Das DLR-Institut für Raumfahrtsysteme mit Sitz in Bremen hat ein neues Labor in Auf- trag gegeben, welches für die Erforschung und die experimentelle Verifikation neuer An- sätze in der Lageregelung gedacht ist. Das Labor mit seiner speziellen Einrichtung wurde Facility forAttitudeControl Experiments (FACE) benannt. Der zentrale Teststand ist ein Luft- lagertisch und wurde von der Firma Astro- und Feinwerktechnik Adlershof GmbH gefertigt. Das Luftlager besteht aus einer festen Schale in der eine Halbkugel liegt. Zwischen bei- den Elementen wird mit Druckluft ein Luftkissen aufgebaut, welches es ermöglicht die Halbkugel sehr reibungsarm zu bewegen. An dieser Halbkugel ist eine Tragplattform be- festigt auf der Experimente zur Lageregelung montiert werden können. Diese Plattform kann sich beliebig weit um die Hochachse drehen, ist in den beiden anderen körperfesten Achsen aber begrenzt durch mechanische Anschläge bei Auslenkungswinkeln von unge- fähr 20 ◦ . An der Plattform ist eine akkubetriebene Stromversorgung angebracht, die die
For the validation of the formation control strategy, the Test Environment for Applications of Multiple Spacecraft (TEAMS), a test facility forsatellite formations and swarms based on air cushion vehicles, at the Institute of Space Systems of the DLR in Bremen, Germany, is used. Two 5-DoF vehicles are floating on a granite table with a total experiment area of 5m x 4m. Each vehicle has a thruster system, a reaction wheel system and its own onboard computer running the control algorithms. A laser, attached to the first vehicle, is pointed at a fixed target on the other vehicle to demonstrate the telescope application. The Clohessy-Wiltshire equations are used for the guidance and feedforward control of the satellites position in orbit. A second controller is used to calculate the correct pointing angles for the laser and the target. The control law is based on the relative distance and velocity of the two satellites. To achieve the optimal control of the linear system, LQR controllers are used together with feedforward for optimal tracking and to cover the nonlinearities. The characteristics of the facility and its devices are introduced, the control algorithms for both satellites are explained and simulation and HIL test results are presented.
Since GRACE Follow-On is a very close rebuild of GRACE, the attitude determination and controlsystem will be almost identical to GRACE, with only few upgrades. One of the upgrades will be mounting of three SCA heads onboard each spacecraft, which is very beneficial for reasons discussed in the previous section. At the moment, no further information about the sensor performance and mounting geometry has been published yet. Also, no information about the in-flight and on-ground SCA data processing have been provided yet. However, it is expected that the performance will be slightly better than the performance of the GRACE SCAs. Similarly to GRACE, using magnetic torquers and cold-gas thrusters forattitudecontrol, the inter-satellite pointing will be maintained within a deadband of at most a few mrad. Such pointing precision will be sufficient for the microwave ranging, but surely the pointing requirements of the laser ranging interferometry will not be fulfilled. For this reason, the LRI features its own pointing controlsystem. This is accomplished by using a quadrant photodetector, which allows for measuring tilt angles between the local oscillator wavefront and the received wavefront by using differential wavefront sensing (DWS) (Anderson, 1984). This tilt information is processed by a digital control loop and fed back to the steering mirror (SM) electronics (Schütze et al., 2014a).
The second main factor to be taken into account when choosing the orbit is that, in order for the magne- torquer to be able to properly control the attitude of the satellite, the direction of the magnetic field should not be steady. As pointed out in previous sections, the exclusive use of magnetorquers to control the atti- tude of the satellite leads to a locally underactuated spacecraft. In order to be able to minimize the effect of this local phenomena, the orbit of the spacecraft needs to be such that the magnetic field varies throughout the orbit allowing exerting control torques in any direction. Taking into account that Earth’s magnetic dipole (simplification of the actual magnetic field, which is considerably more complex) has an inclination of around 11 degrees, it can be deduced that orbits with low inclinations may encounter this kind of problem. In this orbits the magnetic field remains close to an axis normal to the orbital plane, imposing a hard limitation in the torque applicable in that direction, making the spacecraft less controllable or even uncontrollable. In ap- pendix D, several analyses over a range of orbital inclinations and right ascensions of the ascending node are run. These analyses confirm the lack of controllability (with the controller developed) of the systemfor some orbits with low inclinations.
The performance of the onboard system was good over 2 years until the gyroscope assembly began delivering permanently misleading data to the onboard attitudecontrolsystem. As a consequence the orders sent to the reaction wheels were based on erroneous attitude estimation and were not related to the actual situation. The reaction wheels ran then to their limit rotation rate and stayed in this state. This happened during a non coverage period, leading to the complete depletion of the batteries… The board software entered a self-sustained cycle: the empty batteries triggered the board to boot itself, which consumed the remaining power available. A spin motion prevented a long exposure of the solar panels to the sun, rendering communication with the satellite very difficult. Finally, when the gyroscope assembly could be switched off, there was only one operational wheel left. Fortunately a complete check of the BIRD onboard systems and the payload showed good performances.
2.3 ADR Scenario Definitions and Assumptions
This paper focuses on studying the attitude motion of the chaser-tether-target systemfor the critical parts of a tethered ADR mission, which are during and following a deorbit burn applied by an active chaser. A feedback control is developed to stabilize the tether and the target separation distance in a towing ADR mission. This improves safety substantially by reducing the danger of chaser-debris collision and thus fragmentation. The paper assumes that the challenging rendezvous and target capture has been already performed. The space debris is a passive, uncooperative satellite in low Earth orbit. The chaser is equipped with a large main engine (2000 N) for performing the deorbit burn and reaction controlsystem (RCS) for applying force and torque corrections. The main engine is of an on-off type and the RCS can deliver variable thrust. Following the deorbit burn, the chaser activates two closed-loop controllers. First, the relative distance between the end bodies is controlled to maintain a small tension in the tether, and second, the chaser’s orientation is controlled to ensure the correct attitude of the chaser. The end bodies are connected by a discretized viscous-elastic tether. Based on the attitude motion analysis [18,19] of the inactive Envisat satellite, which is the focus of the e.Deorbit mission , a representative ADR mission scenario is considered which accounts for small, residual initial angular rates of the target prior to the deorbit burn. Furthermore, no input shaping of the deorbit burn is used. The main thrust is modeled as a step function, which accounts for simplified and worst-case approach. However, the reader should note that a combination of discrete deorbit burn shaping and closed-loop control following the deorbit burn may improve the performance. In the proposed analysis, the emphasis is placed on avoiding the tether tangling around the target which can result in tether rupture or debris collision and thus debris fragmentation.
The Technology Experiment Carrier TET-1, launched on 22 nd July 2012, is part of the FireBird (Fire Bispectral InfraRed Detector) constellation. Originally it was flown to qualify new technological solutions for their application in space projects in course of the “On- orbit Verification of new techniques and technologies” (OOV) program. Since 2016 it forms the FireBird mission together with the identically constructed BIROS (Bi-spectral InfraRed Optical System) satellite, launched on 22 nd June 2016. The goal of the FireBird mission is to monitor and detect high-temperature events from space.
Figure 4 shows a result of the complete chain of the initial acquisition phase. The x-axis range is four hours which is the allowed time until the spin axis and with it the solar panels shall point toward the sun. The first di- agram shows the current angular rate, the second the an- gle between solar panel normal and sun direction and the last diagram is the current dipole command for magnetic torque actuation. During the whole sequence the satellite is already considered as deployed. At first the B-dot con- troller is active which can be seen directly as the magnetic dipoles are driven to saturation. When the angular rates are damped the spin controller takes over which generates a spin around the major moment of inertia axis. After this the precession controller is switched on in parallel to the spin controller. The result is that the solar panel normal is going to align with the sun vector.
Electromagnetic actuators are a particularly effective and reliable technology for the attitudecontrol of small satellites. Such actuators operate on the basis of the interaction between the magnetic field generated by a set of three orthogonal, current-driven coils and the magnetic field of the Earth and therefore provide a very simple solution to the problem of generating torques on board of a satellite. More precisely, mag- netic torquers can be used either as main actuators forattitudecontrol in momentum biased or gravity gradient attitudecontrol architectures or as secondary actuators for momentum management tasks in zero momentum reaction wheel based configurations. The use of electromagnetic actuators, however, gives rise to a number of difficulties when it comes to con- trol system design. The torques which can be gen- erated in this way are instantaneously constrained to lie in the plane orthogonal to the local magnetic field
The latest generation of GPS satellites, termed IIF for “follow-on” is built by Boeing. Between 2010 and spring 2015, a total of nine Block IIF satellites has been added to the GPS constellation. The orientation of the spacecraft coordinate system is shown in Fig. 4 based on design draw- ings contained in Fisher and Ghassemi (1999) and Dorsey et al. (2006) . The direction of the positive x-axis can most easily be identiﬁed from the placement of the large auxil- iary payload receive antenna, which extends from the x-face. Furthermore, a þx-oﬀset of the navigation antenna relative to the center of the Earth-facing þz-panel may be recognized. The manufacturer-speciﬁc axis designa- tion is compatible with the IGS convention, as indicated by the (positive) x-oﬀset of the antenna derived by Dilssner (2010) , who adopted an IGS-style yaw-steering attitudefor the PCO and PCV estimation of the IIF spacecraft. Speciﬁc aspects of the IIF attitudecontrol law during the eclipse season are likewise discussed in Dilssner (2010) .
Scenario 4 (SFM): Strictly speaking, SFM has the same formulation for both control laws and hardware requirements; as for CAM. The main difference lies however in the fact that both modes are modelling two completely different scenarios. SFM’s primary goal is to stabilize the satellite under any initial conditions; this means, it covers the requirements of all three previous modes (see Table 6.5). They are, the initial quaternion (transition from any mode), the initial angular rate with a maximal magnitude of 10 deg·s −1 and random direction (simulating a transition from
In order to prove that the algorithm can also handle the frequent changes of the GPS satellite constellation typical for space-borne applications, numerical simulations have been executed. During these simulations, the C/A-code and carrier-phase measurements of the GPS system on the Flying Laptop satellite have been created using realistic attitude scenarios. In the first attitude scenario, the nadir-pointing-mode, the satellite orbits the earth with a constant angular velocity vector in nadir orientation. In this attitude mode, the assumed system model for the Kalman Filter comes close to the real world dynamics, thus the filter does not suffer performance problems. The second attitude mode is the target pointing mode. During this mode, the satellite is pointed to a target fixed on the earth’s surface and thus continuously controlled by the attitudecontrolsystem, which applies control torques to keep the satellite aligned with the reference frame during its pass over the ground target. Due to the applied torques, the angular velocity vector of the satellite changes and the assumption made for the system model in the Kalman filter is no longer true.
The satellite platform BIROS is the second technology demonstrator of DLR’s ‘FireBIRD’ space mission aiming to provide infrared remote sensing for early fire detection. Among several mission goals and scientific experiments, to demonstrate a high-agility attitudecontrolsystem, the platform is actuated with an extra array of three orthogonal ‘High- Torque-Wheels’. However, to enable agile reorientation, a challenge arises from the fact that time-optimal slew maneuvers are, in general, not of the Euler-axis rotation type; espe- cially whenever the actuators are constrained independently. Moreover, BIROS’ on-board computer can only accommodate rotational acceleration commands twice per second. The objective is therefore to find a methodology to design fast slew maneuvers while con- sidering a highly dynamic plant commanded by piecewise-constant sampled-time control inputs. This is achieved by considering a comprehensive analytical nonlinear model for spacecraft equipped with reaction wheels and transcribing a time-optimal control prob- lem formulation into a multi-criteria optimization problem. Solutions are found with a direct approach using the trajectory optimization package ‘trajOpt’ of DLR-SR’s opti- mization tool, Multi-Objective Parameter Synthesis (MOPS). Results based on numerical simulations are presented to illustrate this method.
Nachdem die Simulationsumgebung entwickelt ist, wird sich auf das Gebiet der Lagebestim- mung konzentriert. Hierbei wird sich auf sogenannte KALMAN Filter konzentriert. Da der Satellit ein nichtlineares System ist, wird zunächst ein Extended KALMAN Filter (EKF) en- twickelt. Dieser EKF schätzt den eigentlichen Zustandsvektor des Satelliten mit Hilfe eines Fehlerzustandsvektors, wodurch eine präzise Lagebestimmung ermöglicht wird. Als Vergleich zum EKF wird eine relativ neu Erweiterung innerhalb der KALMAN Filter betrachtet. Dieser Unscented KALMAN Filter (UKF) ist in der Lage Nichtlinearitäten eines Systems besser abzu- bilden und somit eine höhere Genauigkeit zu erreicht. Sowohl der EKF wie auch der UKF werden theoretisch erläutert und anschließend in die Simulationsumgebung integriert. Die ab- schließenden Vergleichstests beider Filter zeigen, dass der UKF den EKF übertrifft und tatsäch- lich bessere Ergebnisse liefert.
launched to space in 2019 and demonstrate the latest generation of the optical space infrared downlink system developed by the German Aerospace Center together with the industrial partner Tesat-Spacecom. The retroreflector is optimized to allow for a coarse verification of the satellites attitudecontrolsystem. By analyzing the returning photon count during satellite laser ranging (SLR) when the satellite is operated in station pointing mode attitude information is obtained. To achieve this goal, the entrance face of the retroreflector is recessed by a circular tube-shaped aperture. Due to this recession, the signal reflected from the retroreflector falls off rapidly when the retroreflector is tilted away from the SLR station. From measure- ments of the retroreflectors far-field diffraction pattern and calculations, we believe that it should be possible to determine the orientation accuracy of the satellite to within ± 2°. The proposed method is an effective and cheap way for coarse attitudecontrol, e.g., for satellites of mega-constellations with any existing satellite laser ranging ground station.
The magnetic coil system is an attractive method of desaturation for low Earth orbit (LEO) satellites due to the relatively high magnetic field intensity at lower elevations. This system also has high reliability since it includes only simple static devices (a magnetometer, a signal processor, and three coils). Other advantages are that it does not depend on a fuel supply and is much lighter than the simplest low specific impulse thruster system. Disadvantages of this system are that it may require significant amounts of power at higher altitudes. Also, coil commands may last over a large fraction of the orbit (or over several orbits) to reach desaturation. Magnetic systems may also interfere with the operation of certain payloads. The other well-known technique used for desaturation are mass expulsion torqueing. Thrusters are usually used to desaturate the momentum storage systems. In operation, a jet is fired to produce a torque opposite to the direction of the accumulated angular momentum while the satellite is commanded to maintain its attitude. This results in a wheel acceleration that counteracts the applied torque. For a desaturation system using body-fixed offset roll/yaw thrusters, the efficiency of the system can be defined as the ratio of the daily secular momentum increase to the angular impulse provided by the thruster. A reasonable design value for the efficiency is about 80 percent .
In addition, we describe the test workflow as well as the software technologies used for the implementation. This paper ends with first reports of experience from the tests of the attitudecontrolsystem software of the compact satellite BIRD. At least the paper shows an outlook for the use of the upcoming OOV TET platform.
where M þ d is the pseudo-inverse of M d , which is not invertible due to its dimension and is rank-deficient in hover mode since the elevator effectivity is zero. ! is the pseudo-control input for the INDI control loop and is a vector of commanded angular accelerations. The prospect of INDI as described in Smeur et al. 5 is that closed-loop dynamics from i to _ X i equals the corresponding actuator dynamics, where i corresponds to the moment axis (i ¼ x; y; z). That is, the roll accel- eration loop has the dynamics of the aileron actuators. The actuator dynamics are modelled for each actuator individually as first-order lags and are denoted as A i ðzÞ. We note that the pitch axis is actually over-actuated during the transition and fast-forward flight, since both the elevator and the auxiliary motor thrust are effective. For the purposes of designing the attitude controller, we use the slower dynamics of the auxiliary motor to derive attitude controller gains. Figure 4 shows the structure of the attitude controller. To close the attitudecontrol loop, we use a traditional cascaded approach, where both the angular rates and the attitude angles are fed back using proportional (K _x) resp. proportional derivative (PD) controllers. To generate the wing-fixed commanded angular rates from attitude angle errors, the commanded attitude rates _ U c ; _H c ; _W c are first multiplied by the inverted atti- tude angle dynamics T X (18) and then transformed into the wing-fixed coordinate system using the transforma- tion matrix T wb (1).
DLR/Moraba has a long history in the development of Rate Control Systems (RCS) which have been flown for several programs for research in weightlessness in Europe and Brazil. They had been used in various European µG programs such as MAXUS, MASER and TEXUS. Moraba has now developed a RCS for 14 inch payloads. This RCS uses a standard interface and can be easily joined to a REXUS service module. This combination (RCS and REXUS service system) came into operation for the MAPHEUS maiden flight.
Besides specified advantages, there are however several disadvantages, which makes the realization of TPMS for traffic data detection challenging. First of all, TPMS com- munication is not standardized, which leads to numerous implementations of TPMS based on the the demands of the producer itself. This means that future system needs to be able to detect all possible sensor types. However, certain efforts for standardizing this communication exist . Next, EMC (Electromagnetic Compatibility) regulations for vehicles allow sensors to radiate relatively low power level. The power of the transmitted signal is additionally influenced by the metal rim and the car body , . Further, the characteristics of the communication channel change dynamically due to the rotation of the wheel and put serious demands on the dynamic range of the receiver . The system should be able to cope with highly redundant data, present due to the fact that every vehicle has more than one TPMS sensor. The scope of this paper will be oriented more towards evaluating potentials for using TPMS in deriving traffic information, while the other addressed problems will be the topics of further research.