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In the past few years, a new spirit for the exploration of the Solar System spread among the space community and reaching the Moon has been set as the first step of this new program. In this frame, going back to the Moon is needed to familiarize with a new way of living in a different environment, adapting to it and testing new technology. It is also true that, at the rate we are consuming the terrestrial resources, we will soon run out of them. This will put us in front of a dramatic change in our life style. Moreover, it is not unlikely that an asteroid could impact the Earth, causing extinction of many species and difficulties for survival. Then, these unpredictable reasons increase the importance of exploring and adapting to new extraterrestrial environments. Therefore, a feasibility study of a mission to the Moon to set up a permanent base has been carried on. The first part is concerning the delivery of the lunar payload into a LEO parking orbit. For this, the analysis of the capabilities of existing launchers is performed. The ascent trajectory problem is tackled by formulating it as an initial value problem (IVP), in which, given the launcher’s initial conditions, the state vector is propagated following a control law optimized to give the largest payload mass. Moreover, the launcher is subject to constraints dictated by the mechanical and thermal properties of the launcher itself. The optimal control law is sought by means of a Particle Swam Optimization method, which simulates the behavior of a flock of birds searching for food. Single and multi-objective optimization is performed. Single-objective optimization aims at maximizing the payload mass satisfying path constraints and the boundary constraints dictated by the orbital elements of the final orbit. Multi-objective optimization aims at maximizing the payload mass and minimizing the error on the final orbit simultaneously. Other experiments include the optimization of the two aforementioned objectives and the minimization of the violations of the path constraints. It has been found that, to fulfill the requirements of the lunar campaign, a very tight schedule and international cooperation is needed. Yet, existing launchers can be used for this mission for cargo expeditions. However, it is strongly suggested to commence development of a manned launcher and a spacecraft capable to land and host astronauts for multiple days on the Moon.

Understanding the Sun's natural processes helps us to understand how the Earth responds to the Sun's variations and helps improve our ability to predict its behavior. A probe in polar orbit around the Sun would be able to help provide a more complete picture of the Sun by studying coronal features from various angles, by linking particle and field observation to images of the Sun, by determining the magnetic structures and convection patterns in the polar regions of the Sun, and by following the evolution of solar structures over a full solar rotation. Previous theses have tackled the transfer from Earth to Solar Polar Science Orbit by using solar sails to propel the spacecraft. An alternative is a spacecraft using solar electric propulsion. This offers the advantage of a spacecraft that is more maneuverable, less mechanically complex, and is based on proven technology. Unlike the solar sail spacecraft however, this requires the spacecraft to carry propellant onboard to perform the transfer. An almost entirely analytical implementation is constructed, featuring the use of the shape-based approach (exponential sinusoid) to model the transfer. The implementation connects multiple shapes by way of gravity assist at specified planets, and uses Edelbaum's theory to model the remaining necessary inclination change, to investigate a transfer from Earth to a solar polar orbit at 0.4 AU distance from the Sun. An optimal solution is sought using an optimization strategy consisting of a random (Monte Carlo) method to perform an initial exploration of the search space, a global (Genetic Algorithm) method, and a local (Nelder-Mead) search method to refine the optimum located by the global method. The goal is to minimize propellant mass and transfer time, while maintaining a realistic transfer that does not violate various imposed constraints (such as departure and arrival velocity). Two solar electric propulsion options are offered (one performing a swing-by at Venus and another performing a swing-by at Jupiter), and compared to a reference solar sailing design. After experimentation it was found that the model is useful for simpler transfers, but a more involving implementation (using multiple separate coasting and thrust arcs) is required to give it the means to tackle more intricate problems. In addition, a more in-depth exploration of the inclination change maneuvers would be beneficial (especially with regards to a distance constrained maneuver).

On March 17, 2009, a rocket named Stratos was launched, by Delft Aerospace Rocket Engineering, or DARE in short. The purpose of this rocket was to break the European altitude record for amateur rockets which was 10.7 km at that time. A design of a small sounding rocket, such as Stratos, is a good example of an interdisciplinary challenge. An optimal design is a combination of structures, manufacturability, propulsion, aerodynamics, electronics and many other aspects The Stratos rocket achieved an apogee altitude of 12.551 meters at the Esrange Space Centre, thereby setting a new European altitude record for amateur rockets. A detailed flight trajectory reconstruction is done, whereby differences between simulation and reality are explained. The objective of this thesis is to give a detailed insight in the design, the simulation tools, the production process and the results of the mission. The design philosophy of the Stratos rocket, simulation software and the production planning concept are valuable tools in the development of newer and more powerful rockets, which could ultimately result in a successor of the Stratos rocket.

Four decades after Neil Armstrong's 1st steps on the Moon the time has come to take it one step further, the permanent occupation of humans on lunar soil. Permanent occupation of humans will revive the dream of exploring space and will open the doors for space flight to Mars or even further. For this reason this project investigates the possibilities of placing a base on the South Pole of the Moon before July 1, 2020. The South Pole has been chosen for several reasons, although the main reason is the recently discovered presence of water by NASA's LCROSS mission in 2009. The base itself consists of four modules made of an inflatable structure to minimize for weight. The following research question was set at the beginning of this study, Is it possible with the current technology to establish a permanent manned base on the South Pole of the Moon before July 1, 2020? This project was setup by ESA and is a joined effort of three students of the Delft University of Technology, Antonio Pagano, Valentio Zuccarelli and myself. The mission itself is therefore also divided into three parts, the Earth ascent, the Earth-Moon trajectory and the descent to the lunar surface. This study will focus on the last part, the design of the lunar descent trajectories. The goal of this study is, Design of the optimal lunar descent trajectory with the goal of constructing a permanent base on the lunar South Pole. This project assumed that the launchers Ares-I and Ares-V are available for the lunar base mission, as well as Constellation's lander Altair. The descent trajectory is optimized for maximum payload that can be brought to the lunar surface. The descent starts at the end of the Earth-Moon trajectory, a 100 km altitude, 89.8 deg inclination circular parking orbit. The following three different methods are used to study the descent trajectory and are compared at the end: dividing the descent into intervals, indirect method and gravity-turn trajectory. By using the indirect method the biggest payload mass can be brought to the lunar surface. With the results a launch schedule is proposed, including 19 launches, that make it possible to build the lunar base before July 2020.

A transition through the Earth's atmosphere is inevitable if it is desired to bring or return something useful from space. The transition is also known as the atmospheric entry, which is characterised by a highly energetic vehicle. Three types of entry can be distinguished: ballistic, glide and skip entry. In the hypersonic transition of a glide entry, the heating and structural loading can become severe enough to damage the vehicle and/or the crew. The attitude of the vehicle has to be controlled to avoid this. In addition, the vehicle should target an interface with the TAEM phase. During which the vehicle is aligned with the runway. The process of determining the attitude throughout entry is called mission planning. For on-board mission planning, simplifications need to be made on the vehicle, its environment and the flight dynamics to achieve an acceptable computation speed. As a consequence, the real trajectory will deviate from the planned trajectory. The trajectory tracker has the task to steer the vehicle towards the planned trajectory. The combination of a planning and tracking algorithm forms a guidance algorithm. The main question of this thesis work is formulated as: Is it possible to design an on-board executable guidance algorithm, for the hypersonic transition phase, which safely targets the TAEM interface? Four tasks have been derived from this question: simulator development, planning algorithm design, tracking algorithm design and guidance algorithm testing. The simulator serves as a test bed for the guidance algorithm. In the development of this software, a systems-based approach is taken with respect to the vehicle and its environment modelling. The advantage is that the software has a clear modular structure and it becomes easy to extend the simulator with new capabilities. The design of a trajectory planner has been decomposed into an angle of attack and bank angle planner. The angle of attack planner operates on the assumption of an equilibrium-glide trajectory. The bank angle planning algorithm is centered around an iterative search for a drag profile that corresponds to an estimated trajectory length travelled between entry and TAEM interface point. By incorporating the bank angle planning and tracking algorithms in one guidance algorithm, the main question is answered positively. The algorithm executes fast enough for an on-board implementation.

Saturn’s moon Titan is unique in that it is the only moon in the Solar System with a significant atmosphere. Its atmosphere is thicker than the Earth’s and it contains several layers of hazes. Knowing the properties of the hazes is crucial for understanding the existence and evolution of this atmosphere and the dynamical processes that take place. For a spacecraft instrument, the highest haze layers hamper the view on the lower layers. Instead, a hot-air balloon floating through Titan’s atmosphere, as proposed in the Titan Saturn System Mission (TSSM) to ESA, appears to be the ideal in-situ element for the desired science goals. The haze particles can be studied by an instrument on the balloon’s gondola that measures sunlight that has been scattered by the haze particles. We present SPEX, the Spectropolarimeter for Planetary EXploration, as an instrument to retrieve information on the number density, composition, size and shape of the haze particles. As payload on the TSSM balloon, SPEX would measure simultaneously the radiance and degree and direction of linear polarization of scattered sunlight with wavelengths from about 0.4 to 0.8 µm. In particular the degree of linear polarization of this light is known to be very sensitive to the microphysical properties (size, shape, composition) of the scattering particles. SPEX has originally been designed as Mars orbiter payload, with 7 fixed downward- and 2 fixed limb-viewing apertures.

The Global Trajectory Optimization Competition (GTOC) is a competition that seeks to stimulate improvements in the fields of mission analysis and optimization. This research focuses on the problem posed in the second version of the competition, GTOC2. The problem concerns an asteroid tour mission, where the asteroids are subdivided into 4 groups, and one asteroid from each group needs to be visited. Minimization of final mass over time of flight is sought. Previous attempts of the Mission Analysis Department at the Faculty of Aerospace Engineering to solve this problem using conventional methods were unsuccessful. The goal of this research is to investigate alternative methods to model and solve the GTOC2 problem. The assignment has been divided into three consecutive parts. The first part concerns the asteroid selection and sequencing, the second part assesses the phasing characteristics of the selected sequences and the last part concerns the exact trajectory determination. For the asteroids selection and sequencing two methods will be investigated, being the Nearest Neighbor Heuristic and the Branch- and Bound Method. Both methods will be applied to three different search spaces. For the second part, an orbit model based on exponential sinusoids will be developed. Each part of the trajectory between two different bodies will be modeled by an individual exponential sinusoid. Unfortunately, the timeframe of this research and the quality of the results of the second step did not permit an analysis of the third part of the assignment. Nevertheless, an advice can be formulated for future analysis of the GTOC2 problem. Although the best results are obtained by the Nearest Neighbor Heuristic, results also indicate that an improved Branch- and Bound Method will most likely provide better results. The continuous solution method, based on exponential sinusoids, proved to be unable to perform an accurate phasing assessment of a given asteroid sequence.

Europa, the smallest of the Galilean satellites, is characterized by a young water-ice surface showing a rich structure of cracks, ridges and chaotic terrain. Below the ice shell a subsurface water-ocean might exist, which in Europa’s case would be in direct contact with the silicate mantle. This direct contact is important from an astrobiological perspective, as living organisms might have developed at places where a hot part of the mantle contacts the ocean. As a consequence, many scientists consider Europa as the planetary body with the largest probability to harbor life within our solar system. It is therefore of utmost importance to find out whether a subsurface ocean exists underneath the ice shell. One plausible method to determine the existence of an ocean is based on the measurement of the radial deformation induced by the eccentricity-tide acting on Europa. If an ocean is present in the interior, the radial deformation at the surface will be one to two orders of magnitude larger than in the case that an ocean is absent. Such a large difference in deformation can be detected from measurements performed by a dedicated orbiter. As a result, a mission to the Jovian system would be required. An alternative method to determine whether there is a subsurface ocean underneath the ice shell is based on establishing a connection between the shape of the lineaments observed on the surface and the acting stress fields. Stresses at the surface might be induced by several different mechanisms, from which only the two most important will be analyzed in this research: the eccentricity-tide acting on Europa and non-synchronous rotation (NSR) of the ice shell. The first mechanism, i.e. the eccentricity-tide, induces a highly variable stress field that explains the formation of cycloidal features even without taking into account NSR stresses. The second mechanism, i.e. NSR, induces a nearly static stress field that explains the formation of slightly-curved or global lineaments. As both types of lineaments exist on Europa’s surface, the strength of NSR stresses should have changed throughout the geological history of Europa. Such a change can be driven by a variable rate of non-synchronous rotation, which can be the result of thickness variations in the ice shell. One important result obtained in this research is that tensile stresses at the surface of models without a subsurface ocean are too small to originate a crack at the surface if NSR is not taken into account. If NSR stresses are added to the modeling of the stress field, tensile stresses only become large enough to break ice when the stress field is practically static. As a result, the existence of cycloidal features strongly suggests the existence of a subsurface ocean underneath the ice shell of Europa.

The steadily growing amount of orbital debris increases the probability and amount of collisions between two objects in orbit about the Earth. These collisions in turn create even more debris, and it is therefore important to keep track of future conjunctions. The U.S. Space Surveillance Network (SSN) uses ground- and space-based sensors to observe and track objects of about 10 cm and larger, of which the orbital information is coded in Two-Line Element (TLE) sets and listed in a catalog currently containing about 20,000 objects, which is partly distributed to the public. Using this TLE data, the Simplified General (and Deep-space) Perturbations 4 (SGP4/SDP4) analytical propagator is used to propagate the orbits of these objects, and includes secular, long-period and short-period perturbations due to the Earth's gravity field including J2, J3 and J4 and resonance effects for 12- and 24-hr orbits, as well as perturbations due to atmospheric drag, solar radiation, and gravitational attraction of the Sun and the Moon. The propagated orbits are used to predict conjunctions of pairs of objects. However, due to the large and increasing amount of objects in the catalog, numerically analysing all pairs would be too time-consuming. Therefore, numerous fast filters and sieves were designed to limit the search space of conjunction analysis, by discarding object pairs that are proven to never be able to conjunct. Four implementations of the classical perigee-apogee filter, next to six sieves with a new fine conjunction detection method, were analysed, implemented, and tested in terms of performance. The filter makes use of the altitude band of an object, and can be applied pre-hand. A method based on minimum and maximum radius determined from ephemerides, was found to be the most accurate and reliable in long-term application, while being able to be fine-tuned to the performance needs of a conjunction analysis process. Increasingly complex sieves are subsequently applied to the ephemerides at each time instance in a time interval, in order to efficiently discard object pairs. Eight possible improvements to the underlying theory and application of these sieves were made, resulting in one optimal combination of these improvements, and eventually resulting in a conjunction analysis system that is almost three times as fast as the best found reference method.

The Cryosat-2 mission is an European Space Agency mission with the main objective to measure and monitor the variation in sea-ice and main ice sheets on Earth, located at Greenland and Antarctic. The satellite's main instrument is a state of the art altimeter, which can measure the distance between the satellite and the ice surface with extreme accuracy. To extract the variation in height of the ice surfaces from these measurements, it is important to know where the satellite was at the time of the measurements. The method that is used to obtain this is called Precise Orbit Determination (POD). POD combines accurate measurements together with physical models related to the satellite, in such a way that it can determine accurate orbit solutions for the satellite. This report is a discussion about the validation and improvement of the orbit determination of the Cryosat-2 mission. The current orbit determination process that is available is validated on three different subjects. First, a new data format for Satellite Laser Ranging (SLR) measurements was investigated, which is called the Consolidated Ranging Data format. SLR measurements are accurate range measurements that are used to validate the orbit of Cryosat-2. A converter is written that could convert the new data format, in such a way that the current orbit determination software could use it. Several features of the new data format were examined, for potential improvements. The second investigation was about the generation of the Doris beacon coordinates and their effect on the POD. The Doris system is the main orbit determination system onboard the satellite. Due to different abrupt and slow motions of the Earth surface, the Doris beacons moves with respect to the defined reference frame. A new definition of the Doris beacon coordinates, called DPOD2008, is used in the Cryosat-2 orbit estimation. The solution of the coordinates sets were validated and other coordinate sets were used, to conclude that DPOD2008 was generating the best orbit solution for Cryosat-2. The final investigation was on the solar radiation pressure modeling of Cryosat-2. Currently, a 6-panel box model is used as defined by ESA. In the empirical residual accelerations a signal is visible which has correlation with the solar radiation pressure. It was decided to use micro models generated by the University College London at the research group of Prof. Marek Ziebart for the solar radiation pressure computations. Two micro models are constructed, although the implementation of the UCL models in the GEODYN software is delivering unsuspected results. Several bugs are fixed or bypassed in the implementation of the models, but still clear differences between the internal GEODYN model and the UCL model implementation are visible. Recommendations are given for further research in the micro model investigation for the solar radiation pressure computation.

Space debris is becoming an increasing problem requiring a low cost propulsion system capable of disposing non operable satellites and spent rocket stages. To this end electrodynamic tethers have been proposed as a viable alternative to conventional propulsion. Aside from a de-orbit system electrodynamic tethers have a wide range of applications as a power generator or propulsion system being ideally suited to orbital environments having strong magnetic fields for example Jovian orbits. Previous experiments performed in the development of electrodynamic tethers used insulated wire tethers and separate current collection devices, either large passive spheres or active plasma contactors. The subject of this thesis is to design and size an electrodynamic tether experiment for use onboard the Delfi-1 University satellite. The main objectives of the experiment are to provide a proof of concept of bare electrodynamic tether propulsion and deploy and operate a tape tether. A tape tether design has been selected having a favourable geometry for current collection and survivability in the micrometeoroid and orbital debris environment in comparison with ‘traditional’ wire tethers. The tape tether also has favourable drag properties allowing for a rapid deorbit when cut from the main satellite decreasing the risk of collision with operational satellites. A secondary objective for the experiment is to determine if the release of a neutral gas can enhance the current of the tether and the deorbit performance. This experiment is fundamental in the development of a fully passive deboost capability for tether equipped systems. The limited amount of storage and mass, power and data rate available when using a micro-satellite platform for the experiment combined with low complexity and cost requirements inherent to the Delfi-1 project drives the design to a bare minimum required for performing the primary scientific objectives. The baseline design for the experiment is characterised by a passive bare floating electrodynamic tether deployed in nadir direction using a passive deployment mechanism.

55 years have passed since the launch of the first man-made space vehicle, Sputnik 1. While spaceflight has changed significantly in many ways, further improvement in rocket technology is still being pursued. Launch-assist systems, which give the launch vehicle a combination of initial velocity and initial altitude, could create such an improvement. There are two factors that come into play when determining if this might be true. Firstly, the cost of building and operating both the assist platform and the launch vehicle and secondly, the total system performance. This thesis focuses on the performance of the entire system. The definition of the performance of a system utilizing the assisted launch technique was chosen to be the payload mass-to-initial mass ratio of the launcher. This was maximized by optimizing various parameters of the launch vehicle and the launch trajectory using a differential evolution algorithm. Optimization was performed for launchers using either kerolox propellant or hydrolox propellant using a variety of initial altitude and initial velocity combinations. To show continuous trends of the vehicle design parameters through the whole range of simulated launchers, all launchers consist of only one stage. The end result of the thesis provides insight of the relationships between the performance, the launch vehicle design, the trajectory profile and the magnitude and type of the assist. It also provides a comparison between the performance of launchers using low specific impulse and high density propellant, such as kerolox, and high specific impulse and low density propellant, such as hydrolox. The results are intended to be used as a tool to base design decisions on during future concept studies.

There is a renewed interest, expressed by ISECG, to return to the Moon and establish a longer human presence on the lunar surface. Long term missions require supplies, which can be transported to the Moon by a new promising solution: the Solar Electric Propulsion (SEP) stage. The goal of this study is to create a conceptual system design (phase 0) of such a SEP stage for Earth-Moon cargo transfer missions. The design is driven by the propulsion and electric power system, which are the main focus of the study, although all other subsystems are also covered to obtain a holistic system design. Designing a SEP stage is a multi-disciplinary task, in which the trajectory analysis, propulsion system and electric power system are tightly coupled. A mission analysis program is created for simulation of the spiral transfer, while flexible design tools are developed to create the conceptual design. The tools provide the possibility to quickly evaluate a different mission scenario, such that the most suitable scenario, in consultation with the customer, can be selected. Aconceptual design is created that meets the mission objectives and requirements. During the design phase it was identified that the concept of a solar electric propulsion stage, accommodates some critical issues and technological challenges. Especially the high power demand, leads to the usage of highly conceptu l power conditioning techniques, which still have to be proven in space, possibly by a precursor mission. It was also found that the SEP stage, compared to a chemical rocket, is capable of transporting 32% more payload, while having the same initial mass. This number can even increase to 90% in case a different mission scenario, with a longer transfer duration, is selected.

In recent years, space missions to asteroids have been a focus of research. This type of mission has introduced a new challenge in space mission analysis: target selection. Indeed, given the large number of small Solar System bodies and their various orbital and physical characteristics, selecting a specific target is a rather complex task. This problem of target selection was given to the participants of the second Global Trajectory Optimization Competition (GTOC2) in 2006. The GTOC2 problem relies on the design of a low-thrust mission with four consecutive asteroid rendezvous. The main challenge in finding good solutions for GTOC2 stems from the vast search space and stringent constraints set out in the problem description. At the time, a TUDelft-led team of students participated in GTOC2. Since then, there has been a concerted effort within the Astrodynamics & Space Missions (A&S) research group to further investigate feasible solutions and develop new methods to achieve better results. Building on the lessons learned from these previous efforts, we design and implement an algorithm aiming at reducing the pool of candidate targets in a computationally fast manner. This is achieved by decoupling the integer and continuous aspects of the problem. The combinatorial problem is modelled as a static, single-source Shortest Path Problem, tackled with a multi-path greedy algorithm. The continuous aspect is optimized on a leg-per-leg basis with two different optimization techniques. With computational efficiency in mind, the individual transfers are modelled as bi-impulsive, high- thrust arcs, computed with an efficient Lambert solver. The resulting combinatorial tool builds on and contributes to the existing capabilities of Tudat, an Astrodynamics toolbox developed within the A&S Department. The greedy searches in the complete asteroid pool are performed with different parameters and different directions, the sensitivity to the bounds of the continuous optimization problem is analyzed, two different costs functions are investigated. The computational times achieved with the designed greedy algorithm are very attractive. For the largest time intervals considered, where the impact of phasing is reduced, the sequence corresponding to the GTOC2 runner-up is present within a reduced set of asteroid combinations. This points to a fast and accurate pruning of the initial integer search space.

Using satellite gravity data between February 2003 and January 2008, we examine changes in Greenland's mass distribution on a regional scale. During this period, Greenland lost mass at a mean rate of 179 ± 25 Gt/yr, equivalent to a global mean sea level change of 0.5 ± 0.1 mm/yr. Rates increase over time, suggesting an acceleration of the mass loss, driven by mass loss during summer. The largest mass losses occurred along the southeastern and northwestern coast in the summers of 2005 and 2007, when the ice sheet lost 279 Gt and 328 Gt of ice respectively within 2 months. In 2007, a strong mass loss is observed during summer at elevations above 2000 m, for the first time since the start of the observations.

The island of Sulawesi, eastern Indonesia, is located within the triple junction of the Australian, Philippine, and Sunda plates and accommodates the convergence of continental fragments with the Sunda margin. We quantify the kinematics of Sulawesi by modeling GPS velocities and earthquake slip vectors as a combination of rigid block rotations and elastic deformation around faults. We find that the deformation can be reasonably described by a small number of rapidly rotating crustal blocks. Relative to the Sunda Plate, the southwestern part of Sulawesi (Makassar Block) rotates anticlockwise at ∼1.4°/Myr. The northeastern part of Sulawesi, the Bangai‐Sula domain, comprises three blocks: the central North Sula Block moves toward the NNW and rotates clockwise at ∼2.5°/Myr, the northeastern Manado Block rotates clockwise at ∼3°/Myr about a nearby axis, and East Sulawesi is pinched between the North Sula and Makassar blocks. Along the boundary between the Makassar Block and the Sunda Plate, GPS measurements suggest that the trench accommodates ∼15 mm/yr of slip within the Makassar Strait with current elastic strain accumulation. The tectonic boundary between North Sula and Manado blocks is the Gorontalo Fault, moving right laterally at about 11 mm/yr and accumulating elastic strain. The 42 mm/yr relative motion between North Sula and Makassar blocks is accommodated on the Palu‐Koro left‐lateral strike‐slip fault zone. The data also indicate a pull‐apart structure in Palu area, where the fault shows a transtensive motion and may have a complex geometry involving several active strands. Sulawesi provides a primary example of how collision can be accommodated by crustal block rotation instead of mountain building.

One of the major problems one has to deal with when working with Gravity Recovery and Climate Experiment (GRACE) data is the increasing error spectrum at higher degrees in the provided Stokes coefficients, appearing as unphysical North‐South striping patterns in the maps of equivalent water height (EWH). This phenomenon is commonly suppressed by application of a Gaussian smoothing filter, which unfortunately causes loss of signal and leakage between basins. In this paper we show how a significant amount of the striping can be removed by making use of the temporal characteristics of the error spectrum. The Stokes coefficients are decomposed using empirical orthogonal function analysis and the individual modes are tested for temporal noisiness. After filtering, maps of EWH are largely free of striping. Tests on simulated EWH estimates show that our filtering technique has a marginal effect on the predicted geophysical signal.

With the help of four years (2002–2005) of CHAMP accelerometer data we have investigated the dependence of low and mid latitude thermospheric density on the merging electric field, Em, during major magnetic storms. Altogether 30 intensive storm events (Dstmin <−100 nT) are chosen for a statistical study. In order to achieve a good correlation Em is preconditioned. Contrary to general opinion, Em has to be applied without saturation effect in order to obtain good results for magnetic storms of all activity levels. The memory effect of the thermosphere is accounted for by a weighted integration of Em over the past 3 h. In addition, a lag time of the mass density response to solar wind input of 0 to 4.5 h depending on latitude and local time is considered. A linear model using the preconditioned Em as main controlling parameter for predicting mass density changes during magnetic storms is developed: ρ =0.5Em+ρamb, where ρamb is based on the mean density during the quiet day before the storm. We show that this simple relation predicts all storm-induced mass density variations at CHAMP altitude fairly well especially if orbital averages are considered.

At the Delft University of Technology, the Delfi-C3 3-unit Cube Sat was finalised for launch on PSLV C9. An overview is presented about manufacturing, assembly, integration and verification activities on subsystems and Proto Flight Model (PFM) until delivery for launch. Delfi-C3 is the first Dutch university satellite and the fourth Dutch satellite. It was designed and built by the faculties of Aerospace Engineering and Electrical Engineering, Mathematics and Computer Science. Industry partners provided payloads: Autonomous Wireless Sun Sensors and Thin Film Solar Cells. Attention is paid to the experiences and lessons learned in this student project. The test plan and activities may serve as a valuable scheme for the next university satellite and other cube sats. During the hardware phase of the project, firstly development tests were done (e.g. deployment mechanisms, antenna pattern), followed by functional performance tests of subsystems (electronics, measurement systems, payload interface electronics, power system, communications system and attitude control behaviour). The complete satellite was subjected to a test program, consisting of vibration-, thermal vacuum-, deployment-, communication- and performance tests. With regular health and functional checks, the satellite status was verified for changes during testing.

The Delfi-C3 nanosatellite successor, Delfi-n3Xt, is currently under development at Delft University of Technology and scheduled for launch in the first half of 2010. This improved three-unit CubeSat platform allows novel technology qualification for future small satellites and innovative scientific research. The platform is improved by implementing a high-speed downlink, three-axis stabilization and a single-point-failure free implementation of batteries in the electrical power subsystem. Apart from giving a description of the three main advancements, this paper also gives an overview of the five payloads: A microsystems technology based cool gas micropropulsion module, a multifunctional particle spectrometer, a set of hydrogenated amorphous silicon solar cells, an efficient transceiver module for nanosatellites, and a memory unit based on low-cost commercial grade ash memory cards with innovative radiation protection electronics. Although the accommodation of the five payloads in a nanosatellite of only (10 x 10 x 34) cm is ambitious, this paper shows feasibility and proves that nanosatellites are powerful instruments for the qualification of novel technology and innovative scientific research.