Analysis of interplanetary solar sail trajectories with attitude dynamics

We present a new approach to the problem of optimal control of solar sails for low-thrust trajectory optimization. The objective was to find the required control torque magnitudes in order to steer a solar sail in interplanetary space. A new steering strategy, controlling the solar sail with generic torques applied about the spacecraft body axes, is integrated into the existing low-thrust trajectory optimization software InTrance. This software combines artificial neural networks and evolutionary algorithms to find steering strategies close to the global optimum without an initial guess. Furthermore, we implement a three rotational degree-of-freedom rigid-body attitude dynamics model to represent the solar sail in space. Two interplanetary transfers to Mars and Neptune are chosen to represent typical future solar sail mission scenarios. The results found with the new steering strategy are compared to the existing reference trajectories without attitude dynamics. The resulting control torques required to accomplish the missions are investigated, as they pose the primary requirements to a real on-board attitude control system

V3C: Kontrollzentrum auf einem Laptop

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

Particle-based modeling of the unsteady flow in a high-test peroxide catalytic chamber

A particle-based mathematical model is proposed. It contains a one-dimensional approximation of the flow with heat transfer in the chamber wall and the catalytic material. It is designed for packed bed reactors with spherical granulated catalytic materials and is customized for high-test peroxide as fluid. The unsteady flow in the cold start phase can be analyzed. The temperature values increase over the time and the mass fraction of the chemical components are useable for a better understanding of the process inside and a precise design of new catalytic chambers with an improved functionality

The EnMAP imaging spectroscopy mission towards operations

EnMAP (Environmental Mapping and Analysis Program) is a high-resolution imaging spectroscopy remote sensing mission that was successfully launched on April 1st, 2022. Equipped with a prism-based dual-spectrometer, EnMAP performs observations in the spectral range between 418.2nm and 2445.5nm with 224 bands and a high radiometric and spectral accuracy and stability. EnMAP products, with a ground instantaneous field-of-view of 30m×30m at a swath width of 30km, allow for the qualitative and quantitative analysis of surface variables from frequently and consistently acquired observations on a global scale. This article presents the EnMAP mission and details the activities and results of the Launch and Early Orbit and Commissioning Phases until November 1st, 2022. The mission capabilities and expected performances for the operational Routine Phase are provided for existing and future EnMAP users

The EnMAP imaging spectroscopy mission towards operations

EnMAP (Environmental Mapping and Analysis Program) is a high-resolution imaging spectroscopy remote sensing mission that was successfully launched on April 1st, 2022. Equipped with a prism-based dual-spectrometer, EnMAP performs observations in the spectral range between 418.2 nm and 2445.5 nm with 224 bands and a high radiometric and spectral accuracy and stability. EnMAP products, with a ground instantaneous field-of-view of 30 m x 30 m at a swath width of 30 km, allow for the qualitative and quantitative analysis of surface variables from frequently and consistently acquired observations on a global scale. This article presents the EnMAP mission and details the activities and results of the Launch and Early Orbit and Commissioning Phases until November 1st, 2022. The mission capabilities and expected performances for the operational Routine Phase are provided for existing and future EnMAP users

Mission Analysis of Robotic Low Thrust Missions to the Martian Moons Deimos And Phobos

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 Mission Operations at GSOC

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

Preparation, Handover, and Conduction of PRISMA Mission Operations at GSOC

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

Flight Times to the Heliopause Using a Combination of Solar and Radioisotope Electric Propulsion

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"