photoelectric conversion element array developed by Hamamatsu Photonics K.K.. The MPPC has 128 x 128 pixels and each pixel consists of multiple sub-pixels. The sub-pixel is a Geiger mode Avalanche Photodiode (APD), thereby enabling high sensitivity capable of photon counting and realizing a long measurement range with relatively low energy density light. Moreover, the intensity resolution of the Geiger mode ADP is generally 0 or 1, but multiple sub-pixels enable the intensity resolution of each pixel of the MPPC to be 1 or more. The ROIC is a readout IC designed with radiation resistance by JAXA for the MPPC. The ROIC simultaneously reads out and outputs the light intensity image and distance (time of flight) image of the target in one shot. MPPC-ROIC development has been completed up to the prototype, which has 32 x 32 pixels. Following its feasibility study and MPPC-ROIC prototype development, JAXA initiated its development of Flash LiDAR for an improved H-II transfer vehicle tentatively called HTV-X as its first demonstration model (hereinafter, FL). We have completed the breadboard model (BBM) design, manufacturing, and evaluation test under the limited conditions, as well as the preliminary design of an engineering model (EM) and a flight model (FM) of FL. One of the main points of FL development is the design of operation scenarios and modes. The targets of FL are multiple reflectors mounted on the ISS. Although the reflector looks like a single point at a long distance, it appears separately upon approaching. Therefore, operation modes (i.e., number of visible reflectors) must be designed according to the distance. And in order to detect only the reflector (as the intensity of the background object’s reflected light is set less than the detection limit), a scenario and the function of adjusting the intensity of projection light must be designed according to the distance. For the measurement performance of FL, we conducted simple geometric analysis by considering the number of pixels, field of view angle, pixel quantization error, and random noise of the time of flight. We consequently confirmed that the measurement requirements from the HTV-X docking mission are satisfied. This paper describes the technical features of FL and the MPPC-ROIC, and presents the preliminary design and simple analysis results of FL. It also describes the applicability of FL to future space missions.

RVS®3000-3D LIDAR – Gateway Rendezvous and Lunar Landing Schmitt C1, Dochow S1, Windmüller M1, Both J1, Mongrard O2 1Jena-optronik Gmbh, 2European Space Agency The return of human presence in cis-lunar space and on the surface of the Moon for missions of increasing durations will be a key milestone towards the ultimate goal of manned missions to Mars. The assembly, operation and supply of the Gateway, representing the necessary human outpost in cis-lunar orbit and a key node in the lunar transportation architecture, will therefore be one of the major key challenges in the upcoming years.

For autonomous rendezvous and docking with the Gateway intelligent relative navigation sensors are required. Jena-Optronik’s new 3D LIDAR called RVS®3000-3D represents a solution to this challenge via the combination of a high resolution scanning LIDAR with robust pose estimation algorithms. The new generation LIDAR benefits from the legacy of 48 delivered RVS® sensors which all flew flawlessly to the International Space Station on board of ATV, and HTV spacecrafts. The RVS®3000-3D LIDAR hardware successfully reached TRL9 via its maiden flight to ISS in 2019 on Cygnus NG-11 and several more units are under contract and even already delivered for the upcoming missions. In the paper we present 6DOF pose estimation performance estimates of the RVS®3000-3D vs. the International Docking Adapter (IDA), which will be used on ISS for the crew commercial program and is also foreseen as the standard docking interface for the Gateway. The simulations are based on experience and data gathered with RVS®3000 Engineering Model in several ground tests, e.g. vs. IDA FM3 at the .

In parallel to the establishment of the Gateway station in lunar orbit, a series of robotic mission to the lunar surface are foreseen, paving the way for human return. For autonomous and safe descent high resolution terrain mapping is required to detect and avoid hazards, especially in the more challenging polar regions of interest. For this application also the RVS®3000-3D is an excellent solution since it was designed to address long range and uncooperative targets. In the paper test results obtained with a lunar mockup up to 1000m will be presented outlining the RVS®3000-3D’s imaging capabilities vs. lunar regolith. Finally intelligent algorithm solutions for dense hazard map generation and safe landing detection will be presented.

FaintStar - An Intelligent Single-Chip Sensor Head for Star Trackers - Evaluation Results Ogiers W1, Dendoncker M1, How L2, Kowaltschek S3 1AMS Sensors Belgium, 2AdvEOTec, 3ESTEC Since 2003 ESA has been roadmapping and sponsoring the development of CMOS Active Pixel Sensors (APS) dedicated to optical navigation. This resulted in chips, such as the STAR1000 and HAS2, that enjoyed a very successful commercial life. The latest result from this roadmap is FaintStar, created by CMOSIS (now AMS Sensors Belgium) under ESA contract 4000110482 "2nd Generation APS improvements for flexible low cost and mass sensors".

FaintStar (FS) is a radiation-tolerant one megapixel image sensor with 10µm pixels and rolling shutter. All interaction with the user is over a single 80 Mb/s SpaceWire link. FS offers readout modes such as full frame, windowed, dual rolling shutter, with 12 or 11 bit per pixel. The on-chip signal processing encompasses 'pixels-to-centroids', including bad pixel replacement, background estimation and subtraction, spike filtering, bright object extraction, and photometric barycenter calculation. Its algorithms are generic, non-reliant on third-party IP, yet highly user-configurable. Raw and JPEG-compressed image output is also supported. The chip contains two 1.8V regulators, power-on-reset, temperature diode, low-rate analogue inputs with their own 10 bit ADC, and digital GPIO. While the intended application is star tracking, high-accuracy sun sensing and rendezvous, navigation or rover cameras are also possible.

The concept for FaintStar dates back to 2007. Initially it was seen as a successor to the Low Cost and Mass Sensor (LCMS) demonstrator chip (FillFactory, 2004), sharing its 512 x 512 pixel format, but executed in an advanced silicon technology, with increased on-chip processing capabilities, and with a truly hermetic package. Its target market was low-cost applications. The CMOS technology of choice was the UMC 180 nm process, facilitating the then-emerging pinned diode pixels with correlated double sampling. UMC was selected because of its space heritage, including the ESA-funded DARE radiation-hardened logic cell and memory libraries (IMEC). As CMOSIS had no experience with UMC, a long preparatory period of pixel development was called for, using self-funded test chips as well as piggybacking on other ESA contracts (e.g. Sun Sensor on a Chip, 4000108300).

The FaintStar contract started in 2012. In Phase 1 a prototype chip, FaintStar1, was to be developed and characterised, followed by a period of independent evaluation by the European STR industry. Phase 2 would combine the lessons of these activities into FaintStar2, which would be manufactured and then subjected to an ECSS evaluation campaign, followed with commercialisation.

During FS1 detailed design market forces quadrupled the focal plane array (FPA), to one megapixel, the minimum deemed acceptable for future generations of star trackers. At that time, after trying tens of pixel designs, five viable pixel candidates remained. The FPA was made to contain all five types, allowing characterisation with valid statistics. This was the only concession to FS1's prototype status: all other blocks and functions were designed-in from day one, and the design flow was fully product-worthy.

In the meanwhile extensive package derisking and development was performed, resulting in a full-custom 100-lead ceramic quad flat pack with seam-welded lid and sapphire window. It obeyed all initial requirements. When presented to the industry players it was rejected for reasons of cost, bulk, and handling ease. An updated version of the package traditionally used for star sensors was then adopted: a J-leaded ceramic chip carrier with glued coated BK7G18 glass. This is a non-hermetic solution ('semi-hermetic'), with an in-flight performance that is known to be trouble-free.

FaintStar1 silicon became available in June 2016. Fully functional, it was characterised for electro-optical performance, total ionising dose, protons, and heavy ions. A partial report was presented at GNC 2017. The optimal pixel could then be selected, as a trade-off between end of life (EOL) dark current performance and quantum efficiency. This preference was shared by a number of STR manufacturers.

New insights showed that the sensor would initially be used mainly as a raw image provider, i.e. without much pixel processing. FaintStar2 was subsequently designed as a copy of FaintStar1, with optimisations of its image data format and interfacing to make it faster, and with an all-new low-power mode for non- processing scenarios. FS2 is an electrical and physical drop-in for FS1.

Production of 12 wafers was ordered in January 2018, and FaintStar2 silicon arrived in April, finding a company in transition from a specialist design house to a consumer-oriented production enterprise with little bandwidth for niche developments. This posed a serious challenge, resolved with a partnership with opto- electronics test house AdvEOTec (), who would perform the characterisation, evaluation, and screening. AMS would remain responsible for the wafer and device testing, assembly, functional validation, overall supervision, and sales.

After the initial validation of early samples the Flight Model assembly and test flow was set up. In November 2018 seven wafers were tested on the wafer prober, at temperatures of 25C and 60C, including full functionality, power, electro-optical parameters, and optical defects. 174 dies on 4 wafers were selected for the evaluation campaign and were shipped for assembly. All assembled devices then underwent dimensional metrology, electrical and EO testing, leak testing, and visual inspection.

From this lot were picked 55 devices for evaluation testing and 54 devices for screening. Of the screened devices 24 would be used for evaluation testing at qualification levels, with the remaining 30 chips available for immediate sales.

Visual inspection revealed no large problems. There were smaller issues like mild particle contamination and glass lid damage that impacted yield and that will be addressed in a future revised assembly and test flow.

Screening, including stabilisation bake, temperature cycling, burn-in (240h), leak, X-ray and visual inspection was finished in October 2019. Burn-in brought no drift of parameters.

When writing this abstract the evaluation was still on-going on unscreened samples, with good results already obtained for high temperature operating life (HTOL, 2000h, +125C), single event effects/heavy ions, protons, total dose, vibration, moisture resistance, constructional analysis. By the time of the conference the results for mechanical shock, thermal cycling, ESD, as well as selected tests on screened devices will also be available.

The Terma T1 Star Tracker Qualification Test Results and New Developments Davidsen P1, Mikkelsen O1, Bohn P1, Pedersen J1, Petersen C1, Hansen D1, Kaas K2 1Terma, 2Space Inventor 1. T1 Star Tracker Qualification Test Results

The design and qualification of the Terma T1 star tracker Electronics Unit (EU) and associated software supported by ESA GSTP has been completed with key results presented in the paper: • EU environmental test campaign • Software tests comprising acquisition and tracking robustness with and without heavy transient radiation events • Software tests verifying autonomous update and robust use of a full sensor dark current map • Software tests verifying overall dynamic attitude accuracy • Test using proton irradiated Faintstar-2 sensor • Results from night sky testing

The first flight model production of the complete T1 STR including design and test of a new 26 deg SEA baffle will also be presented.

2. New Developments A number of new developments expanding the market for the T1 product with new optional optics, baffle and COTS computer will be presented.

In order to address the marked for small size star trackers, an 18 mm aperture aspheric optics has been designed and tested. The small optics allows for a very compact baffle suitable for nano- and micro satellites.

A miniature Optical Head integrated COTS computer has been developed in partnership with Space Inventor. The computer fits within an envelope of only 57x57x9 mm3 and has been integrated together with the T1 Faintstar-2 PCB forming a very compact star tracker assigned the model name T3. The properties of the T3 star tracker including the small aperture optics will be presented in the paper along with EQM test results obtained during an ARTES study.

The algorithmic part of the star tracker software has been enclosed in a separate Star Tracker Library (called STRLib). The paper will present how this library can be used for hosting star tracker software on an AOCS OBC using a time-space partitioning based software architecture. A prototype star tracker software, which can be integrated in a dedicated star tracker partition, has been developed as part of an ARTES study and tested on an evaluation Zedboard.

NEWTON CMG PACKAGE : Making CMG cluster control as easy as wheels Dupuis A1, Pareaud T1, Dupuis A1, Armand S1, Fruchard D1 1Airbus Defence and Space S/C manoeuvrability is a key design aspect of agile S/C missions, which mostly relies on large torque and angular momentum actuator capacity, such as CMG actuators. Since the early 2000s, Airbus-DS SAS developed with CNES and ESA support, a family of CMG for a large range of space applications, mostly targeting agile missions. Today, we are presenting our new product that simplifies integration and use of CMG cluster at spacecraft level: NEWTON package. Contrary to a classic CMG cluster, NEWTON package directly includes steering laws controlling CMG, increasing the level of service and reducing the risk for customers. With it, it becomes easy to transition from a wheel-based, standard AOCS satellite to a CMG- based, agile AOCS. The proposed paper and presentation targets three aspects: main principles, equipment architecture and system integration and operations. These three items are detailed furthermore below. Newton main principles Newton uses a roof array CMG cluster configuration. This configuration is classical, and is constituted of two pairs of CMG with parallel gimbal axes. Each pair provides angular momentum and torque capacities in the plane orthogonal to the gimbal axes. Newton X axis is defined by the intersection axis between the two planes. The plane elevation angle influences the capacities distribution over the cluster Y and Z axes. This angle can be tuned to balance torque capacity among Y and Z axes, allowing the tuning of the actuator capacity to a large set of missions. The use of CMG to produce the expected three dimensional torque require to overcome well-known issues, the main ones are: - Saturations: the Airbus-DS SAS CMGs are constituted of a gimbal providing a rotation axis perpendicular to the constant wheel momentum, and enabling reorientation of this momentum in S/C frame. The gimbal motor introduces two main physical limitations, first on the acceleration capacity (due to torque capacity), and second to the maximum rate. - Non-linear actuator: contrary to a wheel, the torque direction and norm produced by a CMG depends on its current state (wheel momentum, gimbal orientation and rate). - Singularity: it is defined by the loss of torque capacity along one or more directions due to some relative geometrical configuration of the CMGs. The Newton package concept consists in operating CMG cluster as a wheel array benefiting from the high torque and angular momentum capacity of the CMG actuator. The customer directly commands a three dimensional torque, while the Newton package control laws solve the different issues. With this configuration, the maximum produced torque is 32Nm along Xaxis and maximum momentum is 43 Nms along axis with stability better than 22mNms through a typical AOCS closed loop. The paper and presentation will detail the equivalent wheel array abstraction introduced for Newton.

Newton architecture The Newton assembly is composed of three different units: - A CMG unit which is the wheel/gimbal mechanical assembly and is referred as CMG-M (Control Moment Gyro Mechanism), - Driving electronics, which command the CMG-M and is referred as CMG-E (Control Moment Gyro Electronics), - A CMG On-board Computer, which is in interface between the satellite OBC and the CMG E and contains the steering laws. This unit will be referred as CMG-O (Control Moment Gyro OSCAR). CMG-M and CMG-E come from already presented, and assumed well-known, Airbus DS CMG 15-45 cluster, each CMG providing maximum of 15Nms of angular momentum and 45Nm of torque. CMG-M, CMG O and CMG-E units are interconnected through an internal and dedicated interface cables. Newton CMG package is connected at CMG O level only to the spacecraft On Board Computer through a redundant 1553 MIL bus. Thermal control is fully tuneable at satellite level: each mechanism can be thermally regulated independently thanks to thermistors and heaters. All Newton units are separately powered by the main non-regulated power bus from the satellite. Dedicated power protections shall be implemented outside of NEWTON CMG package.

Newton package integration and operation at AOCS level The paper will present four important aspects of system integration with the AOCS: - On-board operational logic, and AOCS close loop concept: Newton integration in AOCS closed loop is eased by an “autonomous torque box” concept. The mode logic targets initialization, nominal operations and contingency investigation. Initialization of the CMG cluster is performed automatically by NEWTON package in order to reduce disturbance torques. After the initialization of the CMG cluster, AOCS shall send periodically a 3D torque command at 16Hz. This command is processed by NEWTON to drive the four CMG M with the correct angles computed by the embedded steering laws. The CMG O sends to the spacecraft On Board Computer the information corresponding to the angular momentum of the cluster (direction and value). - FDIR concepts: a key concept of the FDIR is that Newton functionally protects the physical limitations of CMG whatever the input commands are. In addition, Newton control laws work in 4 but also 3 CMG configuration to tolerate one CMG persistent failure. FDIR concept falls into two parts: o Protections requiring quick reaction that can only be done at Newton level. o Other protections, that are delegated at system level (monitoring and recovery actions to be implemented and tuned by customer, to fit its own logic and need). - End-to-end validation: great efforts were made to define a validation logic that ease integration into the S/C, providing early testing capacity trough functional simulators (Matlab functional model for AOCS design and testing, C-library for integration in customer validation benches and HW/SW simulator for flatsat activities). In addition, Airbus-DS developed a validation facility to valid Newton Hardware through a macro- dynamic table. - Newton package content and engineering: the Newton package is delivered with usual ICD (thermal, mechanical, electrical) and a user manual. Some domain specific models can be delivered upon request.

Newton available hardware options Newton package is also available with all size of Airbus-DS CMGs: CMG 4-6, CMG 15-45, CMG 40 60 and CMG 75 75. In addition, TM/TC interfaces with Newton package are the same whatever the CMG capacities. Reaction Wheels with Internal Speed Control Antunes Ferreira R1, Kocman T2, Serafin M3 1Bradford Engineering, 2SYDERAL Polska, 3SYDERAL Swiss In 2006 the Reaction Wheels technology from Airbus Defence and Space (Mechanism Product Group, Stevenage) was transferred to Bradford Engineering. These Reaction Wheels have an extensive heritage with long operation in missions such as, among others: SOHO, XMM, INTEGRAL, Rosetta and AEOLUS. Part of the transfer process to Bradford, the design was modernized and qualification was reached in 2010, followed by industrialization and successful delta-qualification concluded in 2015. In 2017, a dedicated activity for technology consolidation related to micro-vibration stability and characterization of re-lubrication mechanism was successfully completed. Since October 2018, Bradford Reaction Wheels are flying on board of the BepiColombo spacecraft. Several future space missions have very stringent pointing accuracy requirements. S/C pointing is influenced by disturbances applied externally to the S/C, e.g. solar wind, as well as disturbances occurring internally. On the latter, the disturbances generated by the Reaction Wheels have a significant overall impact on the pointing accuracy of the S/C. Those disturbances are applied in all 3 directions in force and moment and are known as micro-vibrations. From the S/C side, usage of dampers and operating at lower speed ranges are some of the strategies applied to reduce the micro-disturbances. However, disturbances originating from torque irregularities from the Reaction Wheel are known to have a low frequency content which are difficult to be damped. With a Reaction Wheel internal speed and net torque controller, those disturbances can be internally minimized, resulting in a lower impact on pointing accuracy at S/C level and improving the performance of the entire system. Under ESA CTP (Core Technology Program), Bradford Engineering together with SYDERAL Polska and SYDERAL Swiss are currently designing a controller with updated Wheel Drive Electronics (WDE) using as much as possible the current Reaction Wheels Assembly (RWA) design and heritage. The program is divided into two phases: in the first phase, the current WDE and RWA are evaluated with detailed trade-offs to support the internal controller optimization, digitalization of several circuits and overall reduction of recurring costs. This intended preliminary design will have the least impact on the RWA design in order to maintain the current qualification status. In the second phase, the WDE design and controller are completed. An EM WDE will be produced and verified with different size of flight representative RWAs, with the goal to reach TRL 5. During the first phase, it has been concluded that the required changes on RWA are minor, only the resistor- capacitor network of the tacho signal and the shutter ring tolerances, responsible for the RWA rotor speed measurement, will need to be improved. On the WDE side, the internal clock will be improved, the measurement time improved from 10 Hz to 16 Hz compatibility and the Least Significant Bit will be reduced to 0.02 RPM, which is a two orders of magnitude improvement compared to the current 2.72 RPM. The torque rise/fall time will be reduced from 40 ms to 20 ms. A PID controller has been selected with tuneable parameters via telecommand, making the operation easier, both on-ground and during flight. The controller would be implemented in FPGA with soft-core processor and the usage of external memory will be used during EM development and re-evaluated for final design. In this paper the first phase of the program is presented including the preliminary design and trade-offs as well as the targeted improvements, and findings obtained until now. Magnetorquers design for improved demisability in uncontrolled re- entry LEO platforms Santos A1, Meyer J1, Pereira J1, Soares T2, Beck J3, Palombo E2 1Lusospace, 2European Space Agency, 3Belstead Research Ltd The increasing Space Debris population, particularly of man-made origin at Low Earth Orbits, puts in risk the sustainable use of the space environment. There is a trend for countries and agencies to set requirements that impact design at system and subsystem level, particularly for LEO platforms. During uncontrolled re-entries, some elements have a higher probability of surviving re-entry and hit the ground. This is due to different factors such as design, material, spacecraft positioning and others. Magnetorquers (MTQ) are found to be one of such elements; ergo, their development shall aim for an improved demisability performance to mitigate casualty risk estimate. Magnetorquers are relatively simple actuators that can generate a torque by interacting with Earth’s magnetic field. They are used to provide control to the spacecraft attitude. This paper starts by showing a MTQ demisability assessment: computer simulation and test validation campaign. Previous experimental data had been already collected by the authors and is hereby consolidated. The second part of the paper reveals how the outputs of the demisability assessment are used iteratively to design an MTQ with improved demisability features and performance. Data-driven conclusions and overall recommendations are provided at the end of the paper.

AQUILA: development of a European high-accuracy accelerometer for space applications Torasso A1, Beitia J1, Torasso A1, Loisel P1, Vandersteen J2 1Innalabs Ltd, 2European Space Agency Since the 1950s, accelerometers are used in Space applications to perform several different functions in satellite and vehicle AOCS and GNC. Accelerometers are now in the baseline of a variety of science and exploration missions, in particular for missions leaving Earth Orbit.

High-accuracy accelerometers delivering sub-0.1mg accuracy are typically required to measure the alteration of spacecraft trajectories, which translates, ultimately, into velocity changes (also called delta-V), for navigation and orbit insertion purposes. Also, in some applications, those high-precision accelerometers would be required for angular momentum monitoring or vibration monitoring. To address these performance requirements (as well as other key functional requirements such as measurement range and bandwidth), the Dry Pendulous Servo Accelerometers technology would be the most widely used today, whereas MEMS technology is seemingly a good fit to lower grade functions (e.g. >100µg to few mg).

InnaLabs Ltd, a privately held limited company incorporated in Ireland in 2011, is producing European ITAR-free Quartz Pendulous accelerometers since 2013, achieving top-end navigation performance with its AI-Q-2000 series of accelerometers. Thousands of those accelerometers have been since integrated into a wide range of land, marine, and aerospace platforms. In 2019, InnaLabs has been selected by ESA to develop AQUILA, a European Space Navigation-Grade Dry Pendulous Servo accelerometer meeting the requirements for both institutional and commercial Space applications.

The AQUILA development contract was kicked-off in January 2019 with the primary goal of developing an alternative to non-EU Navigation Grade Space accelerometers currently used on board ESA missions and other European Primes’ platforms (with similar fit, form and function). Performance specifications include, among others, stringent bias and scale factor stability (i.e. 0.1µg short term stability, 2µg over 1 hour, 50µg over 1 month, and scale factor stability over 1 month of 100ppm) under full operating conditions (i.e. temperature, time, launch environment), including more than 15-years exposure to Space radiations (100krad goal). AQUILA is re-using the electro-mechanical cell used in the InnaLabs AI-Q-2000 series of COTS accelerometers which are proven insensitive to radiations and robust to stringent mechanical shock and vibration levels. Therefore, the main development activity is focused on the accelerometer electronics to achieve radiation hardness and single event immunity, but also a demonstration of performance under operating conditions. In particular, a hybrid electronics component populated with thoroughly-selected high- reliability EEE parts is being designed.

This paper describes the technology of the InnaLabs AQUILA Dry Pendulous Servo accelerometer and will provide information on its design. Key specification parameters are presented together with breadboard test data supporting compliance to ESA requirements.

A new iXBlue Fiber Optic Gyroscope offering accuracy and reliability at a fraction of their historical cost Bonnefois J1, Ferrand S1, Lecamp G1, Ollivier C1, Ustaze S1 1Ixblue We present a fully European new gyroscope product dedicated to modern space needs: mass produced, multi- role and without compromising on reliability.

The Fog40NS is a three-axis fiber optic gyroscope. It is the child of both in-house iXblue heritage: space equipment production, including the Astrix product line, and naval sensors mass production. iXblue gyroscope has a legacy of more than 300 years cumulated time in-orbit without incident or even significant ageing and almost twenty thousand gyroscopes delivered around the world. In 2018 iXblue started the development of a new space ITAR-Free gyroscope in line with the ongoing newspace revolution. Fog40NS is a 100x100x95mm component weighing 0,8 to 1,5 kg according to its desired radiation shielding. Its short production lead time can address the large numbers required by constellation market. It is suited for the LEO, MEO and GEO orbits with a 15 years guaranteed lifetime. It is designed to be suitable with Electric Orbit Rising (EOR) through the radiation belts without compromising the 15 years GEO warranty. It fits both AOCS propagation during an extended attitude sensor blackout and high speed fine guidance of agile telescopes. Its rotation rate dynamic range and absence of saturation makes it a great choice for planetary exploration and atmospheric re-entry when coupled with three accelerometers. Being instantaneously at full performance when switched on, a perk of vacuum, allows energy conservation strategies for low power satellites. The documented absence of significant FOG ageing or drift allows the satellite integrator to dispense with any complex or costly recalibration pre or post flight. The Fog40NS maintains the numerous benefits of fiber optic gyroscope: with 40mm fiber coils the ARW is below 5m°/√h for the mainstream product. No non-linearity around zero rotation, a high sensor bandwidth above 1 kHZ, no startup delay, no speed saturation as FOGs are only limited by angular acceleration. As this solid-state technology does not drift, the only documented ageing impact is a rise of power consumption which fits the theoretical prediction and a 7 Watt End Of Life power warranty.

Connections to the satellite platform are done through a RS-422 bus. Power supply is possible at both unregulated 28 Volts or regulated 50 Volts. With possible adaptation up to 100 Volts. To slash the costs a COTS strategy has been performed for the electronic design.

Automotive COTS parts qualified in batch have been preferred to HighRel components. The Automotive Components class can ensure a good traceability, compatible with our product assurance. High Fog40NS production volume mitigates the cost of individual batch qualification, especially SEE and LET tests. In the end, reliability and FITs are not inferior to a classic HighRel product: a FIT rate below 1000 is expected.

The Fog40NS, to be rolled out in 2021, will be produced in a dedicated fab room, adjacent to the terrestrial production room of its sister products. Standardization of the fab tools and procedures and mass production will be a leitmotiv for an important part of the price reduction

ARIETIS and ARIETIS-NS, new highly competitive gyro solutions for space applications Torasso A1, Beitia J1, Torasso A1, Potel N1, Kowaltschek S2 1Innalabs Ltd, 2European Space Agency In the last few years, InnaLabs has realized the potential for its Coriolis Vibratory Gyroscope (CVG) technology for space applications and has since gained extensive interest from the European Space Agency (ESA) and other space prime contractors. Following the successful adoption of Innalabs CVG technology on a Earth Observation constellation, two products are being developed with the support of ESA, namely ARIETIS, with the support of ESA CTP, and ARIETIS-NS, co-financed with ESA GSTP, to provide highly competitive solutions to Primes and satellite manufacturers acting in LEO, MEO, and GEO as well as scientific and exploration space applications.

ARIETIS and ARIETIS-NS are two gyroscopes (rate measurement units). ARIETIS is an ITAR free, Rad- Hard 3-axis gyroscope whose main applications are Telecom (15+ years GEO), Navigation, Earth Observation, as well as Science missions. It uses class 1 EEE components and will be fully qualified to ESA ECSS standards. ARIETIS-NS is an ITAR free, rad and Single Event Effects (SEEs) tolerant 3-axis gyroscope derived from ARIETIS which use up-screened automotive quality COTS EEE to withstand space environment in Earth Observation, telecom, and re-entry vehicles missions. ARIETIS-NS is intended for commercial satellites (both LEO and GEO), constellations and mega-constellations. It is developed and qualified as per a mix of ECSS and IPC standards, allowing InnaLabs to adapt to customer Product Assurance requirements.

ARIETIS and ARIETIS-NS are both based on InnaLabs proprietary Coriolis Vibratory Gyroscope (CVG) technology which is currently used in Space on board LEO constellation for which InnaLabs CVG gyros have already cumulated more than 1,100,000 hours, thus confirming the suitability of the technology to space environment. Other applications of the technology also include commercial products for land, marine, and aerospace.

A number of key developments have been introduced by InnaLabs in both products and are presented in this paper. This includes a new generation of a CVG sensor referred to as CVG2, smaller and lighter compared to the existing InnaLabs off-the-shelf CVG, and upgraded by the implementation of a dynamical balancing process improving the bias and noise performance under static and dynamic conditions. Another novelty is the signal processing of the main gyroscope functionalities, i.e. the control loops and temperature compensations, which is an FPGA-based digital design, compared to analogue loops implemented in InnaLabs COTS gyros. Such a design allows for more flexibility and precision and hence guarantees better performance in terms of bias stability, scale factor stability and noise mitigation.

As part of the products development, several breadboards have been built and tested to validate the new digital electronics and complete the compensation process. Preliminary results will be presented in this paper, all within specification and proving that the designed digital gyroscope satisfies a very low Angular Random Walk (ARW) (less than 0.005 °/√hr), a bias stability over the full temperature range better than 1°/hr, and bias instability in the order of 0.02°/hr. This paper also presents ARIETIS and ARIETIS-NS gyroscopes specification, design, functional architecture, and discusses the test and verification approaches followed by InnaLabs.

Evolving SiREUS Utton M1, Durrant R1, Whitley E2, Kowaltschek S3 1Thales Alenia Space UK Ltd, 2Silicon Sensing Systems Ltd , 3European Space Agency The initial exploitation of MEMS devices within the AOCS arena has been based on migrating the use of terrestrial MEMS technology to the space domain and has culminated in the move away from expensive bespoke designs to the use of a standard mass produced MEMS based sensor head. This has addressed part of the cost element, and more importantly provided an order of magnitude improvement in performance from the original designs. The development and qualification of the SiREUS NG10 Inductive MEMS gyro is the forerunner to the planned use of this new sensor head and is capable of utilising future evolutions of the device.

The SiREUS NG10 builds on the original SiREUS gyros that have flown on Cyrosat, Sentinel 3 and MTG missions, and in doing addresses the lessons learnt from the first development programme. The unit has been developed to provide a next generation radiation hardened capability for telecom and EON missions that require a coarse rate sensor.

Further development of the gyro unit is based on the migration of the architecture to a software based design that utilises the emerging micro-controller capability to address key cost reduction requirements, whilst maintaining unit performance. The availability of key components in rad-hard and rad-tolerance form provides flexibility to meet the New-Space market demands.

This paper presents the SiREUS NG10 gyro product, its design characteristics/performance, and provides details on the next phase of the development into a reduced cost gyro module; the SiREUS GM20.

IBIS and MAUS, the Jalapeno’s under the Sunsensors. Leijtens J1 1lens research & development The IBIS Sunsensor is a small digital Sunsensor currently under development in frame of an ESA Artes contract. Based on an application specific CMOS sensor, the sensor is to combine albedo insensitivity with low power operation and a high radiation tolerance. As such, it is expected to be a small but very potent sensor to which a parallel with a Jalapeno pepper can be drawn.(small but potent) In frame of the IBIS developments, some demonstrators had to be designed. As these demonstrators were based on the use of a nano-D connector, this led to a very low-profile setup which for the time being has been transformed into a low profile analogue Sunsensor. This MAUS Sunsensor is small enough to be accommodated on a cubesat but reliable enough to survive any other cubesat component. The presentation will focus on the development status of both products.

Challenges of LEO missions

Utilizing Low Earth Orbit for a Space Interferometer Laboratory by Three Satellites Ito T1, Kawano I2, Funaki I1, Sakai S1 1Japan Aerospace Exploration Agency, 2Japan Aerospace Exploration Agency Spacecraft formation flying is a promising technology that can overcome the limitation of spacecraft’s physical size and realize flexible and advanced observations such as long focal length telescopes and space in-terferometers. Space missions by formation flying do not necessarily increase spacecraft size to achieve the higher performance, whose feature is suitable for the s-class and m-class categories of space science missions. The Japanese ETS-VII (Engineering Test Satellite-VII) successfully demonstrated the autonomous formation flying and rendezvous in the late 1990s, followed by the technological demonstration or earth observation mis-sions such as DARPA’s Orbital Express, German TerraSAR-X/TanDEM-X, and Swedish PRISMA.

Space missions by precision formation flying have been also proposed and ongoing in the world. Some of them require not only “precision” but also “autonomy” in its formation flying. One representative of them is ESA’s PROBA-3 (to be launched in 2021), which will demonstrate several formation activities (from coarse to fine one), its autonomy and safety management, in addition to sun coronagraph observation for a scientific objective. PROBA-3 is capable of resizing its baseline length between 25 m and 250 m in a Highly Elliptical Or-bit (HEO) while maintaining sub-millimeter- and arc-second-order relative displacement and pointing accuracy. PROBA-3’s mission concept demonstrates an excellent balance between the scientific purpose in solar physics and technology acquisition for future precision formation flying missions like XEUS (X-ray telescope) in a small-class category.

In Japan, after ETS-VII no formation flying missions were launched by JAXA, but recently an increasing attention has been paid again, particularly for space science applications. The first version of space science and technology roadmap (hereafter “RM”) was published in March 2019. It mentions formation flying as one of the key technologies for realization of a “space interferometer”, particularly for gravitational-wave and infrared astronomy.

Japan has been proposing a space gravitational-wave observatory: B-DECIGO, whose target frequency of gravitational-wave sources (0.1-10 Hz) is complementary to those of the other missions such as ESA’s LISA and China’s TianQin. The three satellites are placed in a triangle topology keeping its baseline around 100 km. The Fabry-Perot laser interferometer measures the displacement between each of the in-ter-spacecraft’s floating proof-masses extremely precisely. To make the Fabry-Perot laser interferometer being operational, each displacement has to be suppressed much less than the laser’s wavelength (e.g., < 1 nm if the lasers’ frequencies are fixed and stable). On the other hand, the infrared interferometer requires at least three spacecraft, one of them has an infrared interference detector and the two of them have a mirror to reflect the infrared sources from the sky to the detector satellite. The target baseline length is around 100 m for a near- infrared wavelength and 1000 m for mid/far-infrared wavelength. To make the infrared interfer-ometer being operational, the difference of the distance between each arm has to be sup-pressed less than the infrared wavelength (e.g., < 1 μm).

The RM shows a typical technological acquisition scenario as a “mission roadmap” that formation flying technology to be acquired by 2030, gravitational-wave observatory and infrared interferometer to be demon- strated in the JAXA’s M-class (< 15 billion Japanese Yen) mission category in 2030s, in order to realize space gravitational-wave observatory and infrared interferometers in 2040s which can achieve the long-term space scientific goals.

Clearly, the common feature of the two different missions is that they are realized by (at least) three satel-lites and have to maintain their optical path lengths in a very high precision (e.g., < 1 μm). To answer the vision of the RM, the authors have identified the required technologies for formation flying by “three satellites” from the viewpoints of “autonomy” and “precision” and started to seek the best opportunity to acquire/demonstrate these technologies in space. The space environment in Low Earth Orbit is a strong candidate for this purpose in terms of mainly a reasonable access cost to space and ease of spacecraft operations, provided that the baseline length of spacecraft formation maintains a short distance (in the order of 100 m) to reduce the relative accelera-tion disturbed by the J2 term of the earth gravity. In fact, the relative acceleration in Low Earth Orbit (altitude: e.g., 500 km) with a baseline length of 100 m has the same-order magnitude in Geostationary Orbit with that of 1000 km. A space environment suitable for testing the high-precision formation flying technologies could be created in Low Earth Orbit, and even the verification of the design sensitivity of the space interferometers (end-to-end demonstration) would be possible.

This paper analyzes the possibilities on utilization of the Low Earth Orbit as a “laboratory” for the -mum technological acquisition to make a steady step toward the future space interferometer missions. The re- quired technologies and its uniqueness for those future space interferometer programs are described first, then the achievable technologies in the low earth mission are investigated.

GNC design solution for the deployment of BIOMASS large deployable reflector PAREAUD T1, Escudier J1, Pelletier E1, Puchot A1, Watt M2, Deslaef N3 1Airbus Defense and Space SAS, 2Airbus Defense and Space Ltd, 3European Space Agency BIOMASS was selected as the 7th Earth Explorer mission in Spring 2013 at the User Consultation Meeting in Graz, Austria and will be the 4th Core Earth Explorer Mission. The overall objective of the mission is to reduce the uncertainty in the worldwide spatial distribution and dynamics of forest biomass in order to improve current assessments and future projections of the global cycle. BIOMASS will enable the first global scale, systematic measurement of forest biomass. This objective is achieved by the implementation of a P-band Synthetic Aperture Radar (SAR) instrument, providing global maps of forest biomass stocks, forest disturbance and growth.

In May 2016, it was announced that Airbus Defence and Space will build the Biomass satellite. The space segment is comprised of a single low Earth orbit satellite platform carrying a fully polarimetric P-band Synthetic Aperture Radar (SAR) instrument. The P-band SAR is the major driver for the system design, requiring the use of a large aperture antenna in order to achieve the required coverage and performance. The antenna is a large deployable reflector (LDR) of 12 m diameter which is stowed during launch and is deployed as part of the LEOP (Launch and Early Operation Phases) activities. The LDR deployment is a key phase for BIOMASS AOCS design

The LDR deployment needs to be performed in a staged approach allowing operational breakpoints where detailed examination of the performance of individual steps in the sequence will be possible. For each stage, there are thermal and power constraints at the beginning of the deployment step implying constraints on attitude pointing. During the whole deployment step, there are also mechanical constraints on the reflector and in particular on the three boom hinges implying constraints on satellite rate and angular acceleration. All these constraints have been considered to define the best solution for the LDR deployment baseline.

The successive steps of the (nominal) LDR deployment sequence show a spectacular evolution of the geometry and MCI properties of the spacecraft, which is a major constraint for the design of the AOCS control architecture during this sequence. The evolution of the inertia products of the complete spacecraft along the LDR deployment based on the multi-body LDR deployment dynamic simulator model is given, clearly showing the huge variations of the total inertia matrix. If we consider that an anomaly may occur at any time during the deployment, it means that the AOCS control loops active during this phase have to be robust to a very wide range of dynamic conditions in terms of global MCI and flexible modes and to the non- linear effects introduced by the unlocked hinge.

This paper will present the AOCS solution to ensure a robust control during the LDR deployment phase and satisfying the thermal, power and mechanical constraints necessary for the safe deployment of the reflector. In particular we will present: - AOCS design for LDR deployment. This section will in particular highlight how the Astrobus avionic product of Airbus Defence and Space was adapted to reach this goal with minimum design modification. For example the same mode (Acquisition and Safe Hold mode) is used for initial acquisition, for acquisition after a deployment step and for safe mode anytime during the satellite lifetime.

- AOCS tuning strategy for each phase of the LDR deployment. This section will present how the different AOCS modes involved during deployment are tuned depending on the LDR configuration (either in one of the 4 locked nominal configurations or in an un-locked state). We will also explain how the robustness of the tunings has been assessed.

- AOCS impact on the operations of the LDR deployment This section will add some perspective on the operational aspect of the LDR deployment with the AOCS timeline and how it was optimized to reduce the total duration of the boom deployment.

- AOCS performance assessment during the LDR deployment. This last section will focus on the design validation and on the performance assessment strategy for each phase of the deployment. In particular we will explain how the LDR cinematic during each deployment step was modelled to determine its huge impact on the spacecraft dynamic, either in the case of a boom deployment step or the last reflector deployment step (“the blooming”).

Lightweight algorithms for collision avoidance applications Gonzalo Gomez J1, Colombo C1 1Politecnico di Milano Low Earth orbit is becoming an increasingly congested region, partly due to the flourishing of a new space economy. Bolstered by cost reductions in hardware and launch services, several companies are proposing and deploying new constellations for diverse space-based services such as communications or Earth monitoring. A significant example is the space-based-internet constellation by SpaceX, currently composed of 180 satellites but approved by the US Federal Communications Commission for more than 10,000. Furthermore, NASA and the US military are actively studying synergies with the private sector, seeking for benefits in terms of cost, deployment times, and robustness. In this increasingly congested scenario, collision avoidance activities are expected to increase both in frequency and complexity. The most recent warning call came on 2 September 2019, when ESA had to manoeuvre its Aeolus winds monitoring spacecraft to avoid a potential Close Approach (CA) with a Starlink satellite. Notably, SpaceX declared they were unaware of the increase in collision risk due to a bug in their paging system. This event underlines the need for improvement not only in tracking capabilities, but also in procedures, communications, and the efficient management of increasingly large sets of data. Among the different strategies to tackle this last issue, the availability of fast algorithms for the analysis of CAs and the design of Collision Avoidance Manoeuvres (CAMs) can prove a very useful tool.

For the last two years, the European Research Council-funded COMPASS project [1] has being developing new approaches and tools for the analysis and design of CAMs, with a focus on computational efficiency. These algorithms rely on analytical and semi-analytical methods for the modelling of changes in the orbital elements due to different types of forces, such as impulsive [2] or low-thrust CAMs [3]. The changes in orbital elements are mapped into displacements at the nominal CA through relative motion models, and the resulting displacements are analysed in the b-plane to separate the effects of phase change and geometry change. Different criteria for the optimization of the CAM are considered, mainly maximization of the miss distance and minimization of the collision probability. For fully analytical models the optimization problem can be reduced to an eigenvalue/eigenvector one [4,5], while the high computational efficiency of semi- analytical models allows for their use in parametric analysis or in combination with optimisation algorithms. The evolution of uncertainties before the CA has also been considered [6], particularly regarding the contribution of drag and Solar Radiation Pressure (SRP). The models allow for the computation of State Transition Matrices (STMs), which can be used for the propagation of covariances or design of several types of CAMs. A practical application example was the evaluation of CAM capabilities through area-to-mass modifications by a deorbiting sail, performed within the ESA-funded study “Environmental aspects of passive de-orbiting devices” [7]. Gathering all these developments, the Manoeuvre Intelligence for Space Safety (MISS) software tool has been recently introduced to provide an integrated approach to SSA-related operational activities [8]. The high computational efficiency of this tool makes it suitable for several applications, such as on-board and autonomous operations or in combination with Artificial Intelligence (AI) algorithms (e.g. to accelerate training).

This paper presents the latest developments in the underlying algorithms of MISS, with focus on the modelling of new perturbing forces and improving computational efficiency. The latter is achieved through the introduction of new analytical models, to complement or replace previous semi-analytical ones. The capabilities for the computation of STMs are also expanded, to better incorporate the effect of perturbations like drag and SRP. Particularly, by defining an extended state vector including the area-to-mass ratio and the drag and reflectivity coefficients, it is possible to directly apply the resulting STM to the design of CAMs by sails based on the modification of the effective area-to-mass. Additionally, the use of the b-plane projection to predict whether new CAs between both objects may occur later in time and how the planned CAM affects them is explored. Several test cases assessing the accuracy and numerical performance of the models are presented, quantifying the gains in computational times due to the use of analytical approaches compared to semi-analytical and fully numerical ones. Finally, the possible use of these algorithms for autonomous operations and their synergies with AI approaches is discussed.

References: [1] https://www.compass.polimi.it/ [2] J.L. Gonzalo, C. Colombo, and P. Di Lizia, “Analysis and Design of Collision Avoidance Maneuvers for Passive De-Orbiting Missions,” 2018 AAS/AIAA Astrodynamics Specialist Conference, Snowbird (UT), 19 - 23 Aug 2018. No AAS 18-357 [3] J.L. Gonzalo, C. Colombo, and P. Di Lizia, “A semi-analytical approach to low-thrust collision avoidance manoeuvre design,” 70th International Astronautical Congress, Washington, D.C., USA, 21-25 October 2019. IAC-19-A6.2.3 [4] B. A. Conway, “Near-optimal deflection of earth-approaching asteroids,” Journal of Guidance, Control, and Dynamics, 24(5):1035-1037, 2001 [5] C. Bombardelli, and J. Hernando-Ayuso, “Optimal impulsive collision avoidance in low earth orbit”, Journal of Guidance, Control, and Dynamics, 38(2):217-225, 2015 [6] J.L. Gonzalo, C. Colombo, and P. Di Lizia, “Drag- and SRP-induced effects in uncertainty evolution for close approaches,” 4th International Workshop on Key Topics in Orbit Propagation Applied to Space Situational Awareness (KePASSA), Logroño, Spain, 24-26 April 2019. [7] C. Colombo, A. Rossi, F.D. Vedova, A. Francesconi, C. Bombardelli, M. Trisolini, J.L. Gonzalo, P. Di Lizia, C. Giacomuzzo, S.B. Khan, R. Garcia-Pelayo, V. Braun, B.B. Virgili, H. Krag, “Effects of passive de- orbiting through drag and solar sails and electrodynamic tethers on the space debris environment,” 69th International Astronautical Congress, Bremen, Germany, 1-5 October 2018. [8] J.L. Gonzalo, C. Colombo, and P. Di Lizia, “Introducing MISS, a new tool for collision avoidance analysis and design”, First International Orbital Debris Conference, Sugarland, TX, USA, 9-12 December 2019. AstroBus product AOCS architecture for Airbus LEO missions Riant P1 1Airbus Defence And Space More than 10 years ago, Airbus started the definition of the AstroBus product line for LEO Missions. Today, this product line covers a large number of mission classes, including LEO optical Earth observation missions, Earth observation missions with radars or more specific instruments, and also LEO science missions. AstroBus platform has been today delivered, or will be delivered, to many customers, including ESA, ESA / Eumetsat, CNES, DLR, German and French MOD, Export customers, and also Airbus internal customer for Airbus own earth observation missions. The presentation will explain how this product is managed since more than 10 years, and how technical features of the AOCS have been added or adapted step by step to cover new types of missions. It will give an overview of ongoing and future evolutions of this product.

Microsatellite AOCS design for the agile mission MICROCARB Delavault S1, Rineau G2, Génin F1 1Cnes, 2CS Comunication&Systemes Global warming has become a major environment concern. It is driven by greenhouse gases, such as (CO2) which is the most important greenhouse gas produced by human activity. Therefore, limiting the impact of human CO2 emissions is now a crucial international challenge. In this context, France has decided to launch the mission MICROCARB, a satellite able to monitor the carbon dioxide cycle at a global scale. The project is led by CNES in collaboration with French scientific laboratories. It will improve the understanding of the exchanges between CO2 sources and sinks, their variability and sensitivity to climatic events.

The cost and planning constraints have naturally led the CNES to choose a MYRIADE generic platform for MICROCARB. The initial MYRIADE platform has been designed in the late 1990’s for the 150 kg class satellites. The generic Attitude and Orbit Control System (AOCS) allows the satellite orientation around 3 axes with demonstrated performances in nominal mode better than 0.005° for pointing accuracy and better than 0.02° for pointing stability. Initially, the agility was not a main constraint so the actuators have been chosen on geocentric mission hypothesis.

However, in order to meet the scientific objectives of the mission, specific pointing modes have been designed to measure the spectral radiance of the solar radiation reflected by Earth in different direction (at nadir, at glint point, on calibration stations, …). These innovative pointing modes require a challenging agility for a microsatellite. The most demanding mode for AOCS is the Fixed-Target used for the instrument calibration on TCCON (Total Carbon Column Observing Network) station. In order to make relevant measurements, the satellite has to maintain the instrument line of sight toward the TCCON station during several minutes, which requires an angular velocity ten times higher than the ordinary geocentric pointing. The probationary CITY mode is also very demanding in terms of agility as it consists in a scan resulting of three chained acquisitions over a 40km wide zone. It will demonstrate the capacity to map CO2 emissions at regional and city scales.

The MYRIADE platform has already successfully carried more than 15 satellites. Nevertheless, the generic bus needed to be upgraded to meet MICROCARB specific needs. This paper focuses on how the agility requirement has driven the design of the Attitude and Orbit Control System for MICROCARB. First, we will briefly present the mission and the satellite with a focus on the AOCS sub-system. Then we will explain how the MYRIADE AOCS architecture has been upgraded in order to be able to follow the demanding guidance profiles. Indeed, all the AOCS components have been improved from the equipment accommodations to the attitude estimation and control algorithms: - The wheels configuration has been optimized for the Fixed-Target mode which requires a high angular velocity around the pitch axis. A pyramid configuration oriented towards the pitch axis with a 45° tilt angle was chosen. This configuration results in an increase of the capacity by a factor 2.8 compared to a one- wheel configuration on the same axis. Moreover, the four wheels pyramid configuration for a three axes control offers a degree of freedom which has been used by the AOCS algorithms to optimize the available torque and angular momentum capacity. - To meet the performance requirements, the generic Myriade control algorithms also needed to be upgraded. As the controller structure is an inheritance from Myriade, the structured H-infinity method has been used to synthetize the MICROCARB controller with a higher band-width. In addition, in order to reduce the pointing error induced by the controller response time coupled with the high dynamics of the guidance profile, a feed-forward command has been implemented. - Moreover, in addition to wheels configuration modification, it has been decided to move towards a gyroless AOCS in order to manage gyro obsolescence and save space inside the platform. Then, in order to meet the stringent attitude knowledge error required by the mission, a dynamic filter has been implemented to propagate the attitude. The performances obtained with this design will be illustrated with highly representative simulation results for the Fixed-Target and City modes. At the end, the paper will conclude on how these AOCS improvements allow the MYRIADE bus to be compliant with MICROCARB mission requirements.

Passive rate damping of non-operational satellites in Low Earth Orbit to enable Active Debris Removal BENOIT A1, SOARES T1, ODDENINO D1, Van den BROEK M1, SANTOS A2, FRANCO N2, BRANCO J3, REIS G3 1Esa/estec, 2Lusospace, 3GMV

The growing number of Space Debris has become a major risk for the sustainable use of space and in particular of the LEO protected region (0 to 2000 km altitude). The global application of Space Debris Mitigation requirements is paramount and they are now strictly implemented by ESA, with the atmospheric re-entry of satellites in the LEO protected region within 25 years after their end-of-mission and the passivation of electrical and propulsion systems.

Unless the satellite altitude is sufficiently low to allow the orbit to decay naturally within the 25-year limit, orbit manoeuvres have to be performed, requiring the necessary on-board resources to be still operational. Although ISO 24113 and derived ECSS requirements impose a certain probability of success for the post- mission disposal manoeuvres, catastrophic failures could let a non-controlled, non-passivated satellite in the LEO protected region for a very long duration, as is the case with Envisat.

ESA under the umbrella of the Clean Space Initiative has then promoted a number of activities in the area of Active Debris Removal (ADR). First efforts have been devoted to the design of a chaser which would perform the rendezvous and capture with the orbital debris. However, it has become clear that the removal of unprepared satellites was extremely risky and complex. One of the main challenges driving the complexity of the rendezvous and capture of a debris is that its angular rates can build-up when the satellite becomes non-operational. Observations of several non- operational LEO satellites often show angular rates above 2 deg/s. The prediction and estimation of the angular rates and attitude of non-controlled satellites to be captured is crucial for the design of the ADR vehicle and to de-risk these critical operations. Unfortunately, the long-term dynamics evolution of non- controlled satellites in Low Earth Orbit had not been studied much so far by AOCS engineers.

In order to define how future satellites could be prepared for removal, a series of activities were initiated on the so-called Design for Removal (D4R) approaches to study different methods on the side of the target satellite to de-risk an ADR mission: passive rate damping and devices to facilitate tracking from ground and in-orbit rendezvous and capture. Such elements have already been included in the Space Segment Requirements Document of the six High Priority Copernicus Missions Phase B2/CD Requests for Proposals issued end of 2019.

This paper first presents analytical findings by ESA governing the dynamic evolution of the angular momentum (satellite with a dominant spin rate) under the main influences of the Gravity Gradient torque and the Magnetic torque generated by the satellite residual dipole. In order to damp the angular rates, kinetic energy dissipation means are necessary. Eddy currents in the structure are quite limited. A simple and elegant solution, taking advantage of the equipment already on board of LEO satellites was found through the short-circuiting of the satellite magnetic torquers (ESA patent pending Ref EP19182205). The motion of the satellites short-circuited magnetic torquers within the Earth magnetic field results in the induction of currents in their coils, dissipating energy by the Joule effect. Damping and tilting torques can be analytically derived and angular rates damping time constants roughly evaluated using some simplified assumptions as presented in the paper.

Results from a comprehensive set of High Fidelity simulations of the attitude motion in orbital space environment, obtained through a specifically designed simulation framework, are also presented and compared with the analytical framework above. Such mutual support between analysis and simulations is instrumental. The analytical results help to identify the parameters driving the dynamic evolution (e.g. satellite inertias and orientation versus the orbit pole and the Earth magnetic pole and characteristics of the magnetic torquers) and allow to extrapolate the findings of a set of simulations performed by industry to a variety of satellites inertia configurations and dynamic conditions.

Although the magnetic torquer detailed design is not the purpose of this paper, some real tests results are presented in order to validate important parameters driving the complex electromagnetic interactions between the rotating short-circuited magnetic torquers and the Earth Magnetic Field (electromotive force, induced current, energy dissipation and magnetic torque).

The preliminary list of general requirements and passive devices which have been elaborated by ESA for the six High Priority Copernicus Missions (Phase B2/CD KO around mid 2020) to facilitate Active Debris Removal (ADR) activities is then presented.

Preliminary conclusions and recommendations are finally proposed.

High Performance Pointing Systems

High accuracy image stabilization system for GEO High Resolution Missions Guercio N1, Durand G1, Christy J1, Perez Gonzalez A1, Sechi G2, Boquet F2, Deslaef N2, Aigouy G3, De Lepine X3, Massotti L4 1Thales Alenia Space, 2European Space Agency, 3CEDRAT Technologies, 4RHEA BV for ESA The demand for higher image resolution of Earth Observation and science missions is driving complex designs not only at payload level but also at platform level, with very accurate sensors and expensive means to compensate the platform perturbations. This is becoming very critical for up-to-date missions, and will become a show stopper for future very high demanding missions. High resolution Earth observation missions from geostationary orbit are a good example of such design complexity. Indeed, the telescopes require very large apertures and focal lengths in order to comply with the desired resolutions from such distances. Working very close to the diffraction limit, which allows reducing the overall size of the telescope, while providing high image quality, is one of the most important challenges of those systems. The image quality depends mainly on the scene-focal plane relative motion during the exposure time and on wavefront errors, which can be static (AIT, launch and gravity effects) or slowly time varying (thermo-elastic) and dynamics (structure and optical bench vibrations). Several solutions have been used in the past to improve the dynamic performance by constraining the platform with quiet mechanical environment and high accuracy attitude control systems. This work focuses instead on the design and performance assessment of an image stabilization system, through the development of a high fidelity line-of-sight simulator and using modern control approaches. Eventually, the development of an image stabilization breadboard is foreseen to increase the maturity of the solution using representative off-the-shelf components. The work presented has been performed in the frame of two studies: “High accuracy image stabilization system for GEO High Resolution Missions - Preliminary design and performance assessment” and “Fast Loop Image Processing for Line of sight Accurate Pointing” funded by ESA.

In the first section, the paper will present a literature survey to assess past or on-going missions that use the line of sight stabilization system for Sun observation (, HINODE, and HMI), inter-satellite telecommunication (Silex), Earth observation (ABI) and science (JWST). The pointing requirements and the solutions adopted are studied and compared, highlighting that the most demanding ones uses the main instrument optical path for sensing, adding dedicated image sensor (CCD) or a set of photodiodes. The motion compensation depends on required pointing accuracy and reactivity, and small fast steering mirror with voice-coil or piezoelectric actuators are generally selected. Several sensing and actuation means will be presented, taking into account the main drivers: - Line of sight measurement: observation of all line of sight modes, measurement accuracy, frequency, latency, impact on the platform and maturity level. Line of sight actuation: accuracy, control bandwidth, correction amplitudes, exported torques, impact on the platform and maturity level. After a trade of on actuation and sensing means, a specific fast steering mirror design and a selection of the best available image detector is discussed. An important task of the studies was the development of the high fidelity simulator, representative of the expected line of sight motion. The model includes a high frequency opto-mechanical dynamics, from finite element models and optical transfer functions, the AOCS model which includes a high fidelity reaction wheel model, the line-of-sight actuator and image processing equivalent model, and the representative disturbances. Modern robust control approaches, such as the H- infinity synthesis, have been used for the design of controllers with a definite structure dealing with model uncertainties. The procedure is applied to the three loops involved in the image stabilization system: fast steering mirror internal loop, line-of-sight stabilization loop based on image processing, and attitude control system. The performance campaign consisted in the validation of the simulation environment, a simplified robustness verification with worst case parameters and the assessment of the LOS stabilization performance in different operational conditions, such as the illumination conditions and the scene types. In conclusion, the closed loop simulations confirmed the clear benefit of the line of sight stabilization system for long exposure time acquisitions, even with bad SNR and low contrasted scenes, enabling the observation of a larger Earth surface (latitude/longitudes) and extending the mission availability. Eventually, the paper will present the prototype of the image stabilization system for the GEO-HR Earth observation mission that has been developed to increase the TRL of the proposed solution. The breadboard is composed by non-space qualified components, with functional and performance representativeness of the considered mission. The objective is to reproduce and stabilize a disturbance time series generated by the spacecraft high fidelity simulator, correlating the results with the expected model in the loop performance.

Reduction of flexible modes excitation during slews Garus A1, Eymard J, Labroquère J, Perez Gonzalez J, Rouziers F 1Thales Alenia Space The determination of an attitude guidance law and the tuning of a controller is a key component for the achievement of the mission and the management of satellite resources. In preliminary mission analysis context or during the system development phase, the definition of guidance laws and AOCS controller is essential to ensure the proper functioning of the satellite and mission performances. Indeed, the duration of the maneuvers impact the mission availability as the time required to rally from attitude A to attitude B cannot be used for the mission. Two main phases are present during an attitude rallying maneuver: • A phase of rallying from an attitude A to an attitude B, • A tranquilization phase once attitude B is reached. Reducing the duration of the rallying or tranquilization phase reduces the total rallying time and therefore increases the availability of the mission. Tranquilization phase has two different purposes: • Allowing controller convergence, • Allow sufficient damping of flexible modes of the satellite. This paper presents methods to reduce mode excitation during the maneuvers, leading to an overall improvement of mission availability.

Characterization of SADM induced disturbances and their effects on spacecraft pointing errors Alazard D1, Cumer C2, Sanfedino F1, Oddenino D3, Brugnoli A1 1ISAE-SUPAERO, 2ONERA, 3ESA-ESTEC This paper presents the modeling and analysis of the induced oscillations due to the Solar Array Drive Mechanism (SADM) on the pointing performances of Earth-observation satellites. First the SDT (Satellite Dynamics Toolbox) is used to obtain a Linear Parameter Varying model of the spacecraft fitted with a flexible solar panel parameterized according to its angular configuration. Then a detailed model of the SADM including the stepper motor and the gearbox is included using the TITOP (Two-Input Two-Output Ports) approach. Such an approach allows to build the dynamic model of the structure just by assembling the sub- structure models. TITOP models consider the 6 d.o.f. (degrees of freedom) to take into account all the geometrical, kinematic and dynamic couplings required to capture the propagation of the perturbations inside the overall flexible structure. Moreover these models allows the user to evaluate the pointing errors on the 3 axes. Parametric robustness analysis tools are then used to isolate the worst harmonics of the various SADM imperfections and the solar array angular configuration from the pointing performance point of view.

On-board Real-time Calibration of Non-synchronous Attitude Sensors and Gyros Curti F1, Toglia C2, Perez Gonzalez J3, Sechi G4 1School of Aerospace Engineering, Sapienza University of Rome, 2Thales Alenia Space , 3Thales Alenia Space, 4ESA ESTEC - AOCS & Pointing Systems Section The aim of this paper is to present the contents and major findings of on-board calibration filters developed in the framework of the ESA study “Future AOCS Enabling Technologies – FEAT”. The idea behind the proposed methodology is to design filters that, while performing the canonical gyro-stellar estimation of attitude and angular rates, can estimate also attitude sensors and gyros parameters that are directly calibrated in flight. The attitude system configuration of this study is characterized by attitude sensors with at least one Star Tracker (STT) and a Fine Guidance Sensor (FGS), and one FOG redundant IMU with 4 gyros. The FGS is more accurate with respect to a STT (of the order of milli-arcsec), but it can only operate in a quasi-inertial attitude at very low angular velocity (< 1 arcsec/s) and low rate (0.5 Hz). A calibration procedure is proposed in order to estimate the mounting misalignments of STTs and gyro sensing axes, and the biases and scale factors of each gyro channels. The approach proposes two subsequent calibration phases to exploit the FGS performances: • Calibration Phase 1 (CP1): the FGS is chosen as Master Sensor (reference sensor) and the attitude is quasi-inertial (angular rate within the FGS working range). This phase aims at estimating the static misalignments of each STT starting from an a priori knowledge of its mounting matrix, e.g. provided during satellite AIT/AIV. They are modeled as three small Euler angles between the nominal and real STT mounting matrix. • Calibration Phase 2 (CP2): the STTs misalignments are known from CP1; a STT is selected as a Master Sensor (reference sensor) and the FOG parameters are estimated during suitable calibration maneuvers (“corkscrew” attitude maneuvers). This phase aims at estimating: 4 Symmetric Scale Factors (SSF), the misalignments relative to each axis of FOG sensing unit (2 parameters for each FOG channel), and FOG biases (4 for each channels). The filters’ design is based on the Extended Kalman filter (EKF) approach, which gives a recursive estimator that can easily be implemented on-board. In the CP2, the four-axes FOG introduces redundant parameters for biases, SSFs and sensing unit misalignments. As a result, the filter needs to manage this redundancy. This problem has been addressed in previous works by Pittelkau (see as examples Ref. [1], Ref. [2] and Ref. [3]), carrying out an EKF structure based on the range space and the null space of the physical parameters. This structure allows reducing filter dimensions and design complexity. Moreover, in order to ensure the observability of the calibration parameters, Pittelkau proposes persistent excitations of the filter through suitable maneuvers by defining the dynamic observability. Sensor devices introduce delays due to the acquisition and processing of the measurements. As result, the classical estimation approaches need to be re-formulated to include those delays through a fusion of the measurements. In the present study, the assumption is to have non-synchronous measurements, which leads to model the estimator managing known delays among the available sensors. In order to cope with non- synchronous measurements, the filter is designed by introducing extended state and outputs, which incorporate the delays. This work shows the design of a multi-rate EKF calibration filter (MECF) based on the range-null approach able to manage the redundancy and known delays of the avionics. A sensitivity analysis is carried out in order to establish how the filter performances change as a function of unpredicted covariance values. The results show that gyros angle random walk noise and the STT noise mainly affect the filter steady-state performance. In addition, a suitable analysis is conducted to ensure the dynamic observability of MECF. The analysis shows that the sinusoidal calibration maneuvers exciting the satellite during CP2 need to have different frequencies in the components in order to get the dynamic observability. The comparison of the test cases shows as large values of the amplitude reduces the time response of the filter. Nevertheless, large values of angular rate reflect in the drawback of large values of the control torque. On the other hand, the angular acceleration increases with the frequency, thus the torque rate capacity of the actuators needs to be considered. The results of pseudo-random maneuvers are comparable with the behavior of sinusoidal maneuvers with different frequencies. As a result, the dynamic observability is obtained with periodic maneuvers with different frequencies in its components or a band-limited pseudo- random maneuvers, whatever is the amplitude of the angular rates. The calibration procedures are tested on a satellite high fidelity simulator. The high fidelity simulator is composed of a nonlinear satellite plant undergoing environmental disturbances and equipped with a set of fully representative models of attitude sensors (three star trackers and one FGS), one FOG rate sensor (integrated gyro with 4 channels), one set of 4 reaction wheels. The avionics algorithms are composed of sensors preprocessing algorithm, reference generator, controller, and the calibration filter to estimate calibration parameters and provide filtered attitude and rate. A set of simulations are carried out to test the performance of the calibration algorithms under two different scenarios: a nominal case and a degraded case. Such scenarios reflect the operational modes of reference missions and cover different reference attitude and sensors availability.

References 1. Pittelkau, Mark E. Kalman Filtering for Spacecraft System Alignment Calibration. Journal of Guidance, Control, and Dynamics. 2001, Vol. 24, No. 6. 2. Pittelkau, Mark E. Advances in Attitude Determination with Redundant Inertial Measurement Units. AAS/AIAA Spaceflight Mechanics Meeting. Tampa, Florida, 22–26 January 2006. 3. Pittelkau, Mark E. Everything Is Relative in Spacecraft System. Journal of Spacecraft and Rockets. 2002, Vol. 39, No.3

Spacecraft Line-of-Sight Jitter Mitigation and Management Lessons Learned and Engineering Best Practices Dennehy C1, Wolf A2, Swanson D3 1NASA, 2Jet Propulsion Laboratory/California Institute of Technology, 3Raytheon Space and Airborne Systems Predicting, managing, controlling, and testing spacecraft line-of-sight (LoS) jitter caused by micro-vibrations due to onboard internal disturbance sources is a formidable multidisciplinary engineering challenge, especially for those missions hosting high-performance, vibration-sensitive optical sensor payloads with stringent pointing stability requirements. Clear trends exist within NASA and ESA toward planning technically aggressive spaceflight missions that include ultra-high-performance optical payloads with delicate, highly vibration-sensitive scientific and observational instruments. The missions include JWST, EUCLID, LUVOIR, HABEX, OST, MTG, and GOES. To successfully meet these challenges, NASA and ESA will need to leverage and build upon their collective past experiences and lessons learned in addressing micro-vibration problems. Looking back, one sees that both space agencies, together with our industry partners, have a long, technically rich, and impressive history of solving the difficult engineering problems associated with managing, controlling, and testing spacecraft jitter and micro-vibrations.

To identify lessons learned and best engineering practices NASA sponsored a two-day Spacecraft LoS Jitter Workshop on October 23–24, 2019, near Goddard Space Flight Center in Columbia, Maryland. In addition to NASA and JPL participants, the ESA was actively engaged in this workshop in the GNC and mechanism areas. Representatives from NASA’s industrial partners, independent consultant subject matter experts, and members of academia also participated in the workshop. Collective NASA-ESA experiences with undesirable jitter perturbing payload instrument pointing and pointing stability have taught us the imperative of having a multidisciplinary perspective on this problem. The workshop’s goal was to provide a multidisciplinary forum to elicit deeper understanding of the issues related to solving the spacecraft LoS jitter/micro-vibration problem. A primary objective was to identify and share best practices, rules of thumb, and options for jitter- related activities in the following technical areas: spacecraft pointing system architecture trades and definition, requirements definition and flowdown, modeling and simulation tools and techniques, subsystem and component characterization testing, spacecraft system-level end-to-e testing, and overall jitter/micro- vibration risk-reduction approaches and techniques. Participants shared their experiences and lessons learned across a spectrum of relevant engineering and technology disciplines. The attendees included specialists in GN&C, pointing systems, mechanisms, structures, finite element modeling, isolation systems, system engineering, and system testing. The workshop was also successful in identifying technical and programmatic (contractual) issues, barriers, and challenges related to solving the spacecraft LoS jitter/micro-vibration problem. This paper will further describe the motivation for the NASA Spacecraft LoS Jitter Workshop and summarize the findings, observations, and recommendations identified during the workshop.

Several lessons learned and engineering best practices identified at this multidisciplinary workshop are expected to improve the engineering practices for solving the jitter/micro-vibration problem. For example, one recommendation emerging from the Spacecraft LoS Jitter Workshop is the need for NASA and ESA (either independently or collaboratively) to create a “Jitter Engineering Guidelines” handbook capturing the most relevant lessons learned and best practices for the community of practice. As envisioned by the workshop participants, this handbook would also document key jitter-related nomenclature so as to establish a common engineering lexicon, thus improving effective technical communications across the multiple disciplines and within the community of practice. Another recommendation is for NASA and ESA to develop and provide jitter diagnostic in-situ instrumentation to their programs and projects. Most importantly, flying such an in-situ diagnostic package would permit the capture of in-flight jitter-related data to support the process of post-launch validation of pointing system models used for pre-launch jitter predictions. This diagnostic instrumentation package would also provide critical on-board performance data useful for jitter- related anomaly resolution and/or the tuning of overall pointing system parameters. Regarding ways to overcome non-technical (i.e., business and or contractual) issues, one idea surfaced at the workshop was the use of generic state-space dynamic models, to allow analysts to avoid proprietary data concerns by our industrial partners. As mentioned above, this paper will capture all the findings, observations, and recommendations identified during the workshop last October.

In-Orbit Experiences and Demonstrators

In-flight Experience on GNC Design Challenges for BepiColombo Steiger C1, Belien F2, Fugger S2, Casasco M3, Altay A4, Montagnon E1, Budnik F1 1ESA/ESOC, 2Airbus Defence & Space, 3ESA/ESTEC, 4Terma GmbH BepiColombo is an ESA cornerstone mission to Mercury in collaboration with the Japan Aerospace Exploration Agency (JAXA), with Airbus Defence & Space as prime industrial contractor for the ESA contribution. The mission’s two scientific orbiters –ESA’s Mercury Planetary Orbiter (MPO) and JAXA’s MIO orbiter– were launched together in Oct 2018 as a single composite spacecraft, including a MIO sunshield and a module with electric propulsion (EP) to support the 7-years cruise phase with planetary flybys at Earth (1x), Venus (2x) and Mercury (6x). Orbit insertion in December 2025 by weak-stability capture will be followed by a series of manoeuvres and module separations to inject the two spacecraft into their respective orbits. The scientific phase of the mission will last one Earth year, with an extension possibility of another Earth year. Flight operations of the composite spacecraft and of the MPO at Mercury are performed by ESOC Darmstadt, Germany.

Following launch in Oct 2018, near-Earth commissioning took place until late 2018. The first two electric propulsion “thrust arcs” –periods when EP is used continuously to adjust the trajectory– were performed from Dec 2018 to Feb 2019 and from Sep 2019 to Nov 2019. Operations in 2019 also included activation of a new platform on-board software. The S/C is now bound to return to Earth in April 2020 for the first planetary swingby, after which the S/C will travel to the inner solar system, with a at Venus in Oct 2020.

Key challenges driving the BepiColombo GNC design include S/C modularity, severe pointing constraints in the harsh environment at close Sun distances, and EP operations in cruise. As described in the GNC 2017 papers on S/C modularity and GNC FDIR and autonomy challenges, this leads to a complex GNC design and operations. Taking into account the wealth of in-flight experience gained, the following aspects are of particular interest: - S/C modularity: the BepiColombo GNC is required to deal with different S/C configurations and the consequent complex ground-provided context information, with several sensors and actuators duplicated for different configurations. Following thorough validation of operations in all configurations in pre-launch testing, the S/C is now in cruise configuration. Units not used in this configuration have been checked out as far as possible prior to their operational usage at Mercury. - Electric propulsion operations: the EP system employs 4 thrusters (QinetiQ T6 gridded ion thrusters) mounted on pointing mechanisms. Two thrusters can be used in parallel, with a power demand of up to 11kW under maximum thrust. Each thrust arc requires a custom spacecraft attitude, with the attitude and solar array guidance profiles commanded by ground. Weekly interruptions of EP operations are needed for orbit determination, collecting radiometric data when not thrusting. Ground coverage in the thrusting attitude using one of the moveable antennae is not guaranteed, and needs to be analysed on a case-by-case basis for each thrust arc. By now two thrust arcs with a total thrusting time of over 4 months have been completed, using both dual and single thruster configurations. - Solar array guidance and power constraints: despite approaching the Sun, the mission is highly power- constrained. Below 0.62 AU Sun distance, the solar arrays have to be offpointed to avoid overheating, while still extracting maximum possible power in these conditions. The available power drives the commanded thrust levels during EP arcs, in turn affecting the trajectory. This requires a complex planning process to get the most of the system, de-risking the thrust arcs as much as possible. While not yet at Sun distances where solar array offpointing to avoid overheating is required, the overall process was already successfully run during the first two EP thrust arcs. Outside of thrust arcs, power demand is much reduced, allowing to offpoint solar arrays to limit degradation. - Failure Control Electronics (FCE): the strict S/C attitude and pointing constraints at Mercury and during most of the cruise phase require a hot redundancy approach to guarantee attitude control. A second on-board computer –the FCE– was implemented, running a simplified GNC software, ready to take over attitude control temporarily in case of transient outage of the prime on-board computer at safe mode entry. So far, three safe modes were entered, with the FCE intervening as expected. - Guidance for launch: due to the strict S/C attitude constraints and the required accuracy of ground-provided ephemeris data, default guidance settings are complex, consisting of >4000 parameters dependent on the launch date. This made generation of these settings on industry side impractical – they were instead generated by ESOC for upload on the launch pad, requiring to put in place a custom interface and process for provision of this data from ESOC to the industrial prime, and eventually to the AIT team at the launch site. - Attitude and antenna guidance in cruise: the original baseline for cruise was to have the S/C rotating around the sun line with a 3h period to save fuel. As this would have led to antenna movements beyond the allocated budget as well as interruptions during ground contacts, a dedicated “flip” attitude strategy was designed. The S/C completes a full rotation around the sun line in typically 24h, spending extended amounts of time in attitudes differing by a 180 deg rotation around the sun line, with fast slews in between. This approach has been used extensively, allowing to cover ground contacts without interruptions, while retaining the fuel- saving aspects of not staying in a fixed attitude throughout.

Starting with an overview of the mission, the GNC subsystem and operations concept, GNC-related in-flight experience of the first 1.5 years of flight will be presented, with a special focus on the return of experience for mission-specific GNC aspects. An outlook on upcoming GNC activities will be provided as well. Solar Orbiter AOCS in flight return of experience Cantiello I2, Monroig G1, Palombo E2, Pereira V2 1Airbus Defence And Space Ltd, 2European Space Agency (ESTEC) ESA Solar Orbiter’s journey around the Sun starts in the early morning of 10th of February CET, launched from Cape Canaveral on V 411 rocket supplied by NASA, for an unprecedented mission to study our star up-close, reaching a perihelion of 0.28AU and a solar inclination at mission end of 33deg. This paper presents an overview of the AOCS operations and in-flight results in the first 4 months of the mission, covering LEOP, Near Earth Commissioning Phase (NECP) and beginning of Cruise Phase (CP). The LEOP is defined as lasting 3 up to 7 days from T0 (launch) until the completion of the initial trajectory correction manoeuvre, which marks the beginning of the NECP. NECP is planned to last up to 3 months from launch, before entering in Cruise Phase, which is mainly dedicated to S/C monitoring and payload performance verification. The activation of the AOCS units required for the initial Sun Acquisition and Survival Mode (SASM), i.e. Gyro, Fine Sun Sensors and CPS Thrusters, takes place as part of the LEOP auto-sequence triggered immediately after the SW start up and separation detection. The first Rate damping and Sun Acquisition performance is analysed, alongside the following Solar Array deployment phase and entry to Earth Strobing mode, with a rotation around the Sun line, marking the end of the auto-sequence. Further commissioning activities during LEOP are presented, including the commissioning of the AOCS units and their performance assessment (in chronological order Fine Sun Sensors, RW Checkout, Star Tracker checkout), first entry in Wheel Safe Mode, Wheel Offloading, entry in Normal Control Mode, and deployment of Radio Plasma Wave (RPW) Antennas and I-boom. The performance of the Control Mode, entered for Launch Correction Manoeuvre, is also illustrated, together with the activation and calibration of the Accelerometers used for Delta-V operations. The results from the Gyro calibration process and redundant STR commissioning, planned in Near Earth Commissioning Phase, are also presented. In conclusion, the key AOCS Subsystem and Hardware unit in flight behaviours are summarized with an overview on the status of the S/C to date.

Visual Sensor Suite – Flight Experience on MEV Griebel M1, Schmidt U, Würl R 1Jena-optronik Gmbh Jena-Optronik GmbH, located in Jena/Germany, has profound experience in designing and manufacturing star trackers since the early 80ties. Today the company has a worldwide leading position in supplying geo- stationary and earth observation satellites with robust and reliable star tracker systems. Embedded in a development contract (17317/2003/F/WE) from the European Space Agency, CMOS Active Pixel Sensors were introduced for advanced star tracker technologies. Using these technologies Jena-Optronik GmbH developed a multi-purpose visual sensor suite (VSS) consisting of different optical heads and an advanced electronic controller box. The optical head uses FaintStar the 2nd generation of CMOS active pixels sensor developed within ESA funded programs. The sensor can be combined with several optics for narrow and far field applications. The e-box is equipped with a LEON 3 processing unit and can be extended to full redundant operation. The wide range of configuration enables the system to different applications such as Rendezvous and Docking (RvD), Space Situational Awareness (SSA), Surveillance and Star Tracking as well. Fields of view of 19deg and 68deg with corresponding baffles of 26 and 85 SEA are available, others on request. Typical configuration is a narrow field of view (NFOV) optical head stereo camera pair for far range object detection from 45.000m down to 10m combined with a wide field of view (WFOV) optical head stereo camera pair for close range object detection from 0.5 to 100m. So that there is a range overlap of 90m. A further pair of WFOV optical head is installed for redundancy and additionally surveillance in close range. The development and qualification period were performed in about 2 years of concentrated work in closed collaboration with the first customer Northrop Grumman Innovation System to provide the Visible Sensor Suite for the first Mission Extension Vehicle – MEV1. Short time after delivery of the first flight set it was launched in October 2019 and successfully switched ON 5th of December 2019. The paper will present detailed information about the qualification phase and the achieved key parameters. We will give mission experiences and outlook to ongoing developments for high performance star trackers and solutions for constellation approaches.

Flight of a New Miniaturised and Integrated Spaceborne GNSS Receiver Unwin M1, Palfreyman A1, de Vos van Steenwijk R1, Miles S1, Rawlinson J1 1Surrey Satellite Technology Ltd The NewSpace Cubesat movement has set standard platform requirements for integration of modules in a compact Cubesat platform. Often the available GNSS receivers have not been developed specifically for Cubesats and so need to be hosted on a separate interface and motherboard, leading to a deeper module with a higher power than necessary.

GNSS receivers have been produced for small satellites by SSTL for over 20 years; the SNAP-1 nanosatellite predated Cubesat standards but was only 7 kg, and was launched in 2000 and used an SGR-05 miniaturised GPS receiver to monitor orbit manoeuvring. The SGR-05U has subsequently been flown on a number of Cubesats, and other larger SSTL GNSS receivers have been used for large missions, attitude determination, high altitudes, and for GNSS Reflectometry.

Two needs arose at SSTL, one for a GNSS receiver for a Cubesat mission called -1, and secondly for a GNSS receiver that would be integrated into an avionics module for an efficiently compact single board for a satellite that can accommodate somewhat larger payloads than possible on a Cubesat. The GNSS receiver must be low cost to manufacture, take low power, and there was a need for the use of GPS, Glonass and in future, Galileo and Beidou signals. A second front-end would sometimes be required to support inertially pointing missions. As a result, the SGR-Ligo was developed, and produced in two variants – Cubesat and integrated avionics. The core design of the SGR-Ligo was based upon a low-power COTS-based FPGA with soft-core processor and proprietary GNSS correlators derived from SGR-ReSI and SGR-Axio receivers. There are mitigations against radiation through TMR of registers, EDAC of memory, and a fast acting current detection switch protecting against SELs. It has two front-ends that can support GPS and Glonass (at a slightly different frequency to GPS), or be configured to operate from two GPS antennas on either side of the platform. Low power operation with passive antennas is possible, and the power can be reduced through shutdown of unwanted circuitry, achieving continuous positioning at power of lower than 0.5 watts. Preliminary experimentation on the TDS-1 mission achieved Galileo signal acquisition and tracking in orbit, verifying the correlators’ suitability for multi-constellation applications.

The Vesta-1 mission was a Cubesat mission commissioned to claim frequency filings in orbit, and was developed in SSTL using a mixture of SSTL and bought-in modules. The SGR-Ligo was embarked on Vesta- 1, and included a bare passive patch antenna on the space-pointing facet that could be accommodated within the Cubesat launch envelope. The satellite was launched in December 2018, along with 64 other satellites. All satellites are tracked by NORAD and allocated orbital elements, but a common problem is that elements can be allocated to satellites incorrectly, and it can take days or weeks to correct this. It became clear after some patchy communications that VESTA-1 did not have the correct elements. Fortunately the SGR-Ligo was operated shortly after launch, and its data was used to identify the correct NORAD orbital elements.

The first implementation of the integrated SGR-Ligo is on 17kg technology demonstration satellite DoT-1, which was launched in summer 2019 with the intention to de-risk future SSTL missions using this architecture. The SGR-Ligo has been tested successfully on this mission. Now the data from the two front- ends is being diverted to a separate processor to permit the implementation of a new demonstration of GNSS- Reflectometry, collecting the reflected GNSS signals to take geophysical measurements off the Earth’s surface. This is a stepping stone towards a candidate ESA mission called HydroGNSS that will sense soil moisture and other essential climate variables using reflected GNSS signals – this payload design is derived from the SGR-Ligo technology.

The SGR-Ligo is currently being prepared for launch on two more satellites – one small target satellite due for launch in 2020, for use as a demonstration of orbital debris removal. Secondly it is in the integrated coreDHS avionics of a satellite being built for a customer engineer at SSTL. In future, the SGR-Ligo will be at the heart of every coreDHS flown on SSTL’s upcoming constellations, but for smaller and alternative form-factor missions, such as Cubesats there will remain a requirement for the stand-alone SGR-Ligo space GNSS receiver.

Auriga CP and Auriga SA flight heritage Piot D2, Charavel R1, Gelin B1, Maksimous M1, Charavel R1, Lieutaud A2, Vignon B2, Miguel R3, Charront Y4 1Sodern, 2Airbus OneWeb Satellite, 3Héméria, 4CNES On February 27th 2019, twelve Auriga CMOS star trackers were successfully launched aboard the inaugural six satellites for the Airbus OneWeb Satellites constellation. The next launch for this constellation, scheduled in early 2020 will bring into flight tens of Auriga. In addition, the stand alone version with the optical heads equipped with an electronic unit was launched on December 18th 2019 aboard CNES Hemeria 12U nano-sat ANGELS. By mid-2020, hundreds of Auriga will be produced offering unusual large quantities for production and exploitation trend analysis. The Auriga was originally designed for constellation applications and is capable of delivering accurate & robust attitude solutions in a compact, lightweight form factor with short lead times. This paper presents the results of this new and revolutionary star tracker product line including attitude accuracy, stray-light levels, sun exclusion angle, operation with Moon in the field of view and also manufacturing return of experience for this star tracker mass production.

An overview of the missions where Auriga-CP and Auriga-SA are on board is presented.

The Auriga family in its two form factors is described: AURIGA-CP: a centralized version comprising up to three AURIGA Optical Heads (OH) and software hosted on the On Board Computer (OBC), which computes the quaternion, AURIGA-SA: a stand-alone version comprising up to three AURIGA Optical Heads (OH) and a dedicated Processing Unit (PU), which computes the quaternion.

The global excellent behavior of the star tracker is shown on some examples of typical results obtained for several orbits with the good operation of the multiple heads management based on the same principle as HYDRA which provides the same benefits in terms of availability and resilience. Lost-in-space acquisition is required on only one OH, the successful OH allowing self-designation for the other ones. Blinding and occultation of one OH is autonomously managed; the unavailable OH is kept in tracking mode thanks to the other ones eliminating the need of lost-in-space reacquisition. The use of the blended solution computed over the total number of stars for each OH allows accuracy improvement with the same performance around all axes. It also improves the robustness when one OH is looking along a line of sight with only faint stars. An additional feature is line of sight correction to remove offset between multiple heads. The straylight properties (Sun Exclusion Angle, straylight levels) are verified for several optical heads thanks to the information provided in the telemetries. The nominal behavior of the star tracker with the moon in the field of view is also presented. The attitude accuracy is analyzed through Power Spectral Density to be able to transform it into the standard error terms classification, as described in the ECSS-E-ST-60-20C on stars sensors terminology and performance specification, in Low Frequency Spatial Error (or FOV errors), High frequency Spatial Error and temporal noise. The results are compared to the expected values of the performance assessment. This method allows a detailed and deeper validation than a standard method that gives only the temporal noise. Single star accuracy is also presented as it provides an even more precise verification because the averaging effect at quaternion level can mask a discrepancy on one star for example. Analysis of the discarded stars is also presented for parametrization checking of the thresholds used for the coherence tests. A few parameters tuning is proposed to optimize the measurements of the star tracker thanks to the lessons learn brought by this detailed analysis.

The imaging LiDAR of the in-orbit demonstration RemoveDebris Ummel A1 1CSEM SA Started in 2013, RemoveDebris aimed at demonstrating in-orbit active debris removal (ADR) enabling technologies. Net and harpoon were tested end of 2018 and early 2019 as capture means. Inflated structure and dragsail were evaluated to change the form factor of a target and to slow down the 100 kg main satellite. For the purposes of the mission, two 2U CubeSats were ejected by the main satellite and used as targets instead of real space debris. This paper is focused on the flash imaging LiDAR that was assessed as vision-based sensor allowing debris tracking, monitoring and capture under the supervision of the GNC ADR computer. RemoveDebris was special, it disrupted our traditional development approach for the LiDAR. Unconventionally, within the same project we started at TRL2 (paper concept) based on previous flash LiDAR developments, and ended at TRL8 with an in-orbit demonstration. The development and budget risks mitigation started right at the beginning of the project with the choice of the components. None of the ones used is space qualified. The cost of such component simply does not match the budget available. Moreover, the key components such as the time-of-flight detector and compact illumination source do not exist yet in the form of ruggedized versions. Instead, we used commercial off-the-self components available in 2013. We designed the LiDAR with these components and assessed the unit according to test plans and methodologies grouped now under the umbrella of the “new space” trend. It survived in LEO for 1.5 years providing the consortium a large set of color and 3D images taken in-orbit in real space operational conditions. The most challenging operation parameter was to cope with the background sunlight either reflected by the Earth or the target. Following several trade-offs, in-orbit experiments time windows were chosen to favor strong illumination of the target by the Sun. For sensors with their own illumination source like a LiDAR, this is not the preferred operational situation. In addition, the settings defining the image dynamic range had to be defined on Earth several days before the experiment took place with no mean to adjust them in real- time. The effect of the strong sunlight is visible in the images. For some of them, some areas are saturated while some others provide good signal allowing to measure the satellite-target distance for these parts. From these data several secondary figures mandatory to capture debris can be calculated such as: 3D target image, relative attitude, velocity and rotation of the target, etc. The RemoveDebris experience is unevaluable. It shows what it takes in a real situation to capture good quality 3D images with flash imaging LiDAR, especially when mission objectives are at stake and images must be captured outside ideal conditions. Nevertheless, the LiDAR captured images of the target when it was only a few pixels large in the field-of-view at distance larger than 100 m.

Aeolus AOCS In-flight Performance Davies A1, Chapman P1, Hyslop A2 1Airbus DS, 2Vitrociset/Leonardo, for ESA The Aeolus spacecraft is a demonstrator for a novel technique of measuring wind speed profiles from space using a Lidar, with the aim to provide additional data to improve the accuracy of weather forecasting. It is the 5th mission in ESA’s Earth Explorer program. The spacecraft underwent a long development led by Airbus (spacecraft and AOCS prime), and Aeolus was successfully launched into a Sun-synchronous -dusk orbit by a Vega launcher in August 2018. In early 2020 its data has started to be used in regular weather forecasting.

The AOCS is a bespoke design to both provide fine pointing for the payload and satisfy multiple safety requirement constraints for payload exclusion zones. The paper will describe the AOCS design, the development challenges for the AOCS, and the inflight behaviour.

Aeolus hosts a single payload, the Atmospheric Laser Doppler Instrument (ALADIN), which determines Doppler shifts of the backscatter from ultra-violet laser pulses fired at an angle toward the Earth, allowing the wind component along the horizontal line of sight to be determined. The height can be deduced from the time delay between emission of the laser pulse and reception of the backscatter signal. The payload has multiple safety pointing constraints for the AOCS, which evolved during project development as the payload was developed: the telescope has a lifetime time limit for exposure to the atmospheric flow (Aeolus operates at unusually low altitude, and ATOX may damage optical surfaces), and the telescope has short-term time limits for sun-pointing (it may sun point for short intervals, to limit heating). The instrument pointing constraints had a significant impact on the AOCS design. The instrument requires fine pointing for normal operations, with a steering law guidance profile that nominally provides zero Doppler shift for ground reflections of the laser (considering the local velocity of the earth surface).

The AOCS contains 4 active modes: Initial Acquisition Mode (IAM), Normal Mode (NM), Thruster Control Mode (TCM), and Safe Mode (SM). IAM performs rate damping & attitude control using a Coarse Earth and Sun Sensor (CESS) & IMU for attitude & rate measurement respectively. Thrusters (RCS) & magnetorquers (MTQ) supported by magnetometer readings (MTM), are used for actuation. NM attitude control uses an Autonomous Star tracker (AST) & IMU for inertial attitude & rate measurement respectively. GPS measurements are used as inputs to the guidance function to generate the required pointing. Reaction wheels (RW) are used for actuation, wheel momentum being managed using MTQ, supported by an onboard magnetic field model. TCM performs delta-v for station keeping or orbital debris collision avoidance, and momentum offloading, and uses the same control concept as NM, but using RCS for actuation. Safe mode performs rate damping & attitude control using the same control concept as IAM, but with redundant hardware and the IMU is replaced by the Rate Measurement Unit (RMU).

For a LEO spacecraft relatively large momentum size wheels (45 Nms) were selected due to the low altitude. The CESS was selected to provide simultaneous omni-directional earth and sun vector measurements with eclipse detection, and with in-built redundancy using triple majority voted sensors, enabling it to be safely used for control in the IAM loop, in FDIR for failure detection of an anomalous attitude, and for control in safe mode.

This paper will discuss the more challenging aspects of the AOCS design and the in-flight performance of those elements, units and overall AOCS.

In-Orbit Experiences with the Spin Stabilized Attitude Control System of Eu:CROPIS Heidecker A1, Schlotterer M1 1German Aerospace Center (DLR) The German Aerospace Center (DLR) launched on 3rd December 2018 the Euglena Combined Regenerative Organic food Production In Space (Eu:CROPIS) mission. Eu:CROPIS is a mission of DLR’s compact satellite class with the objective to test several biological experiments at different levels of artificial gravity. A dedicated attitude control system (ACS), which utilizes a spin stabilized concept, is part of the satellite to achieve this objective. The following paper reflects the challenging aspects of the attitude control system design and wraps up one year of flight experiences related to the ACS.

The satellite design for the mission Eu:CROPIS is driven by two factors. On one hand all primary and secondary scientific payloads shall be accommodated. On the other hand the satellite bus development shall be performed in such a way that DLR accumulates engineering knowledge and prepares the way for further scientific compact satellites.

There are four payloads integrated on board of Eu:CROPIS. The primary and name giving payload Eu:CROPIS is provided by the Friedrich-Alexander-University of Erlangen-Nürnberg and DLR Institute of Aerospace Medicine, Cologne. The second payload PowerCell is delivered by NASA AMES Research Center. A third payload RAMIS (RAdiation Measurment In Space) is provided by DLR Institute of Aerospace Medicine, Cologne. It records the radiation environment in low earth orbit. The fourth payload is SCORE (SCalable On-boaRd computing) which is a technology demonstrator from DLR Institute of Space Systems, Bremen.

The driving requirements for the satellite design are to expose first and secondary payloads to different levels of gravity that are representative for Mars and Moon environment. Therefore, the whole satellite is put into a constant rotation between 5 to 31.16 rpm and the payloads are oriented in such a way that the centrifugal force is simulating gravity.

The satellite has a mass of 226.62 kg and a cylindrical shape with a diameter of ~1 m and a height of ~1m. Four solar panels are unfolded after launch and initial stabilization. The nominal power consumption is about 200 W.

The final orbit for Eu:CROPIS was selected in such a way that a piggyback ride share option is available and DLRs ground station in Weilheim can be used for nominal operation. Additionally, it is a hard requirement that the altitude does not drop below 550km at mission end out of the increased atmospheric disturbance torques. Therefore, Eu:CROPIS was placed on the 3rd December 2018 by a Falcon-9 rocket into a circular sun synchronous low earth orbit.

The Eu:CROPIS spacecraft utilizes a nowadays rarely used attitude stabilization concept. Its scientific payloads require a continuous rotation rate and therefore active spin stabilization is selected. The spin is controlled purely by the interaction of magnetic torquers with the geomagnetic field. No other actuators are available. In addition to the three orthogonally arranged magnetic torquers the spacecraft carries , sun sensors, gyroscopes and two GPS receivers. The on board attitude determination algorithm is an Unscented Kalman Filter (UKF). It estimates angular rate, magnetic field and sun vector in body frame. During nominal operation the satellite orients its solar panel normal vector directly to the Sun and spins around it. The operators can select any spin rate and a deviating sun angle which allows offsetting the solar panel normal from the Sun.

Within in its first year of operation several anomalies were recognized by the ACS. All of them were recovered and none led to a mission critical situation. The first anomaly occurred directly after separation from the launch vehicle, which released the satellite with a much higher angular rate than expected. It was inherently solved by design although the design was never tested for such a situation. Another anomaly occurred when the two on board computers were operated at the same time. This led to the temporary loss of all angular rate gyroscope signals which had to be solved by the attitude determination algorithm. Additionally, several radiation events occurred which lead to the temporary drop out of sun sensors and single gyroscopes.

Beside the anomalies the nominal operation of the ACS was executed. This includes an initial detumbling and acquisition which is followed by the solar panel deployment. After the full reorientation to the Sun, the satellite is spun up to 5 rpm, which initiates the scientific phase. In the scientific phase the satellite is spun up/down to different rotation rates and maintained it for several months. This sets the environmental condition for any scientific experiment which required the simulated gravity.

This paper will start with the general overview of the Eu:CROPIS mission and its scientific experiments. Within the following it will concentrate on the attitude control system. It will elaborate on specific design aspects of a spin stabilized satellite and focuses on the experiences gained within one year of satellite operation. Additionally it will look into the ACS related anomalies which occurred in orbit and explain how they were handled.

AOCS support for optical communication experiments with the small satellite BIROS Terzibaschian T1, Raschke C2, Bärwald W2, Schultz C1, Maibaum O3, Lauterbach A1, Schrandt F1 1German Aerospace Center (DLR), 2Astro- und Feinwerktechnik Adlershof GmbH, 3German Aerospace Center (DLR) Optical communication from space to ground is a real challenge for the attitude control of a small satellite. The paper presents results which were generated in space. It is focused on attitude control. 2016 BIROS (~60x60x80 cm3, 120 kg) was launched into a Sun synchronous Earth orbit. The satellite bus is a reuse of the TET-1 bus (launched 2012). BIROS project is an DLR R&D project with additional funding from the German Federal Ministry of Education and Research. This opened the opportunity to support other technological DLR experiments in space beside the main task of the FireBird science mission. One experiment belongs to optical communication from space to ground with finally 3 different optical communication systems (OCS) on board of BIROS. The systems have the requirement that the optical axes of the OCS on board have to point exactly to the optical ground station (OGS) and have to keep this pointing while the satellite is approaching to and passing the OGS. TET-1 and BIROS are equipped with different attitude control modes which are defined by target orientation and target rate profiles, the minimum set of required AOCS devices, active software modules and some specific system states – bus states and attitude control system (AOCS) states as for example the availability of a true inertial attitude knowledge. The “Target Pointing Mode” is one of these attitude modes. It controls a freely selectable satellite fixed axis pointing to a commanded target on ground. which is defined by its x,y,z WGS 84 coordinates. The AOCS is based on state estimation and state space control. The target attitude defines the “mean” value of the commanded kinematic target attitude (orientation and rate). This is affected by an unavoidable jitter motion resulting from a noisy attitude determination and a noisy attitude control. The AOCS of the bus is designed for a control accuracy of better than 5 arc minutes (1 σ) according to the optical main payload’s requirements. This corresponds to ~750 m deviations on ground. The real flight results showed 1σ values of 2 arc minutes attitude error (~300 m) for TET-1 and BIROS. One optical communication system was designed with a “field of view” of <30 arc minutes. The two other optical communication systems have remarkable smaller FOV values. That’s why an additional uplink beam (beacon laser) and an additional 4 quadrant detector on board were required. A special air bearing AOCS test stand was equipped with an additional beacon laser and a 4 quadrant receiver in order to allow realistic AOCS tests on ground. This additional system did extend the “Target Pointing Mode” from TET-1 by an additional sensor system including the necessary software modules for BIROS. The next technical challenge is the alignment between the satellite’s body fixed coordinate system with the bore sight frames of the 3 different OCS and the 4 quadrant detector for the beacon laser beam. The mechanical design has to limit any relative motion of all 4 optical devices during the launch. The satellite frame is physically presented by an optical alignment cube stiffly connected with the two star tracker cameras. All payload instruments on board of BIROS are finally aligned with respect to this optical cube. But the star cameras are mounted on their own mounting bracket while the optical payloads are mounted on a common plate of the payload segment. It was known from vibrational tests that launch vibrations can cause deviations in the order of some arc minutes. Vibration tests showed that the relative orientation of all OCS was stable within main requirement to keep the parallel alignment of all three optical axes of the OCS within the FOV of the 4 quadrant receiver. This assures that optical communication can be done while the beacon laser can be received (due to geometry of all 4 rays). This was verified by the vibrational tests on ground. But it was expected that the relative orientation of all payload devices in the satellite frame will be changed after launch. While camera systems did find the outer orientation by using photogrammetry the OCS need a “search strategy” by doing some very precise scan motions over the OGS. The paper presents this search strategy by using the real flight data. Unfortunately this search strategy was stopped before completing the search. Instead of this it was demanded to use another secondary optical camera payload on board for the verification of the target pointing. This is a CCD matrix camera. Due to the extremely low priority of this camera for the mission there was no calibrated orientation in the satellite frame measured before launch. The after launch knowledge was a general “mounting accuracy” assuming skilled mechanics (~ 0.5 deg, verified by tests) plus the changes due to launch vibrations. By using again photogrammetry it was finally possible, to estimate a rough orientation of the matrix camera. This could be verified by several pictures of the Moon taken in the AOCS mode “Inertial Pointing Mode”. The matrix camera was aligned with a direction in an inertial fame pointing from BIROS to the Moon. The images verified the optical camera axes orientation and illustrate the remaining real jitter motion. The Moon is a diameter of ~ 30 arc minutes (comparable with the FOV of the 4 quadrant detector for the laser beacon from ground) and the sequence of images from different days shows a “motion” of the Moon within the image corresponding to the 2 arc minutes control accuracy. After knowing the optical axis of the matrix camera it was possible to point this axis to a target on ground. The sequences of images show a stable target pointing as required and within the known control accuracy which was stable and accurate over several minutes - in accordance to the standard AOCS telemetry data. The search strategy was not completed up to now but interrupted by changing the priorities within the FireBird mission. Therefore there are no results of any optical communication available up to now but promising results of the BIROS AOCS.

Sensors Data Fusion and Autonomous Navigation

GENEVIS: Generic Vision-Based Navigation for Descent & Landing Djafari-Rouhani D1, Duteïs P1, Brochard R1, Tiberio S1, Sanchez Gestido M2 1Airbus Defence and Space, 2ESTEC, European Space Agency GENEVIS is an ESA-co-funded Airbus Defence and Space project whose objective is to specify, develop, test, validate and integrate Vision-Based Navigation (VBN) processing solutions on a space-grade processor. As part of the project, a reference scenario was derived to analyse close rendezvous strategies, and another to analyse descent and landing strategies. The flight assessment focuses on the descent and landing scenario, for which the final objective is a precision lunar landing. This study case is comparable to the on-going PILOT project with ESA, where a VBN design is to be embarked on-board the Russian lander.

The VBN solution associates image processing algorithms with navigation algorithms. The image processing part itself is made of two subsystems: a relative feature tracking algorithm and an absolute landmark matching algorithm. Both subsystems were developed by Airbus Defence & Space and aim at covering multiple use cases. The relative feature tracking algorithm is an in-house variant of the Kanade-Lucas-Tomasi tracker optimized for low-power embedded platforms and scenarios with challenging image motion such as aircraft landing. The absolute landmark matching algorithm is a landmark-matching method which is able to compensate image distortions due to terrain elevation. This method does not rely on specific ground features such as craters and, therefore, is not limited to the characteristics of the lunar terrain. The image processing outputs are provided to the navigation algorithms where they are first pre-processed to produce a measurement depending on the available information. This process, called Ultra-Fast SLAM (UFS), deals with the non-linearity induced by image features prior to entering the filter in order to mitigate consistency issues and increase robustness [1]. The outputs of this process are the measurement and its associated error covariance which are passed on to the navigation filter. Fusion is then performed together with the measurements coming from other platform sensors, such as star trackers and IMUs, by a Schmidt EKF implemented using computationally-efficient UD factorizations. The output navigation estimates are used for platform guidance and control and thus verify specific requirements.

Simulations are performed to generate sensors outputs, including camera synthetic images, which are provided to the VBN solution in order to assess its performance in terms of positioning accuracy and robustness. In particular, the SurRender® software [2] is used to simulate images that achieve a very high level of physical representativeness and accuracy: the tool, optimized for space scenarios, can accurately model a real camera with its typical properties, such as noises, motion, projection, distortion and PSF. The consistency of state updates is verified through covariance and Monte Carlo analyses, where the overall performance of the estimation is evaluated. The analyses’ results confirm the robustness of the implemented design and the sharp enhancement provided by VBN over pure inertial navigation during the main braking phase.

The complete navigation design, including image processing, is implemented in real time on one core of a LEON4 CPU embedded on a GR740 development unit to perform processor-in-the-loop (PIL) tests. Sensor data are simulated on a separate computer in real-time and are sent over an Ethernet link. A second breadboard is made in order to perform hardware-in-the-loop (HIL) tests and demonstrate the developed solution on actual flight data. This prototype is realised by assembling small form-factor sensors on a platform mounted on a helicopter. All sensors are components off the shelf (COTS) chosen to present the same characteristics as the lunar lander ones. The helicopter-flown trajectory also aims at being representative of the relative dynamics of the lunar landing scenario. An additional, high-accuracy navigation solution, based on RTK GPS, is embedded on-board to assess the real-time performance of the VBN solution against the computed reference.

The results of the PIL tests demonstrate that the full-software design proposed can be successfully ported to a space-grade processor with sufficient processing margins, without resorting to FPGAs, and thus encourage the use of the proposed design for precision lunar landing. The HIL tests demonstrate the correct timestamping of the various measurements and allow the acquisition of actual sensor outputs. These are used to check against the models and synthetic images used for the implementation and for the simulation assessment of the landing VBN solution. Replays of the acquired helicopter flight data indeed demonstrate the functionality of the designed vision-based algorithms with real data.

[1] A. Robin, B. Polle, P. Vidal et al., From space exploration to UAV: development and real-time performances of a robust, autonomous vision-based navigation, 3rd CEAS Specialist Conference on Guidance, Navigation and Control, Apr. 2015.

[2] R. Brochard, J. Lebreton, C. Robin et al., The SurRender software, 69th International Astronautical Congress, Oct. 2018.

Computational Guidance & Navigation for Bearings-Only Rendezvous – methods and outcomes of GUIBEAR Branco J1, Lourenco P1, Branco J1, Briz Valero J2, Mammarella M2, Peters T2, Manara S3, Bernardini D3, Bemporad A3, Foglia Manzillo P4, Witteveen J4, Cropp A5 1GMV, 2GMV, 3ODYS, 4cosine Remote Sensing B.V., 5ESA-ESTEC GUIBEAR is an on-going activity supported by the European Space Agency aimed at the demonstration of the feasibility of bearings-only (B-O) far-range rendezvous in near rectilinear Halo orbits (NROs). The studied scenario is the far-range rendezvous of the HERACLES Lunar Ascent Element (LAE) with the Lunar Orbital Platform – Gateway (LOP-G) that is orbiting the Moon in an NRO. This activity advances the readiness level of the applicable technologies through a solution based on Computational GNC, structured around a Model Predictive Control framework and tools; and an approach design, development, verification and validation based on iterative prototyping which integrates high-fidelity hosting considerations into the design from a preliminary level: concurrent GNC and software architecture design with hosting, runtime assessment and optical high-fidelity simulations. This paper: presents the context – mission, system and hosting, and algorithm-level challenges and proposed concepts; details and justifies the architecture and design solutions; outlines the methodologies employed for prototyping verification and validation; provides conclusions on the obtained results.

Estimating the range to target is the first order challenge of bearings-only rendezvous: the relative position is not instantaneously observable and bearings measurements are nonlinear, which are a source of inconsistency, over-conservatism or divergence of estimates. Observability analysis shows that manoeuvring between at least three non-coplanar (or four coplanar) measurements is necessary to resolve the full relative position and velocity observable. As a result, estimation and control/guidance problems are inherently connected, especially in an autonomous setting. A bearings-only navigation filter for rendezvous has to cope with the non-linearity of the measurement model and the instantaneous non-observability, while including knowledge of manoeuvres through thruster acceleration observables that introduce a significant source of uncertainty. Given all the challenges the navigation filter is exposed to, careful planning of the trajectory in order to maximize the information gain it provides is extremely important. Effectively correcting in closed loop the results of unmodelled dynamics, particularly manoeuvre dispersion and knowledge, favour autonomous on-board guidance functions that present a challenge both in terms of computational load, fuel- cost, estimation performance, and safety. NROs present significantly different orbital environments to existing Earth orbit B-O rendezvous heritage. The nature of the 3-body problem and the absence of complete analytical closed-form solutions are important challenges, but extensive study of this kind of orbits shows that, for the scenarios considered and short timescales, closed-form approximation are applicable. The distances introduce significant challenges on target detection and tracking as the target is at sub-pixel resolution, such as increased exposure times. This purely radiometric detection relies on the capability of disentangling the target contribution to the pixel readout from the background contribution due to cosmic background radiation, background signals, and to instrument noise. Several other hosting challenges are present, eg, effectively modelling the accelerometer delta-V measurement errors - from drifting biases to multiplicative cross-couplings. Finally, the online nature of the GNC solution also presents significant challenges in terms of runtime constraints and processing architectures. The on-board computational power has to be able to solve the constrained optimization problem that underlies the guidance function in real time.

These challenges are addressed through the design of an algorithmic or computational Guidance and Navigation architecture. The computational efficiency, accuracy and robustness of the trajectory generation are at the foundation of the design, as the numerical generation of guidance commands relies extensively on on-board computation. The autonomous guidance function is based on an optimization problem with the objective of 1) maximizing the observability of the range and 2) minimizing the fuel consumption. This optimization problem bridges the gap between estimation and control, thus approaching the potential of a truly autonomous system: the navigation filter informs the computation of the trajectory, which in turn is optimized to improve the navigation performance. In order to close the loop, the optimization problem is solved online, and its solution is updated whenever the uncertainty of the navigation filter decreases. Adding the capability to enforce input (saturation, maximum ΔV, minimum acceleration necessary for detection) and output (safety and final conditions) constraints to the optimization of the performance index lead to the formulation of the multi-objective optimization problem in a model predictive control (MPC) framework. Ultimately, this provides a process to reduce the navigation errors systematically throughout the mission, while taking advantage of this same reduction to improve the commanded plan. The bearings-only estimation filter is tailored for monocular vision-based navigation, taking into account the particularities of the NRO environment and the difficulties of navigating when the target is always at sub- pixel resolution. A weighted least squares batch estimator is employed, improving the accuracy and outlier rejection while dealing with varying camera integration times and computational cost. This filter is complemented by a higher frequency state estimator including a propagator for when measurements are not available, and combines the higher quality information from the batch filter. The resulting architecture combines sequential and batch processing in high and low frequency processing tasks. This exploits the structure of the mission scenario with the objective of improving the navigation performance, as it minimizes sensor measurement errors, saves computational time, and minimizes the real-time operational constraints. The proposed Computational G&N architecture is verified and validated in simulation, taking as baseline the studied scenario. The scenario tackled in this activity is the approach and autonomous injection of the LAE in the LOP-G’s NRO, leaving the former in handover conditions for the start of rendezvous proper, at 100 km from the latter.

The feasibility of performing far approach through these mass and cost-saving camera-only GNC techniques is established and justified through prototyping and extensive testing in high-fidelity simulations. Autocoding and Software-In-The-Loop testing are performed as a bridging step towards the subsequent Hardware-In-The- Loop campaigns A novel method employed for benchmarking of runtime of the auto-coded software is described: preliminary non-realtime PIL assessment employed at algorithm development level. This method includes a step of experimental compilation and running of a segment with a number of GNC cycles in a flight-representative processor thus allowing adjusting GNC parameters and runtime latency emulation blocks in the functional simulator.

Distributed Cooperative Visual Odometry for Planetary Exploration Rovers Bahraini M1, Aouf N1, Beauvisage A1, Govindaraj S2 1Department of Electrical and Electronic Engineering, City University of London, 2Space Applications Services Abstract Planetary robotics navigation has attracted great attention of many researchers in the recent years. Localisation is one the most important problem for robots in another planet in the lack of GPS. The robots need to be able to know their location and the surrounding map in the environment concurrently, to work and communicate together in another planet. In the current work, a novel algorithm is designed to cooperatively localise a team of robots in another planet. Consequently, a robust algorithm is developed for Visual Odometry (VO) using optical flow KLT feature tracker to localise each single robot in a planetary environment while detecting both intra-loop closure and inter-loop closure using previously observed area by the robot and shared area from other robots, respectively. To validate the proposed algorithm, a comparison is provided between the proposed approach and the conventional KLT feature tracker. Accordingly, a planetary analogue real dataset is used to investigate the accuracy of the proposed algorithm. The results promise the concept of distributed cooperative VO to increase the accuracy of localisation.

Introduction For the last few decades, mobile robots and autonomous systems have become a hot research topic resulting in major advances and breakthroughs. Currently, mobile robots are able to perform complex tasks autonomously in various domains including military, medical, space, and commercial applications that can operate on the ground, at sea and in space. Deploying a team of robotics to lunar for construction purposes in the next few years is one of the main plans for planetary and space application. In those applications, mobile robotic platforms should perform complicated tasks including navigation in complex dynamic environments. A robotic platform needs to be self-localised, if it is designed to autonomously navigate itself in its environment without any human intervention or even in the absence of Global Navigation Satellite System (GNSS). Readers may refer to survey papers that outline recent developments in Simultaneous Localisation And mapping (SLAM). Visual cameras are the most commonly used sensors due to low power requirements, availability and highly physical compactness, which can be used for space application to perform VO. To improve the accuracy of single platform SLAM/VO, the literature examines the cases of cooperative SLAM/VO in centralized and decentralized manners which refer to whether all platforms send their information to a central SLAM/VO algorithm on board of one platform (centralized) or SLAM/VO is computed on multiple platforms (decentralized). Distributing the system across a network of vehicles (decentralized/distributed SLAM/VO) allows for efficient processing in terms of both computational time and estimation accuracy.

Distributed Cooperative VO Approach In this work, a robust stereo-temporal matching approach is designed to prevent from losing feature points during motion estimation process. The image is divided to some buckets to propagate the feature points through the whole image. A border is set around the image to avoid from generation feature points in this region which are not reliable due to movement of the mobile robot. Optical Flow, KLT feature tracker is used to track feature points in the stream of stereo images. Additionally, the M-estimator SAmple Consensus (MSAC) algorithm is used to eliminate spurious correspondences in this algorithm. Also, PnP algorithm is applied for 2D-3D matching in motion estimation. The selected keyframes including feature points and their descriptors along with their 3D position are stored on a server and can be uploaded by each robot to detect any overlapping fields of view for two or more robots. The optimisation process triggers after finding a number of shared areas between any robots involved in the loop closure (inter-loop closure) or even any previously seen area by the robot itself (intra-loop closure).

Results and discussion There are many challenges in the space environments in terms of existing features, brightness and illumination. In order to validate the proposed approach and compare the results with available existing methods in literature, a planetary analogue real dataset is used. Firstly, the robustness and accuracy of the proposed VO algorithm is investigated. Herein, two robots are started to move from different initial points. The trajectory of the robots is estimated using proposed VO. In addition, the estimated trajectories of the robots obtained by applying conventional KLT-VO algorithm are plotted and compared with the single version of the proposed algorithm. It should be noted that KLT-VO algorithm is using KLT for feature tracking between temporal images and BA for localisation. Also, the obtained trajectory from applying the mentioned algorithms can be compared with the ground truth (GT). It can be seen that the accuracy of the single version of the proposed algorithm is higher that the conventional KLT-VO.

Also, the proposed algorithm for distributed cooperative VO is applied on the real dataset to improve the accuracy of the proposed VO algorithms. In this scenario, the Robot 2 has detected part of the observed environment by Robot 1. Thus, the optimisation process triggers for Robot 2. The error between the estimated trajectory and the GT is calculated and compared with single VO algorithm. Clearly, the accuracy of localisation is improved in steps nearby 2200 and 3300 which the Robot 2 has detected a previously seen area by Robot 1.

A deterministic and high performance parallel data processing approach to increase guidance navigation and control robustness. Ghiglino P1, Harshe M1 1Klepsydra Technologies GmbH New generations of spacecrafts are required to perform faster onboard processing. Space exploration, rendezvous services, space robotics, etc. are all growing fields in Space that require more sensors and more computational power to perform these missions. Furthermore, new sensors in the market produce better quality data at higher rates while new processors can increase substantially the computational power. Therefore, near-future spacecrafts will be equipped with large number of sensors that will produce data at rates that has not been seen before in space, while at the same time, data processing power will be significantly increased.

In regards to guidance navigation and control applications, vision-based navigation has become increasingly important in a variety of space applications for enhancing autonomy and dependability. Future missions such as Active Debris Removal will rely on novel high-performance avionics to support advanced image processing algorithms with large workloads. Even more complex is the case of vision-based precision landing, where there needs to high rate processing can be the tipping point of a successful mission.

This new scenario of advanced Space applications and increase in data amount and processing power, has brought new challenges with it: low determinism, parallel data processing, cumbersome software development, etc. For that ESA, is promoting the use of MATLAB / Simulink autocoding tools in order increase software quality, data modelling and development process. For large and parallel data processing, like vision base navigation, however, MATLAB is not well suited to produce code that is optimal or deterministic to the destination hardware environment.

In this paper we present a novel toolset for SIMULINK that can produce deterministic and optimal code for scenarios likes the one presented above. Based cutting-edge software engineering techniques, this solution is used in aerospace and robotics applications producing data processing rates that are substantially faster than other available products. The experimental setup consists of an on-board sensor data fusion simulation with a Zynq board as processor-in-the-loop connected to the simulation via CAN-Bus, which is a technology currently promoted by ESA. The presented results show that our implementation is not only high performance and scalable but also has a friendly interface that can reduce the development time.

GPS Constellation Modernization Impact on GPSR at GEO Ramsey G1, Chapel J1, Freesland D2, Krimchansky A3, Chu D3 1Lockheed Martin Space, 2ACS Engineering, 3NASA/Goddard Space Flight Center The Geostationary Operational Environmental Satellite-R program (GOES-R) has launched two of the next generation geostationary weather satellites, both of which are now fully operational. GOES-16 launched in November 2016 and GOES-17 launched in March 2017. The GOES-R spacecraft provide dramatic improvements in GEO weather observation capabilities over the previous generation.

The GOES-R series GPS Receiver (GPSR) system consists of a single Rx antenna, bandpass filter and Low Noise Amplifier (LNA) serving to provide input to a 12-channel, single frequency (L1) coarse acquisition (C/A) GPSR. The GOES-R Vehicles process the collected pseudo-range and Doppler data in the GPSR which provides a Kalman filter output Earth Centered Earth Fixed (ECEF) position of the satellite for mission processing of collected science data. The GPSR system was designed and tuned in order to facilitate tracking the extremely weak GPS signals, dominantly tracking side-lobes at GEO on the order of 10e-18 Watts. Use of a GNSS system on a spacecraft is desirable for three main reasons 1) Position, velocity and timing (PVT) are improved, 2) demand upon ground support is reduced, 3) having real-time PVT available to the Flight Software increases automaton.

GOES-R is the first operational satellite to utilize GPS sidelobes for navigation at GEO, which is the key factor in the systems highly accurate, robust and continuous navigation solution. However, this also renders a distinct sensitivity to changes in the GPS signal pattern in the sidelobe regime. Any satellite intending to maximize GPS navigation performance at GEO will need to implement a GPSR system that tracks sidelobes. Any such program would have a vested interest in characterizing the evolution of the constellation’s transmit patterns to ensure the system continues to meet requirements for the life of the system.

A variety of improvements to the system have been implemented to improve the navigation performance and robustness of the system as a whole: including mitigating thermal sensitivity, tuning the internal Kalman filter and accounting for EOP format changes. However, the performance implications of the upcoming transition to a modernized GPSIII constellation and the associated Earth serving Transmit pattern on this side-lobe capable GPSR at GEO has yet to be quantified and characterized. This is the primary topic of this paper. This paper will address several distinct topics: 1) GOES-R GPSR performance change regarding a fully modernized constellation via hi-fidelity simulation, 2) GOES-R GPSR acquisition and tracking characterization regarding the first operational GPSIII vehicle, 3) heritage to modern GPS transmit pattern comparison and 4) general insight of relevant signal requirements for the next generation of GNSS as it pertains to including GEO GPSR facilitation.

The current constellation consists of IIR, IIF and IIR-M block-types, each with a distinctive transmit pattern which is most variant in the side-lobe regime. The GPSIII antenna pattern is depicted, clearly visible are the highly dynamic side-lobe regions with all the hills and valleys. Note that 90 degrees off the GPS boresight translates to about 40 degrees off Nadir for GEO satellites. The GPSIII Tx pattern structure is most similar to the IIR-M pattern but with a higher main-lobe power and deeper side-lobe valleys which will impact tracking performance and resulting navigation quality for GEO GPSR.

This GPSR system has proven itself to have great coverage capabilities under nominal Nadir pointing circumstances, tracking over 11 satellites on average. The system is also highly robust to off-nominal circumstances including off-Nadir pointing, maneuvers and (simulated) degraded constellation due to the side-lobe capable aspect of the system. The overall performance exceeds expectations and provided significant operational leeway, off-nominal robustness and dynamic capabilities. The current GOES-R on- orbit received signal carrier-to-noise spectral density (C/N0) depictions clearly show the high-power GPS main lobe. The primary receive gain (near 20 deg off-boresight) which corresponds with the first side-lobe regime is also accentuated. Another bump in received power can be seen in the second side-lobe regime. These signal quality metrics will change as the constellation modernizes, this paper will provide expectations for how these metrics will change and the implications thereof.

The GOES-16 GPSR was initialized on December 2016 in less than 5 minutes and it has since exceeded expectations. GOES-16’s 3D position accuracy is better than 15.2 meters 99% of the time and the total velocity accuracy for 99% probability is found to be 0.52 cm/s, this includes station-keeping maneuvers. Additionally the clock accuracy can be assumed to be consistent with ground testing, near 70 ns PPS accuracy RMS. During even the longest duration station keeping maneuver we expect position solution accuracy on the order of 30 meters or better depending upon the axis. It is worth noting that our primary instrument, the Advanced Baseline Imager, is sensitive to position error in the in-track and cross-track axes, where position error translates to pointing error, whereas radial position error would translate to pixel size error. Thus, the intent of improving performance was first to improve in-track and cross-track position solutions then radial position solutions followed by velocity. This on-orbit performance exceeds all position, velocity and timing requirements by an order of magnitude. The average tracking of 11 GPS and an excellent average DOP of 7.9 shows we have significant margin to maintain this excellent performance. These GOES- 16 performance results show that the final frontier of GPS at GEO, is now a proven operational capability. All future GEO Satellites must consider the addition of a GPSR in their spacecraft design, otherwise they are sacrificing spacecraft capabilities and accuracy along with incurring increased and continual demand on ground support. All future GEO satellites using GPSR need to take into considerations changes in the constellation makeup and the resulting performance implications as we will here.

Future work on the topic of GPS modernization includes on-orbit performance analysis as more GPSIII satellites become operational, a more comprehensive assessment and statistics of vehicle to vehicle signal and ephemeris quality variation. Other future work on the topic of GPS at GEO includes definition of variation of metrics in the side-lobe regime such as group delay and analysis on repeatable error sources like multipath.

Current Space Mission Validation & Verification

SmallGEO Product Line FlatSat - A powerful tool for AOCS verification Brito A1, Brito A1, Neumann N1, Behrmann N1, Kröger L1 1OHB System AG The SmallGEO geostationary telecommunication platform represents OHB’s entry in the telecom satellite market. After the successful launch and delivery of the first two SmallGEO spacecraft, ’s H36W-1 and EDRS-C (respectively in 2017 and 2019), OHB is developing two more missions based on the same platform: Electra and Heinrich Hertz.

The advantages and use cases for a representative satellite engineering model (EM, known internally as FlatSat) became evident during the verification phase of the AOCS design of the first two projects. A representative FlatSat in terms of hardware, software and ground support equipment (GSE) allows for early testing and verification of functional chains, something that would otherwise be possible only on the flight hardware.

Amongst other advantages, the FlatSat has a realistic representation of the timing and delays introduced by the real hardware and its interaction with the software, which is critical to simulate the correct behaviour of the AOCS control loops. Furthermore, the FlatSat architecture is scalable: units can be represented by engineering models, when a more realistic behaviour is needed, or by emulators, when a lower fidelity can be accepted: this enables testing of specific test cases (e.g. failure cases) on the real hardware, rather than simulating them in pure software environment.

This cost-effective approach, including in-house developed emulators, allows a high degree of customization of the SmallGEO FlatSat concept to each mission, enabling the continuous development and improvement of the common platform while maintaining the basic common elements. It allows also to replace units in the course of time, which might become obsolete by state of the art units or to introduce new developments. Lastly, it allows also having a configurable reference platform for verification and validation of satellite software updates in case software modifications with hardware-software interaction impact are planned on in- orbit satellites.

This paper, divided in two main sections, will present the FlatSat’s architecture for the SmallGEO product line and its role in the verification of the AOCS design. Starting from an overview of the AOCS development and test activities, the first part will make a strong case for a representative FlatSat, with a focus on the lessons learnt in the first two SmallGEO missions. The second part will detail the FlatSat’s architecture for Electra and Heinrich Hertz (see annex), showing how the common architecture has been tailored to the different needs of the two missions. The paper will also give a snapshot of the most relevant AOCS testing activities currently ongoing in both projects.

Independent V&V of the ExoMars 2020 Cruise GNC, through analytical techniques, simulation-based analyses and tests Recupero C1 1Deimos Space S.l.u. Abstract

Elecnor DEIMOS, both in Portugal (Deimos Engenharia) and Spain (DEIMOS Space) contributed to the EXOMARS 2020 program from preliminary design to post flight analysis in phases C/D, in multiple fields such as the EDL (Entry, Descent and Landing) Mission Analysis, the V&V of the Cruise GNC, and the AOC GNC SCOE (Special Check-Out Equipment) activities. This paper presents an overview of DEIMOS Space’s activities, with SENER as subcontractor, within the Cruise GNC activities of the ExoMars 2020 program. DEIMOS Space is the responsible of the independent V&V activities, supporting the analysis and verification of the Cruise Guidance, Navigation and Control (GNC). It is an activity aimed at validating the AOCS algorithms in an independent way, hence providing a different and new independent perspective when compared to the traditional SW-based only approach via ISVV. The Validation & Verification (V&V) campaign, in fact, shall ensure, to the maximum extent possible, that the Cruise GNC can adequately control the SpaceCraft (S/C), and hence that the Descent Module (DM) accomplishes its objective of reaching the separation point with the required accuracy. The outcomes of the independent analysis, materialized as simulation reports and performance assessments, allows the customer to early identify possible bugs and unexpected problems, timely taking the necessary actions to ensure the adequate quality of the GNC product. Finally, the Software (SW) simulation framework (Cruise GNC Functional Engineering Simulator) enabled the efficient selection and investigation of test cases whilst maximizing functional coverage, to support the definition of the test cases to be executed, as part of the V&V activities, on the System Avionics Test Bench (ATB).

Introduction

The ExoMars 2020 mission, co-managed by ESA and , consists of a Spacecraft Composite (SCC) to be launched in July 2020 by a Proton-M Launch Vehicle (LV) and is led by TAS-I. The mission will demonstrate key flight and in-situ enabling technologies in support of the European ambitions for future Mars exploration missions, and will pursue fundamental scientific investigations. The mission objective for Europe is exobiology using a drill system and an elaborated laboratory that will sever 3 payload instruments (Micromega, MOMA, Raman, for a total of 9 payload instruments embarked on the Rover), while for Russia the interest is to have a static scientific platform on Mars to monitor the planet external and surface environment. The SCC includes the Carrier Module (CM, led by OHB Bremen) and the Descent Module (DM), which accommodates the Rover Module (RM, led by TAS-I with the Rover Vehicle led by Airbus UK). The CM will transport the DM during the cruise phase, covering the transfer orbit form Earth to Mars, starting at the separation of the SCC from the Launcher, and ending at the separation of the DM from the CM, a few days prior to the Entry, Descent and Landing (EDL) phase of the DM in Mars. Once the DM has landed on the Martian surface and after the egress of the RM the landed part of the DM will remain operational as a long-living Surface Station, while the RM will perform its nominal surface science operations. DEIMOS is the prime contractor of the independent V&V activities for the Cruise GNC. The activity supports the full analysis, verification and validation of the Cruise Guidance, Navigation and Control (GNC) algorithms, against the applicable requirements, during the phases C/D of the project. The paper will describe the activities performed to ensure the full V&V of the Cruise GNC algorithms reaching a proper Performance Assessment: • Analysis and analytical verification of the Cruise GNC algorithms, including activities of prototyping, unit testing and robust control techniques, to ensure the full understanding of all functionalities that have been included; • Preparation and maintenance of a pointing budget, including the analysis of the Cruise GNC pointing error sources, non-linearities and non-idealities, considering the SC, its units and environment, and the GNC algorithm properties; • Construction and validation of a Functional Engineering Simulator (FES), specifically designed to support the verification of the Cruise GNC algorithms, and implementing high-fidelity models of the SC, Actuators and Sensors; • Development of a Cruise GNC Verification Plan and the execution of the verification campaign in the FES; The process has been iterated during the consecutive design loops, while the algorithms and the available data were progressively refined, starting from data retrieved by the units specifications, then using the declared units performance, and finally the measured units performance. A representation of the work logic of the V&V campaign and its loops is shown. The last activity of the Cruise GNC Verification and Validation, consists in the • Support the GNC verification on Avionic Test Bench (ATB), defining ad hoc procedures and tests to be executed on the ATB with the GNC SW in the loop, that will be the most realistic simulation environment available to the project. This has not been iterated, being the final step of the whole process and preliminary to the flight, and consists of supporting the test plan and specification preparation, as well as processing the test results and review them, to provide the final assessment of compliance status.

Automatic code generation of the GNC software for Spacebus Neo Dandré P1, Grossingho T1, Chevallier M, Lopez Negro P 1Thales Alenia Space The aim of this paper is to present the experience feedback of the automatic code generation of the GNC software on the development of Spacebus Neo. The auto-code framework is based on MathWorks Simulink development environment and internal guidelines derived from experience in previous programs and tradeoff studies about different architectures. This framework has enabled to meet the timeline for the GNC software delivery with almost no bug at VLV and FCV, and is retained as the standard for the new GNC developments at Thales Alenia Space.

The automatic code generation perimeter on Spacebus Neo includes the GNC manager (TM/TC definition and execution, phase definitions and transitions management) and GNC functions. This framework has been used on Spacebus Neo for the development of four GNC applications and a library of around 100 GNC functions.

The to the auto-coding process for the design of the GNC software was decided at the end of 2016 in a context of low maturity with respect to the schedule and insufficient flexibility in the existing process of GNC software conception to face the challenge.

The key drivers for the decision were the flexibility of the new process enabling to manage late inputs and requirements, its capacity to reduce the GNC design length from specification to binary delivery and the ability to develop in large team integrating with co-engineering of GNC and OBSW team. The process enabled an early debug of the GNC software, simplified the GNC simulators management merging all the evolutions in a single simulator and an efficient and friendly development environment with quick feedback loop and good traceability.

The process enables the usage of existing manual code and the re-use of flight code of the OBSW which makes the process interesting for new developments. This functionality is used on Spacebus Neo at elementary function level, GNC function level or application level.

Additionally is has enabled to meet the following specific needs : develop a GNC mode re-using existing GNC functions library that can be de-activated, develop a functionality that is an option on the product line, manage the evolutions on the product line.

Several tools have been developed along this auto-coding process framework in order to remove the activities with low added values and minimize the risk of non-quality: automated generation of a GNC function in the auto-code format, automated generation of the code documentation with a content equivalent to the manual code specification document, automated export to database.

A LISA mission simulation environment: TAS NUMES simulator framework a powerful tool for future missions Dionisio S1, Cesare S1, Basile F1 1THALES ALENIA SPACE ITALIA SPA ESA’s future mission LISA, Laser Interferometry Space Antenna, will detect and observe gravitational waves that are emitted during the most powerful events in the universe. LISA will detect gravitational radiation from astronomical sources, observing galaxies far back in time and testing the fundamental theories of gravitation. Its observatory will consist of three spacecraft traveling in near-circular heliocentric orbits in a triangular formation with arm length of 2.5 million of kilometers. Each spacecraft will contain two Moving Optical System Assemblies (MOSA), each one composed by a Telescope, an Optical Bench (OB) and a Gravitational Reference Sensor (GRS), the latter containing in turn a Test Mass (TM). Each LISA satellite can be described by a complex dynamic which accounts for 20 degrees-of-freedom (DoFs); 6 DoFs for the spacecraft translation and rotation, 6 DoFs for each TM translation and rotation and 2 rotational DoFs for the MOSA articulation. In science, 18 of the aforementioned 20 DoFs have to be controlled by the DFACS, Drag-Free and Attitude Control System, in order to keep stable the inter-spacecraft interferometer lasers relative pointing at micro- radian level and let the test masses free-falling along the LISA interferometer arms. An high-fidelity simulation environment is therefore deemed necessary to both design and, in future, verify the mission and satellite performances.

In the frame of one of the ESA contracts called “LISA Phase-A System Study for a Gravitational Wave Observatory” of which TAS is prime, a LISA's DFACS functional simulator has been set up exploiting the TAS simulation framework called NUMES, New Mission End-to-End Simulator, that enables to easily build- up specific E2E, End-to-End, software environments for different kind of missions. The TAS’s NUMES framework is constituted by a validated collection of C/C++, Fortran source files and Matlab/Octave macros running on different environments (e.g. Linux, Windows). They constitute a tool capable to numerically simulate, with the appropriate level of fidelity, the dynamics of one or more satellites subject to the external environment (solar radiation pressure, magnetic field, etc.) and to the action computed by the board controller for various mission phases and operational modes. They include models of the sensors and actuators entering in the control loop, which for several missions, as for LISA one, comprises also elements of the payloads (e.g. accelerometers feeding the drag-free control system). Each new specific simulator is built starting from the available ‘bricks’ of the NUMES environment. This library has been built up by TAS based on its heritage in the designing and validation of E2E simulator environments for flight missions and preparatory studies (e.g. GOCE, Exomars, NGGM). Indeed, part of the NUMES ‘bricks’ have been validated with internal TAS research founds, through on flight results of missions in which TAS has been involved. NUMES can interact with external models (e.g. ephemeris) and can support the implementation of ad-hoc models developed for the specific project purposes as for the LISA case (e.g. control algorithms, payload elements feeding the controller, dedicated actuators, etc.). The NUMES version for TAS LISA Phase A study, in science mode instantiation, includes; one full- dynamics LISA spacecraft accounting for all the 20 DoFs, all sensors and actuators models (including payload used as DFACS equipment) defined by mass properties, size, field of view/ranges, mountings and the model characteristics, environmental disturbances (e.g. Self-gravity, Stiffness, Solar Radiation Pressure), DFACS controllers and orbital ephemeris to simulate incoming laser beam directions spacecraft relative pointing. In particular, the 20 DoFs full dynamics can simulate dynamically the cross-couplings due to the telescope and spacecraft jitters which is one of the most critical aspects of the LISA mission to be assessed. Indeed, the dynamics model is constituted by the motion equations of the entire multibody system composed by the LISA spacecraft, the two moving telescopes and the two test masses. This complex model represents the plant to be controlled by the control system. It is a large, non-linear system whose inputs are forces (both disturbances and control actions at given stations) and whose outputs are the state variables (positions and velocities) at some selected points (e.g. spacecraft center of mass, telescope mountings, etc.). Each body is connected to the others with different types of constraint (free - F, locked – L, constrained – C or transition - T). The nominal constraint scheme of the kinematic chain is the one called “Science Mode Configuration”. In this configuration the spacecraft is free in position and rotation, the MOSAs are fully constrained in position and rotation, but for the rotation around their pivot axis. At each simulation step the entire kinematic chain is automatically calculated by the motion equations: therefore the cross-couplings between spacecraft and MOSAs attitude are dynamically computed during simulation.

The dynamics model code of the generated equations is exported and included, as for the other NUMES bricks (sensors/actuators models), in Matlab / Simulink. The TAS LISA simulator allows both control performances and system/mission performance verification. Indeed, its outcomes includes system information such as the propellant and power consumption due to the required thrusts, as well as all the state variables on which the control performances need to be evaluated. The following paper briefly describes the LISA scientific mission challenges and the TAS LISA simulator environment. It shows as well some preliminary results relevant to the LISA science phase.

Future Space Missions

GNC design for Rendezvous Autonomous CubeSats Experiment (RACE) Mission Bidaux-Sokolowski A1, Kicman P1, Kłak M1, Mujdjei L2, Gogu D2, Cocco F3, Ankersen F4, Pirat C4, Walker R4 1GMV Innovating Solutions Sp. z o.o., 2GMV Romania, 3Gomspace, 4ESA/ESTEC The design presented in this paper is a part of ESA’s RACE (Rendezvous Autonomous CubeSats Experiment) mission. The mission expands state of the art for small platforms in terms of GNC autonomy. Two 6-unit CubeSats equipped with comprehensive set of sensors and actuators are due to accomplish a set of proximity operations. Two baseline experiments are Close Fly-Around (CFA) and Rendezvous and Docking (RVD). The GNC sensors and actuators choice is challenging regarding the power, dimension and mass constraints of a 6U CubeSat. Both satellites share most of the GNC equipment: 3-axis gyro, 3-axis magnetometer, fine and coarse sun sensors, GNSS receiver, magnetorquers, reaction wheels and propulsion system. The difference is that one of the spacecraft is equipped with a Narrow Angle Camera (NAC) and a star tracker while the second satellite has a Wide Angle Camera (WAC) for close range navigation. Both spacecraft use also different LEDs patterns for visual cooperative navigation. The propulsion system consists of 12 cold-gas thrusters located so that thrust and torque can be simultaneously executed in any direction. These micro thrusters (1 mN) are prone to a large execution error, which is another complexity for the design. Another difficulty comes from the fact that there is a limitation on the number of thrusters that can be fired simultaneously (due to the peak power restrictions of the platform). Inter Satellite Link (ISL) is used actively in cooperative experiments while during non-cooperative phases it exchanges only data used by FDIR. During RVD experiment both spacecraft cooperate to perform a fully autonomous docking sequence. The target controls only its attitude whereas the Chaser controls both its attitude and position. In this part of the mission the satellites exchange their telemetry through the dedicated ISL. The autonomous RVD phase starts at -30m along V-bar axis in LVLH frame, in an along-track formation. Then a series of ΔV pulses transfer the chaser to -10 along-track formation. The sequence of pulses is computed onboard based on linear relative dynamics – well known Hill’s equations. This multi-impulse transfer algorithm combines high precision of maneuver with propellant use much lower in comparison to continuous control. When the transfer termination is confirmed, continuous control forced-motion is initialized and the Chaser follows approach profile. The spacecraft accelerates to relative velocity of 1cm/s along the V-bar and keeps the approach velocity until contact of docking mechanisms is detected. This forced-motion final approach phase is where 6DoF robust controller is applied. For other phases, where the translation is uncontrolled a 3DoF attitude controller was designed. The relative 6DoF navigation system continuously use at least one of the cameras. Since WAC is used at close distance, the spacecraft with WAC acts as the Chaser in this experiment. Such solution is crucial due to minimizing the delay from input to actuation. The navigation system provides Port- to-Port (P2P) state estimate to the controller: position of Chaser’s docking port frame with respect to Target’s port in Local Vertical Local Horizontal (LVLH) reference frame as well as relative quaternion. The docking mechanism’s construction entails that the lateral docking error is kept under 5mm with 100% of margin. This accounts for both estimation and control error. Robust control synthesis and analysis framework is used to address the experiment requirements simultaneously taking uncertainties into account. The uncertain parameters are in particular the mass, inertia and flexible solar panel natural frequency. For control design both H∞ and μ-synthesis frameworks have been used. H∞ was applied as primary method for fast tuning. Once a satisfactory result has been achieved, μ-synthesis was used for final tuning guaranteeing both robust performance and stability. The H∞ synthesis utilizes Riccati-based algorithm whereas μ-synthesis is based on D-K iteration. Since the GNC loop operates at 1Hz all the synthesized robust controllers are discretized using this sample time. The CFA is an uncooperative experiment. One of the spacecraft (Target) is tumbling and the telemetry sent to the Chaser is not used by the GNC, just by the FDIR to eventually stop the experiment in case of non- nominal behavior. The second spacecraft (Chaser) flies in close formation in a representative manner as would be a mission to inspect a malfunctioned satellite or debris. The spacecraft equipped with NAC acts as a Chaser in this experiment with the unique relative sensor being the narrow angle camera. The CFA starts when the two satellites are in along-track formation. When the Target tumbles itself to a given angular rate, the Chaser applies a sequence of ΔV pulses to enter a passively safe orbit with a V-bar drift. Then no more translation maneuvers are executed, only the attitude is kept so the target is kept in the field of view of the Chaser’s camera. The most challenging part in this phase is the bearing-only navigation based on the centroid measurements. Since the CFA lasts many orbits, the formation repeatedly goes through the eclipse when no relative measurement is available and the GNC can rely only on the state propagation. The robust control techniques used to synthesize and analyze the controllers both for CFA 3DOF controllers and for 6DOF coupled controller will be presented in detail in the article. The general architecture for the synthesis of the controller is presented. The tuning of the various weighting functions, the LFT models used for the plant, the synthesis techniques (H∞, and μ-synthesis) and synthesis results will be detailed in the article. The linearized control loop including state observer (Kalman filter) was positively analyzed in terms of μ bounds lower than one. Additionally, classical gain and phase margins were calculated using one loop at a time method, cross checking the consistency of the conservative robust techniques with the classical margins. The integrated GNC system has been verified and tested in high fidelity simulator. Monte Carlo simulations proved that the system maintains stability and performance of the system under any expected parameters deviation docking within requirements by a safe margin.

Mars Sample Return – Test campaign for Near Range Image Processing on European Proximity Operations Simulator Burri M1, Benninghoff H1, Rems F1, Risse E1, Kanani K2, Falcoz A2, Masson A2 1Deutsches Zentrum Für Luft- Und Raumfahrt (DLR) / German Aerospace Center, 2Airbus Defence and Space NASA and ESA signed a letter of intent in April 2018 to pursue a Mars Sample Return (MSR) mission. The goal is to return soil samples from Mars to Earth until 2030 for detailed analysis in sophisticated laboratories far too heavy to transport to Mars. In a first step Mars2020 Rover will retrieve 36 samples on Mars. To minimize the risk of complete loss, all samples filled in tubes will first be left back on the Martian surface. The Sample Fetch Rover will then collect the tubes and load them into a container within the Mars Ascent Vehicle. ESA’s Earth Return Orbiter (ERO) will then catch the orbiting sample container (OS) in low Martian orbit. The wide-angle camera is used as rendezvous sensor. Sealed in a biocontainment system, the sample will travel back to Earth with ERO and land in the United States. The OS has no attitude control system and will be tumbling in orbit with potentially quickly changing light conditions. The image processing for rendezvous guidance navigation and control (GNC/IP) has to work on space qualified computing units hardened against radiation, but with limited computing power. To verify the critical function for MSR-ERO, a system developed with generated synthetic images is cross validated by experiments with representative hardware on a robotic testbed under real time conditions. The rendezvous image processing used during the tests was an improved version of the system developed during the GENEVIS study . The image processing is a Shape Matching technics which requires a database generated offline using the 3D model of the OS and the knowledge of the mission scenario. The algorithm first segments the image to extract the region of interest around the target. Then a reference view is obtained over a request in the database. The final pose in the camera frame is obtained by geometrical transform compensation between the database sample and the image. The GNC/IP demonstration system consists of a commercial-off-the-shelf camera as representative model of the wide-angle camera, a 52MHz LEON3 on which image processing is implemented, and a 3D printed mockup of the OS. This system is mounted on the EPOS test bench to simulated image acquisition in representative conditions (OS tumbling, MSR-ERO motion and Sun illumination). The EPOS facility integrates two modified 6 axis industrial robots and a 25m linear rail. The relative pose can be set with 250Hz and with submillimeter accuracy. The first robot on the rail transports the IP demonstration system, while the mockup was mounted on the rotating plate on the second robot. This plate as well as the robots and the background are draped in black for high optical contrast to the white sample container. The illumination with spectral properties similar to Sun light is provided by a large mobile spot light. The 3D model of the OS designed by the Jet Propulsion Lab is used to print mockups of the OS. Four different 3d-printed and painted sample container mockups exist. Printing them at scale 1:4 and 1:1 allows to perform an 80m approach on a 25m linear rail. And two different location of the support fixation allow rotation of the sample container around different axis without visible fixations. The experiments represent the baseline scenario of an approach trajectory from 80m to 0.8m distance between the spacecraft with a velocity of 3.5 cm/s. The experiments covered open loop approaches with various chaser and sample container attitude motions, and considered 3 different illumination scenarios. The Shape Matching shows good results on EPOS images, with a range estimation of typically 2% of range at very close distance (2-4m), and a lateral error less than 1cm. The performances observed during the robotic campaign are a bit higher than the ones on simulated images. It is explained by calibration issues (defocus especially) which could not been done because a lack of time, and by a ground truth accuracy limited to misalignment residuals of OS and camera on EPOS robots. To show the suitability of the Shape Matching solution with real MSR-ERO mission, it has been implemented on LEON3 processor (@52Hz). The observed CPU time (on thousands of images) is 400ms, which is sufficient and keeps comfortable margins for a GNC frequency foreseen at 0.5Hz. In summing up, a representative model of a close range GNC/IP was built. It could be successfully demonstrated that the image processing runs on representative flight computer and that the performance is sufficient for a capture manoeuver. For distances below 20m, the pose estimation shows to work with at maximum 1cm lateral deviation and 2.5cm deviation along the imaging axis; enough for a safe capture manoeuver. LICIA GNC baseline for DART-Didymoon impact tracking Capannolo A1, Zanotti G1, Lavagna M1, Simonetti S2, Zannoni M3, Dotto E4, Pirrotta S5 1Politecnico Di Milano, 2ArgoTec, 3Università di Bologna, 4INAF - Istituto Nazionale di Astrofisica, 5ASI - Agenzia Spaziale Italiana In the contest of planetary defence, the DART mission will achieve many different records, among which being the first full-scale kinetic impactor to be used as demonstrator. Another achievement of the DART mission is connected to the release of a witness of such impact, the CubeSat LICIA (Light Italian CubeSat for Imaging Asteroid). LICIA is an ASI () mission that, in collaboration with NASA, will vastly expand the scientific outcome of the DART mission, with the exclusive possibility to acquire images of the outcomes of artificial impact by flying by Didymoon. The spacecraft is a 6U CubeSat that will be equipped with two different imaging cameras, used to acquire relevant images of the newly formed crater, of the generated ejected particles, and of the opposite side of the moonlet, not visible from DART view. The challenges posed by the limited capabilities of the miniaturised spacecraft are reflected in particular on the GNC architecture, that shall be able to satisfy stringent requirements for the sake of science. The GNC subsystem consists of reaction wheels and a delta-v thruster as actuators, and a sensors suite composed by the payload cameras, sun-sensors and a star tracker. This paper will present the process followed to design the GNC architecture for what concerns the attitude of the CubeSat. A particular focus is posed on the presentation of the Guidance and Control. A brief description of the implemented Navigation strategy is also given. The main mission phases that drove the design of the guidance are the approach phase (from release by DART spacecraft, to its impact on Didymoon's surface) and the science phase, core of the mission where the takes place, of the duration of few minutes. Each phase is characterised by different requirements and associated performances. The approach phase is characterised by several slews due to the alternation of Sun and Earth pointing, but the dynamics of the bodies of interest is such that very slow relative motion is experienced, and steady state pointing strategies is a viable option. A set of slew strategies have been proposed, based on the motion of two axes of the spacecraft, containing the full set of instruments to be pointed. All proposed strategies have the objective of satisfying the pointing of the body axis of interest (camera, solar panels, antennas), called primary axis for convenience, to the desired target. The individual differences are instead related to the objective for the secondary axis (the axis not used in the current mode). The combination of primary axis pointing and secondary axis motion define the complete rotation that the actuators have to follow. Each strategy is evaluated in terms of minimisation of the magnitude of the cumulative attitude rotations, to reduce control's convergence time and propagation of actuators' errors, and of other properties related to manoeuvres execution time and transition to science phase. In contrast, the science phase is characterized by a constant pointing of the targeted asteroid, and no mode change is planned. However, the high flyby speed requires a high performance tracking, capable of reducing the load on the reaction wheels mounted on the CubeSat, while ensuring a constant presence of the target within the camera's field of view and a contained pointing oscillation frequency. Again, different strategies are evaluated, to specify the orientation of the spacecraft during flyby. Their performances are evaluated in terms of visibility of target's surroundings (useful to maximise ejecta imaging coverage) and of rotation axes coupling (which affects oscillations and convergence rate of the controller). The navigation strategies that are employed during the different phases of the mission are defined by the suite of sensors that are going to be used. Indeed during the approach phase the distance from Didymos system is too high to be able to resolve the asteroids with the camera, meaning that only star tracker and sun sensors can be used for the attitude reconstruction, while dedicated DSN tracking windows will be exploited to perform on-ground orbit determination. The science phase, instead, is more critical and the need of fast tracking demands an autonomous navigation. The smaller distance enables the exploitation of the embarked payload for the navigation, that exploits peculiar Image Processing algorithm to retrieve the attitude state of the CubeSat. For what concerns the control design, a continuous control algorithm is required, being the reaction wheels block the only available attitude actuator. Due to the computational simplicity, a simple proportional derivative controller has been studied as basis. The design of the PD controller followed different tuning strategies, that are dictated also by the phase of the mission, due to the different navigation strategies that are employed. For what concerns the approach phase, the tuning has been done taking as reference a single slew manoeuvre that switches between Sun and Earth pointing. The PD gains have been optimised by minimising the convergence time of the slew, imposing constraints on the overshoot of the response and on the maximum accumulated angular momentum, in order to avoid wheels saturation The science phase results as more critical, due to the stringent requirements in terms of pointing errors and actuation synchronisation. Different tuning strategies have been added due to the possible guidance scenarios. A first tuning was performed to minimise the control gains themselves, with the objective of avoiding an overbearing control action. A different strategy was followed by minimising the control energy of the system. In all cases, constraints on the maximum angular momentum accumulated in the wheels and the maximum pointing error are imposed. The simulations with the optimal gains show good performances in terms of pointing errors, but introduce some small amplitude and high-frequencies oscillations that are detrimental to the scientific images acquisitions. Other more complex control strategies are explored to assess a performance improvement to the quality of the observations.

The LISA DFACS: overview of the design activities for the drag-free mode Vidano S1, Novara C1, Grzymisch J2, Pagone M1 1Politecnico Di Torino, 2European Space Agency This paper summarizes the activities carried out during the LISA Drag Free and Attitude Control System (DFACS) preliminary prototyping study under the ESA Technology Development Element (TDE) program. This activity aimed at the mathematical modelling of the LISA spacecraft and test mass (TM) dynamics; at the design of high performance controllers for the science phases of the LISA mission. In particular, the activities related to the control design for the drag-free mode are here presented.

The idea of a space-based interferometer for the measurement of gravitational waves in the 0.02 mHz - 1 Hz bandwidth had its origin in the 90's, since ground-based interferometers are generally affected at low frequencies by the Earth's seismic activity and consequently cannot detect the gravitational waves emitted by some astrophysical objects. LIGO and Virgo ground-based interferometers were built in the mid 90's and had been improved in the following decades, achieving the first observation of a gravitational wave in 2015.

The overall complexity of such kind of space mission and the uncertainties regarding the current state of the art of technology, persuaded ESA to develop a pioneering mission in the mid '00. LISA Pathfinder was launched in 2015 and performed in-orbit tests until 2017, demonstrating key technologies required to satisfy LISA performance requirements. The success of this mission, together with the first observations of gravitational waves by means of ground based interferometers, ultimately resulted in the selection of LISA as the third large class mission in the next ESA Cosmic Vision program. In 2018, ESA started several technology development studies for the LISA mission, such as the system Phase-A and the LISA DFACS prototyping, which this paper addresses. This study began with a review of the LISA concept in order to identify the bodies involved, the degrees of freedom, the kinematic chains, the actuators and sensors characteristics, and derive a mathematical model for the GNC design. The current LISA concept consists of a constellation of three spacecraft travelling on different inclined heliocentric orbits. This results in a Sun-facing spinning triangle with a nominal side length of 2÷5e6 km that follows the Earth at an average distance of 60e6 km. Due to the orbital dynamics, the inner angles of the triangle change periodically between 59° and 61°. Each spacecraft consists of a science module that carries two moving Optical Assemblies (OA) whose purpose is to track the breathing angle of the constellation. Each OA is composed of a telescope, an optical bench for laser interferometry and an electrostatic suspension system (also known as Gravitational Reference Sensor), which houses a suspended cubic test mass (TM). The GRS was tested in-orbit by LISA Pathfinder and will be inherited by LISA. The spacecraft is also a well-balanced system given its geometrical symmetry and mass distribution, such that the overall center of mass is close to the barycentre and there is little spacecraft induced gravity on each test mass. Since the control problem is based on nanoscopic scale quantities, some disturbances that affect the system dynamics are not negligible. For instance, even if the test masses are apparently unconnected and suspended inside the spacecraft, the local electromagnetic and gravitational fields determine low-scale couplings that act as virtual springs. The spacecraft/test mass stiffness problem has been also addressed in LISA Pathfinder and in some other preliminary studies about LISA. To conclude, a single LISA spacecraft is a multi-body system characterized by 20 degrees of freedom: (i) 6 DoFs for the external body; (ii) 1 DoF for each optical assembly; (iii) 6 DoFs for each test mass. After a literature review, suitable reference systems were assigned to the plant and a nonlinear model of the dynamics was obtained by means of a Newton-Euler approach. Then, the analytical nonlinear model was linearized and written in state space form, obtaining a 20x17 MIMO system. All the resulting 340 transfer functions are unstable having a double pole at the origin. In order to validate the model, the frequency and time responses have been compared with a SimScape multibody benchmark model. SimScape is a CAD-like software based on Matlab/Simulink, allowing to build models of complex physical systems using a set of elementary blocks, avoiding the use of mathematical equations. Although the SimScape model is accurate in simulating the spacecraft dynamics, it does not provide a mathematical description of the system that is being modeled.

The central part of this study was concerned with control design, where the objective was to satisfy several nanoscopic-scale requirements in the frequency domain for each output variable. If these requirements are not fulfilled, gravitational waves cannot be observed. Therefore, it was necessary to design a control system able to manage the plant instability, to compensate for the environment, actuation, sensing noises and to fullfil very demanding performance requirements. In the last few years, thanks to LISA Pathfinder, it has been possible to characterize many of the LISA noise contributors in the frequency domain. For this reason, by knowing the frequency noise shapes and the plant transfer functions, a very promising drag-free control technique to be used is the mixed sensitivity h-infinity. Moreover, since the B matrix of the linearized state space model is full rank, it is possible to perform a full decoupling of the system. Hence, the final control architecture is based on the decoupled plant, obtaining a parallel of h-infinity SISO controllers. The mixed sensitivity technique relates the weighting functions to the frequency requirements and the noise shapes of sensors and actuators. Then, a solver finds the optimal controller able to stabilize the system and satisfy the constraints of the weighting functions. Finally, the obtained controllers have been tested in a Monte Carlo campaign in order to evaluate robustness against parametric variations, such as the spacecraft mass and inertia or the spacecraft-test mass stiffness matrix. Investigation of Multi-Body/Multi-Actuator Modeling Techniques for Applicability to Future Space Observation Missions Ponche A1, Marcos A1, Ott T2, Schleicher A2 1University Of Bristol, 2Airbus Defence and Space GmbH The science goals of future observation missions become increasingly demanding in terms of availability, precision (e.g. of the line-of-sight (LoS) pointing) and spacecraft agility requirements. These demands are driven by the fact that maximization of operational time, telescope aperture (especially for missions with multiple instruments) and scientific image resolution would lead to an increased scientific return. The currently envisioned design solutions rely on the use of multi-body and multi-actuator spacecraft that offer the possibility to perform interconnected manoeuvres between the main (rigid and flexible) spacecraft body and the scientific payload/instrument (such as large mirrors of up to 1/3 the total weight). Such spacecraft architectures are promising as they can improve the pointing performance, but represent a challenge to the attitude control design. Modern robust control design techniques have the capability to address the multi- body/multi-actuator (MB/MA) challenges as they are specifically developed to handle multi-input-multi- output (MIMO) systems as well as robustness issues. In order to develop a robust attitude control design for this type of missions, it is first necessary to obtain adequate models. In turn, the modeling step can be addressed by studying first the attitude control design criteria (as adapted to MB/MA spacecraft) and second, by investigating the possible MB/MA dynamics modeling options. One example for this type of MB/MA systems is the future Advanced Telescope for High-Energy Astrophysics (), which is composed of a main body and a significant appendage mass mounted on a hexapod mechanism, which enables the mass to move with six degrees of freedom relatively to the main body.

For this study, and without loss of generality, a two-body system connected by a six-degree-of-freedom actuator in free space is considered. In addition, the system will include non-collocated attitude actuator/sensor pairs, formed by one of the bodies containing the attitude actuators while the other (considered as an appendage) contains the attitude sensor and moves relatively to the other body. This moving body implies a change in the global spacecraft attitude and thus, it is important to study whether the control system can cope with this change and/or allow both bodies to move in parallel in order to optimize the full spacecraft motion.

This article presents a study on the most appropriate spacecraft dynamics modeling techniques required to model MB/MA spacecraft systems. The study started by developing a two-body model of sufficient detail to simulate and quantify the impact of a body motion on the overall spacecraft attitude. A first modeling method led to an analytical dynamical formulation of the two-body system, but the resulting model is not easily extendible (to other bodies and/or flexible appendages), since it requires a new formulation and verification for each added component. Therefore, other MB/MA modeling approaches have been investigated to formulate the dynamics of the initial two-body model in a form more amenable to expansion while retaining the physical nature of the components. The resulting global system dynamic model can be used to derive uncertain models (in the robust modeling paradigm of a linear fractional transformation or LFT) and subsequently to design robust controllers.

The results presented here are the outcome of the first year of a PhD project conducted by the University of Bristol, UK, and Airbus Defence and Space GmbH in Friedrichshafen, Germany. The results show a promising research direction to model future MB/MA spacecraft towards increasing their agility and operational availability, as well as to improve pointing performance through robust attitude control architectures. This research on relative motion of interconnected rigid or flexible bodies aims at enabling similar technology for any future MB/MA space systems such as the Geo HR and ARIEL missions (using a multi-actuator line-of-sight control) or the TerraSAR-X and NFIRE missions (using interconnected body motions through laser telecommunication terminals).

Closed-Loop Guidance for Low-Thrust Interplanetary Trajectories Using Convex Programming Hofmann C1, Topputo F1 1Politecnico Di Milano A closed-loop guidance scheme for low-thrust interplanetary transfers is developed. The algorithm is based on convex programming and repeatedly recomputes reference trajectories in certain time intervals. The state of the spacecraft is propagated during these periods using the obtained controls until a new trajectory is to be calculated. A mesh refinement procedure adjusts the number of nodes based on the linearization error. The effectiveness of the approach is demonstrated in several numerical simulations. The proposed method is a promising step towards autonomous guidance in real space missions due to its rapid speed and excellent robustness. Autonomy, Fault Tolerant Control and Operations

Autonomous Guidance for Electrical Orbit Raising Lagadec K1, Locoche S1, Erb S2 1Airbus Defence And Space, 2ESA/ESTEC The use of electric propulsion for orbit raising is a key factor for reducing the cost of access to space, for telecommunication satellites going to geostationary orbit (E172B, SES14, SES12) as well as for LEO satellites climbing to their operational altitude. The increased time-to-orbit is however a downside, with additional operation costs from ground station usage and personnel. Moreover, when envisaging large LEO constellations, the number of satellites that must be simultaneously operated might result in operation bottlenecks.

Increasing autonomy is therefore an important objective for reducing EOR costs, and the anticipated generalization of on-board GNSS receivers is a key element towards that goal. We present the results of a 2- year R&D study (ARTES funding) which looked into guidance strategies that would make autonomous EOR possible while being compatible with on-board CPU limitations.

The study investigated two major classes of solutions:

- semi-autonomous solutions are an extension of what Airbus is currently implementing for telecom missions with EOR: the use of on-board navigation data and smart compression allows to extend the validity of the ground-optimized guidance profile from a few weeks to a few months, thus reducing ground intervention by 80% with only minor changes to the current architecture.

- fully autonomous solutions represent a significant step further, where the on-board system can autonomously compute an optimum guidance scenario (thrust profile and complete 3-axis attitude trajectory), based on the latest GNSS data, thus limiting ground intervention to health-checks and collision-risk monitoring. It is this latter class of solutions which we present here.

After comparing several alternatives for the fully autonomous guidance approach, a best candidate solution was selected and implemented in detail. Functional simulations over a complete transfer showed that despite the disturbances and uncertainties, the performance in terms of optimality was remarkable (propellant expenditure was within 1% of the optimum, as computed by Airbus's OptElec tool).

The guidance software prototype then went through the same formal code-generation process as for real on- board software developments, and tested on a LEON3 processor. Preliminary results from the Processor-In- the-Loop test campaign show that the peak combined CPU load for the 2 asynchronous tasks plus the cyclic task amounts to less than 1%: this demonstrates that the fully autonomous guidance algorithms can indeed be implemented on board.

Measuring Resilience of Autonomous Controllers to Spacecraft Missed Thrust Events Rubinsztejn A1, Sood R1, Laipert F2 1The University Of Alabama, 2Jet Propulsion Lab As spacecraft adopt low-thrust electric propulsion to reduce propellant mass, they will need to contend with new challenges inherent in these technologies. Electric propulsion’s susceptibility to missed thrust events is especially challenging due to their stochastic nature and possibly mission-ending nature. While most research focuses on designing baseline trajectories resilient to missed thrust events, a new paradigm has recently emerged, autonomous recovery. With autonomous recovery, spacecraft do not need to wait for the flight controllers to design a new trajectory after a missed thrust event but can do so on board, saving time and allowing a faster response. This paper establishes metrics to compare different autonomous approaches and their responses to missed thrust events. Additionally, motivated by ESA’s Earth Return Orbiter in the Mars Sample Return mission concept, this work compares four different autonomous recovery methods for a sample Mars to Earth return trajectory.

GNC for In Orbit Robotic Operations

Modelling and attitude control design for autonomous in-orbit assembly CUMER C1, Rognant M1, Biannic J1, Roos C1 1ONERA One of the key challenges for future space missions is the autonomous in-orbit assembly of large structures, when they cannot be self-deployed as a single piece. The H2020 project PULSAR (acronym for ”Prototype of an Ultra Large Structure Assembly Robot”), which started on February 2019, aims at developing and demonstrating key technologies for in-space assembly of the primary mirror of a large telescope thanks to three demonstrators. Even if a specific application has been identified for this project, some developments could extend to other needs, such as the assembly of solar panels for power plants, light sails to reach the outermost regions of the solar system. This is particularly the case of the modeling part of the global satellite system: the idea is to obtain a model sufficiently representative of the inertia changes during the assembly phase and to make a pre-study of the achievable performances of both the deployment system and the attitude control system in the presence of disturbances and depending on the available actuators.

The modeling tool must be able to represent the various dynamic couplings in both translation and rotation which exist between the subsystems of the satellite, namely: - the main hub, which is assumed to be rigid, - two flexible solar panels, - a sunshield, which is represented through its four support beams; these flexible beams are cantilevered on the four sides of the main hub, - 36 segmented mirror tiles (SMT): they are first stored in 6 stacks in a SMT container directly attached to the main hub; once assembled, their connection interfaces induce flexibility, - a manipulator arm, which ensures the assembly task, by moving along a rail; the modeling tool must take into account the selected assembly scenario, which consists in pre-assemblies of 5 tiles in an intermediate plane, before integrating them on the mirror under construction; a first pre-assembly of 3 tiles constitutes the base of the mirror.

Physical properties of these subsystems are deduced from a technology review and a mission analysis (https://www.h2020-pulsar.eu/listings), with a clear reference to similar missions like the James Webb Space Telescope (launch is currently scheduled in 2021). In this context, two modelling tools are used: - the first one is the Satellite Dynamics Toolbox (SDT - https://personnel.isae-supaero.fr/danielalazard/ matlab-packages/satellite-dynamics-toolbox.html), based on a multi-body modelling approach. This toolbox allows to build generically a parametric linear model of a satellite composed of a rigid hub, rigid appendages and/or flexible appendages, - the second one is the open-source NPS-SRL/SPART (https://spart.readthedocs.io/en/latest/ ), which can derive easily the dynamics of the robotic manipulator. It is explicitly dedicated to kinematic trees composed of rigid links and joints. The implementation of these tools on the PULSAR use case will be presented in the first section of this paper.

In the second section, our main concern will be the design of efficient attitude controllers for the deployment phase and the observation phase. During the deployment phase, the main issue consists in preserving a reasonable pointing accuracy of the satellite to maintain the link with the ground station, despite the slow evolution of the inertia and frequency modes of the system and the torque perturbations that are generated by the robotic arm motions. During the observation phase, the inertia matrix stops varying significantly. However, the desired pointing accuracy becomes more stringent and the fully deployed primary mirror tends to generate badly damped and rather low frequency torque perturbations.

To address these two phases a modular attitude controller structure, will be used. It consists of six blocks: 1. Gain-Scheduled PID controller with Anti-Windup: This is the central element of the control system. In nominal conditions, the PID controller ensures good stability properties during the deployment phase with a low contribution (PD type behavior) and enhanced performance properties (pointing accuracy) during the observation phase with higher integral action (and possibly lower stability margins). Mainly during the deployment phase, the gains will be adapted to the varying inertia of the system. Mainly during the observation phase, the integral gain will be adapted to the control activity by an anti- windup device to account for the limited torque capacity of the reaction wheels. 2. Robotic Arm System and Segmented Mirror Torque Perturbation Estimator: Since the main perturbations generated by the robotic arm system, the solar panels and the deployed mirrors essentially act as matched disturbance torques, they can be efficiently compensated by the proposed structure if a good estimation of these torques is available. The objective of this block is thus to provide such an estimation with the help of a robust observer. 3. Input Shaping Filter: In case of possibly too aggressive maneuvers (delivered by the AOCS reference block which is linked to the mission) flexible modes might be unduly excited or limited torque production capacities of the reaction wheel system might be exceeded (limited amplitude and limited kinematic momentum beyond which the torque cancels out). In such cases, the anti-windup device will be helpless. The role of this filter is then to shape the reference signal according to the current control activity, the inertia of the system and the approximate frequency of the bending modes. 4. Sensors Fusion: The current attitude and angular velocity of the satellite are provided by sensors fusion of the Star Tracker and Gyroscope. 5. System Inertia and Primary Frequency Estimator: This block plays a keyrole in the structure since it provides on-line estimations of the varying inertia (which is rather slow but significant during mirror deployment) and varying bending modes frequencies. The first information regarding inertia is essentially used to adapt the PID gains while the second information regarding frequency is mainly used by the input shaping filter. 6. AOCS Reference: This block is outside the scope of the attitude control system. It provides the reference attitude trajectory to be tracked: desired attitude angles, angular velocities and accelerations. Finally an implementation of this attitude controller structure will be presented in a third section. EROSS Project – Coordinated control architecture of a space robot for capture and servicing operations Dubanchet V1, Casu D1, Rekleitis G2, Paraskevas I2, Papadopoulos E2, Andiappane S1 1Thales Alenia Space, 2National Technical University of Athens (NTUA) The on-going developments of On-Orbit Servicing operations are massively investigating the usage of space robotics to realise precise tasks in an autonomous way. Such technologies raise the issue of properly coordinating the motion of a robotic arm mounted on a floating platform in order to ensure the tracking and/or capture of a target object. The Guidance, Navigation and Control (GNC) architecture of such a system, often referred to as “space robot”, is presented in this paper, with an emphasis on the control algorithms and how these are being implemented using the platform and the robotic arm processing units to account for the capabilities and limitations of the space hardware at hand.

The work to be presented in this paper is related to the H2020 project “European Robotic Orbital Support Services” (EROSS), which receives funding from the European Union’s Horizon 2020 research and innovation programme under grant agreement No 821904. The paper will be organised as follows: a first section will cover the EROSS servicing mission, with a particular focus on the design of the space robot used for this operation; then a second section will describe the overall GNC architecture with the platform and robot sharing, and will introduce the theoretical foundations of the coordinated control. A third section will provide simulation results of the reachable performances for the berthing with a collaborative spacecraft. Next, a short section will tackle the capabilities of the processing units being developed to run such controllers on-board, before providing conclusions and recommendations on the hardware implementation of these technologies.

The EROSS mission scenario focuses on the last steps of a traditional rendezvous in space between a Servicer spacecraft chasing a Client spacecraft to be serviced. After the orbit injection by the launcher, a first set of orbital manoeuvres puts the Servicer within the orbital plane of the Client, then a second set of manoeuvres brings it closer to the Client, either behind or above it [1]. The final forced motion allows the Servicer to perform a straight line relatively to the Client in order to capture it by the robotic arm. The coordinated control of the Servicer space robot is then designed to allow a smooth and safe deployment of the robotic arm while the Servicer platform is moving towards the Client, and also to perform the final capture with a Servicer platform floating, for safety purposes.

To that end, the Servicer vehicle is seen as a traditional platform when its embedded arm is stowed. During that time, the platform controller is based on traditional techniques developed within Thales Alenia Space to ensure the proper thruster pointing during the orbital manoeuvres, to maintain the solar panels illumination for power generation, while also keeping track of the Client spacecraft within the relative sensors field of view when the relative navigation starts. On the other hand, as soon as the robotic arm is deployed during the forced motion, the Servicer turns into a “space robot” [2-3-4] with multiple Degrees-of-Freedom (DoF) which moves and aligns the arm end-effector with a target point on the Client spacecraft.

Many techniques were developed in the past to control such a system. The National Technical University of Athens (NTUA) played a key role in these developments, and it is in charge of the robotic control in the EROSS project. Different GNC techniques are summarized in [5] for a space robot, with the main hypothesis of controlling or not the platform when the arm is moving. In the first case, the space robot is said to be “free flying” when the platform controller is actuating the thrusters or reaction wheels to maintain a given pointing while compensating for the disturbances coming from the robotic arm motion. From a practical point of view, this method induces vibrations from the platform to the robotic arm since its structure is very light and sensitive to any disturbance exciting its low frequency flexible modes. On the other hand, the platform controller can also be completely switched off to prevent the transmission of any residual vibration from the platform to the arm: the space robot is then said to be “free-floating”. This method also presents the additional advantage of saving fuel and power on the platform side, but at the expense of a reduced workspace for the end-effector of the robotic arm, and of the loss of direct position and attitude control of the Servicer base, which may rise safety issues at such close proximity. The proper trade-off must then be made to select the relative distance between the Servicer and Client platforms at which the robot switches from a free-flying to a free-floating mode in order to capture the Client with a maximum motion accuracy and robustness.

The control architecture designed for the EROSS mission will be evaluated and validated through numerical simulations using a high-fidelity simulator, by coupling traditional Attitude and Orbit Control System (AOCS) simulators from Thales Alenia Space with multi-body modelling tools to account for the dynamics of the space robot during the capture of the Client. The performance will be presented in terms of relative position and attitude tracking error and attitude between the Client capture point with respect to the end- effector.

In addition to this kinematic figure of merit, another performance index will be derived by investigating the complexity of such algorithms and by analysing their compatibility with the current and on-going developments of space processors. This last step is necessary to bridge the gap between the theoretical works available in the literature and their implementation on a real space robot for a short-term mission.

The resulting performances and trade-offs will eventually be reviewed to propose improvements and recommendations for future space robotics developments from an industrial and practical point of view.

[1] Fehse, “Automated Rendezvous and Docking of Spacecraft”, 2003. [2] Flores-Abad et al., “A review of space robotics technologies for on-orbit servicing”, 2014. [3] Yoshida and Wilcox, “Space Robots and Systems”, 2008. [4] Yoshida, “Achievements in Space Robotics”, 2009. [5] Moosavian and Papadopoulos, “Free-flying robots in space”, 2007.

Results of the COMRADE project: Combined control for robotic spacecraft and manipulator in servicing missions: Active Debris Removal and Re-fuelling Colmenarejo P1, Santos N2, Serra P2, Telaar J3, Strauch H3, De Stefano M4, Giordano A4,7, Mishra H4, Ott C4, Henry D5, Visentin G6 1GMV, 2GMV Skysoft, 3Airbus Defence and Space, 4Deutsches Zentrum für Luft- und Raumfahrt (DLR e.V.), Institute of Robotics and Mechatronics, 5Bordeaux University - IMS-LAPS, 6European Space Agency, 7Technische Universität München The current space debris environment poses a safety hazard to operational spacecraft as well as a hazard to public safety and property in cases of uncontrolled re-entry events. The accidental 2009 satellite collision between Iridium-33 and Cosmos-2251 led to a debris cloud with up to 823 new large debris object catalogued and many others not catalogued. Among the different possible remediation actions to the debris increase it is the Active Debris Removal (ADR). COMRADE activity pertains to the particular ADR technologies that involves a chaser equipped with robot(s) being operated in tight coordination with the chaser platform motion. Historically, decoupled control and collaborative control have shown that solutions based on them seem to work for some simulated cases (e.g. German DEOS mission and ESA e.Deorbit targeting ENVISAT satellite). However, concerns are raised by the fact that the attribution of control authority to the robot and platform is arbitrary, tailored to the specific situation and may not be adequate if the actual situation in the ADR mission differs from that modelled ahead of the mission. In addition, the rigid attribution of control authority limits the possibility to treat failures and anomalies with a coordinated approach that may use the highly redundant motion ability of the chaser (possibly higher than 12 degrees of freedom) to continue the operation unaffected by the contingency. COMRADE objective has been to design, develop, and test the control system of a robotic spacecraft (i.e. a servicing spacecraft equipped with a manipulator) tasked to perform an Active Debris Removal (e.Deorbit mission used as reference) and a refuelling mission (ESA ASSIST reference mission). This control system provides for combined control and management of the whole chaser, during approach, grasping, stabilisation and hold of the debris with the aim to perform the controlled de-orbit. COMRADE control system has been designed as a multi variable combined (12-13 Degrees of Freedom) control system using modern robust multi variable synthesis methods able to handle the uncertainties of the system using integrated design methods leading to that and realistic state-of-art equipment. The GNC and Avionics architecture for COMRADE is common to both mission scenarios. Since the ADR/e.Deorbit scenario is more challenging (i.e. uncooperative and tumbling target), this is the sizing case. The robot arm is equipped with 7 joints, a gripper and a vision system for relative navigation between gripper and grasping point. The combined controller for chaser platform and robot arm issues force and torque commands for the platform, which are translated into thruster opening times by the thruster management function, and joint torque commands which are realised by the inner joint control loop. The inner joint control loop is closed by the joint torque sensors. It commands a motor torque to establish the desired joint torque level. The combined control loop also needs the joint position measurement as input. - Full scenario phases have been covered: approach and synchronization phase, reach/capture/rigidization phase and stabilization/detumbling phase. - The Control System includes Guidance, Navigation, Controller, FDA/FTC and Modes Manager functions, with special emphasis on Controller function design, analysis and validation. - Approach/synchronization phase has considered robust H∞ 6DOF controller over a rigid body with sloshing and flexibility (solar arrays, stored robotic manipulator) effects as main perturbations. The linear (µ-analysis) and non-linear (Monte Carlo with non-linear MIL simulator) results have demonstrated the required performances (including robust performance and stability). - Reach, capture and rigidization phase has considered a dual approach and implementation: ­ Robust H∞ 13DOF controller over a multi-body system composed by the spacecraft platform plus a robotic manipulator with 7DOF (and grasping/re-fuelling end-effector at the end). ­ A compliance/impedance 13DOF controller over the same multi-body system as for the robust H∞ controller. - Stabilization/detumbling phase has considered robust H∞ 3DOF attitude controller over the full composite (chaser spacecraft + target spacecraft + rigidized robotic manipulator joining both vehicles) with sloshing and flexibility (solar arrays, stored robotic manipultor) effects as main perturbations. Obtained results have demonstrated the required performances (including robust performance and stability). - Advanced FDA/FTC techniques have been also considered as an additional Failure Detection and Accommodation layer on top of the nominal control design. ­ During the synchronization phase, the FDA/FTC has concentrated over thrusters failures. The FDI unit is based on a residual generator based on a bank of nonlinear unknown input observers + a decision making rule that uses a set mechanism based on the residual. The FTC unit is based on the NIPC technique ­ During Reach/Capture phase, the FDA/FTC has concentrated over robotic arm joints failures. The applied FTC solution are: At control level: schedule the NDI inner loop by zeroing the adequate column of the Jacobian matrix. At guidance level: if the fault can be accommodated at the robot level, zeroing the adequate column of the Jacobian matrix and tuning the developed singularity avoidance algorithm; otherwise, same as above, and perform a new attitude/position trajectory to track for the platform while maintaining the gripper trajectory, considering collision avoidance with the target.

An integrated, coherent and incremental DDVV approach concept based on the chain MIL ==> Autocoding ==> SIL ==> PIL T/B ==> HIL T/B have been used. The PIL test bench has been used to measure the execution time of the generated code so as to guarantee compatibility with realistic space-representative processor capabilities and increase the Control maturity and readiness towards its flyable version. Finally, a HIL-based verification/validation campaign has been performed using HW (robotic manipulator, gripper, refuelling device, vision-based camera/Image Processing) in the loop to validate the Control design and performances in presence of real HW equipment. The HIL testing has been split between GMV’s platform-art test facility where the PIL has been reused completely as integral component and realistic camera/Image Processing has been added and between the DLR’s OOS-Sim test facility, the latest focused on testing including contact. Critical GNC Aspects for ADR missions Branco J1, Colmenarejo P2, Serra P1, Peters T2 1GMV, 2GMV In the past decade, GMV has been, both as a participant and a reviewer, at the forefront of ESA’s efforts in identifying and responding to criticalities in GNC technologies for future ADR and in-orbit servicing missions. This paper compiles in a critical analysis the state of the art of ADR mission design from a GNC perspective, informed from GMV's recent experience in the topic. In addition to outlining some of the main challenges and describing the state-of-the-art of GNC responses, it outlines and describes novel methodologies for design, development, verification and validation to advance of their readiness. This review of results and extraction of conclusions is framed around the efforts to develop a cost-driven solution to perform the removal from LEO orbits, particularly SSO, of ESA-owned satellites and demonstrate capabilities and technologies for in-orbit servicing. GNC Critical Aspects within ADR missions covered are: concepts of operations; guidance - angular synchronization, passive safety trajectory planning; control – coupled, compliant and combined control , precise attitude control during boosts, 6 DOF force motion control during synchronization phases, visual servoing and compliant 10+DoF control during reach, phase rate damping and MCI estimation; capture devices; camera-only-based navigation – image processing from acquisition and tracking, through bearings- only navigation, to model-based pose estimation; operations control authority and autonomy levels, modes management; software and GNC architectures and DDVV cycles; FDIR. The know-how to inform the critical was acquired in mission development activities, from flagship initatives to large targets to CubeSat in-orbit demonstrators; and specific element technological development programmes , particularly to investigate and mature the complex couplings between the different control systems (GNC including image processing and robotics) for autonomous capture, by deriving required algorithms and performing HW-in-the-loop end-to-end demonstration. Recent focus has been applied to fully combined control for the operations of grasping, stabilisation and hold of the debris with the aim to perform the controlled de-orbit, as an alternative to decoupled, tele-op, collaborative options - to overcome the problem of arbitrary, case-tailored control authority. Usage of the non-visible spectrum to tackle challenges like eclipses and early detection has also been a recent subject of investigating , as well as the topic of de- tumbling strategies, particularly robust control for robotic arm capture but also contactless detumbling. This paper is a next step in a series of iteration of application of lessons learnt in a proposed small scale missions focusing on the feasibility of removing small debris objects while testing servicing and capture techniques such as nets and a robotic arms. Attention is given to system-level and mission-level aspects and scalability. Collision-free guidance: Rigid-link capture strategies may require the chaser to synchronize its attitude and translation motion with a tumbling target. In this paper a first-order logic and guidelines are establish to justify guidance strategies able to guarantee a collision free approach and capture wrt to the shape (particularly protruding elements and location of access point), tumbling state of the target and type of capture mechanism. GMV advances Poinsot’s description of the motion to guarantee that the approach over the angular momentum vector is collision-free. Contacteless detumbling: The propulsion system of the servicer spacecraft may not have enough control authority to effectively perform the synchronization manoeuvre at the desired angular rate. GMV proposed a method to contactless detumble prior to capture, based on plume impingement. The method consists of using the gas jet exhaust from a thruster mounted on the chaser to influence the target. Advanced Control Techniques The chaser spacecraft often includes flexible elements and fuel sloshing with uncertain properties (the latter are particularly challenging to determine in advance of the operation). These, together with elements such as robotic manipulators, end-effectors and clamping mechanisms, introduce couplings between translation and rotation states and the mathematical model is only partially known. Moreover, proximity manoeuvring impose stringent performance requirements to the control system that must be considered at synthesis level. Advanced robust control techniques that allow to formulate disturbance and noise rejection problems in the frequency domain were applied by GMV to the synthesis and analysis of controllers for the most critical phases. Visual servo and Compliant Control The problem of capture the target with a robotic arm, after the synchronization manoeuvre has been tackled using independent controllers for the chaser base and robotic manipulator and by applying different compliant control techniques. The paper will address different solutions: basing the manipulator on visual-servo control decoupled from a translational robust 6DoF controller for the platform; using visual servoing to guide the robotic arm to the capture position by means of 3D model matching; and full DoF robust combined H∞ control, including in the synthesis-level weights that allow to limit the joint loads and issuing commands to the robotic arm joints and to the chaser propulsion system. Nonlinear compliant control techniques will also be reviewed – these ensure that during the motion of the arm the forces and torques (especially acting on the gripper) are kept within a certain limit and that the gripper, the robotic arm and the clamp are not damaged. Coupled-control After the capture with a robotic manipulator or clamping mechanism, if the coupled system is tumbling, the chaser must keep actuating to avoid large loads on linking system. This problem has been addressed in the referred studies with specially designed GNC mode designed to slow down and deorbit the coupled system while limiting the loads.

Real-time combined control for active debris removal for a satellite with a robot arm implemented on a SBC connected to a detailed multi- physics and VR simulation Reiner M1 1German Aerospace Center (DLR) Within a recent project of ESA together with DLR-SR an implementation for a real-time combined control setup for a satellite with a robot arm was developed which is running on a single board computer (SBC) connected to a detailed multi-physics and VR simulation. The combined controller, which simultaneously controls both the 7-axis robot arm as well as the 6 DOF satellite, is a further development from a previous project to be able to run in real-time on a relative weak SBC as an intermediate step to increase the technology readiness level (TRL) of the control system. The controller is implemented as an advanced robust discrete H_inf controller und combined with an online trajectory planning for use within an active debris removal (ADR) scenario. For the tuning of the controller a multi-stage optimization is used to optimize the weighting filters of the loop shaping design. The optimization takes linear criteria such as stability, phase margin and roll-off as well as nonlinear system performance criteria into account. The real-time implementation also includes the inverse kinematics for the 7-axis robot arm and a lot of changes were necessary to make the system real-time capable on the SBC. The SBC is coupled to a strong desktop PC where the main “model in the loop” (MIL) simulator as well as a virtual reality (VR) visualization are running. The main model of the ADR scenario is a detailed multi- physical simulation model based on the Modelica modelling language. It uses advanced libraries developed within the DLR for satellites, the orbit environment as well as robots. The MIL includes a detailed orbit simulation including physical effects such as sloshing and detailed thruster models for the chaser satellite and the flexibility of the solar panel for the target satellite (Envisat). The 7-axis robot arm on top of the chaser satellite is also implemented in great detail and includes models for the robots flexible powertrains and friction. The MIL simulator was heavily modified from a previous implementation to be able to run in real-time. Some approximations and simplifications of components with a large demand of processing power were necessary to achieve this. The coupling of the MIL, SBC and VR environment is done using a UDP communication setup, as well as using a dedicated network card for the connection of the SBC to the desktop PC. This setup has inherent delays and multiple different discrete sample times which made the implementation difficult. The controller had to be re-tuned to be robust against these varying delays while still keeping a good and robust performance. The original optimization and synthesis setup for the combined controller had to be changed to take this into account. In addition the robustness was also checked using a large Monte Carlo simulation which takes parameter variations for the physical systems as well as different possible communication delays into account. To give the viewer an immersive experience and to be able to look at the ADR scenario from any angle, the MIL simulator is coupled to a detailed VR simulation of the scenario, also running in real-time. Within the VR simulation the user can use an IMU (inertial measurement unit) based joypad to move around the two satellites in three dimensions, to get a view from any position. The VR headset allows the user to look in any direction. The detailed VR environment not only includes both satellites but also the orbit with the earth, sun and moon implemented in a physically correct way (not as simple background images). The sloshing as well as the elastic deformations are also visible within the VR environment. The overall system can give the user a great and detailed insight for the most critical part of an ADR scenario, were the chaser satellite synchronizes with the spinning target satellite and the robot arm is used to grasp the target satellite and then de-tumbles it before a controlled de-orbiting. GNC for Planetary Exploration

Development, tests and results of onboard image processing for JUICE Jonniaux G1, Regnier P1, Brochard R1, Gherardi D2 1Airbus Defence and Space, 2ESA JUICE (Jupiter Icy Moon Explorer) is the first Large Class mission of ESA’s Cosmic Vision program, to be launched in 2022. It includes flybys around Europa, Ganymede and Callisto as well as an orbit insertion around Ganymede.

This mission has stringent pointing accuracy requirements during the Jovian moons fly-bys, phases in which a classical navigation based on Doppler measurements and on-ground image processing with a data cut-off a few days prior the fly-by Closest Approach does not permit to meet the science pointing requirements. Ground-based navigation indeed cannot accurately predict in advance the spacecraft position with respect to the science observation targets on the surface of the moons at the time of the Closest Approach due to moons ephemeris and spacecraft late targeting manoeuvres uncertainties, hence a large attitude guidance error.

In order to enhance the mission performance in this regard, therefore the science return of the mission, Airbus Defence and Space has designed and implemented an ambitious unique autonomous vision-based navigation solution named EAGLE (for Enhanced Autonomous Guidance through Limb Extraction), which will efficiently complement the navigation capabilities of the ground segment.

This paper first presents the Jovian visual environment specificities an image processing algorithm will have to face. It then describes how the JUICE Navigation Camera (NavCam) has been finely modelled to allow highly realistic simulations of the images. The image processing itself is then presented, and finally the validation plan executed to increase the TRL of the algorithms is detailed. In particular, it presents the closed- loop GNC simulations, the hardware-in-the-loop, the real sky tests and the in-orbit experiments that were and will be carried out to ultimately assess the performance of EAGLE.

The validation campaign includes tens of thousands of different cases leading to millions of images both obtained from simulated images and from the actual NavCam engineering model; and only a fraction of those images can be considered as “nominal”. The main challenge for any image processing algorithm that has to run autonomously on-board a spacecraft is not only the nominal performance, but is also its behaviour in all other off-nominal cases. Achieving a good nominal performance is easy compared to achieving high robustness, and the validation plan has been defined accordingly. All specified parameters have been pushed several times beyond their expected worst case and stress case values and new parameters were added that could lead to an off-nominal image. The envelope of correct behaviour has then been estimated, and an auto-check has been implemented in both EAGLE’s image processing and navigation filter so that it can automatically detect whether the image processing output is relevant. It is now known that not only the requested performance is achieved in nominal cases, but it is demonstrated that the AOCS behaviour is robust to EAGLE’s off-nominal situations. In parallel to all those images either directly simulated or obtained through the stimulation of the NavCam EM, and in order to get rid of any potential bias induced by the synthetic nature of the images, tests were conducted on the real sky where nominal and off-nominal images could be taken with the NavCam EM as well.

Application of Advanced Navigation Techniques for Lunar and Mars Pinpoint Landing Hormigo T1, Hormigo T1, Oliveira J1, Câmara F1, Dubois-Matra O2 1Spin.Works S.A., 2ESA-ESTEC In the scope of future planetary missions, a key enabling technology will be the ability to land in a precisely defined area, to allow reaching either small areas with high scientific interest, or pre-existing ground assets (in support of a manned or sample return mission). This work presents an integrated solution for the trajectory design, integrated 6DOF G&C algorithms, navigation filters, sensors and actuators that demonstrably enables pinpoint landing on the Moon and Mars. The proposed solution corresponds to the output from the recently completed ESA activity "Advanced Navigation Techniques for Pinpoint Landing - ANPLE", which was led by Spin.Works with the support from DLR. The core of the work consists of a set of covariance analyses that indicate how various advanced navigation solutions could enable - or not - pinpoint landing at the Lunar south pole and on Mars. To this end, different data filter algorithms are reviewed, image processing algorithms are surveyed as means to acquire surface relative and absolute navigation information, and innovative sensor technologies and processing units are discussed. A feasible trigger (using only data that is known to be available onboard) and a measurement rate are defined for each sensor + mission profile. A trade-off between different navigation solutions is then performed, based on extensive covariance analysis of potential sensor + navigation filter + image processing combinations. Finally, and in light of the European Space Agency current and upcoming missions, three navigation solutions are drawn: near-term missions to the lunar south pole; mid-term missions to Mars with a parachute-based descent phase; and long-term missions to revisit the Moon and Mars (with a purely retropropulsive descent phase). For a Lunar Descent and Landing (D&L) targeting the south pole, it is clear that pinpoint landing requires overcoming significant challenges, especially in the case that vision-based relative navigation algorithms are used in the difficult local lighting conditions. We assume that, starting from a highly inclined (L2 South Near Rectilinear Orbit) orbit corresponding to that of the Lunar Gateway, a de-orbit manoeuvre takes place, followed by a trim manoeuvre to correct for dispersions, and a continuous propulsive braking phase until touchdown. Initial knowledge and dispersion are derived assuming ground-based navigation until the de-orbit manoeuver, and fully autonomous flight (including image-based navigation) thereafter. Apart from the navigation algorithm, we show that a judicious use of currently available guidance and control algorithms is capable of achieving the necessary performance to enable pinpoint landing. With respect to the Mars Entry, Descent and Landing (EDL) mission scenario, and in order to improve the current entry GNC architecture, a lifting (and guided) entry is seen as a natural way forward. After the entry phase, however, we propose two alternative architectures for the descent and landing - one in which a parachute-based descent phase is followed by a propulsive terminal descent (near-term architecture), and another where the entry phase is directly followed by a retropropulsive phase for the last few kilometres (and 100s of m/s) down to the Martian surface (long-term architecture). The rationale is that the latter would allow skipping the parachute phase entirely - along with the associated drift caused by atmospheric uncertainty and winds during this phase -, both eliminating system complexity and ensuring high control authority throughout the EDL sequence of a Mars mission, thereby increasing the likelihood that true pinpoint landing is achievable with incremental technological advances as opposed to new, major investments. The baseline trajectories derived for the Mars case include the interplanetary transfer, entry, descent and landing phases, which are designed together primarily to ensure that the dynamic conditions at the entry point (where 6DOF simulations start) result in entry peak decelerations comparable to those of past missions, that the dynamic conditions at parachute deployment (where applicable) are within existing system operational constraints, and that the propulsive phase minimizes the propellant spent. Initial knowledge and dispersions (at the entry point) are, as in the Lunar case, derived from ground-based navigation up to a few hours before entry, with an autonomous GNC taking over after that point. In both the Lunar and Mars cases, we assume that a Hazard Detection and Avoidance system will be used, and will operate during the last few hundred metres prior to touchdown. Several details are, as a consequence, taken into account for trajectory design purposes (e.g. optimized sensor placement + attitude, restrictions on flight path angle and observation geometry, plume avoidance angle, appropriate time intervals to account for all onboard processing tasks). The culmination of this work corresponds to the implementation of end-to-end 6DOF GNC algorithms (including image processing algorithms), for both the lunar D&L as well as the Mars EDL, on a high-fidelity simulation tool, and the execution of Monte Carlo simulation campaigns with a realistic set of mission, environment and vehicle uncertainties where the proposed navigation solutions are used. The results of this campaign demonstrate that the selected GNC architectures indeed enable true pinpoint landing (safe touchdown with a 3-sigma less than 20m of the intended target, for both Lunar and Mars missions). Furthermore, they strongly suggest that a relatively wide set of achievable technical solutions are within reach that would, if developed, enable pinpoint landing missions in the scope of future planetary exploration.

Autonomous Moon Landing Guidance, Navigation and Control Systems Development and Validation Hamel J1, Neveu D1, Nagaty A1, Minville M1, Mercier G1, Sobiesiak L1 1NGC Aerospace Missions to the Moon are now back in the priorities of governments and private organisations. Many private companies are currently working on the development of Moon landing systems with the objective of providing commercial landing services to space agencies, research institutes and private organisations. In order to enable the customers to land at all lunar sites of interest, it is required to overcome the challenges of providing global access, including reaching the desired area with great accuracy, and landing on various types of terrains, potentially hazardous for the Lander.

Therefore, future landing systems need to be equipped with high-accuracy navigation systems to accurately reach the target and Hazard Detection and Avoidance (HDA) systems to detect and avoid surface hazards at touchdown. High-accuracy optical navigation and HDA have been identified as key enabling technologies by private and governmental organisations willing to explore or commercially exploit Moon surface. However, as of today, these systems are still not commercially available and have very limited or no space flight heritage. These systems are being actively developed by ESA through its PILOT program, by NASA and by private organisations.

A particular challenge with autonomous optical navigation and HDA landing systems is validation. This validation must guarantee the system can operate autonomously in real time and is robust to a specified envelope of dynamical, environmental and lighting conditions. However, these systems operate in a dynamic environment which is difficult to emulate on Earth.

With the new paradigm of future lunar missions being led by commercial organisations, lunar transportation service providers look for affordable and commercially available sensor and processing hardware, in order to reduce overall system cost. Expensive technology development programs and test campaigns requested by risk-adverse space agencies do not economically make sense for commercial service providers. Development and validation costs have to be lower and turn-around time has to be faster.

The paper thus presents validation infrastructure and facilities built for the testing and validation of autonomous landing GNC systems, especially tailored towards validation of systems for commercial landing services. This infrastructure includes high-fidelity software simulation environments, dynamic scaled test facilities and UAV systems for full-scale testing.

The paper also presents the latest results of autonomous Moon landing GNC subsystems using this infrastructure. Hardware-in-the-loop (HWIL) demonstrations of HDA, optical relative navigation and optical absolute navigation will be discussed. These results, obtained by reproducing scaled landing trajectories over surface mock-ups, demonstrate the real-time execution of these systems.

The HDA results have been obtained in the context of the ESA PILOT program, aiming at the development of a Visual Navigation and HDA system for Moon landing applications. The main objective of the HDA function is to detect surface slope, roughness and shadow hazards on the surface during the descent based on Lidar and camera measurements and to designate the coordinates of the safest landing sites. The HDA software is composed of functions dedicated to command the Lidar, process the raw Lidar measurements, perform motion compensation, reconstruct the surface topography, process the raw camera images, ground- reference the resulting Lidar and camera images, compute hazard maps and then designate the safe landing site coordinates on the surface. Two of the most important challenges for the HDA function are to 1) reconstruct the terrain topography from Lidar scans despite Lander translational and attitude motion and 2) process Lidar and camera information to provide a safe site recommendation to the Lander platform within the allocated time. In the context of the Phase B of the PILOT program, the performance of the HDA function of the PILOT system has been validated in a high-fidelity closed-loop software simulation environment and in a HWIL dynamic laboratory environment with a PILOT Lidar development model. Dynamic HWIL testing has successfully demonstrated one of the critical functions, which is the capability of the system to compensate for the Lander motion during the scanning phases. This requires real-time adaptation of the Lidar scan pattern based on the estimated states to maintain the desired resolution and coverage, and real-time ground-referencing of the measurements to reconstruct the terrain topography.

The optical absolute navigation is a crater-based detection and matching technique providing orbiter or Lander position measurements by matching craters seen in the camera image with a reference crater database built prior to landing. This fully-autonomous technique processes camera images during the descent, extracts craters, and matches the detected craters with a pre-stored reference crater map. The reference map is built prior to landing based on a Digital Elevation Model (DEM) of the Moon surface or orbital imagery. The integration of this function with optical relative navigation and a navigation filter has also been demonstrated in real time in the HWIL demonstration environment with scaled landing trajectories. These results demonstrate the feasibility of estimating the Lander position using crater-based navigation and the Lander velocity from optical relative navigation concurrently with a single navigation filter, even with a large uncertainty on the initial conditions.

Guidance, Navigation, and Control System for NASA’s Mars 2020 Mission Brugarolas P1 1NASA Jet Propulision Laboratory The paper will present an overview of the Guidance, Navigation and Control (GNC) system for the Mars 2020 mission. The Mars 2020 mission will launch in July-August 2020 and land at the Jezero carter (Mars) in February 18th of 2021.

The GNC system for the Mars 2020 mission is composed of three systems: - Cruise Attitude Control System - Entry Descent and Landing (EDL) GNC system - Surface GNC system

The Cruise Attitude Control System is the attitude control system for the cruise stage that will point the vehicle in its trip from Earth to Mars. The cruise stage is a spinner with a nominal spin rate of 2 RPM. It uses a star scanner and digital sun sensors for attitude determination and a propulsive RCS blowdown system (4.5 N) with 8 thrusters for attitude and maneuver control. This Cruise ACS architecture has a strong proven record. It was initially designed for the in 1997, then used for the Mars Exploration Rovers (Spirit and Opportunity) in 2003, and then enhanced with the use of the EDL IMU for the (MSL) in 2012. The Mars 2020 system is inherited from MSL with minor changes to improve performance or fix small problems. The main objectives of the cruise ACS system is to point the spacecraft during the 7-month cruise phase to Mars and to execute trajectory correction maneuvers. The main performance driver is the EDL attitude initialization knowledge requirement (0.15 deg, 3 sigma), which was tightened since MSL (0.25 deg, 3 sigma) to partially mitigate the MSL touchdown velocity anomaly. To achieve the improved performance, the cruise attitude determination algorithms were modified to optimize the processing of the sun sensor measurements.

The EDL GNC system is a 6 DOF control system that will land the vehicle at the Jezero Crater in Mars. This system will perform a guided entry, modulating the lift vector achieved via balance mass deployments through bank commands, to deliver the vehicle to the parachute deployment target conditions. During entry the closed loop system will use the IMU as a sensor and RCS thrusters (250N) for attitude control. Then, command the opening of a supersonic parachute using a new trigger, which by using downrange it allows to reduce the downrange errors dispersions and it is capable to reduce the landing ellipse to an approximate 8 km circle. This provides a reduction in landing ellipse area of about 40% compared to MSL. It will eject the heatshield and acquire the ground with its radar (6 beam system). While still on the parachute, it will perform terrain relative navigation to localize itself with respect to an onboard map using the Lander Vision System (LVS), and then perform the Safe Target Selection (STS) to find a safe landing target. The LVS system is composed of a camera and vision compute element, which between 4200-2200 m over the ground generates map relative localization measurements. STS collects and saves the best LVS measurement, and at 2200m over the ground will use the best LVS measurement to perform a search for the safest target within a reachable region from an onboard map of safe targets. The GNC system will then trigger the power flight phase, where it will perform a 6 DOF guided profile to the selected safe landing target using the IMU and the radar as its sensors and 8 Mars Lander Engines (throttleable up to 3200N) as its actuators. The GNC powered flight algorithms and software is inherited form MSL. It will land the rover using the skycrane maneuver. In addition to the changed mentioned above, two other changes were implemented to partially mitigate the MSL touchdown velocity anomaly, (i) a new local gravity correction term in the navigation filter and (ii) tighter filtering of the radar measurements during the skycrane phase to reduce the sensitivity of the navigation filter to dust/radar interactions. After landing the rover, the descent stage will perform a flyaway maneuver to dispose the descent stage at a safe distance.

The Surface GNC system will provide GNC capabilities to pilot the rover on Mars. In particular, it will provide (i) attitude determination from sun centroiding using the Navigation Cameras and the rover IMU, (ii) HGA pointing commands from inertial vector services, (iii) science mast pointing using vision closed loop in support of science acquisition using the navigation cameras or the RMI sen-sor and a 2-DOF guidance approach to mitigate backlash in the science mast pointing actuators, and (iv) autonomous rover driving using the navigation cameras, the vision compute element to perform stereo vision, and advanced path planning and execution algorithms.

This paper will describe these systems, their capabilities and innovations.

Acknowledgments: The research described in this article was carried out in part at the Jet Propulsion Laboratory, California Institute of Technology, under a contract with the National Aeronautics and Space Administration.

Copyright (2020). California Institute of Technology.

Control Challenges for NASA’s Mars Helicopter Grip H1 1Jet Propulsion Laboratory 1 Introduction In 2020, NASA will launch a drone-sized helicopter to Mars as part of its next rover mission. The helicopter will perform a series of experimental test flights to demonstrate the readiness of helicopter technology for Mars exploration. The Mars Helicopter is an unusual vehicle, having the ability to perform autonomous flights in an atmosphere with only 1% the density of Earth’s atmosphere, while surviving in a self-sufficient manner in the inhospitable environment of Mars. Among the many challenges in developing this vehicle was the design of a robust flight control system, capable of safely flying the vehicle without human intervention or external navigation aids. In this paper we will discuss how this was accomplished through extensive analysis, careful algorithm design, and a unique testing and validation program.

2 Control Challenges The Mars Helicopter project was conceived in 2013 as a potential add-on to a future Mars mission. Early development efforts focused heavily on the challenge of producing sufficient thrust in the low-density atmosphere of Mars, but analysis and initial testing quickly revealed challenges associated with flight dynamics and control of the vehicle. These challenges are due to a combination of properties inherent to helicopters, and properties specific to helicopters in the Mars environment, including: _ Relative Scaling of Forces: The low density on Mars, combined with lower gravity, fundamentally alters the relationships between aerodynamic, gravitional, and inertial forces on the vehicle, making it dynamically dissimilar [1] to both large- and small-scale helicopters on Earth. Of particular importance is the low proportion of aerodynamic to inertial forces in the rotor flap dynamics, which alters the helicopter’s low-frequency response to both control inputs and environmental disturbances, and introduces poorly damped high-frequency rotor-fuselage modes that complicate both the control problem and the vehicle design. _ Periodic dynamics: As for any two-bladed helicopter the dynamics is time-periodic; however, the effect is more significant for the Mars Helicopter, due to the extremely large rotor compared to the fuselage, which results in large periodic variations in vehicle inertia. _ Unstable dynamics: As is typical for helicopters, the dynamics is unstable in open loop, with increasingly fast instabilities in forward flight. This places fundamental restrictions on the ability to robustly stabilize the vehicle. _ Cross-Axis Coupling: Helicopters exhibit cross axis coupling, which varies across the flight envelope; thus, control robustness must be evaluated in a multiple-input multiple-output (MIMO) sense. Additional challenges of the Mars Helicopter include the difficulties of testing in relevant environments, and limited availability of information in flight – there are no global navigation aids, no significant magnetic fields, and no practical way to sense airspeed.

3 Topics of The Paper In this paper we will discuss the control-related challenges of the Mars Helicopter in detail, and present the approach taken to overcome these challenges, including: _ Modeling and Simulation: A custom simulation environment called HeliCAT was developed using the Darts/Dshell multibody engine developed at JPL. This environment was used for flight dynamics analysis, control design, and end-to-end V&V with flight software in the loop. This modeling revealed the fundamental differences in flight dynamics between Earth and Mars, which influenced both the design of the control system and the physical vehicle. _ System Identification: An extensive system identifiation program was developed to confirm or update crucial aspects of the dynamics model. The system identification took place in Mars atmospheric conditions inside large vacuum chamber, with the vehicle in various configurations, including fixed on force-torque sensor, on a movable arm, on a gimbal, and in front of a large array of fans to simulate forward flight. _ Analysis Toolchain: A toolchain was developed to synthesize models appropriate for control design based on the full multibody dynamcis model, including linearization, conversion of periodic dynamics into a linear time-invariant model, and model reduction. _ Robust Control Design: An iterative control design strategy was developed that consisted of approximate decoupling of channels via static mixing matrices, design of individual loops using single-input single-output (SISO) tools, and evaluation of robustness using MIMO disc margins. _ Flight Testing: Multiple flight tests were conducted in Mars atmospheric conditions within a vacuum chamber. In many of these tests, partial gravity offloading was necessary to account for the difference between Earth and Mars gravity. Nichols chart of the pitch control loop for the Mars Helicopter for two different flight regimes. The chart shows a tightening of the constraints on the gain margin due to increased frequency of the unstable eigenvalues in forward flight. Although the main focus of the paper is controls, guidance and navigation, as well as GNC-related hardware, will also be briefly discussed.

[1] B. Mettler. Identification Modeling and Characteristics of Miniature Rotorcraft. Kluwer Academic Publishers, 2003. doi: 10.1007/978-1-4757-3785-1.