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Im Rahmen von "Responsive Space" wird das Missionsbetriebssystem V3C (Verlegefähiges Compact Control Center) entwickelt. V3C ist auf handelsüblicher mobiler Hardware integriert, wird autark betrieben und kann schnell an verschiedene Orte verlegt werden.Wir zeigen Design, Implementierung und Bereitstellung von V3C, berichten über den erfolgreichen Demonstrationsbetrieb des BORIS-Satelliten und beleuchten verschiedene Betriebskonzepte, insbesondere die Koppelung mit einem GEO-Relais-Link.

To fulfill the objectives of deep space missions, such as in situ measurements at an outer planet’s moon or investigations at main belt asteroids, spacecraft must be provided with sufficient energy to get to these distant objects. This energy, expressed with the so-called ∆V -budget, which is the sum of required velocity changes along a spacecraft’s trajectory. As today’s and future deep space missions are infeasible using chemical propulsion alone, their trajectories involve one or more close flybys at mass-rich celestial bodies to gain additional orbit energy. These maneuvers are called gravity assists and depend on the relative positions of the assisting body and the respective target. Due to the orbital motion of both bodies, the required constellation may however repeat only every few decades. This constrains both trajectory and mission design and small launch windows can be the result. Any project delay hence threatens an on-time launch and thus potentially puts an entire mission at risk. Mitigation of that risk is possible through using low-thrust propulsion which can provide the required ∆V of a deep space mission without gravity assists. Contrary to chemical propulsion, having thrust values up to kilo-Newtons at specific impulse (Isp) values of 300-400 s, low-thrust propulsion currently offers only approximately one Newton at maximum. This thrust is achieved either through the ejection of carried-along particles, which are accelerated to very high velocities, or the reflection of sunlight photons. High exhaust gas velocities and very low propellant consumption make the respective Isp of low-thrust propulsion one magnitude higher than for chemical propulsion. The low-thrust propulsion concept of solar sailing even utilizes the solar radiation pressure for the generation of thrust, making it independent on any propellant. The different characteristics of low-thrust propulsion and chemical propulsion result in different trajectories. Therefore the methods for the optimization of trajectories of chemically propelled spacecraft are of limited use for the optimization of low-thrust trajectories. New methods were developed for this purpose, and one of them is Evolutionary Neurocontrol. This global optimization method combines the two biology-inspired mechanisms artificial neural networks and evolutionary algorithms. Called neurocontrollers, the artificial neural networks are used for spacecraft control. The optimization capability of evolutionary algorithms is used for the training of neurocontrollers. Contrary to other optimization methods, Evolutionary Neuroncontrol does not require an initial guess solution to work, which increases its usability for non-experts in optimal control and optimization. Evolutionary Neurocontrol was applied successfully in the past to various low-thrust transfer problems. Each of those problems, however, consisted of only one single heliocentric transfer from one celestial body to another. The problem of global optimization of multiphase low-thrust trajectories remained unsolved. This thesis describes how Evolutionary Neurocontrol can be extended to multiphase low-thrust transfers. An existing implementation was revised and complemented with new capabilities, concepts, and functionalities. Examples of the new features are a generic multiphase simulation framework, the support of nonheliocentric transfers, and third-body perturbation. The resulting method has been validated on various complex low-thrust transfer problems, which included two-phase transfers, like Earth-Moon-transfers, or heliocentric rendezvous missions with multiple targets or multiple propulsion technolgies. If available, the results were compared with published reference solutions. Finally, Evolutionary Neurocontrol was successfully applied to the design of a trajectory for a so-called Interstellar Heliopause Probe mission. Including a close flyby at Jupiter and using two different propulsion technologies, the resulting transfer brought the spacecraft to a heliocentric distance of 200 AU in less than 25 years.

Dieser Bericht gibt eine Zusammenfassung des PRISMA Missionsbetriebs durch das GSOC sowie eine Übersicht über die erzielten Ergebnisse. In Kapitel 3 werden zunächst die Zielsetzungen des Vorhabens sowie die Rahmenbedingungen beschrieben, unter denen das Projekt durchzuführen war. Anschließend gibt Kapitel 4 eine Übersicht über den Projektverlauf und beschreibt die wichtigsten Ereignisse. Kapitel 5 beschreibt die wichtigsten Ergebnisse des Vorhabens und schildert gemachte Erfahrungen.

The experimental satellite project PRISMA was initiated in 2005 by Sweden, France, Denmark, and Germany, with the Swedish Space Cooperation (SSC) as the project lead. The purpose was the demonstration of necessary techniques and the validation of the respective sensor technology for future missions that involve close formation flight and rendezvous in space. At that time, the German Aerospace Center DLR was not only involved in providing satellite GPS hardware and navigation software components but also as one of the experimenters for GPS-based navigation and autonomous formation flight. The idea of also conduction a part of the flight operations phase from Germany came into discussion at the end of 2009, with the purpose of sharing mission operations cost. This was agreed by Sweden and Germany shortly before launch of the two PRISMA satellites, which took place in June 2010. Nine months later, mission operations were handed over from SSC’s control center in Solna, Stockholm, to the German Space Operations Center (GSOC) in Oberpfaffenhofen, Germany. After successful operations by GSOC, the re-hand over of the mission back to Solna was performed in August 2011. The baseline concept for the German PRISMA ground segment foresaw cloning of the Swedish ground segment developed by SSC at GSOC to minimize the development and test effort, but specific adaptations were needed to integrate PRISMA into GSOC’s multimission environment. Furthermore, the original station network, which consisted only of the Kiruna ground station in North Sweden, was extended by two additional DLR ground stations in Weilheim, Germany, and in Inuvik, Canada. That extension proved especially beneficial to the shift concept. Another important aspect was the training of the German operations personnel in a short time. This was realized by training on the job concept, which kept the additional workload for teaching and training on acceptable levels and at the same time supported the Swedish flight operations team during their operations phase. This paper gives an overview of the GSOC ground segment and the flight operations activities. It reflects the challenges with regard to personnel and to the technical implementation of PRISMA flight operations at GSOC with limited available time. It also summarizes the lessons learned after five months of successful flight operations.

We investigate the interplanetary ight of a low-thrust space probe to the heliopause, located at a distance of about 200AU from the Sun. Our goal was to reach this distance within the 25 years postulated by ESA for such a mission (which is less ambitious than the 15-year goal set by NASA). Contrary to solar sail concepts and combinations of ballistic and electrically propelled ight legs, we have investigated whether the set ight time limit could also be kept with a combination of solar-electric propulsion and a second, RTG-powered upper stage. The used ion engine type was the RIT-22 for the �rst stage and the RIT-10 for the second stage. Trajectory optimization was carried out with the low-thrust optimization program InTrance, which implements the method of Evolutionary Neurocontrol, using Arti�cial Neural Networks for spacecraft steering and Evolutionary Algorithms to optimize the Neural Networks' parameter set. Based on a parameter space study, in which the number of thrust units, the unit's speci�c impulse, and the relative size of the solar power generator were varied, we have chosen one con�guration as reference. The transfer time of this reference con�guration was 29.6 years and the fastest one, which is technically more challenging, still required 28.3 years. As all ight times of this parameter study were longer than 25 years, we further shortened the transfer time by applying a launcher- provided hyperbolic excess energy up to 49 km^2/s^2. The resulting minimal ight time for the reference con�guration was then 27.8 years. The following, more precise optimization to a launch with the European Ariane 5 ECA rocket reduced the transfer time to 27.5 years. This is the fastest mission design of our study that is exible enough to allow a launch every year. The inclusion of a y-by at Jupiter �nally resulted in a ight time of 23.8 years, which is below the set transfer-time limit. However, compared to the 27.5-year transfer, this mission design has a signi�cantly reduced launch window and mission exibility if the escape direction is restricted to the heliosphere's "nose".

The Martian moons Deimos and Phobos are interesting targets for exploration missions, especially within the frame of a crewed Mars orbit mission. To minimize the risk to a crew and also to support EVA site selection, a robotic precursor mission should investigate both moons in advance. The focus of this study is on mission analysis of such a precursor mission that utilizes low-thrust propulsion, in particular Electric Propulsion, for the transfer to the Martian system. We assumed a launch by a Soyuz Fregat in 2018 and a direct injection into an escape trajectory with a hyperbolic excess velocity v∞ = 0 km/s. The spacecraft uses electric propulsion for the interplanetary transfer to Mars and also for spiraling down from an elliptic capture orbit to its destination. The mission analysis comprises dedicated missions to either Deimos or Phobos, and combined missions with Deimos as the primary target and, during a possible mission extension, Phobos as the secondary target. We used two different electric engine types for this study, which represent a wide range of specific impulse Isp. The employed thruster types were the Snecma PPS®1350-G with Isp = 1,650 s and the Astrium RIT-22 in two configurations having Isp = 3,704 s and Isp = 4,763 s. Within the analysis, we varied the number of engines and the available electrical power, followed by a down selection of a system design. In the second part of this study we investigated the implications of transfer time and thruster count on the mission itself caused by permanent degradation of the power and propulsion subsystems. Therefore we selected 15 state vectors of the nominal transfer trajectory and, starting from each of these 15 new initial states, optimized new minimum-duration transfers under the assumption of permanent engine failures or degraded solar cells.

On-Orbit Servicing (OOS) is not only an option for the repair and the upgrade of space assets that suffer from technical failures but might also be a promising business case, e.g. extending the lifetime of geostationary communication satellites. Another example of robotic service operations would be the servicing of low-earth orbit satellites. Currently, this expertise is being developed at GSOC in order to operate the Orbit Life Extension Vehicle (OLEV) and the German Orbital Servicing Mission (DEOS). The multinational commercial mission OLEV will achieve this task by docking a smaller servicer spacecraft to the apogee engine nozzle of a regular geostationary communication satellite and performing the station keeping task for a number of additional years in service and a de-orbiting at the end. The mission features a novel space-application of interesting technologies like high level autonomy and robotics. Contrary to the other mission phases, a different network architecture must be used because of the challenging telepresence (short latency) requirement. This requirement holds also true for the technology demonstration mission DEOS, as its robotic docking, berthing and de-berthing operations of the two spacecrafts are to be carried out by video-based telepresence. Due to the different orbit heights of both missions, the resulting requirements on the network differ significantly. Whereas in DEOS both spacecrafts (servicer and client) are controlled by one team at GSOC, OLEV is characterized by operating the satellites with very close coordination between GSOC and the client control center. This paper presents the special features and resulting distinctive challenges of these missions and how GSOC copes with them during the mission preparation, i.e. how they are reflected in the ground system design as well as the flight operations of the respective phases.