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The 7th International Conference on Astrodynamics Tools and Techniques (ICATT) is an event organized by the European Space Agency (ESA), the National Aeronautics and Space Administration (NASA), the Japan Aerospace Exploration Agency (JAXA), the Deutsches Zentrum für Luft und Raumfahrt (DLR), the Centre National d’Études Spatiales (CNES) of France, the Agenzia Spaziale Italiana (ASI), the Tsentralniy Aerogidrodinamicheskiy Institut (TsAGI) of Russia, the United Kingdom Space Agency (UKSA), and the Romanian Space Agency (ROSA).
ICATT aims at providing agencies, companies, organizations, universities and research institutes with a forum of excellence in the area of astrodynamics and space flight mechanics. Participants are invited to showcase their latest tools and techniques so as to promote the creation and exchange of ideas and the identification of new trends and required developments: challenges in the field of astrodynamics and orbital mechanics, current status of tools, their pros and cons, visions for the future, etc.
In addition to the latest theoretical advances in the field of astrodynamics, ICATT is especially devoted to astrodynamics software tools. Demonstrations and short tutorials are welcome. Furthermore, as in previous editions, ICATT offers a series of keynote lectures. These lectures are delivered by experts from specific astrodynamics fields.
The 7th edition of ICATT is dedicated to the exploration of the Moon and beyond.
Tutorial at the 7th ICATT
Tutorial at 7th ICATT
PTScientists is bringing down the cost of exploration and building reusable space infrastructures so more people benefit from access to space - because space belongs to everyone.
ALINA is designed to be compatible with all major commercial launch vehicles (including SpaceX’s Falcon 9, and ISRO’s PSLV), allowing invaluable flexibility when contracting a launch. This means that customers can choose their launch vehicle based on price and availability, rather than being restricted by type.
Once inserted into an Earth-orbit, ALINA will use its own propulsion to transport the craft 380,000 km to the Moon, and into lunar orbit. Payload racks on either side of the ALINA vehicle can be used to house orbital payload (if a lunar landing is not required), and ALINA can also dispatch CubeSats into lunar orbit.
From lunar orbit, ALINA will perform an accurate soft-landing on the Moon at a designated landing site. With each flight, ALINA can deliver up to 100 kg to the surface of the Moon. For Mission to the Moon this includes carrying two Audi lunar quattro rovers.
Break for lunch
Tutorial at 7th ICATT
optimization methods and related modelling related to spacecraft design
Opening speech of the 7th ICATT conference by Dr. J. Bals from DLR
Key note speech by ESA about the exploration of the Moon
Ascent trajectories for expendable and reusable launch vehicles; computation of payload injection and deployment; branching and abort trajectories; launcher separation and boosters come back; safe trajectories and splash down of rocket stages; ascent from planets and moons;
Future launch vehicle concepts and technologies for expendable and reusable launch vehicles are currently investigated by the DLR research project AKIRA, focusing on vertical takeoff and horizontal landing (VTHL), as well as horizontal takeoff and horizontal landing (HTHL) concepts.
Dedicated developments of multidisciplinary frameworks for launch vehicle modeling and preliminary design optimization have been presented in the relevant literature. Moreover, it is common that these activities are performed by several independent, discipline-specific tools. With such an approach, only a limited amount of interactions of the involved disciplines with the overall system dynamics can be accounted for.
This paper focuses on the multidisciplinary launch vehicle guidance, control, and dynamics modeling framework that has been developed at DLR-SR in support of the aforementioned reusable launch vehicle design activities taking into account highly interconnected disciplines (propulsion, aerodynamics, and GNC, amongst others) and changing environmental conditions. The modeling framework is based on the object-oriented, multi-disciplinary, and equation-based modeling language Modelica. In this paper, dedicated 3-DoF and 6-DoF models, covering the kinematics and dynamics formulation, environmental effects, aerodynamics, and propulsion models for system dynamics analyses, trajectory simulations and GNC design are presented.
In particular, we showcase the advantages of using nonlinear inverse models obtained automatically by Modelica. This method establishes a direct connection between 3-DoF and intermediate 6-DoF models considering trajectory optimization results provided by the DLR Trajectory Optimization Package `trajOpt'. With this approach, angular rates and resulting moments can be obtained by the intermediate 6-DoF model for subsequent controllability analyses. The benefits of our modeling framework are discussed in terms of future GNC design and trade-off studies.
This work presents the new set of equations of motion incorporated in ASTOS release 9.3.
The new implemented feature, based on DCAP multibody software, provides the building blocks to simulate a complete launcher scenario considering vehicle flexibility, sloshing effects, stages separation, engine pressure oscillations and complex aerodynamic loads distribution.
The interaction between those phenomenon and the ascent flight control logic could turn the entire vehicle dynamics unstable. The new functionalities allow ASTOS software to simulate and predict such catastrophic scenarios.
The multibody equations of motion feature let DCAP compute the entire system dynamics while ASTOS provides all the external forces such as aerodynamics, gravity accelerations and actuator output.
Five major features are organically embedded in the new MBS capabilities:
This work details on each of the above mentioned new functionalities, showing user input and results taken from typical example scenarios.
Low Earth Orbits (LEO), Medium Earth Orbits (MEO), High Earth Orbits (HEO), Geostationary Orbits (GEO); station-keeping; optimization of loitering arcs; computation of drag-free orbits; circular and elliptical orbits around planets and Moons; resonant orbits and fuel-efficient trajectories;
Poincare is a modular trajectory design tool based on a catalog of three-body science orbits and a differential corrector to compute connecting transfer arcs between orbits in multi-body systems. Poincare attempts to offer a unified approach, i.e., an "all-in-one" integrated search within one interface and setup in MONTE (JPL's signature astrodynamic computing platform) The Science Orbit Design Tool -- first module --facilitates rapid and well-informed decisions regarding the selection of periodic orbits for a particular mission and enables the simultaneous study of various orbit alternatives. The second module -- the Reference Trajectory Design Tool -- allows the user to calculate optimal transfer paths from a departure orbit to a science orbit via dynamical systems structures, resulting in an end-to-end reference trajectory.
In the 1960's, the application of insight from the circular restricted three-body problem (CR3BP) moved into the `space age' when a mission to the Lagrange points was considered for NASA's Apollo program. Since then, many of the structures that emerge in the CR3BP have been more actively exploited in trajectory design. Consequently, successful missions to the vicinity of the Lagrange points have since been launched, including the International Sun-Earth Explorer-3 (ISEE-3), the Solar Heliospheric Observatory (SOHO), the Advanced Composition Explorer (ACE), and the Microwave Anisotropy Probe (MAP). Parallel to the development of these mission concepts, the possibility of applying dynamical systems techniques to the design of these types of trajectories was also being considered. In fact, in the 1960's, Conley had investigated low energy transfer orbits to the Moon using dynamical system techniques. In the 1990's, the use of invariant manifolds in the design process to construct pathways between the Earth and the Sun-Earth libration points was finally applied in an actual trajectory: the trajectory supporting the Genesis mission.
Human activity in the space has caused the growth of a very large population of resident space objects (RSO). More than 19,000 objects are currently catalogued by 18th SPCS (former JFSCC) with sizes starting around 10 centimetres in LEO and around 1 metre in GEO. Space debris has nowadays become a very important thread to space operations as high-risk collisions are predicted daily between operational spacecraft and space debris objects.
Most space agencies have their own programs to deal with this thread, both from a mitigation point of view (IADC guidelines implementation, spacecraft design, active space debris removal), and from an operations point of view (e.g., space surveillance and tracking, collision avoidance). On the other side, the space private sector has been developing and using its own solutions to tackle the problem.
One of the key aspects to implement such measures is the availability of a catalogue of RSOs, not only characterising the properties of the objects, but also providing precise ephemerides that allow the prediction of high-risk collision events accurate enough and time in advance.
To build-up and maintain a RSO catalogue, in addition to the required sensors (radars, telescopes, SLR stations), it becomes necessary to have the required ground-segment infrastructure able to process efficiently all the data provided, in form of observation tracks, from those sensors.
Since 2007 GMV has developed and used methods to identify, track and catalogue RSOs. The SST Catalogue Maintainer Software (catmai) is GMV’s software capable of maintaining a catalogue of man-made Earth orbiting objects and their orbital information through the processing of measurements from a pre-defined space surveillance network of sensors.
catmai is composed of an initial orbit determination tool, an orbit determination module, a track-to-orbit correlator, a track-to-track correlator, an orbit-to-orbit correlator and a catalogue post-processing component for the analysis of the cataloguing performances.
This paper will focus on the methods implemented in GMV’s cataloguing solution and their performances in terms of success rate and false positive detection for the following processes:
Initial orbit determination, to obtain the first estimation of the orbits from very few observations and with no a-priori information. A set of initial orbit determination methods are available for different number and type of measurements.
Orbit determination, to improve the first estimation of the orbits by considering all available data, via both sequential and batch least-squares approaches. These methods are used during the correlation processes as well as for updating catalogued orbits.
Track-to-orbit correlation, to correlate uncorrelated tracks (UCTs) with already catalogued objects for catalogue maintenance. Correlation is performed in the measurements domain, i.e. synthetic measurements are compared against real ones provided by the sensor network. The figure of merit considered is the distance between real and synthetic tracks (Euclidean, Mahalanobis or Bhattacharyya). The track-to-orbit correlation algorithm involves: Synthetic tracking generation, pre-filtering, Synchronisation, Residuals computation and Correlation statistics computation.
The performance of the track-to-orbit correlator has been evaluated on both optical and radar cataloguing maintenance scenarios. It is able to provide success rate around 99.5% (true positives), around 0.6% of false negatives while avoiding false positives.
Track-to-track correlation, to associate and correlate UCTs among them in order to identify new objects not previously catalogued (i.e. catalogue build-up).
The main concept behind this track-to-track association method is a multi-step filter that sequentially applies IOD and simple OD methods to all possible combinations of uncorrelated tracks from survey activities.
The performance of the track-to-track correlator has been analysed on radar cataloguing build-up scenarios, leading to success rates around 99% (true positives) and the false positive rates lower than 0.1%, while keeping a high track usage rate (close to 99%).
Orbit-to-orbit correlation, to correlate objects of the catalogue with those from external catalogues such as Space-Track’s public catalogue. The correlation information of all objects of the involved catalogues is maintained from one analysis to another. This history of the correlation process is stored and used to ensure that two objects that used to correlate keep on correlating even if there is a manoeuvre not detected in one of the orbits through an outlier detection process.
The performance of the orbit-to-orbit correlator has been investigated by correlating a precise catalogue with the Two-Line Element Sets (TLE) catalogue for more than 15,000 objects. Results after one month of analysis reveal a success rate of more than 99.9% (true positives), around 0.08% of false negatives and only less than 0.01% of false positives.
The experience on these subjects gained by GMV with its own software solution for catalogue build-up and maintenance will also be described in the paper, using data from real sensors: more than 30 telescopes, radars and SLRs in five continents, covering all SST telescopes in Spain (OAM, TFRM, TJO, IAC, IAA), Airbus’s GEOtracker telescopes, SpaceInsight telescopes, AIUB telescopes in Switzerland, ESA’s OGS, Russian ISON telescopes network, and radars such as TIRA in Germany, Chilbolton in UK, ESA’s MSSR, and the Spanish Navy SLR station, among others.
As it’s known that our Sun rises always in the east and sets in the west, similarly, all the stars in the sky also rise in the east and set in the west. But day time, due to sun's brightness, we would not able to see the stars in the sky. Apart from this, there are many other parameters that restrict the observation these celestial objects. A few of those are : Time period of rising and setting of an object, latitude, Magnitude of the object, altitude of the location of observation and limitation of the telescope covering the sky in all angular directions. Then the question arises, can we really covering up the complete sky in all angular positions for observation with the availability of telescopes and also with respect to its situated positions for observations? There is always the possibility that we could miss some part of the sky. This aspect is the motivation that prompted to take this topic of research. This is just a questionnaire to start up research. Research has to be yet startup, with the small experiments plot. To carry out this research the sample is taken first for Chennai and then applied for all other available Indian telescopes. This work further can be extended to all other telescopes situated around the world with different geographic locations and plotting observational coverages over the night sky and looking for the list of possibility of unobserved part of the night sky with the sufficient exposure and resolution.
Tools using any open source license; use and perspectives; core repositories and code re-use for astrodynamics computations; free use of astrodynamics code; code repositories; astrodynamics applications and astrodynamics code running on smartphones and tablets;
For numerous years, CNES Flight Dynamics teams have made freely available some astrodynamics tools and libraries as MSLIB library. Nevertheless, these tools, essentially coded in Fortran language needed different versions of compilation depending on used platforms (Solaris, Linux, Windows …) which didn’t ease its installation and therefore limit their dissemination. Some years ago, CNES astrodynamics subdirectorate made the decision to switch to Java language in particular to insure portability whatever the target machine was. As a consequence old generation astrodynamics tools were translated to Java and improved on the process. Moreover, and as a consequence of the new language, these new tools (or new versions of tools) became more easily exportable keeping them available as freely available tools and libraries. The translation and improvement effort includes both low-level libraries as PATRIUS or GENIUS and more sophisticated tools with their own Graphic User Interface (GUI) as PSIMU.
This paper will describe these different tools and libraries always linked to Flight Dynamics applications, their interaction and dependency as well as their dissemination mode (open source, freeware).
Initially we will describe low-level libraries as PATRIUS uniquely devoted to Flight dynamics aspects and GENIUS for scientific GUI development. Secondly we will also present GENOPUS library which is based on both previous ones and allows providing “intelligent” widgets as the one used for defining orbit parameters.
Then, we will present some tools based on these building blocks as PSIMU (for any kind of trajectory extrapolation around Earth) or MIPELEC (optimization of low thrust propulsion). We will also give as example, tools used in operational contexts as ELECTRA.
To finish, means to get and use these tools will be described via the CNES Web site, their licenses, Wikis (including tutorials and Javadoc) or even training course.
Space objects catalog maintenance demands an accurate and fast orbit determination (OD) process to cope with the ever increasing number of observed space objects. The development of new methods, that answer the two previous problems, becomes essential.
Presented as an alternative to numerical and analytical methods, the Draper Semi-analytical Satellite Theory (DSST) is an orbit propagator based on a semi-analytical theory allowing to preserve the accuracy of a numerical method while providing the speed of an analytical method. This propagator allows computing the mean elements and the short-period effects separately. We reproduced this architecture at the OD process level in order to be able to return, as desired, the mean elements or the osculating elements. Two major cases of use are thus possible: fast OD for big space objects catalog maintenance and mean elements OD for station keeping issues.
This paper presents the different steps of development of the DSST OD included in the Orekit open-source library. Integrating an orbit propagator into an OD process can be a difficult procedure. Computing and validating derivatives is a critical step, especially with the DSST whose equations are very complex. To cope with this constraint, we used the automatic differentiation (AD) technique. AD has been developed as a mathematical tool to avoid the calculations of the derivatives of long equations. This is equivalent to calculating the derivatives by applying chain rule without expressing the analytical formulas. Thus, AD allows a simpler computation of the derivatives and a simpler validation. AD is also used in Orekit for the propagation of the uncertainties using the Taylor algebra.
Existing OD applications based on semi-analytical theories calculate only the derivatives of the mean elements. However, for higher accuracy or if the force models require further development, adding short-period derivatives improves the results. Therefore, our study implemented the full contribution of the short-period derivatives, for all the force models, in the OD process. Nevertheless, it is still possible to choose between using the mean elements or the osculating elements derivatives for the OD.
This paper will present how the Jacobians of the mean rates are calculated by AD into the DSST-specific force models. It will also present the computation of the state transition matrices during propagation. Both mean elements and short-periodic derivatives are developed. The performance of the DSST OD is demonstrated under Lageos2 and GPS orbit determination conditions.
Keywords: Orbit determination, Automatic differentiation, Semi-analytical theory, DSST, Open-source, Orekit.
Low Earth Orbits (LEO), Medium Earth Orbits (MEO), High Earth Orbits (HEO), Geostationary Orbits (GEO); station-keeping; optimization of loitering arcs; computation of drag-free orbits; circular and elliptical orbits around planets and Moons; resonant orbits and fuel-efficient trajectories;
Tools using any open source license; use and perspectives; core repositories and code re-use for astrodynamics computations; free use of astrodynamics code; code repositories; astrodynamics applications and astrodynamics code running on smartphones and tablets;
Visit to DLR facuilities
Interplanetary trajectories and fly-by; rendezvous with asteroids and comets; libration point transfers and orbits; resonant orbits; near-Earth objects trajectories; sample return missions; trajectories involving 3rd-body perturbations; coverage of instruments and ground contacts; maintenance or orbital positions around planetary bodies; planetary tours and encounters;
In this paper, a mathematical model was developed to design and optimize interplanetary trajectories that include gravity assist. The method of patched conics and a solver of the Lambert problem transfers are used to cast the space trajectory design process as an optimization problem, subsequently solved by using MATLAB. This model has been tested to provide an overview of the processes involved in the interplanetary trajectory design and analysis of the Juno mission to Jupiter. The resulting mission analysis compared to the actual data published.
Preliminary design of artificial satellite missions commonly relays on the use of simplified models that comprise the bulk of the dynamics. In the case of the gravitational potential, the amplitude of long-term oscillations of the orbital parameters is roughly one order of magnitude larger than the short-period oscillations. Because of that, dealing with just the few more relevant zonal harmonics of the potential is generally suitable for the initial steps of the procedure. In addition, the long-term evolution of the orbital parameters is customarily investigated through averaging procedures that remove the higher frequencies of the motion, in this way notably speeding the process of mission design.
However, there are cases in which the use of simplified models is not an option and full zonal potential models must be used instead. The paradigm is provided by the moon, where, due to the irregular character of the moon gravity field, mission designing of low altitude lunar orbits needs to deal with tens of, contrary to just a few, zonal harmonics. The analytical approach is still possible, but the requirement of handling huge expressions formally usually discourages mission planners, who then resort to numerical procedures. Still, useful compact recursions for dealing analytically with the problem exist in the literature since many years ago, yet limited to the equations of the averaged flow.
On the other hand, the correct computation of initial conditions requires the mathematical transformation from mean to osculating elements, and vice-versa, which may be crucial for the design of missions dealing with unstable orbital configurations, which is the common case of science orbits under third-body perturbations. Based on Kaula's popular recursions, we derive new formulas for the efficient mean to osculating transformation. While some efforts in providing these transformation equations have been made by different authors in the past, we show that the performance of the new formulas clearly surpass existing proposals in the literature. The new Kaula-type recursions, together with Kaula's original recursions for the averaged zonal potential, provide a compact and efficient way of handling analytical solutions of full potential models. The use of this kind of solution is illustrated with application to the design of low lunar orbits.
It is possible to design heliocentric transfers to Mars culminating in ballistic capture and, with an impulsive-thrust strategy, these have already been studied, but were found to be less fuel-efficient and longer-lasting than Hohmann transfers. The objective of the present work is to investigate the characteristics of Earth-Mars low-thrust transfers to ballistic capture.
Small spacecraft are very mass- and power-constrained, so orbit transfers are challenging for them, especially to interplanetary destinations. To try and shift this paradigm, the study was carried out assuming the spacecraft to be a 16-unit CubeSat. In addition, to improve the validity of the results, ballistic capture was designed using a model that included many perturbing forces, namely third-body perturbations, solar radiation pressure and non-spherical gravity.
Some capture orbits were selected, each with a different arrival date at Mars, and targeted from Earth, on multiple departure dates. It was found that if the spacecraft is given enough time, the low-thrust strategy requires roughly the same fuel regardless of Earth departure or Mars arrival dates. In addition, terminating a low-thrust transfer to Mars in ballistic capture does not carry additional costs, when compared to simply rendezvousing with the planet. With the assumed spacecraft and departure conditions, only around 5 kg of propellant are required to reach Mars and get ballistically captured. Nevertheless, the spacecraft needs to fly for at least 3.5 years, which can be too long for a CubeSat.
Astrodynamics for re-fuelling, payload exchange, and tugging; rendezvous approach (far and close); computation of optimal phasing conditions; docking and mating; contact dynamics; berthing; docking in R-bar and V-bar; optimal ground contacts during rendezvous;
Safe de-orbiting and increasing the life-time of satellites by on-orbit servicing (OOS) will be of high importance in future spaceflight. The rendezvous and docking/berthing (RvD/B) phase is one of the most complex and critical parts of on-orbit servicing and debris removal missions. Several missions and developments have been started like the Restore-L mission of NASA, the RSGS (Robotic Servicing of Geosynchronous Satellites) program of DARPA, the Mission Extension Vehicle (MEV) of Orbital ATK and the ESA Clean Space Initiative (e.Deorbit). Robotic servicing will be of importance also in human spaceflight since rendezvous and docking technology generally plays a major role in all assembly, service and maintenance tasks.
All these missions require robust and reliable guidance, navigation and control (GNC) systems for rendezvous and robotic systems for berthing and maintenance tasks. In a recently started project called RICADOS (= Rendezvous, Inspection, CApturing and Detumbling by Orbital Servicing) the German Aerospace Center (DLR) develops a new on-board inspection, rendezvous and robotic system as well as a ground segment for on-orbit servicing including telepresence capability.
The paper presents the current status of the project and the end-to-end testing environment: The space segment is simulated using two robotic hardware-in-the-loop test facilities at the German Aerospace Center: the European Proximity Operations Simulator (EPOS 2.0) at DLR-German Space Operations Center, where the inspection and rendezvous is tested and demonstrated, and the OOS-Simulator (OOS-Sim) at DLR-Robotics and Mechatronics Center, where the capturing and detumbling are performed. The robots’ motion is generated by a numerical satellite simulator in software based on orbit and attitude dynamics for service and target satellite, simulation of actuators and of the satellites’ environment. The communication path from space to ground and vice-versa is simulated such that different scenarios can be tested: Different channel parameters such as telemetry and tele-command data loss, jitter and delay can be chosen for realistic tests. The ground segment is established as for a real on-orbit servicing mission with dedicated consoles (standard satellite console, rendezvous console and robotic console). In a multi-mission control room, which is used for real missions at the same time, the operators can train and collect experience how to run a real on-orbit servicing mission including the robotic capture via telepresence.
The paper also presents a first reference scenario based on DLR’s satellite strategy: DLR will launch several compact satellites like Eu:CROPIS in Low Earth Orbit in the next years. The reference scenario of RICADOS foresees a service satellite with rendezvous and berthing payload which is able to perform service tasks for a fleet of compact satellites in neighboring orbits. This paper describes the entire RICADOS-concept for the selected reference scenario, from inspection and rendezvous towards final capturing and detumbling of the target satellite, the ground contact concept for the mission, and presents results of the latest simulations and tests.
In the context of future Human Spaceflight exploration missions, Rendezvous and Docking (RVD) operational activities are mandatory and critical for the assembly and maintenance of cislunar structures [1]. The Orion spacecraft is expected to handle cargo delivery and crew exchange missions that will all require RVD. Despite extensive experience in Low Earth and Low Lunar Orbits, no operational RVD has yet been performed in the vicinity of the Lagrangian points.
The scope of this research is to investigate the specifics of orbits of interest for RVD in the cislunar realm and to propose innovative strategies and trajectory designs to safely perform these kinds of operations. With a focus on Near Rectilinear Halo Orbits (NRHO), previous work has investigated close rendezvous relative dynamics using linear and non-linear targeting algorithms [2]. Current research focuses on far rendezvous approaches and the investigation of passively safe drift trajectories in the ephemeris model. The goal is to exhibit phasing orbit requirements, given a prescribed target orbit, that ensure safe free motion and natural approach of a spacecraft near the target while exhibiting low cost transfer capabilties.
Ephemeris representations of NRHOs were derived using time-varying multiple shooting and adaptive Long-horizon targeting algorithms, resulting in orbit maintenance budgets comparable to those found in the literature [3]. The choice of the Long Horizon parameter appeared to be closely related to the properties of the NRHO and the benefits of an adaptive parameterization of the algorithm was discussed. Simulations showed significant drift and overlapping properties for phasing and target orbits of interest. This motivated the research of safe free drift trajectories with NRHO-like motion in the Ephemeris model, using impact prediction strategies derived from debris avoidance analysis. Such strategies include a local-plane crossing algorithm to perform a quick wide search for potential overlapping sections, later refined by means of screening volumes with probability of collision computations using covariance analysis.
References :
[1] Stéphanie Lizy-Destrez, Rendezvous Optimization with an Inhabited Space Station at EML2, 25th International Symposium on Space Flight Dynamics, ISSFD, (2015).
[2] Stéphanie Lizy-Destrez, Bastien Le Bihanc, Sara Manglavitia and Antonino Campolo, Safety Analysis for Near Rectilinear Orbit Close Approach Rendezvous in the Cislunar Realm, 68th International Astronautical Congress, IAC, (2017).
[3] Jacob Williams, David E. Lee, Ryan J. Whitley, Kevin A. Bokelmann, Diane C. Davis and Chistopher F. Berry, Targeting Cislunar Near Rectilinear Halo Orbits for Human Space Exploration, AAS/AIAA Space Flight Mechanics Meeting (2017).
Extended abstract
The present international cooperation scenario for robotic and human space exploration is focusing on mission architectures that revolve around building and exploiting a crew-tended cis-lunar space station, known as Lunar Orbital Platform-Gateway. Candidate orbits for this vehicle are the near rectilinear halo orbits (NRHO). Therefore, the capability to inject in NRHO and perform rendezvous and docking or berthing maneuvers with a station in NRHO is key to many future exploration missions.
The aim of ROSSONERO (Rendezvous Operations Simulation Software on Near Rectilinear Orbit) is to provide a tool for preliminary design and assessment of rendezvous trajectory in NRHO and, more in general, in restricted three-body problem scenarios. The tool is developed in MATLAB\Simulink and dynamics simulation is based on the equations of relative motion proposed in 1. Unlike other sets of equations proposed in the past for relative motion in the restricted three-body problem, the equations describe the dynamics of chaser spacecraft with respect to a target in the local-vertical local-horizon (LVLH), a local frame centered on the target generally adopted for rendezvous analysis [2]. For completeness, in last version of the tool, the rotational dynamic has been added based on the equation proposed in [4].
ROSSONERO rendezvous mission description is based on the definition of a set of waypoints in the LVLH, that the chaser must reach during its approach to the target. Two types of maneuver can be used for transferring from a waypoint to the next one: impulsive or continuous thrust. Maneuvers computation is performed by integration of linear equations of relative motion derived in 1 and in [3]. In these References, two different sets of linear equations are presented, obtained by linearizing the exact relative dynamics and using two different assumptions for the primary bodies motion: elliptic and circular restricted three-body problem. ROSSONERO allows the user to choose between these two sets for maneuver computation at each transfer arc. Overall mission analysis is then performed by means of key performance indexes such as maneuver execution error and propellant consumption. Plots showing the trajectory in the LVLH frame and control evolution are provided as well. An example of the output provided by the tool is shown in Figure 1 and in Figure 2 (see attached document).
Acknowledgments
This work was partially supported by the European Space Agency under contract No. 000121575/17/NL/hh. The view expressed herein can in no way be taken to reflect the official opinion of the European Space Agency.
References
The tentative position of the next habitable space station could be a southern L2 Near Rectilinear Halo Orbit (NRHO) of the Earth Moon System. To bring crew and cargo to the station, a safe and efficient rendezvous methodology has to be established. However, a significant body of work remains to be done on the design of the rendezvous procedure between halo orbits. Given fixed start and end halo orbits, direct transfers (such as Lambert arcs) between the two can produce simple strategies with short transfer time at the cost of relatively high velocity increment Δv, measured in m/s [1].
When longer transfer time is allowed, as with some cargo, lower energy transfers, taking advantage the natural dynamics of the Earth-Moon system, can be used. The most commonly used topological objects are the stable and unstable manifold of a given orbit. These trajectories can lead a spacecraft very far from the original orbit at very low maneuver cost. One strategy is to insert into the unstable manifold of the initial orbit, making the spacecraft leave the starting orbit. After an optimized flight time T_F, a maneuver Δv is performed to insert into the stable manifold of the final orbit, and thus converging to the final orbit [2]. This procedure, however, requires the existence and explicit construction of physical intersections of the stable and unstable manifolds of the original orbits.
In this article, a method is presented to generate the manifolds along with an approximation of the set of their intersections by triangulating the surface and applying a modified Moeller’s method [3]. The most promising points are then further refined into true intersections and the lowest Δv is chosen. The article considers transfer between regular halo to halo, NRHOs to NRHOs and halo to NRHOs. Analysis shows that for several configurations the maneuvers cost can be sizably reduced as compared to fully optimized Lambert arcs.
References:
[1] S. K. Wang, et al. Dynamical Systems, the Three-Body Problem and Space Mission Design Caltech, (2000)
[2] S. Lizy-Destrez, Rendezvous Optimization with an Inhabited Space Station at EML2, 25th International Symposium on Space Flight Dynamics, ISSFD, (2015).
[3] T. Moeller. A Fast Triangle-Triangle Intersection Test, Stanford University, (1992)
The existence of families of solar-sail displaced libration point orbits in the Earth-Moon system has recently been demonstrated. These families originate from complementing the dynamics of the classical Earth-Moon circular restricted three-body problem with a solar-sail induced acceleration. The addition of this acceleration makes the problem non-autonomous, but by constraining the orbital period in a differential correction scheme, closed orbits can be found that are periodic with the Sun's synodic motion about the Earth-Moon system. These orbits can be catalogued into traditional orbit families such as solar-sail displaced Lyapunov, halo, and vertical Lyapunov orbits where different families can be generated for different solar-sail steering laws. Previous work has furthermore demonstrated the applicability of these orbits for high-latitude observation of the Earth and Moon. To not only demonstrate the existence and applicability of these orbits, but also their accessibility, this paper investigates the design of solar-sail transfers to a subset of solar-sail displaced libration point orbits in the Earth-Moon system.
Initial guesses for the transfers are generated using reverse time propagation of the dynamics starting from a grid of state-vectors along the targeted periodic orbits. The backwards propagated transfers are truncated at close approach to Earth. Furthermore, the control is provided through a locally optimal steering law that maximises the solar-sail acceleration component along the inertial velocity vector. These near-feasible initial guesses are subsequently transferred into a highly constrained 12${^{th}}$-order Gauss-Lobatto collocation scheme to improve their feasibility. Constraints are included that ensure linkage between the start of the transfer and commonly used Earth parking orbits, a minimum altitude with respect to the Earth and the Moon, as well as a realistic maximum rotation rate of the solar sail of 20 deg per day.
The paper provides sets of feasible trajectories for realistic- near-term solar-sail technology. In particular, transfers to a solar-sail displaced Lyapunov orbit at $L_1$ and a halo orbit at $L_2$ are provided as well as a two-spacecraft transfer to a constellation of solar-sail displaced vertical Lyapunov orbits at $L_2$. This constellation achieves continuous coverage of both the lunar South Pole and the center of the Aitken Basin, while maintaining an uninterrupted communication link with Earth. The Aitken Basin is of great scientific interest as it is believed to hold clues to the history of the Moon and allows access to the deeper layers of the lunar crust. The lunar South Pole is often mentioned as a potential location for a human outpost because it is an area of near-permanent sunlight, providing access to power, and water ice is most likely present in the continuously shaded areas of the crater interior.
For the two-spacecraft transfer to the constellation of vertical Lyapunov orbits at $L_2$, identical launch conditions for both spacecraft are sought for, such that the constellation can be initiated using a single launch by a Soyuz launch vehicle. The resulting transfers allow two 1160-kg spacecraft to be launched into standard highly elliptical Earth parking orbits from where the solar sail is deployed to transfer the spacecraft to their respective orbits at $L_2$. These transfers take 53.1 and 67.9 days to complete before the spacecraft enter their respective constellation orbits. These results prove the accessibility of solar-sail displaced libration point orbits in the Earth-Moon system, thereby reaffirming the potential of solar-sail technology to enable novel scientific missions in the Earth-Moon system.
Recent mission and system studies conducted for the European Space Agency have involved the design of transfers targeting Earth-trailing or Earth-leading heliocentric operational orbits, in a 1:1 resonance with Earth.
Airbus is currently leading two such studies on behalf of the European Space Agency: the Lagrange (Space Weather) mission targeting the Sun-Earth L5 Lagrange Point, and the LISA (Laser Interferometer Space Antenna) constellation of three satellites, selected as the third large-class mission of ESA’s Cosmic Vision Programme, and whose operational configuration consists in a heliocentric triangular cartwheel formation.
As no spacecraft has ever flown to these destinations, the presentation will focus on the mission analysis techniques that have been used to address this very special class of interplanetary missions, characterised by some unique features and challenges.
In particular, the available injection and transfer strategies will be reviewed and thoroughly traded against the mission objectives and constraints: these include classical direct ascent strategies, but also low-energy escape options via the Sun-Earth L1 or L2 points, as well as advanced strategies involving Earth and/or Lunar Gravity Assists. The benefits and challenges of transfers augmented with Solar Electric Propulsion will be highlighted. A special attention will also be brought to the implications of the selected Launch Vehicle, in particular if characterised by a limited range of available declinations (DLA), the resulting seasonal variation of the transfer problem and the impact of the launch windows definition. Finally, orbital perturbations on station and disposal strategies will be addressed.
While the presentation will focus on the astrodynamics techniques for the relatively generic problem of flying a spacecraft into heliocentric space, the application to the Lagrange and LISA missions will be presented as study cases.
This paper presents the analysis of a low-thrust rendezvous mission to a target non-Keplerian orbit of the circular restricted three body problem (CR3BP) in the Earth-Moon system. The dynamical characteristics of this system are revisited, and some non-Keplerian orbits (e.g., L1 halo orbits, NRO and L2 halo orbits) have been simulated to study their suitability for a rendezvous mission. Starting from analytical approximations, a shooting method has been used for the numerical
description of these orbits. Afterwards, the different monodromy matrices related to the integrated non-Keplerian orbits has been studied to analyse their stability and to describe stable and unstable manifolds of the orbit, i.e., ballistic trajectories that can be covered by a spacecraft without any propellant usage.
The design of a low-thrust rendezvous mission in a non-keplerian orbit is approached as an optimal control problem, in which the solution is the thrust magnitude and direction along the path. The propellant consumption has been set as the objective function to be minimized, and the trajectory is subjected to a set of constraints ranging from thrust limitations and time requirements specified for each mission. The problem to be tackled is a rendezvous mission to a specific target in a L2 Halo orbit. Hermite-Simpson collocation method has been used for the numerical description of the dynamical constraints of the system. Then, the problem has been solved numerically with IPOPT (Interior Point OPTimizer). An unpowered trajectory integrated from the problem initial conditions is used for the solver initialization. The rendezvous mission implementation and the interface with that solver have been developed in Matlab.
Three different cases of rendezvous at the target L2 Halo orbit have been studied: 1) rendezvous from a nearby halo manifold, 2) rendezvous from a close halo orbit and 3) phasing manoeuvre starting from the same halo as the target. All the cases have shown to be feasible for being performed with the use of an Electric Propulsion thruster. The final optimal trajectories for those cases are presented. In addition, it has been shown that propellant consumption can be greatly reduced if the stability properties of the targeted orbits are exploited.
The support from JSpOC to current collision avoidance operations is priceless. As a result, most satellite owners and operators have signed with USSTRATCOM an SSA Data Sharing agreement and have Orbital Data Requests in place in order to have access to JSpOC conjunction assessment and collision avoidance support services.
JSpOC issues CDM messages in case an upcoming conjunction is detected. These messages are the main source of information to satellite operators to proceed with their collision avoidance operations. Additionally, JSpOC supports these operations if the operator provides information on the planned station keeping and collision avoidance manoeuvres by running the screening process also against the operational ephemerides.
The main issues reported by operators in this process are the following:
In order to answer to these issues (and particularly the first one), GMV has reached a SSA Data Sharing Agreement with USSTRATCOM to access the Special Perturbations (SP) precise catalogue containing the US unclassified objects, to exploit it and generate derived products for the provision of commercial services. The SP catalogue is routinely downloaded from space-track site (currently, on a daily basis) and post-processed by GMV for the provision of its commercial collision avoidance services through its dedicated focusoc Operations Centre. Additionally, GMV has reached dedicated agreements with a large number of commercial SST data providers (radar and optical) to integrate their sensors with focusoc to collect data in case of high-interest events in order to perform orbit determination on target and chaser objects and refine the risk assessment in case of need.
GMV’s focusoc Operations Centre integrates tools for conjunction detection, collision risk assessment and collision avoidance (closeap) and for SST data processing and orbit determination (sstod) which use state-of-the-art algorithms, support parallel processing and have been used in real operations by a large number of customers, integrated in operational ground control systems. This ensures the efficiency and reliability of the solutions provided by focusoc.
To mitigate the second issue listed above, GMV has also implemented an automated process to extend the span of the downloaded SP catalogue to make it cover a longer period of time both in GEO and LEO as well as MEO and HEO/GTO. This process consists on the fitting of the available ephemerides (by means of orbit determination and parameters estimation) and their propagation with detailed dynamical models. As a result, upcoming conjunctions can be anticipated before receiving CDMs from JSpOC and the urgency of collision avoidance operations is drastically reduced by counting on additional time for decision-making processes and on the possibility to perform collision avoidance by means of adapting upcoming station keeping manoeuvres, particularly in GEO.
To mitigate the third issue listed above, GMV has automated all the processes for conjunction detection and collision risk assessment. One-vs-all assessments are performed against the operational orbit as soon as provided by the operator (including planned manoeuvres) and as soon as a new catalogue is available from JSpOC. As a result, the operator can make any post-manoeuvre analysis for all manoeuvres performed, not only those aimed at collision avoidance ensuring the safety of all their operations. No manual intervention by any man-in-the-loop is required and therefore no additional latencies are imposed to the spacecraft operations.
The focusoc system description and operational tools used, the nature of the services provided, the process to download, fit and propagate the SP catalogue, the performances achieved and the collision risk assessment and avoidance operations performed, including real cases, will be presented. The applicability to several orbital regimes (from LEO to GEO) and different mission phases (not only routine but also LEOP, re-location, EOL, etc.) will also be presented along with the feedback received by the satellite operators currently making use of the service.
See the pdf file attached.
The design of space missions is generally driven by severe requirements on the Delta-v budget. Navigation is also becoming more and more challenging, asking for the satisfaction of stringent conditions characterized by unprecedented accuracy. As a consequence, an increased complexity in the trajectory design is needed, ultimately leading to employing high-fidelity models already in the early stages of trajectory design.
Flying in highly nonlinear gravity fields allows exploiting unique features, such as libration point orbits, ballistic captures, and low-energy transfers. These features are achieved by exploiting the sensitivity in initial conditions of highly nonlinear environments, and open up new scenarios for spacecraft characterized by very limited thrust authority.
In this talk, the tools developed at Politecnico di Milano for high-fidelity trajectory design will be presented. These include ULTIMAT (Ultra Low Thrust Interplanetary Mission Analysis Tool) for design and feasibility assessment of limited control authority missions, GRATIS (GRAvity TIdal Slide) for the computation of ballistic capture orbits, DIRETTO (DIREct collocation Tool for Trajectory Optimization) and LT2O (Low-Thrust Trajectory Optimizer) for the direct and indirect optimization of space trajectories in multi-body models, respectively.
A renewed vision to send humans beyond Low Earth Orbit (LEO) has given rise to a whole range of studies proposing different destinations and operational orbits for a new crew-tended space station; referred here as Deep Space Gateway (DSG). Near Rectilinear Halo Orbits (NRHO) have been identified as one of the most promising destinations for a DSG, due to the combination of both dynamical properties and accessibility to potentially water-rich regions in the Moon.
The aim of this paper is then to revise the suitability of NRHOs as long-term destinations for this new space station. The NRHO family indeed appears as a continuation of the classical Halo orbits, and, as such, also allows continuous line of sight with the Earth. The paper revises the formal definition and identification of NRHOs, as in the CR3BP model. Dynamical substitutes of the NRHOs are also refined in the Bi-Circular Model (BCM) by means of a multiple shooting method. key features such as lunar south-pole coverage, station keeping requirements and accessibility of the orbit are then analysed.
In particular, nine different station keeping strategies were identified and implemented. These strategies correspond to variations of three different underlying schemes; the cancellation of unstable dynamical modes, the use of the multiple shooting method and, finally, a weighted single-criteria numerical optimization.
The accessibility to and from the NRHO is discussed by computing direct and invariant manifold lead transfers. Due to the dynamical characteristics of the L2 NRHO family, the optimal strategy to transfer from LEO orbit is to target an auxiliary halo orbit first, for a posterior sub-transfer to a member of the NRHO family. A minimum LEO-to-NHRO transfer cost of 3.68 km/s is achieved following the aforementioned transfer strategy. An annual station keeping budget of only 1 m/s would be expected considering 1 km and 1 cm/s 3σ navigation errors, as computed in a BCM framework.
Clean Space astrodynamics tools and techniques; nature friendly techniques; safe trajectories; disposal and recycling; sustainability; disposal of spacecraft; collision warning techniques and tools; debris population models; design for demise trajectories; prediction of debris fall out; footprints analysis; collision avoidance (risk computation, avoidance strategies, delta-v budget estimation); end of life disposal; tools for long-term environment; gravity models and atmospheric models; magnetic models; solar radiation and solar wind pressure models; perturbations; tools and techniques to model perturbations; shielding analysis tools and radiation analysis tools and techniques; meteoroid and space debris terrestrial environment databases;
As French National Space Agency, CNES is in charge of monitoring safety requirements for people and property related to space operation as defined in French Space Operation Act (FSOA). To evaluate these requirements, and in particular to be able to assess the compliance with safety threshold, CNES has developed its own tools, DEBRISK and ELECTRA. DEBRISK software computes the ablation of the satellite and its components all along the reentry trajectory. This object-oriented code combines models for aerodynamic, aerothermodynamic and heat transfer. It supplies a list of surviving objects with their physical on-ground characteristics. This list is then used as an input for the ELECTRA software. ELECTRA software is designed to estimate human casualty risk during launch and controlled or uncontrolled re-entry. Using Monte-Carlo simulations, dispersion of various parameters as for example characteristics of surviving fragments and population grids, Electra computes the probability of incurring at least one victim on ground, the expected value of the number of victims and the impact risk per country. Among the four computation modes of Electra, the RF mode (final re-entry) is dedicated to the computation of risk a few hours or days before the un-controlled reentry of a space object.
One particular case of use of ELECTRA and DEBRISK occurred in 2018, when the Chinese space station Tiangon-1 re-entered the atmosphere and felt in the South Pacific Ocean on April 2nd. With 8,5 tons, this uncontrolled reentry was of large interest, nevertheless not concerned by the French Space Operation Act. In the frame of this high risk reentry, CNES made use of its flight dynamics technical expertise, the observation means activated through national and international cooperation and a variety of tools, both operational and expertise tools, as ELECTRA and DEBRISK, to predict the reentry date and location as well as the on-ground casualty risk evolution inferred by this high risk reentry.
First, this paper presents how the Tiangon-1 space station was modelled and how the ablation phenomena, induced by the atmospheric reentry, was computed using DEBRISK tool. Second, this paper deals with the risk computation using ELECTRA tool via its final reentry (RF) mode. The orbital parameters used as inputs for these tools were provided by OCC division in CNES who’s in charge of operational orbit determination and to operationally estimate the reentry point and date of reentries considered as high risk. The evolution of the orbit and the reentry prediction is presented in parallel with the evolution of the risk assessment.
The European Union is now developing a federated SST system composed of existing sensors and operations centres in Europe through the EU SST Support Framework. Potential future architectures are also being evaluated for the development of new future sensors, including both radar and telescope sensors and both tracking and surveillance sensors. This brings the need to analyse the performances of different sensor network architectures and topologies.
Typically, the performances of those architectures are normally measured in terms of number of observable objects and number of catalogable objects. By definition, an object is considered observable if the sensor network can observe the object at least once and generate the corresponding track. Similarly, an object is considered catalogable if it can be maintained in the catalogue through the update of its orbital information upon the generation of tracks corresponding to the object during survey observation activities. In order to do so, new tracks need to be correlated to the right object. Hence, the catalogability of an object is directly related to the ability of the system to correlate tracks properly. This depends on the revisit time (i.e., frequency of observations) for a given object population and sensor network, and furthermore, it is also driven by the on-ground infrastructure ability to maintain the catalogue to predict the orbits of the objects, depending on the accuracy of the radar measurements and Space weather indicators predictions (solar flux and magnetic field activity, among others).
Many previous studies are based on a rule-of-thumb stating that an object is catalogable if its revisit time is less than 24 hours. However, this assumption is not properly justified. Additionally, apart from the catalogability of an object, another aspect to consider is the accuracy of the orbital information being estimated from the correlated observations. Again, concepts such as catalogued and well-catalogued are normally used based on the revisit time of the objects. In the frame of these studies, two types of analyses can be performed: based on coverage analysis and based on full cataloguing processes. The first are less time consuming but of lower accuracy as they are based on analysing the observability windows of the objects of the population and based on rules-of-thumb for the revisit time (e.g. 24 hours) to determine the percentage of the population that can be catalogued. The second are much more time consuming and provide more insight but are driven by specific implementations of correlation algorithms. Hence, it is normally preferred to use the former for preliminary design analyses.
This paper presents a new methodology suited for Low Earth Orbit and developed to determine through a coverage analysis the population of objects that can be catalogued by a given sensor network, as well as the expectable accuracy of the orbital information generated from observations of the sensor network. The number of catalogable objects is derived from the observable population, assuming that tracks can be correlated to the right object as long as the position uncertainty (i.e., covariance) of the objects do not overlap. The growth of the position covariance of each object is driven by two main uncertainties: the one of the initial estimation of the semi-major axis of the orbit and the one of the drag effect in the object. The former is related to the accuracy of the observations from the sensors (and also to the observation geometry) while the latter is related to the space environment knowledge (i.e. capability to model the atmosphere density). Depending on the object altitude, the drag effect may be dominant with respect to the initial uncertainty of the semi-major axis. Moreover, the longer the revisit time the more relevant the drag effect becomes. On the other hand, the more populated the orbits become, the more important the initial uncertainty in the semi-major axis becomes since the more objects there are, the sooner their position covariance overlap.
The model described is used to characterize the performances of a sensor network composed of a single survey radar and optimize some of its design parameters. The main free design parameters optimized in the analysis are the location of the radar (i.e., latitude) and the elevation of the radar field-of-view (FoV), while other design parameters are kept fixed: the azimuth of the radar FoV is kept pointing southwards at all times, and the size of the FoV. The power of the radar is also a parameter considered in the analysis in order to optimize the location and pointing elevation as a function of the radar power. It is important to note that the radar location and field of view constrain the orbit observability and revisit times. Depending on the location and field of view of the radar, the revisit times of the observable population vary and hence the number of catalogable objects.
Some of the conclusions justified in the paper and related to cataloguing performance that can be derived from the analysis are:
The higher the number of objects, the more relevant the accuracy of the sensor is.
For a given sensor location and accuracy, there is a saturation limit in the number of catalogued objects, even if more objects are observed.
Higher elevations are better in terms of number of catalogable objects, while lower elevations are preferred in terms of revisit time and track duration.
The rule-of-thumb of 1-day revisit time for catalogable objects is too crude as it does not take into account the number of objects of the observable population and the accuracy of the sensor.
One of the deliverables of the ReDSHIFT H2020 project will be a software tool available to the scientific community and the public via a web-based interface. The ReDSHIFT software is thought as a tool for spacecraft operators, space agencies and research institution to design the end-of-life of any Earth-based mission and to study the interaction with the space debris environment.
In this talk the general description of the tool will be given together with a detailed description of the modules currently in the more advanced development state: the disposal module, via impulsive manoeuvres (manoeuvre module) and solar and drag sails (solar dynamics module), the spacecraft population interaction module and the re-entry module.
Given the initial orbit of the spacecraft and the spacecraft characteristics in terms of its cross area and mass, the optimal options for end-of-life disposal are given and compared; namely end-of-life disposal via one or a sequence of manoeuvres, end-of-life disposal through the use of a solar/drag sail or end-of-life through a hybrid sail + manoeuvre approach. This module is based on a study of the natural orbit evolution in the low to medium and geostationary regions that was performed to identify long-term stable orbits or resonance conditions to be used as graveyard or re-entry trajectories. The optimal manoeuvre to reach such re-entry or graveyard conditions is calculated. Moreover, the re-entry can be enhanced through a sail. In this case, different strategies for sail attitude control were previously compared and selected.
The optimal disposal by this module is passed to the space environment module so that the effect of this disposal on the space debris environment is calculated. This is done based on precomputed long-term simulations of the whole space debris environment, under different scenarios, to be used for the computation of the collision risk for the spacecraft in the disposal phase.
In the case the disposal trajectory is a re-entry one, the condition of the orbit at 120 km are used to verify the demisability of the spacecraft. This is done, by default, using some predetermined spacecraft configuration but the external user can also load a preferred configuration.
All the modules are interfaced and linked, in a carefully triggered processing chain, through the openSF simulation framework properly configured and customised to adhere to the needs of the ReDSHIFT SW tool.
It is the aim of the ReDSHIFT tool to contribute in a proactive way to the mitigation of space debris problem via passive end-of-life mitigation.
The research leading to these results has received funding from the Horizon 2020 Program of the European Union’s Framework Programme for Research and Innovation (H2020-PROTEC-2015) under REA grant agreement n. [687500]- ReDSHIFT.
Coffee break
Computation of planetary re-entry trajectories; optimal guidance; skipped and bounced trajectories; aerocapture, aerobraking, and aerogravity assist manoeuvres; descent and landing trajectories; re-targeting guidance; hazard-avoidance trajectories;
The execution of precision landing missions applicable for future Mars missions (in particular, Sample Return Missions and Human Missions) is a major technical challenge that will require the adoption of a set of technologies that have not yet been demonstrated in flight.
In the scope of an ongoing ESA activity related to the development of advanced navigation techniques for pinpoint landing on the Moon and at Mars, an effort was directed towards the trajectory definition component of such missions. For the Mars case in particular, it is found that the enhanced control authority permitted by using a retropropulsive phase directly after the reentry phase (e.g. without the use of parachutes) is very likely to improve the ability to land precisely at a pre-defined site. This requires a sequence of design steps not normally associated with the concept of optimization (such as including specific guidance algorithms in the trajectory design loop, both during the entry and the retropropulsive phases), while simultaneously ensuring that there are specific time windows for the ground observations necessary to acquire sufficient knowledge to enable true pinpoint landing (defined as landing well within 100m of a selected site). The present work describes the mission design process which was followed in the scope of the ESA ANPLE activity for the purpose of demonstrating the feasibility of Mars pinpoint landing missions, from interplanetary transfer to touchdown, and considering the specific application to a NASA MER/NASA Phoenix/ESA Exomars EDM-class vehicle and assuming the availability of highly accurate navigation aids (including vision-aided absolute navigation means and the use of orbiting and surface radio beacons).
The Gala Dinner will take place on THU 8h November 2018 at the traditional Paulaner am Nockherberg famous restaurant in Munich. Buses will depart from DLR at 19:00 to transport the participants to the restaurant.
Computation of low thrust orbits in any given mission arc; optimal trajectories involving low thrust; low thrust transfer to LEO and GEO; comparison of electric and chemical propulsion trajectories; design and optimization of low thrust orbit transfers; low thrust station keeping; low thrust orbital transfers in the Two-Body problem;
The continuing development of solar-sail technology in combination with the rising interest in a mission to the Sun-Earth $L_5$ point for heliophysics and the search for Trojan asteroids, raises the question of using solar sailing as the primary propulsion method to enable such a mission. This paper therefore investigates a range of solar-sail transfers to the $L_5$ point, departing from different invariant objects in the neighbourhood of Earth: natural and solar-sail displaced equilibrium points, families of periodic orbits and their associated stable invariant manifolds. Also the arrival conditions are varied to be either natural or solar-sail displaced periodic orbits around the $L_5$ point. The transfers are obtained using a hybridisation of different trajectory design techniques. First, a multi-objective genetic algorithm is applied to obtain near-feasible initial guesses, which are transformed into feasible transfers using a differential correction method. Through a continuation on the fixed time of flight, the differential corrector is subsequently used to reduce the transfer time. As the differential corrector implements a stepwise constant control of the solar-sail attitude, a pseudospectral optimisation method is finally taken at hand to obtain a smooth, continuous control profile, to, if possible, further reduce the transfer time. This approach results in fast solar-sail transfers of 396 to 1194 days, depending on the departure and arrival configuration and the assumed solar-sail technology. These results can serve as preliminary design solutions for a mission to the Sun-Earth $L_5$ point.
Please find an extended abstract in the attached pdf-file
We consider the problem of optimal low-thrust spacecraft geostationary orbit (GEO) insertion from initial circular orbit with 800 km height and 51.6 degrees inclination. Minimal time for electric propulsion insertion of considered nuclear powered heavy spacecraft [1] is about 117 days. Significant amount of this time (~90 days) the spacecraft spends in regions with harsh space radiation environment inside the Van Allen radiation belts. To reduce an absorbed total ionizing dose (TID) for onboard electronic systems we propose a method of changing shape of the insertion trajectory and examine efficiency of this method. The main idea of the proposed method is to consider TID as a part of state variables set and to add the equation for TID change over the time to the equations of motion. Then if we add to the low-thrust optimal time GEO insertion problem [2] with the new set of state variables a condition of fixed TID at the end of transfer, one could obtain trajectories with lower final TID values. The obtained optimal control problem was solved using the maximum principle for one orbit time-averaged equations of motion. For numerical solution of corresponding boundary value problems we used the numerical continuation method with the predictor-corrector scheme. Numerical integration of considered ODEs was performed using the DOP853 integrator code.
The dose calculation for obtained trajectories we performed with CmdLineAe9Ap9 [3] software and Python scripts using AE8/AP8 MIN, AE8/AP8 MAX and AE9/AP9 electron and proton flux models of the Van Allen Radiation belts. In order to tackle discontinuities in right parts of ODEs we constructed a two-step spline approximation scheme of dose rate function. The first step is cubic smoothing spline approximation for noise level reduction of the CmdLineAe9Ap9 dose rate data. The second step consists of high-order (11th or higher) usual spline interpolation of the smoothing spline values. High-order order spline is needed to meet smoothness conditions of the right parts for Dormand-Prince ODE integration method.
We managed to reduce the final TID values by 25-38% (depending on the flux model) of final TID value on the minimal time GEO insertion trajectory. The duration of GEO insertion transfer is increased by no more than 7% of the minimal insertion duration. The additional required characteristic velocity for obtained trajectories is 320-560 m/s (depending on the flux model) with respect to the minimal time insertion trajectory.
References
[1] Legostaev V. P. et al., Prospects for and Efficiency in Application of Space Nuclear Power Plants and Nuclear Electrorocket Propulsion Systems, Space Engineering and Technology, 2013, No. 1., P. 4-15. (in Russian).
[2] Petukhov V. G. Optimization of Multi-Orbit Transfers Between Noncoplanar Elliptic Orbits, Cosmic Research, 2004, Vol. 42, No. 3, P. 250-268.
[3] https://www.vdl.afrl.af.mil/programs/ae9ap9/downloads.php
Trajectory design and optimization; new mathematical methods and principles; multi-objective optimization; local and global optimization techniques; stability; dynamic systems theory; dynamical models; space flight mechanics mathematical foundations;
The observation of celestial bodies other than Earth can without any doubt benefit from Earth observation satellites. This study, conducted for the latter, can easily be applied to Phobos for example, as part of the MMX (Martian Moons Explorer) project.
Earth observation satellites realize several acquisitions, which are linked together through an acquisition plan. The calculation time needed to construct such a plan is limited; therefore the method usually employed is the greedy algorithm which gives a solution quickly but not optimal. The temporal and kinematical constraints are numerous, making this a high combinatorial problem. The biggest time cost is the management of kinematics, especially checking if the acquisitions can be linked to each other, that is, calculating the minimum duration of the rallying sequence “attitude maneuver + acquisition”. In order to reduce the calculation time, the duration of each attitude maneuver and each acquisition can be calculated approximately, by adding an error margin to the result in order to ensure the feasibility of the acquisition plan. Therefore, a balance must be found between the calculation time an optimization method leaves to the kinematics and the margin associated to the needed approximation. Typically, an exact resolution method which leaves less calculation time for kinematics computation needs a rough approximation, and thus a big margin, which deoptimizes the calculated acquisition plan.
A simplified simulation model is used in order to quickly evaluate the quality of a plan constructed with a method (for example: greedy algorithm, branch and bound, stochastic greedy algorithm, genetic algorithm, simulated annealing, taboo search), as well as the needed calculation time and therefore the time left for kinematics computation. This model is based on a time windows representation.
The originality is to determine these elements according to different margins then determine which model for the calculation of kinematics, with the associated margin, gives the best results.
The results are different depending on the studied context; the goal is to find the method that offers the best compromise between quality of the plan and calculation time left to kinematics.
In order to solve non-linear multiobjective optimization problems, one usually solves multiple scalarized subproblems. This provides a discrete approximation of the Pareto front which gives useful information for the decision maker who, in praxis, has to select one single solution. If the desired solution is not part of the precomputed discrete approximation one needs to apply interpolation techniques.
This contribution shows a method which uses information from parametric sensitivity analysis of the scalarized subproblems in order to choose the stepsize between samples adaptively to obtain a better interpolation between precomputed solutions. The problems are solved with the NLP solver WORHP which provides sensitivity information in an efficient way by reusing the factorization of the KKT matrix of the last optimization iteration. We show the basic functionality of the presented method by applying it to several bi-objective optimization problems. The method can also be used for more than two objectives if one can identify the neighboring precomputed points which are then used for interpolation.
LOTNAV has been for 10 years the ESA reference tool in the design of finite-thrust and ballistic interplanetary spacecraft trajectories and the preliminary assessment of navigation and guidance issues on the computed trajectories. Within DEIMOS Space LOTNAV has recently undergone a considerable update effort to enhance its modularity and flexibility to extend its applicability to new navigation problems. The aim of this paper is then to present LOTNAV capabilities and its most important updated features (i.e. the inclusion of ephemeris data directly from SPICE, new launcher models, the GNSS signal in the navigation analysis, etc.).
The main high-level analyses covered by LOTNAV are:
LOTNAV provides answers to the above aspects in an efficient, modular and integrated fashion. In line with each of the previous functions, a software module was developed:
The Trajectory Reconstruction Module provides with a consistent trajectory definition meeting the finite-thrust mission goals proposed by the user, which will serve for the purposes of the rest of the modules. Then, to analyse the trajectory estimation process, the observables for the system are established and computed within the Measurements Generation Module. Those observables together with the trajectory definition allow carrying out a theoretical assessment by means of Covariance Analysis. Trajectory determination and guidance are dealt together by the Simulation Module, performing a Monte Carlo simulation over the full navigation process thus obtaining empirical statistics of the system knowledge and dispersion in presence of low-thrust guidance.
ELECTRO: a SW tool for the ELECtric propulsion TRajectory Optimization
Low-thrust orbit transfers are becoming increasingly attractive thanks to the mass savings they offer and the maturity of electric propulsion technology. For this reason, there is an interest in developing fast, but still reliable trajectory optimisation methods that can be applied in the preliminary phase of the design of a mission. The tool presented is based on the averaging of the equations of motion written in equinoctial elements over true longitude. The calculus of variations is used to identify the optimal control law. In particular, the indirect optimisation method used here is based on a sequential gradient-restoration algorithm. Perturbations such as zonal gravity harmonics are included as well as shadowing effects, which need to be modelled because electric propulsion is normally switched off during eclipses. For the eclipse detection algorithm, an analytical formulation for the extreme points of the eclipse is mandatory, since the entry to and exit from the eclipse set up the limits of integration for the averaging. Further, low-thrust trajectories require a continuous variation of the thrust direction and this has to be compatible with the capabilities of the attitude control system of the satellite. This constraint can be formulated in terms of maximum angular rate for the satellite axes, maximum angular momentum and/or maximum torque. A discussion on how to cope with these constraints in the optimisation method will be presented. The capabilities of the developed tool are illustrated with examples of transfers to Geostationary Earth orbit (GEO).
Keywords: trajectory optimization, low-thrust, electric propulsion, indirect methods
Traditionally, GTO to GEO transfers using chemical propulsion consist in optimising a rather limited number of manoeuvres. The compliance with the satellite platform and operational constraints are then guaranteed by the launch window design. On the contrary, low-thrust transfers of GEO satellites require very long thrust phases. The complex satellite platform and operational constraints induced by the use of EP and by long transfer durations need to be handled by a dedicated transfer optimisation software. This paper presents the development, validation and various uses of the in-house low-thrust transfer optimisation software, OPTELEC.
The development of OPTELEC was performed to serve a three-fold purpose:
(i) Studies and mission analysis: to be used for system definition, that is transfer ∆V budget and duration estimation, computation of satellite attitude guidance throughout the transfer and optimisation of a selection of injection orbit parameters. Minimum-time and minimum propellant mass low-thrust transfers are covered. Up to three levels of thrust are available (full thrust, reduced thrust and coast arc). The family of injection orbit is not limited to GTO but can typically span from LEO to SSTO.
(ii) Operations: to be used during the Launch and Early Operation Phase (LEOP). This includes the development of high fidelity environment and satellite modelling, along with the satellite platform and operational constraints to be taken into account throughout the optimisation process. Also, real time operations require a fast-running software, without the need of an optimisation technics expertise.
(iii) Versatility: to be developed to handle not only Electric Orbit Raising (EOR) transfers optimisations but also transfers with high thrust propulsion only - using the Liquid Apogee Engine (LAE) - and hybrid transfers including any combination of LAE burns, Reaction Control Thruster (RCT) burns and thrust with EP.
This paper shows how OPTELEC was designed in order to achieve these goals and the applications to a wide variety of missions.
Online optimization and trajectory planning are key aspects of autonomous deep space missions. Taking into account individual target criteria, such as time or energy optimality, any spacecraft maneuver can be traced back to a general problem definition of the form "move the spacecraft from its initial state to a desired final state, while considering a dynamic model and avoiding collisions". This corresponds to an optimal control problem (OCP) and can be solved using WORHP, the ESA solver for non-linear programming (NLP). More specifically, the OCP is transcribed into an NLP by time discretization and the corresponding optimal trajectory - including control variables - is calculated by sequential quadratic programming. In order to obtain a highly efficient solution algorithm, the naturally occurring sparsity of the Jacobian and the Hessian is exploited.
The effectiveness of this approach has already been demonstrated in several DLR projects, such as the deep space missions KaNaRiA and EnEx-CAUSE. In order to make such results immediately available for terrestrial applications, a transfer to current scientific questions is appropriate. Moreover, the transfer would provide a test platform and increase public acceptance. Conversely, the knowledge gained from terrestrial testing can help planning more detailed space missions.
In this work, the DLR project AO-Car for controlling an autonomous vehicle in road traffic is presented as such a transfer. The concept, originally developed in the context of KaNaRiA for trajectory planning and control, is successfully implemented on a research vehicle, a VW Passat. The vehicle is able to explore a parking area autonomously, to identify free parking spaces and to perform a parking maneuver. During the exploration, suddenly appearing objects are recognized. Depending on the scene, a collision avoidance trajectory is computed or an emergency stop is performed. The method presented is based on WORHP and offers a uniform framework for optimal driving maneuvers. It is highly flexible, as reaction speed and passenger comfort can be easily balanced by adaptive weighting of target criteria.
SQP methods for nonlinear programming rely on a quadratic programming solver for computing a search direction in each major iteration. From the start, the large scale NLP solver WORHP has been using the interior point method QPSOL within its SQP framework, which was developed specifically for WORHP. Experience from usage of WORHP in many areas and development of features like sensitivity analysis and feasibility refinement raised interest in a reworked, extended interface between WORHP and its QP solver. Furthermore, additional concepts like multiple centrality correctors seemed promising for improving the overall performance.
Hence, a revised QP solver was designed and implemented. Mehrotra's algorithm that was implemented in QPSOL was extended by Gondzio's multiple centrality correctors and weighting of corrector steps was added. Special care was taken to handle the very general NLP formulation of WORHP efficiently, yielding a very general problem formulation for standalone quadratic programming as well. A clear interface was implemented for retrieving sensitivity derivatives from the quadratic solver directly, allowing WORHP Zen and feasibility refinement procedures easy access to them.
The talk deals with these algorithmic and interface aspects for the development of the new solver within WORHP. Numerical results on the CUTEst test set for nonlinear programming are presented to show the performance improvements over the previous method.
MODHOC (Multi Objective Direct Hybrid Optimal Control) is a toolbox for the design, optimisation and trade off study of space systems and missions.
It solves general nonlinear multi phase optimal control problems, automatically computing a well spread set of optimal trade off solutions. In addition, it is able to handle discrete optimisation parameters.
In order to do so, MODHOC combines a direct transcription method based on finite elements, a global multi objective optimisation algorithm combining evolutionary heuristics and mathematical programming.
MODHOC has been applied to a variety of applications: from the optimisation of launch vehicles and their ascent, abort and re entry trajectories, to the design of the optimal deployment of constellations of satellites, to the design of multi target missions.
In this paper, the main elements of MODHOC are described and the application of the software in space and non space related sample problems is demonstrated.
Preliminary design of artificial satellite missions commonly relays on the use of simplified models that comprise the bulk of the dynamics. In the case of the gravitational potential, the amplitude of long-term oscillations of the orbital parameters is roughly one order of magnitude larger than the short-period oscillations. Because of that, dealing with just the few more relevant zonal harmonics of the potential is generally suitable for the initial steps of the procedure. In addition, the long-term evolution of the orbital parameters is customarily investigated through averaging procedures that remove the higher frequencies of the motion, in this way notably speeding the process of mission design.
However, there are cases in which the use of simplified models is not an option and full zonal potential models must be used instead. The paradigm is provided by the moon, where, due to the irregular character of the moon gravity field, mission designing of low altitude lunar orbits needs to deal with tens of, contrary to just a few, zonal harmonics. The analytical approach is still possible, but the requirement of handling huge expressions formally usually discourages mission planners, who then resort to numerical procedures. Still, useful compact recursions for dealing analytically with the problem exist in the literature since many years ago, yet limited to the equations of the averaged flow.
Based on Kaula’s popular work, we re-derive the long-term potential of the zonal problem in closed form and show that Kaula’s approach in orbital elements provides much more efficient formulas for the construction of the mean elements potential that recent alternative proposals in the literature. The necessity of having available efficient expressions for the long-term zonal potential, from which the evolution equations of the orbit are directly derived, is illustrated with application to the design of low lunar orbits.
Commercial and scientific satellites located in Geostationary Equatorial Orbit (GEO) that are not placed there by the launch vehicle are often injected in a parking orbit. They are transferred therefrom to GEO using their own on-board propulsion system. The classical strategy relies either on Chemical Propulsion (CP) or Electrical Propulsion (EP). The former guarantees very short transfer times, whereas the latter attains propellant savings at the cost of longer on-orbit delivery times. Intermediate performances may be obtained by allowing both propulsion subsystems to coexist on a Combined-Chemical-Electric (CCE) platform.
Therefore, the design and optimization of orbit raising missions has to be treated as a multi-objective optimization problem. The goal is to determine the set of optimal trajectories along with the optimal propulsion subsystem. For such purpose we present the optimization tool MOLTO-OR (Multi-Objective Low-Thrust Optimizer for Orbit Raising). It incorporates models for EP, CP and CCE platforms and realistic effects on the space environment, such as eclipse effects, Earth oblateness perturbations and solar-cell degradation due to passage though the Van-Allen radiation belts. Additionally, complex operational constraints such as slew-rate limitations, avoidance of the Geostationary ring or phasing to a certain orbital slot can be imposed. MOLTO-OR is based on a two-step sequential algorithm. In the first step, the low-thrust control law is derived from a Lyapunov function and the chemical maneuvers are regarded as instantaneous impulses. A heuristic algorithm computes the set of Quasi- Pareto-Optimal solutions trading off propellant mass consumed, time of flight and radiation damage. In the second step, candidate solutions are deemed as initial guesses to solve the Nonlinear Programming Problem resulting from direct transcription of the problem.
A full overview of the capabilities and features of MOLTO-OR will be given. The effec- tiveness of our methodology to generate not only rapid performance estimates for preliminary trade studies, but also accurate calculations for the detailed design, will be highlighted.
Low-thrust propulsion and gravity assists maneuvers are both well known to provide sig- nificant benefits in terms of required propellant mass for interplanetary trajectories. However, the propellant reduction achieved with low-thrust engines, when compared to their chemical counterparts, comes at the cost of a higher transfer time.
Therefore, the design and optimization of interplanetary missions has to be treated as a multi-objective optimization problem. The goal is to determine the set of optimal trajec- tories along with the optimal sequence of planetary flybys. For such purpose we present the optimization tool MOLTO-IT (Multi-Objective Low-Thrust Optimizer for Interplanetary Tra- jectories). It is based on a two-step sequential algorithm. In the first step, the trajectory is assumed to be a Generalized Logarithmic Spiral. A heuristic global search algorithm combined with nonlinear programming are in charge of optimizing the set of parameters defining the spirals, as well as the number, sequence and configuration of the gravity assists. As a result the set of Quasi-Pareto-Optimal solutions trading off propellant mass consumed and time of flight are obatined. In the second step, candidate solutions are regarded as initial guesses for a direct collocation method, where the problem is transcribed into a Nonlinear Programming Problem by discretization, considering the full dynamics and the complete set of constraints.
A full overview of the capabilities and features of MOLTO-IT will be given. In particular, the ability to obtain optimal flyby sequences without a-priory knowledge by the user. Addi- tionally, the effectiveness of our methodology to generate not only rapid performance estimates for preliminary trade studies, but also accurate calculations for the detailed design, will be highlighted.
Bi-lateral discussions between NASA and the European
Space Agency identified the orbiter element as a promising
European-led contribution to a future international Mars Sample
Return campaign. Airbus recently completed the Mars
Sample Return Architecture Assessment Study on behalf of
ESA, with the objective to identify and quantify candidate
mission architectures. The paper describes the mission analysis
that has been conducted to support preliminary system
design, launch mass estimation and mission timeline for the
architectures investigated. It includes the optimisation of interplanetary
transfers, Mars operations including aerobraking
and rendezvous, up to Earth re-entry conditions.
Propagators and integrators; position and velocity prediction; ephemeris computations; conjunctions; numerical integration methods applied to astrodynamics; precise orbit determination for LEO, MEO, HEO, GEO missions; tools and technique for high precision orbit determination for planetary missions; observational data; parameter estimation; orbit determination with multiple tracking techniques;
An algorithm for the computation of distant retrograde orbits is presented. It is based on the computation of an approximate analytical solution of the restricted three body problem in the Hill problem approximation that provides accurate estimations of two basic design parameters. Notably, these parameters can be used for the computation of initial conditions of orbits that are periodic on average, and almost periodic in the original problem. Following application of iterative differential corrections results in the initial conditions and period of a true periodic orbit with the characteristics fixed by the design parameters.
The analytical solution consist roughly of a drifting ellipse, whose guiding center moves around the primary with long-period oscillations, and in which the linear growing of the phase of the satellite is modulated with long-period variations. The analytical solution splits in two parts of different nature. The first one provides the periodic corrections needed for converting osculating elements into the mean ones that describe the long-term evolution of the dynamics. The second part of the analytical solution gives the time history of the mean elements in the form of five Lindstedt series, which are needed for describing:
- the time scale in which the Lindstedt series evolve (1 series)
- the time evolution of the guiding center of the reference ellipse (2 series)
- the linear frequency with which the satellite evolves, on average (1 series)
- the long-period modulation of the phase of the satellite (1 series)
The use of the algorithm is illustrated with different examples, ranging from the typical case of 1:1-resonant distant retrograde orbits, in which the satellite remains always far away enough from the primary, to the challenging case of higher order resonances in which the amplitude of the libration of the guiding center of the orbit can take the satellite much closer to the primary.
The history of the creation of the Draper Semi-analytical Satellite Theory (DSST) started at the Computer Sciences Corporation, with support from the NASA Goddard Space Flight Center (GSFC), in the early 1970s. Then, its development continued at the Draper Laboratory in the 1980s and 1990s. Since 2001, some enhancements to the DSST have been achieved by the technical staff at the Massachusetts Institute of Technology (MIT) Lincoln Laboratory. The original DSST Fortran-77 code exists both as part of the Goddard Trajectory Determination System (GTDS) suite and as the DSST Standalone program. In addition, it can also be executed remotely through the Astrodynamics-Web-Tools service at the University of La Rioja. More recently, DSST has been re-implemented in Java, in the Orekit flight dynamics library. The open-source concept applied to the space domain attempts to provide free access to the software tools needed to operate safely and efficiently in space to all spacecraft and space-system operators. In this work, we address one of the tasks from the list proposed in the paper entitled: ''Open Source Software Suite for Space Situational Awareness and Space Object Catalog Work'' (Cefola et al., 2010): the migration of the original Fortran-77 DSST code to C/C++. In the initial stage, the original design of DSST has been maintained as much as possible, with the aim of taking full advantage of the validation and verification process developed over DSST during the last decades. Then, the parallelization of the semi-analytical theory for multicore and GPU technologies, by using the commonly available parallel programming environments, will be analyzed. We will also consider the development of a parallel DSST orbit propagator based on the Picard-Chebyshev concept. Finally, we will take into account the complexities associated with applying the new DSST C/C++ version to orbits with arbitrary central bodies, including lunar, planetary, natural-satellite, and asteroid orbiters. Part of the validation and verification process of the DSST C/C++ version will be presented here. The resolution of the incompatibilities discovered during this process is an important advance for the following stages of this project.
The Eumetsat GEO fleet has successfully performed its orbital determination using ranging-only data from three different ground-based tracking stations. Data from two different ground stations is used for each of the satellites. For more than 15 years, these results have been used for manoeuvre planning, after manoeuvre calibration and collision risk assessment. Since June 2018, an effort has been undertaken to enhance the orbit determination procedures with optical observations, with the objective of assessing the suitability of using optical data to perform the operational orbit determination, and determining the suitability of routine orbit determination and manoeuvre calibration using mixed data from ranging stations and telescopes. The optical measurements are provided currently by the Deimos Sky Survey (DeSS) telescopes, using additional sensors (coordinated by Deimos Space) as a backup in case of adverse weather conditions or technical issues. The processing of the measurement data is being performed by two separate teams at Eumetsat and at Deimos. This paper describes the processing chain put in place at Deimos for performing the observations and the processing of measurement data, and summarises the findings related to the aforementioned objectives after several months of routine observations.
Each spacecraft in the fleet is observed with optical sensors at least twice every week (in two observation slots), with each slot spanning at least 15 minutes, and with a minimum separation of two hours between slots. A software processing chain based around several Deimos tools has been put in place to plan the observations (SHUX), perform the observations (ITOX), automatically process and resolve the images (TRAX), split tracks (TRACA), and finally perform the orbit determination (TRADE).
The orbit determination is automatically performed weekly with the TRADE by means of a Batch Least Squares approach with a two-week rolling window. In absence of manoeuvres, this allows determining the solar radiation pressure coefficient while maintaining consistency with the previously computed orbits. When a manoeuvre is scheduled, optical observations are taken as soon as possible after the manoeuvre itself. In this case, the paper shows that the orbit determination with optical information only provides results comparable with the nominal range-only orbits.
Finally, the TRADE tool is modified to process range-only measurements from ground-based ranging stations along with optical measurements. A comparison of the quality of the solutions obtained with different combinations of measurements from different sensors is provided.