Flight Dynamics Operations of the TanDEM-X Formation

Since end of 2010 the German TerraSAR-X and TanDEM-X satellites are routinely operated as the first configurable single-pass Synthetic Aperture Radar interferometer in space. The two 1340 kg satellites fly in a 514 km sun-synchronous orbit. In order to collect sufficient measurements for the generation of a global digital elevation model and to demonstrate new interferometric SAR techniques and applications, more than three years of formation flying are foreseen with flexible baselines ranging from 150 m to few kilometers. As a prerequisite for the close formation flight an extensive flight dynamics system was established at DLR/GSOC, which comprises of GPS-based absolute and relative navigation and impulsive orbit and formation control. Daily formation maintenance maneuvers are performed by TanDEM-X to counterbalance natural and artificial disturbances. The paper elaborates on the routine flight dynamics operations and its interactions with mission planning and ground-station network. The navigation and formation control concepts and the achieved control accuracy are briefly outlined. Furthermore, the paper addresses non-routine operations experienced during formation acquisition, frequent formation reconfiguration, formation maintenance problems and space debris collision avoidance, which is even more challenging than for single-satellite operations. In particular two close approaches of debris are presented, which were experienced in March 2011 and April 2012. Finally, a formation break-up procedure is discussed which could be executed in case of severe onboard failures

Collision and evaporation avoidance for spacecraft formation

<p>Formation flying is an extremely promising approach to space operations with the potential to enable new types of missions and providing substantial increase in the performance of future space science and Earth observation applications. To successfully validate formation flying however requires the development of specific technologies and methodologies, which are beyond current state-of-the art in a wide range of diverse fields such as metrology and spacecraft guidance, navigation and control. A number of missions are currently under different stages of development to implement some of these stringent requirements.</p> <p>The paper develops and compares collision avoidance algorithms, demonstrating them within a 6 degrees of freedom, multi-spacecraft environment. At first a number of different collision avoidance scenarios will be identified alongside the triggers that will cause the algorithms to be activated. Once activated the collision avoidance algorithm must ensure corrective action to avoid catastrophic consequences to the mission.</p&gt

Fractionated solar power satellite for regional coverage

This paper presents a preliminary analysis of a fractionated solar power satellite system for regional coverage. The fractionated system is composed of a cluster of satellites, in different possible configurations, that concurrently beam energy to the ground through medium power lasers. The paper presents an analysis of the possible orbit solutions that can be adopted to provide power during the night time to local users in different regions of the world. The system is intended to serve mobile stations or local stations that can be hardly accessed by normal power lines or are cut off during disasters. A preliminary system analysis shows that with a limited number of small size satellite local users can be provided with a few kWh of energy every day

Spacecraft rendezvous by differential drag under uncertainties

At low Earth orbits, differentials in the drag forces between spacecraft can be used for controlling their relative motion in the orbital plane. Current methods for determining the drag force may result in errors due to inaccuracies in the density models and drag coefficients. In this work, a methodology for relative maneuvering of spacecraft based on differential drag, accounting for uncertainties in the drag model, is proposed. A dynamical model composed of the mean semimajor axis and the argument of latitude is used for describing long-range maneuvers. For this model, a linear quadratic regulator is implemented, accounting for the uncertainties in the drag force. The actuation is the pitch angle of the satellites, considering saturation. The control scheme guarantees asymptotic stability of the system up to a certain magnitude of the state vector, which is determined by the uncertainties. Numerical simulations show that the method exhibits consistent robustness to accomplish the maneuvers, even in the presence of realistic modeling of density fields, drag coefficients, the corotation of the atmosphere, and zonal harmonics up to J(8)

Multi-Agent Orbit Design For Perception Enhancement Purpose

This paper develops a robust optimization based method to design orbits on which the sensory perception of the desired physical quantities are maximized. It also demonstrates how to incorporate various constraints imposed by many spacecraft missions such as collision avoidance, co-orbital configuration, altitude and frozen orbit constraints along with Sun-Synchronous orbit. The paper specifically investigates designing orbits for constrained visual sensor planning applications as the case study. For this purpose, the key elements to form an image in such vision systems are considered and effective factors are taken into account to define a metric for perception quality. The simulation results confirm the effectiveness of the proposed method for several scenarios on low and medium Earth orbits as well as a challenging Space-Based Space Surveillance program application.Comment: 12 pages, 18 figure

Extension of the sun-synchronous Orbit

Through careful consideration of the orbit perturbation force due to the oblate nature of the primary body a secular variation of the ascending node angle of a near-polar orbit can be induced without expulsion of propellant. Resultantly, the orbit perturbations can be used to maintain the orbit plane in, for example, a near-perpendicular (or at any other angle) alignment to the Sun-line throughout the full year of the primary body; such orbits are normally termed Sun-synchronous orbits [1, 2]. Sun-synchronous orbits about the Earth are typically near-circular Low-Earth Orbits (LEOs), with an altitude of less than 1500 km. It is normal to design a LEO such that the orbit period is synchronised with the rotation of the Earth‟s surface over a given period, such that a repeating ground-track is established. A repeating ground-track, together with the near-constant illumination conditions of the ground-track when observed from a Sun-synchronous orbit, enables repeat observations of a target over an extended period under similar illumination conditions [1, 2]. For this reason, Sun-synchronous orbits are extensively used by Earth Observation (EO) platforms, including currently the Environmental Satellite (ENVISAT), the second European Remote Sensing satellite (ERS-2) and many more. By definition, a given Sun-synchronous orbit is a finite resource similar to a geostationary orbit. A typical characterising parameter of a Sun-synchronous orbit is the Mean Local Solar Time (MLST) at descending node, with a value of 1030 hours typical. Note that ERS-1 and ERS-2 used a MLST at descending node of 1030 hours ± 5 minutes, while ENVISAT uses a 1000 hours ± 5 minutes MLST at descending node [3]. Following selection of the MLST at descending node and for a given desired repeat ground-track, the orbit period and hence the semi-major axis are fixed, thereafter assuming a circular orbit is desired it is found that only a single orbit inclination will enable a Sun-synchronous orbit [2]. As such, only a few spacecraft can populate a given repeat ground-track Sun-synchronous orbit without compromise, for example on the MLST at descending node. Indeed a notable feature of on-going studies by the ENVISAT Post launch Support Office is the desire to ensure sufficient propellant remains at end-of-mission for re-orbiting to a graveyard orbit to ensure the orbital slot is available for future missions [4]. An extension to the Sun-synchronous orbit is considered using an undefined, non-orientation constrained, low-thrust propulsion system. Initially the low-thrust propulsion system will be considered for the free selection of orbit inclination and altitude while maintaining the Sun-synchronous condition. Subsequently the maintenance of a given Sun-synchronous repeat-ground track will be considered, using the low-thrust propulsion system to enable the free selection of orbit altitude. An analytical expression will be developed to describe these extensions prior to then validating the analytical expressions within a numerical simulation of a spacecraft orbit. Finally, an analysis will be presented on transfer and injection trajectories to these orbits

Satellite formation flying control of the relative trajectory shape and size using lorentz forces

Propellantless control approaches for small satellite formation flying represent a special interest and an important advantage for space industry nowadays. A formation flying control algorithm using the Lorentz force for Low­Earth Orbits to achieve a trajectory with required shape and size is proposed in this dissertation. The Lorentz force is produced as the result of interaction between the Earth’s magnetic field and an electrically charged spacecraft. Achieving the required trajectories represents a challenge since the control is the variation of the satellite’s charge value. This control mechanism simplicity cannot provide full controllability. A Lyapunov­based control is developed for elimination of the initial relative drift after launch and it is aimed for reaching a required relative trajectory with predefined shape and size. The control algorithm is constructed to correct different parameters of the relative trajectory at different relative positions. The required amplitudes for close relative trajectories for in­plane and out­of­plane motion as well as the relative drift and shift of elliptical relative orbits are controllable using Lorentz force. Due to the absence of full controllability, the algorithm is incapable to correct the in­plane and out­of­plane motion phases, once these parameters are defined by the deployment conditions and therefore arbitrary. The proposed control allows the convergence to the trajectory with required shape and size. Centralized and decentralized control approaches are implemented and their performance is studied. The centralized approach considers two satellites formation formed by an electrically neutral leader satellite moving on a circular LEO and a follower which actively controls its orbital motion by changing its charge in order to remain in close vicinity of the leader. Formation flying consisting of more than two satellites with charge­changing capability can also be controlled by the proposed algorithm using a decentralized approach. This work also considers the control of satellite swarm trajectories in a sphere­shaped formation. Numerical simulation of the relative motion is used to study performance of the control algorithm. It implements the model of the geomagnetic field as a tilted dipole. The repulsive collision avoidance control is proposed for the case when the system elements are inside a dangerous proximity area. The convergence time and final trajectory accuracy are evaluated for different simulation parameters and conditions.Métodos de controlo para formações de voo de satélites de pequenas dimensões que não recorram ao uso de combustível representam, atualmente, um interesse especial e uma importante vantagem para a indústria espacial. Nesta dissertação é proposto um algoritmo de controlo que, recorrendo à força de Lorentz em orbitas terrestres baixas (LEO), é capaz de alcançar trajetórias com o respetivo o formato e o tamanho desejados. A força de Lorentz resulta de uma interação entre o campo magnético terrestre e o satélite eletricamente carregado. Alcançar as trajetórias solicitadas revela­se como sendo um desafio visto que o único método de controlo é a variação da carga interna do satélite. Este mecanismo de controlo revela­se como sendo incapaz de conferir controlabilidade total ao dispositivo. Um controlo baseado no método de Lyapunov é desenvolvido com o objetivo de eliminar a deriva inicial do satélite após o lançamento orbital e é destinado a atingir o tamanho e formato predefinidos da trajetória relativa objetivo. O algoritmo de controlo é construído de forma a corrigir os diferentes parâmetros da trajetória relativa em diferentes posições relativas. Usando a força de Lorentz é possível atingir tanto as amplitudes objetivo, considerando ambos os movimentos dentro e for do plano da trajetória, mas também a deriva e o deslocamento relativos da trajetória. Devido à falta de controlabilidade total, o algoritmo desenvolvido é incapaz de corrigir completamente os movimentos dentro e fora do plano da trajetória, visto que estes parâmetros são definidos na sua totalidade pelas condições de lançamento e, como tal, arbitrários. O algoritmo de controlo proposto possibilita a convergência dos valores para o formato e tamanho da trajetória desejada. Ambas as estratégias de controlo centralizadas e descentralizadas são aplicadas e a respetiva performance estudadas. No caso da estratégia centralizada, é considerado um voo em formação composto por dois satélites, onde o Líder se revela como sendo eletricamente neutro enquanto, e prescrevendo uma trajetória terrestre baixa circular, enquanto que o segundo, eletricamente ativo, é capaz de alterar o seu posicionamento relativo requerido, permutando a sua carga interna. Uma formação de voo considerando um número superior a dois satélites, com capacidades de carregamento elétrico, é também controlável considerando o algoritmo proposto. Este trabalho tem também em consideração o controlo da trajetória de um swarm de satélites num formato esférico. Simulações numéricas são usadas como método de análise da performance do algoritmo desenvolvido. Durante o processo de análise é implementado o modelo do dipolo inclinado como forma de simular o campo magnético terrestre. É também aplicado um algoritmo responsável por evitar situações de colisão eminente para casos em que a convergência de movimento dos satélites entra em zonas de proximidade critica. O tempo de convergência e a precisão da trajetória final são avaliadas para diferentes parâmetros e condições de simulação