Nach oben pdf Development of algorithms for attitude determination and control of the AsteroidFinder satellite

Development of algorithms for attitude determination and control of the AsteroidFinder satellite

Development of algorithms for attitude determination and control of the AsteroidFinder satellite

Die vorliegende Diplomarbeit wurde im Rahmen des AsteroidFinder/SSB Projektes des Deutschen Zentrums für Luft- und Raumfahrt e.V. (DLR) erstellt. Der AsteroidFinder/SSB ist ein Kompak- tsatellit und soll zur Entdeckung von Asteroiden eingesetzt werden. Hierfür wird ein Teleskop verwendet, welches die Nutzlast des Satelliten ist. Die Aufgabe lichtschwache Objekte, wie Asteroiden, zu entdecken stellt anspruchsvolle Anforderungen an das Lageregelungssystem. Aus diesem Grund wird AsteroidFinder/SSB durch eine aktive dreiachsige Lageregelung sta- bilisiert. Das vorläufige Design des Lageregelungssystems wurde bereits in den Phasen 0/A des Projektes festgelegt. Aufbauend auf diesen Ergebnissen fokussiert sich die folgende Arbeit auf die Entwicklung der Lagebestimmungs- und Lageregelungsalgorithmen. Dafür wird zunächst eine Simulationsumgebung entwickelt und zwei verschiedene KALMAN Filter untersucht. Die KALMAN Filter sind der Extended und der Unscented KALMAN Filter, welche als Lagebes- timmungsalgorithmen in Frage kommen. Anschließend werden verschiede Flugführungsstrate- gien besprochen, um die Missionsziele zu erreichen. Dies wird gefolgt von der Entwicklung einer Lageregelungsstrategie bei deren Entwicklung Methoden der Linearen Quadratischen Regelung zum Einsatz kommen. Als letztes werden sämtliche Algorithmen durch Simulation validiert.
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Attitude Control System of the Eu:CROPIS Mission

Attitude Control System of the Eu:CROPIS Mission

The Eu:CROPIS (Euglena Combined Regenerative Organic food Production In Space) satellite, scheduled for launch in 2017, is the next mission to be launched as part of the German Aerospace Center’s (DLR) compact satellite program. The mission is currently in Phase C and is being developed jointly by several institutes within DLR. The mission’s focus is to test several biological experiments at different levels of artificial gravity. The payload modules are provided by different DLR institutes, the University of Erlangen and NASA-AMES. The satellite itself has a mass of about 230 kg and includes several subsystems which directly interfere with the attitude control system (e.g. deployable solar panels, liquid pumps and venting devices). This paper provides an overview of the Eu:CROPIS Attitude and Orbit Control System (AOCS). It starts by presenting the design driving requirements and explains how the required g-levels are achieved purely by a magnetic spin stabilization concept. Following this is a presentation of the AOCS modes and a discussion of the chosen sensors and actuators. The attitude determination and attitude control algorithms are then described. Finally, an outlook is given for further development steps of the Eu:CROPIS satellite.
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Using data fusion of DMARS-R-IMU and GPS data for improving attitude determination accuracy

Using data fusion of DMARS-R-IMU and GPS data for improving attitude determination accuracy

The Mobile Rocket Base (MORABA), a division of the Space Operations and Astronaut Training Department of the German Aerospace Center, has developed and flown sounding rocket rate and attitude control systems since 1972. For determining position and attitude, some of the past missions have used a DMARS-R (Digital Miniature Attitude Reference System) roll stabilized platform produced by the Inertial Science Company. The DMARS-R is a high-precision, roll-stabilized IMU (Inertial Measurement Unit) comprising accurate angular rate and acceleration sensors and mounted on a roll-stabilized platform. Standard IMUs are inherently subject to drift in inertial position, velocity and attitude. By the fusion of GPS and with DMARS-R data, one can achieves a long-term drift-corrected IMU enabling for longer duration flight applications, such as satellite launchers, spin stabilized rockets and balloons. The main experiment of the MAIUS mission, initially planned for autumn 2015, requires accurate pointing to the Earth’s gravitational center.
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Efficient and high precision momentum bias attitude control for small satellite

Efficient and high precision momentum bias attitude control for small satellite

three orthogonal axes. For attitude acquisition, 6 solar cells/panels which attaches to each side of the spacecraft are applied as Sun sensor. In the later series, the spacecraft also equips its attitude control system with a magnetometer. These two types of coarse attitude sensor would provide an attitude reference in the acquisition phase while the spacecraft was rotating in the faster speed e.g. after launch separation or in the hibernation mode. Once the reaction wheel could handle the spacecraft rotation below 0.5 deg/s, the star tracker/sensor will establish a more accurate attitude determination, in particular to find the direction of momentum vector. However, operating momentum bias stabilization for small satellites in low Earth orbit requires special provision. Unlike in geostationary orbit, in which the strategy is quite popular, the angular momentum in low Earth orbit is dominantly disturbed by Earth’s magnetic field and the aerodynamic drag. But in the case of LAPAN-A satellite series that orbiting above 500 km, the aerodynamic effect is negligible. Magnetic disturbance could be minimized by compensating the residual magnetic dipole on the spacecraft, which could be done by the 3 air coils after the internal magnetism can be measured. While geostationary frequently use thrusters to maintain the angular momentum and horizon sensor to monitor the direction of the vector, LAPAN-A satellite series use magnetic torquers to generate external torque together with three reaction wheels and utilize star sensor to keep track of the angular momentum vector [46] [57].
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The software architecture for TET and AsteroidFinder satellites

The software architecture for TET and AsteroidFinder satellites

The software architecture for TET [1] and AsteroidFinder [2] satellites is a further step in the development line first used for the BIRD [3] satellite. The new improvements add dependability, flexibility and simplicity to a core avionic system already implemented, simulated and tested in similar ESA and industrial projects, proving the basic concept of the architecture. The new core avionic concept targets the problems of complexity, software- hardware interfaces, and the difficulties of merging many different interfaces into a single system by providing a very simple solution of integrated software and hardware, thus eliminating the barrier between the two. Through this concept both the bus control and payload control can be handled through one system. Implementing a complex parallel system safely requires the composition of a network of simple sequential cooperating applications which can communicate by using well defined interfaces. The basic communication principles common to all target systems are decribed in this paper.
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COORDINATED ORBIT AND ATTITUDE CONTROL OF A SATELLITE FORMATION IN A SATELLITE SIMULATOR TESTBED

COORDINATED ORBIT AND ATTITUDE CONTROL OF A SATELLITE FORMATION IN A SATELLITE SIMULATOR TESTBED

For the validation of the formation control strategy, the Test Environment for Applications of Multiple Spacecraft (TEAMS), a test facility for satellite 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.
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Flight Results of GPS-Based Attitude Determination for the
Canadian CASSIOPE Satellite

Flight Results of GPS-Based Attitude Determination for the Canadian CASSIOPE Satellite

In the last step of the quality control algorithm, time differences of pseudorange and carrier-phase double-difference observations (triple-differences) are formed for each individual satellite. The absolute value of each triple-difference is then compared to a test thresh- old, and the observation is rejected if the threshold is exceeded. This editing procedure is possible due to the high data rate of the receivers and the low angular veloc- ity of the CASSIOPE satellite, which only causes a small effect in the triple-difference due to the geometric change in the baseline. In this editing step, pseudorange jumps and carrier-phase cycle-slips are detected. If cycle-slips are detected, the corresponding ambiguities are newly initialized as float values based on code-carrier differences. The next step is the measurement update. The Kalman- filter processes both pseudorange and carrier-phase observations. It can be configured to use single-frequency measurements from any number of signals. In the case of CASSIOPE, it can process only L1 C/A, only L2 P(Y), or both signals together. No combination is formed when processing more than one frequency. Instead, mea- surements from the different frequencies are treated as independent and complementary observations. The ionospheric delay does not need to be estimated since it cancels out in the differencing over the short baselines. Using both L1 C/A and L2 P(Y) is a particularly interest- ing option, since observations on a second frequency with
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GPS Based Attitude Determination for the Flying Laptop Satellite

GPS Based Attitude Determination for the Flying Laptop Satellite

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 attitude control system, 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.
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The role of attitude determination for inter-satellite ranging

The role of attitude determination for inter-satellite ranging

The first absolute gravity measurements were carried out around the turn of the 17th and 18th century by means of diverse types of pendulum. The first spring-based relative gravimeters were developed in the first half of the 20th century and later on the absolute gravimeters based on rise-and-fall and free-fall methods were built, which were significantly more accurate than the pendulum gravimeters (Torge, 1989). The technology development in the last few decades allowed to build highly precise relative superconducting gravimeters (Goodkind, 1999), and currently absolute quantum gravimeters are under development (de Angelis et al., 2009). With these instruments it is possible to perform the gravity measurements only pointwise. Later, regional measurements have been possible by means of airborne and ship- borne instruments. However, the determination of the global gravity field was not possible before using measurements from space. The very first satellite, Sputnik 1, was launched in 1957 and since then the number of launched satellites has raised exponentially. The satellite orbits are perturbed by the inhomogeneous mass distribution within the central body and so it is possible to recover the gravity field from the orbit tracking data. One of the first global Earth’s gravity field models was presented in 1966 by Lundquist and Veis (1966). The accuracy of the first models was rather low, though.
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Attitude Determination of Geological Layers Using HRSC Data and Orion Software

Attitude Determination of Geological Layers Using HRSC Data and Orion Software

Introduction: Previous work [6, 2] has demon- strated that MOLA and wide angle MOC data can be combined in Orion software to determine attitudes of large scale layering within the chasmta walls of Valles Marineris. Individual layers, between 7.7 km and 83.8 km in lateral extent, were found to have statistically consistent orientations. Layer attitudes indicated largely shallow dips into the canyon and were inter- preted to record the collapse that produced the early ancestral basins [2]. Unfortunately the resolution of the MOLA 1/128° x 1/128° topographic grid file (463 m/pixel) limited the scale of layering that could be in- vestigated. While many more detailed layers are visi- ble in narrow angle images, most narrow angle MOC images only span about 10 DEM pixels, which makes it impossible to obtain quality attitudinal data from them.
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Attitude control analysis of tethered de-orbiting

Attitude control analysis of tethered de-orbiting

In summary, the numerical simulations presented in Figures 6-23 show that the proposed control technique can keep the target alignment angle with a controllable and safe range (less than 90°) for more stiff tethers, which is due to the increased damping of the target attitude motion caused by the restoring torque generated by the tether tension, thus improving safe towing substantially. The results confirm that for safety reasons, elastic tethers should be avoided, because they can lead to target attitude motion excitation and thus debris collision and fragmentation. It is shown, that the tether damping constant has no significant influence on the target rotation. Furthermore, elastic tethers were discovered to be more difficult to control, due to less tether tension required to maintain the same tether elongation. For stiff tethers, a relatively simple guidance and control system is shown to be sufficient to keep the tether tension to reasonable levels, to stabilize the tether and target dynamics, in a towing mission. In off nominal cases where only attitude control thrusters are available, or the control frequency is low, the proposed tethered controller has been shown to stabilize the tethered system thus avoiding a chaotic or dangerous behavior for the chaser, even when the attitude of the target increases substantially, especially about the roll axis.
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Development of an electronic monitor for the determination of individual radon and thoron exposure

Development of an electronic monitor for the determination of individual radon and thoron exposure

In order to cross check any possible issues also an analytic simulation relying on the NIST alpha stopping power and range (ASTAR) database [ 11 ] was developed, which itself is a program based on the ICRU 49 report [ 129 ]. Therefore a comparison of simulation results at STP conditions should yield similar results, thus implying correctness of the Geant4 cal- culations within the boundaries of the introduced assumptions. This verication method is in general suggested for Monte Carlo simulations concerning alpha particle transport [ 47 ]. The analytical program, implemented in Wolfram Mathematica 9.0.1, uses the same start- ing conditions as the Geant4 simulation, with random generation of alpha particle sources in volume and on walls as well as random decay directions into the 2π solid angle facing the detector. However, particle transport is based on an analytic approach, meaning an energy loss is only calculated along the straight line through air and aluminum from the origin to the entry point at the detector. All decays not hitting the detector are not calculated. The kinetic energy loss dE is approximated by linear iterative steps along the ASTAR stopping power dE/dx. Again the simulation is more accurate the smaller the step length, as a constant energy loss per step is required. Using the available NIST values, however, denies the possibility to change the air composition. Only the air density can act as parameter on the stopping power. A total number of n = 150, 000 events per decay energy was simulated and all results were smoothed via Equation 5.17, with ∆E = 150 keV , to account for the Gaussian energy broadening of the electronics.
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From Observational Geometry to Practical Satellite Design: AsteroidFinder/SSB

From Observational Geometry to Practical Satellite Design: AsteroidFinder/SSB

With respect to SSSBs, the innermost solar system still is largely uncharted territory. Information on the orbital parameters and approximate size of previously unknown IEOs and other objects passing through this region is critical for the evaluation of SSSB distribution models. These models are used primarily for two important purposes: In the planetary defence context, they serve to determine the overall risk and frequency of impacts on the Earth and other terrestrial planets, and the size-frequency and relative velocity distribution of the impactors. In the wider scientific context, many of these models are based on the orbital evolution of the solar system as a whole, and modelled SSSB populations serve as sets of test particles that as a whole record and statistically image the integrated influence of various gravitational and non-gravitational effects over time. To determine the relative strength of these effects and their variation over time, observed and modelled populations can be compared at varied parameter settings, and before and after correction for observational biases which may also be determined in the process. The energy-frequency distribution of impactors is also used to determine the age of solid surfaces in the solar system, expanding the relative dating of planetary surfaces from the size-frequency distribution of craters alone. For absolute dating, a reference is required which can only be provided by returned samples that are dated by isotope clocks, such as the Apollo and Luna Moon rocks. The orbital and size-frequency distribution of impactors varies across the solar system. For example, IEOs can presently not reach objects beyond the Earth-Moon-system, and Aten asteroids can not reach the surfaces of main belt asteroids, although both may well migrate over time due to long-term perturbations to hit or become part of either group of solar system bodies. These localized SSSB population differences have to be modelled to determine the absolute age of planetary surfaces outside the Earth-Moon system as long as local surface samples remain unavailable. The high number of observed and modelled bodies enables sound statistical results. Each body adds seven or more parameters to the database; its orbit parameters, estimated size, and occasionally its shape and other physical properties. For example, the model population by Bottke and Morbidelli used in the evaluation of the AsteroidFinder’s performance contains 57649 virtual objects, including 1190 virtual IEOs, down to a limiting absolute magnitude H = 23.0, corresponding to a diameter of about 100 m at an albedo of 0.15. [3]
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Determination of dynamical models for adaptive control systems

Determination of dynamical models for adaptive control systems

It is shown that if the nominal values of all poles of a system are known, and if only one pole changes from its nominal value, then this change may be detected.. It is also demonstrat[r]

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Satellite Selection Methodology for Horizontal Navigation and Integrity Algorithms

Satellite Selection Methodology for Horizontal Navigation and Integrity Algorithms

Satellite selection in general was topic of several papers in the past, providing different algorithms to perform the task. The common goal is in general to find a subset of all satellites currently in view which provides the best (whereby best depends on the application) navigation performance under given side conditions. Simple and computationally efficient approaches using only satellite elevation and az- imuth were presented by Zhang [7], Song [8] and Park [9]. Also the widespread approach of selecting the highest satel- lites, assuming the best signal quality with highest elevation, falls into this category. Another group of algorithms tries to maximize polyhedron volumes or matrix determinants [10–12] as they correlate well with the geometrical dilution of precision (GDOP) of a set. Both of these strategies suffer from the missing possibility to take into account any satellite weighting (e.g. due to signal quality) and focus on 3D po- sitioning, making them less suitable in our context. Phatak presented a method in [13] which allows efficient exchange of single satellites in a set which we will utilize later as well. Beyond that, approaches applying genetic algorithms [14] or artificial neural networks [15] were presented earlier. While all these works show that satellite selection in gen- eral was covered from many perspectives within the last decades, ARAIM in particular has not been addressed so far. As ARAIM comes with several special characteristics, algorithms achieving good selections in terms of DOP are not necessarily suitable for ARAIM without changes. This is especially true for Horizontal ARAIM, focusing on the horizontal performance and therefore requiring appropriate geometries.
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Design and qualification of a recessed satellite cornercube
retroreflector for ground‑based attitude verification via satellite laser ranging

Design and qualification of a recessed satellite cornercube retroreflector for ground‑based attitude verification via satellite laser ranging

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 attitude control system. 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 attitude control, e.g., for satellites of mega-constellations with any existing satellite laser ranging ground station.
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Decision algorithms for modelling, optimal control and verification of probabilistic systems

Decision algorithms for modelling, optimal control and verification of probabilistic systems

provide a value iteration-based decision algorithm to approximate the Pareto set of achiev- able points. Our algorithm iterates over a weighted sum of objectives which are in turn optimized through a value-iteration procedure. It is worthwhile to note that the major difference between the provided value iteration algorithm and the standard value itera- tion is its need to optimize a mixture of unbounded and bounded properties. Moreover, the multi-objective robust controller synthesis for IMDPs cannot be solved on the MDPs generated from IMDPs by computing all feasible extreme transition probabilities and then applying the algorithm in [FKP12]. The latter solution approach could be valid if the two sources of nondeterminisms in IMDPs were resolved cooperatively. As for the sake of ro- bust controller synthesis we need the competitive semantics, one can instead transform IMDPs to 2 1 2 -player games [BKW14] and apply the algorithm in [CFK + 13]. Unfortunately, the transformation to (MDPs or) 2 1 2 -player games induces an exponential blow up, adding an exponential factor to the worst case time complexity of the decision problem. Our pro- posed algorithm instead prevents this difficulty by solving the robust synthesis problem directly on the IMDP so that the core part, i.e., optimizing the weighted sum of objectives, can be solved with time complexity polynomial in the size of the given IMDP model.
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Attitude and Orbit Control of the Grace Satellites at extremely low power

Attitude and Orbit Control of the Grace Satellites at extremely low power

Star Camera and GPS data were handled and processed by the IPU. The part needed by the AOCS was forwarded to the OBDH. Right after launch the IPU was doing well, but already during the commissioning phase the scientists requested the introduction of 1 Hz operations on both star cameras in order to get data to facilitate more accurate attitude determination. The default setting in the IPU was 1 Hz on the prime and 0.1 Hz on the backup camera. Dual 1 Hz operations led to overstrain, which then additionally led to more frequent IPU resets. Also the GPS data showed numerous invalid points which did not have a negative impact on the fuel consumption but regularly increase the time for re-acquisition after an IPU reset. At the beginning of the mission the IPU experienced frequent resets (once or twice per day). Roughly 10% of the resets resulted in a hang-up of the IPU caused by SCA and GPS re-acquisition problems. Additional FDIRs were introduced later in the mission in order to mitigate the problem of hang- ups. In 2006 the IPU software was updated, solving the problems with the Dual 1 Hz operations and invalid GPS data. The interval between resets became then so long, that after roughly 900 hours an IPU reset had to be commanded to keep clock biases within 500 ms (a requirement for science)
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On the Kalman Filtering Formulation for RTK Joint Positioning and Attitude Quaternion Determination

On the Kalman Filtering Formulation for RTK Joint Positioning and Attitude Quaternion Determination

Attitude determination is the process of estimating the ori- entation of a moving rigid body with respect to its environment and it can be considered as a classical problem with a history of more than 50 years. In a multi-antenna GNSS system, one seeks to find the rotation which relates the baseline vectors joining each pair of antenna positions across the body and the ECEF (Earth-centered Earth-fixed) frames. Although GNSS attitude determination provides substantially higher accuracy than other systems, generally based on magnetic effects, its implementation poses some constraints. On one hand, at least a couple of GNSS receivers providing carrier phase are needed. Although receivers capable of providing carrier phase are currently costly, the technology for the production of low-cost receivers is rapidly growing. On the other hand, attitude accuracy is inversely proportional to the separation between the antennas, making this system impractical for small vehicles.
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Development of selective sorbent materials for the determination of polychlorinated biphenyls (PCBs) in the environment

Development of selective sorbent materials for the determination of polychlorinated biphenyls (PCBs) in the environment

MIPs selective to PCBs would contribute positively to their monitoring and especially for selective pre-concentration at trace concentrations, however, imprinting PCBs is challenging. Firstly, some of their congeners are toxic and may therefore not be used as templates. Secondly, PCBs lack functional groups, and therefore do not fully satisfy the requirements for non-covalent imprinting, which favors multifunctional tem- plates. Likewise, the functional monomer bears multiple func- tional groups facilitating the formation of stable pre- polymerization complexes via pronounced e.g., hydrogen bonding or ion-pair interactions. 67,88,89 The only association opportunities for these constituents with functional mono- mers are rather weak π–π stacking, and hydrogen bonding with chlorine atoms. Thirdly, template bleeding is of substantial concern for application of PCB-imprinted polymers in environ- mental trace analysis. Two independent studies have indicated that in molecular imprinting 0.47 –1.38% and 3 ng mL −1 of the template may remain incorporated in the polymer matrix even after extensive extraction procedures. 90,91 Relating this amount of potential residues with the maximum contami- nant level (MCL) of 0.5 ng mL −1 for ∑PCBs in drinking water (as given by the USEPA), template bleeding would signifi- cantly contribute to erroneous results. These challenges have therefore demanded particularly innovative strategies for imprinting PCBs that address these issues, while achieving the desired selectivity and stability properties of the resulting MIPs.
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