The theory of attitude control for satellites is presented. The definition of “attitude” is followed by a description of the several disturbances and of the methods to determine the current status of rotational motion. An attitude prediction into the near future allows for active control, either in one or in three axes, either done autonomously on board or by commanding. The principles of attitude propagation and control are described, as well as the possible types of control mechanism. Comparisons between theory and practice are made and several examples are given from real missions.

In the launch and early orbit phase of a satellite mission, reliable information on the spacecraft’s position is of vital importance for the success of the whole mission. Especially for spacecraft targeting geostationary positions, ground-based orbit information is necessary for planning and calibrating the required manoeuvres to transfer the satellite in its dedicated target orbit. In frame of the ground telescope network SMARTnet, passive-optical observations have been employed during the launch and early operation phase of the geostationary mission EDRS-C. The present study assesses the benefit of employing passive-optical observations on the resulting orbit determination solutions as well as manoeuvre calibration.

The reference orbit implemented for the active TerraSAR-X mission works remarkably well for orbit control purposes, but an unexpected secular drift in the along-track separation between satellite and reference orbits has built up to a 60 s flight-time offset within 10 years of operation. The scope of this work is to understand the origin of the drift and to eliminate the effect for DLRs future repeat ground-track missions EnMAP and Tandem-L. The improved process of reference orbit generation is discussed and the underlying relations for the suggested inclination adjustment are derived. The improved process is successfully validated by means of 1-year numerical orbit control simulation. The presented process is generic and can be applied to any repeat-ground track mission.

DePhine – Deimos and Phobos Interior Explorer – is a mission proposed in the context of ESA’s Cosmic Vision program, for launch in 2030. The mission will explore the origin and the evolution of the two Martian satellites, by focusing on their interior structures and diversity, by addressing the following open questions: Are Phobos and Deimos true siblings, originating from the same source and sharing the same formation scenario? Are the satellites rubble piles or solid bodies? Do they possess hidden deposits of water ice in their interiors? The DePhine spacecraft will be inserted into Mars transfer and will initially enter a Deimos quasi-satellite orbit to carry out a comprehensive global mapping. The goal is to obtain physical parameters and remote sensing data for Deimos comparable to data expected to be available for Phobos at the time of the DePhine mission for comparative studies. As a highlight of the mission, close flybys will be performed at low velocities, which will increase data integration times, enhance the signal strength and data resolution. 10–20 flyby sequences, including polar passes, will result in a dense global grid of observation tracks. The spacecraft orbit will then be changed into a Phobos resonance orbit to carry out multiple close flybys and to perform similar remote sensing as for Deimos. The spacecraft will carry a suite of remote sensing instruments, including a camera system, a radio science experiment, a high-frequency radar, a magnetometer, and a Gamma Ray/Neutron Detector. A steerable antenna will allow simultaneous radio tracking and remote sensing observations (which is technically not possible for Mars Express). Additional instrumentation, e.g. a dust detector and a solar wind sensor, will address further science goals of the mission. If Ariane 6–2 and higher lift performance are available for launch (the baseline mission assumes a launch on a Soyuz Fregat), we expect to have greater spacecraft mobility and possibly added payloads.

The objectives of this document are to investigate the feasibility of HRWS formation flight, the derivation of a flight dynamics operations concept, a DV budget estimate, and a performance estimate for precise orbit determination and relative navigation.

High Resolution Wide Swath (HRWS) is an ambitious Synthetic Aperture Radar (SAR) mission proposed by Airbus Defence and Space, which will potentially exploit formation flight and the novel MirrorSAR concept to achieve unprecedented imaging characteristics by means of fractionated radar architecture, Ref. [1] and [2]. Originally planned as a follow-on of the extremely successful TerraSAR-X project, the HRWS mission and the satellite itself gradually took on a very different form, as the scientific goals grew more and more challenging. Now to be possibly equipped with electric propulsion and more than 3 times heavier than TerraSAR-X, the HRWS satellite is to be maintained within the control tube of merely 100 m (desired minimum) to maximum 250 m radius around the repeat ground-track reference orbit. On top of that, it is planned to augment the mission by 3 to 4 low-cost companion satellites flying in close proximity (down to ~100 m) to one another. Altogether, these challenges call for a very careful consideration of the absolute and relative orbit control concepts in order to develop a safe and precise maneuver strategy.

DePhine - Deimos and Phobos Interior Explorer - has been proposed to ESA as a medium-class mission by a scientific team led by the Institute of Planetary Research of DLR. Planned to begin its science mission in 2033, the spacecraft will fly in quasi-satellite orbits first around Deimos and then around Phobos while collecting remote-sensing data. Motivated by the DePhine proposal, this paper presents the results of a numerical analysis of some quasi-satellite orbits around the Martin moons. The orbits were studied in terms of their size and stability. At the same time, the feasibility of a global ground-track coverage was investigated.

This paper describes a GNC simulation as part of a feasibility analysis conducted by DLR/GSOC for future on-orbit servicing missions in near-geostationary orbit. The simulation addresses a far-range approach from several kilometers down to a few hundred meters, which includes relative orbit determination based on simulated optical measurements, and autonomous maneuver planning for relative trajectory control. One hundred simulation runs are performed with varying initial conditions based on the expected absolute orbit determination errors prior to the approach initiation. The safety of the formation is granted, as the servicer satellite never enters a pre-defined collision-avoidance area around the target spacecraft.The results of the simulation show that a low-cost far range approach based on optical measurements is feasible up to a safe transition to a mid-range sensor. The results of the paper as well as the proposed GNC algorithm itself can find applications in future on-orbit servicing missions in near-geostationary orbit.

A relative motion model for a satellite formation composed of two Earth-orbiting spacecraft located in the geostationary ring is developed taking into account major gravitational and non-gravitational forces. A previously existing model featuring perturbation due to \(J_2\) is enhanced by the perturbations due to solar radiation pressure arising from unequal area-to-mass ratios, as well as the secular and long-periodic gravitational perturbations due to the Sun and the Moon. The extended relative motion model is validated using several typical formation geometries against a reference generated by numerical integration of the absolute orbits of the two spacecraft. The results of this work can find application in future on-orbit servicing and formation flying missions in near-geostationary orbit.

On behalf of possible future missions with electric Propulsion (EP) controlled by DLRs German Space Operations Center (DLR/GSOC) the present operational multi-mission Flight Dynamics System (FDS) is enhanced to support the preparation and operations of such types of project. For designing an easily extendable framework, various low-thrust Scenarios were considered. Each low-thrust phase is modelled by a thrust profile comprising non-equidistant thrust vector and constant thrust level. Based on this design several multimission FD software modules are enhanced, e.g. Orbit Determination (OD) and generation of Orbit Related Information (ORI). The low-thrust transfer trajectories are optimized by means of the software package ASTOS/GESOP [1]. Demonstrating the extended FDS capabilities by means of a GEostationary Orbit (GEO) positioning reference mission shows an excellent consistency between resulting ephemerides by the optimizer in comparison to the FDS, validating the correct processing of thrust profiles within the implemented system.

A relative motion model for a Client and a Servicer spacecraft in near geostationary orbit is developed taking into account various gravitational and non-gravitational forces. In particular, differential perturbations due to Earth’s oblateness, solar radiation pressure and third body gravitational pull are studied, quantified and modeled. The relative motion model is validated using one typical far range approach test scenario against a reference generated by numerical propagation of two absolute orbits.

Any future space debris removal or on-orbit servicing mission faces the problem of the initial relative orbit determination of the servicing satellite to the non-cooperative target. In this work, we analyse the relative navigation accuracy that can be achieved in low Earth orbit, by using ground-based orbit determination from radar tracking measurements for the target, and classical GPS-based orbit determination for the servicing satellite. The analysis is based on the radar tracking measurements obtained from a 10 × 10 × 34 cm small object at an altitude of 635 km. The results show that the relative orbit can be determined with accuracy down to 2 m (RMS) in the semi-major axis, and down to 20 m (RMS) in both the radial and normal separations. From the results, we derive requirements on radar-tracking campaigns.

Mean orbital elements can be used for monitoring the satellite’s long-term behavior and for maneuver planning. In this paper, an analytical algorithm for conversion of osculating orbital elements into mean orbital elements is introduced and evaluated on several satellite missions. The accuracy of the mean orbital elements is estimated and the applications of the algorithm for determination of the relative satellite’s motion are discussed.