The startup Rocket Factory Augsburg (RFA) is developing the RFA ONE orbital launch vehicle. A significant hurdle for its closed-loop guidance is the need to fully deplete the third stage for insertion during the first flight, which results in lengthy burn-arcs and complexity not managed properly by traditional guidance algorithms. Additionally, new testing methods introduce uncertainty and component failure risks, demanding an adaptable guidance system.This thesis therefore aims to enhance RFA's closed-loop guidance system to ensure optimal orbit insertion in all feasible scenarios. This was achieved by redesigning the closed-loop guidance system to make it more robust for extended thrust-arcs.Furthermore an in-flight estimation algorithm for the key performance parameters was implemented. Finally the guidance can now fall back to alternate target orbits in case of insufficient performance margins. It was shown that the new guidance is able to successfully complete the maiden flight with the large amount of expected uncertainty.

In pursuit of cost-effective and sustainable space access, Dawn Aerospace is pioneering innovative launch systems. Their focus lies on a semi-reusable, two-stage-to-orbit launch vehicle, designed for horizontal take-off and landing. While current launch vehicles rely solely on rocket propulsion, the potential of airbreathing engines on the first stage is explored. Airbreathing engines offer increased fuel efficiency and utilize the incoming airflow as oxidiser. This study delves into the feasibility of integrating airbreathing propulsion into Dawn Aerospace's Mk-III vehicle, with a goal to reduce gross take-off mass while meeting mission requirements. Airbreathing engine types are evaluated, with ramjets and turbine engines emerging as primary candidates. Detailed design processes, including vehicle modeling, trajectory analysis, and optimization, are employed. The study reveals the potential of the ramjet concept which could have a lower gross take-off mass compared to fully rocket powered designs. Nevertheless, the application requires an improved vehicle design to obtain a feasible design.

The Deployable Space Telescope (DST) is a telescope that uses foldable optical elements to reach state-of-the-art ground resolutions while minimising the launch volume and total mass of the system. One of the promising concepts that is part of the DST design is the Compliant Rolling-Element (CORE) Hinge. This hinge uses compliant links to let two semi-circular cams roll over each other with minimal friction. For this thesis, a thermo-mechanical analysis is conducted of the DST CORE hinges to determine whether the current design complies with the stability and in-orbit drift requirements of the DST. From this analysis, it can be concluded that this is currently not the case, but that a simple solution is at hand. By elongating the baffle of the DST, which protects the DST from solar irradiation, more solar radiation is blocked and the stability and in-orbit drift are kept within limits.

The demand of small payload launch vehicles has been growing over the past years. Customers base their selection of launch vehicles on cost-effectiveness, flexibility, availability and reliability. A new launch vehicle, the Dawn Aerospace Mk-III, is proposed to be developed, while designed to take into account all these criteria. For flexible and frequently available operations the launch trajectory is integrated in the common airspace. Manoeuvrability is identified as a key capability for safe operations in the common airspace. For high manoeuvrability the first stage is designed as a rocket propelled airplane, a so-called 'spaceplane'. The expendable upper stage is stored internally. After payload injection the first stage returns to the spaceport of take-off. This means the System is a Two Stage to Orbit semi-Reusable Launch Vehicle integrated in the common airspace. Cost-effectiveness is a primary selection criterium in the decision making of customers, which is why cost is included from an early stage in the development. This study shows that for identifying the cost gradient in the design space the total dry mass of the vehicle is sufficient. In this way cost optimality is included, although the Cost per Flight cannot be determined. What is determined is the effect of different technical and operational considerations. Taking into account qualitative cost differences, a selection of cost derived mission scenarios are studied. This includes different Return to Launch Site methods, first stage engine design and lay-out, the prohibition of fairing usage and integrated landing gear for take-off and landing. To analyse and optimize the different designs a Multidisciplinary Design Optimization tool is developed. This tool optimizes the vehicle and the ascent trajectory simultaneously to determine the lowest total dry mass solution meeting all requirements and constraints. To estimate the aerodynamic performance of the first stage the X-34 Advanced Technology Demonstrator is used as a reference vehicle. This means the geometry of the first stage is not optimized, while the size is. The upper stage is modelled as a conventional upper stage, of which the size and geometry is optimized. The study shows that the proposed design is feasible, meeting all requirements and constraints. The result is a vehicle with a total dry mass of 6273.0 kg, Gross Take-Off Weight of 42972.8 kg and a first stage length of 19.4 m. Of the total dry mass 94.5% is reusable. The return of the first stage is driving the trajectory design, as a steep ascent trajectory is required for limiting the downrange of the first stage. This results in 30%-50% more gravity loss in the System, which demands for a high ΔV performance. Due to the size of the first stage and the propellant required for returning, the first stage ascent ΔV is limited. For that reason, the upper stage design has a Propellant Mass Fraction of 0.939, increasing the upper stage ascent ΔV performance. Achieving such a Propellant Mass Fraction is possible, but challenging. The upper stage design is identified as a key element in the System performance for successfully meeting all mission requirements. Three different Return to Launch Site methods are compared. Two methods, in-plane pitch over and aeroturn, are active which requires return propellant. The third method, glideback, is passive. The study shows that glideback can be favourable. The total dry mass increase is 4.4% when compared to in-plane pitch over. The increase in heat load is ~16%, but the heat load in this study is limited with a total heat load of ~2.0 MJ/m^2. However, for glideback an even higher upper stage Propellant Mass Fraction of 0.946 is required. This means that the result of the upper stage design determines if this return method is feasible. Using a shared engine design on the first and upper stage shows promising results. Using a single first stage engine reduces the first stage dry mass by 4.0%. However, using a shared engine design is expected to decrease the development cost drastically, due to the reduced size of individual engines. The prohibition of fairing usage shows a stronger effect on the result. Allowing the use of a fairing decreases the first stage dry mass by a maximum of 9.6%. Fairing usage on the other hand harms the safe operations in the common airspace as the ejection of uncontrolled material requires large safety zones. The penalty on mass is acceptable for allowing integration in the common airspace. A landing gear sized for take-off results in a heavier vehicle. When the first stage is supported by a cart during take-off the first stage dry mass is decreased by 24.6%. When the first stage is air-launched the first stage dry mass decreases by 33.3%.

Single-Stage-To-Orbit (SSTO) launch vehicles are capable of reaching orbit while being potentially fully reusable. A technology capable of enabling SSTO access to space is the air-breathing engine, which has a high specific impulse and low thrust-over-weight (T/W) ratio compared to conventional rocket engines. This necessitates the use of a Horizontal Take-off and Horizontal Landing (HTHL) launch vehicle with a lifting surface, which enables banking maneuvers that can manipulate the orbital Right Ascension of the Ascending Node (RAAN) that the space plane achieves. This means that the space plane is capable of advancing or delaying its time of launch, while still achieving a similar orbital RAAN. This effectively increases the launch window even if a precise orbit insertion is necessary. This thesis answers the question what the fuel-optimal ascent trajectory is for a space plane that would extend its launch window. In order to answer this question, the full six DoF equations of motion have been defined to simulate the ascent trajectory of a space plane. Additionally, a robust Guidance and Control system is developed that can use the nonlinear translational and rotational equations of motion. For several launch times, ranging from two hours earlier to two hours later than the nominal launch time, the ascent trajectory has been optimized that included a banking maneuver. It was found that at the cost of propellant, it is possible to have a launch time delayed are advanced by as much as two hours, which means that the space plane can manipulate the RAAN by ± 30 degrees. The extra propellant cost can be related to increased drag losses during the banking maneuver.

The Mars Ascent Vehicle will bring samples back from the surface of Mars to its low orbit in 2031. Its preliminary design has been performed by NASA, defining it as a two-stage solid propellant rocket. However, a research gap exists in finding what the optimum solid rocket motors geometries are to bring the Martian samples to a target orbit between 300km and 375km above Mars. To this extent, the ascent and propellant burn have been modelled, allowing for variation in the geometry of the motors, but also in the launch and separation angles, and in thrust vectoring control. This model has been incorporated into an optimisation, from which a minimum lift-off mass of 343kg was found, 57kg less than required. In the process, 18 thousand different solutions have been found that satisfy all requirements, would the priority shift from the mass of the vehicle to for instance the burn time, or the final altitude.

Detection of the openings in the Arctic sea ice pack, or leads, allow to sample instantaneous sea surface height (SSH) and this information is crucial for quantifying the impact of sea ice melting. It is therefore important to correctly detect as many leads as possible to obtain more SSH references. This paper studies 12 different classification methods including supervised-, unsupervised machine learning methods and thresholding method, being applied to the Sentinel-3 Synthetic Aperture Radar (SAR) altimetry data collected in March/April of 2017-2020 and June/July of 2020 from areas all across the Arctic Ocean. These are compared and assessed with respect to images taken by Ocean and Land Color Instrument (OLCI), also on board Sentinel-3, ensuring a perfect temporal alignment between the two measurements. The supervised Adaptive Boosting, Artificial Neural Network and Linear Discriminant classifiers showed excellent and robust results in March/April with overall accuracies up to 91.82%. The unsupervised K-medoid classifier produced excellent results achieving up to 91.51% accuracy and it is an attractive classifier as it does not require ground truth data. The classifiers perform poorly in the summer months, as sea ice returns show more ambiguous reflections due to melting. Therefore on summer data, classifications that are solely based on waveform data from SAR altimetry is unsuitable and auxiliary information is required. Furthermore, this paper attempted to identify off-nadir leads (ONL) by adding an extra class in supervised learning methods, intending to reduce the falsely detected leads. Most classifiers failed to detect leads and did not improve their false lead rate. However, as RUS Boost classifier was able to identify 61.6% of total ONLs, this can be used to initially reject these points for more conservative lead detection.

This study addresses the challenge of accurately modeling the multi-regime plasma flow through Ambipolar Plasma Thrusters (APTs), a type of Electric Propulsion (EP) system. Despite their advantages, understanding the behavior of plasma within APTs is complex due to the presence of two different flow regimes: the fluid regime inside the thruster’s source chamber and the kinetic regime inside the thruster’s magnetic nozzle. To enhance the precision of APT modeling, this thesis introduces a novel coupling methodology between fluidic and kinetic solvers, resulting in the MUlti-regime Plasma Equilibrium Transport Solver (MUPETS). MUPETS employs two models, the continuum OpenFOAM code for the fluidic regime and the Particlein- Cell (PIC) Starfish code for the kinetic regime, coupled through a closed-loop iterative coupling scheme. This approach allows for a self-consistent prediction of plasma transport within the entire thruster domain, including the interface between the numerical domains of each model. At this interface, which often coincides with the thruster outlet, the model’s iterative loop self-corrects the interfacing boundary conditions of each numerical domain, removing imposed boundary conditions such as assumed velocities. This has reduced plasma species discontinuities across the models’ interface to less than 4% for electrons and less than 1% for ions. The performance of MUPETS was validated against experimental measures from a laboratory thruster at the University of Padova, showing less than a 19% difference in predicted propulsive performance. The MUPETS code requires minimal alteration of the separate solvers, allowing for the fluid or kinetic solvers to be swapped out with higher-fidelity models or numerically better performing models as required. This flexibility enables the comparison of previously developed and validated models from literature to investigate their suitability for describing the plasma flow across the regime change. The study concludes that the developed coupling method presents an improvement to the state of multiregime plasma flow modeling while providing similar accuracy when used for predicting the propulsive performance of APTs.

The last 20 years has brought with it a surge in efforts toward the Red Planet as the next frontier is human space exploration draws closer. Many concepts have been proposed for a sustained human settlement on Mars, with NASA’s ISRU-to-the-wall campaign identifying the need for a shuttle vehicle between theMartian surface and a station in an orbit around the planet. Two such vehicle concepts have been designed: the Charon by Gaffarel et. al. and Hercules by Komar et. al.. However a Multidisciplinary Design Optimisation (MDO) has thus far not been applied. In this research an MDO is employed for the same mission scenario as the Charon vehicle’s. The vehicle must transport 1200 kg, including 6 crew, to a 607 km circular orbit at 44.96± inclination from the Martian base that is located at 42.5± North and 25.5± East. It must then return to the base, its entry beginning at 80 km altitude at a velocity of 3500 m/s. The MDO in this research is performed by dividing the design of the vehicle and its trajectories into various disciplines, which are optimised in parallel. In reality cost commonly is the dominant factor that drives the design, in this case, the vehicle’s Gross Take OffWeight (GTOW) is taken as the objective. Estimating the cost of the Martian shuttle vehicle within reasonable accuracy is exceedingly difficult, as not only is its realisation still decades away, costs such as the shuttle’s transportation to Mars, its operational costs, and its maintenance costs are very hard to estimate. However, a vehicle’s GTOW is directly influential on its cost, and its reduction is therefore the optimisation’s goal. The disciplines within the optimisation are the vehicle’s geometry, mass, aerodynamics, and ascent and descent trajectories. Constraints are set for both the vehicle and its trajectories, to which the design must adhere, and the performance of the design is determined by a fitness function that ensures the reduction of the vehicle’s GTOW. The optimiser itself is written using Tudat software, a set of C++ libraries developed by the TU Delft. As the same mission scenario is taken for the vehicle in this research as that of the Charon vehicle, the Charon design can be directly compared to that of the optimiser. TheMDO is able to significantly reduce the vehicle GTOW, obtaining an optimum solution of 146.2 tonnes, which is more than 20 tonnes lighter than Charon, at 168.1 tonnes. The greatest reduction in mass is found in the ascent propellant mass, which is the greatest contributor to the GTOW. This reduction is mostly due to the lower maximum Thrust to Weight (TW) ratio used in the MDO. Other scenarios are also investigated and their effects observed. Two other target orbits and rendezvous strategies are tested, namely the same as the Hercules vehicle (108 km pericentre altitude and an eccentricity of 0.0178), and a direct ascent to the orbital node at a circular 500 km altitude orbit. The Hercules vehicle scenario proved to be by far the most GTOW-preferable, with a GTOW of only 103.0 tonnes. The GTOW of the MDO solution found for the Hercules scenario is also less than that of the Hercules vehicle design, which is 162.8 tonnes, however the Hercules vehicle transports a payload mass of 5750 kg as opposed to the MDO’s 1200 kg, therefore the mass reductions cannot be solely attributed to the optimiser performance. It is clear from the results that the altitude of the initial target orbit is the greatest factor contributing to the GTOW, with a reduction in both payload mass andMartian base latitude also shown to reduce the GTOW. Neither an increase in the maximum acceleration constraint, nor a change in ascent thrust profile, were shown to have any benefit on the GTOW. The sensitivity of the optimum design with respect to uncertainties is assessed. The final pericentre and final eccentricity are both most sensitive to the final pitch node value, especially when taking interference with other variables into account. The inclination, however, was almost solely influenced by the first pitch angle value. The latitude and longitude are also almost solely influenced by a single variable, namely the flight path angle, and the final velocity and final pitch angle are both highly volatile to all variables when interference effects are taken into account. TheMDO model as a whole is found to be sensitive to changes in both aerodynamics coefficients and propulsive efficiency; an increase in aerodynamic coefficients adversely affects the GTOW, and an increase in propulsive efficiency benefits the GTOW, and vice versa.

Bottom-up thermal-mechanical design and analysis of a deployable baffle, which serves as a deployable thermal control system for the Deployable Space Telescope. Research being conducted on how to create a stable thermal environment for the mirror support structures so that in-orbit thermo-elastic drift budgets will be met, as well as designing and analysing a feasible deployment system.

With the increase of satellite constellations in orbit, the need for a Space Traffic Management (STM) is becoming undeniable. Within the scope of STM, an impartial and optimized collision avoidance advice based on the perspective of two manoeuvrable satellites could encourage more active collision avoidance and pre- vent on-orbit collisions. Using CDM data and operator input for the setup of the optimization, it was shown that in-track manoeuvres performed by either one or both of the satellites have a minimum usage of ∆V in combination with a small displacement and meet the collision probability threshold. Additionally, it was shown that the use of weights in the optimization cost function can influence the manoeuvre magnitude and the respective satellite that needs to perform the manoeuvre. Combining this optimization with a linearized propagation of the satellite states, a quick recommendation can be generated on the manoeuvre options for both satellites, which could help the development of an STM in terms of an impartial collision avoidance manoeuvre advice system.

Aerosols, also known as particulate matter, are tiny particles or droplets suspended in the atmosphere which interact with solar radiation. These particles are partly responsible for cloud formation and are detected amongst others in (forest fire) smoke, dust, volcanic ash or sulfates. Aerosols have a large influence on the Earth’s climate as well as on air quality. Currently, SPEXone is one of the most promising instruments considering aerosol research due to the versatile nature of measurement methods based on multiple-angle polarization observations. Therefore, the SPEXone instrument will be implemented aboard the PACE satellite, designed primarily for aerosol observations for climate research of the Earth. The PACE satellite will be launched by NASA in 2025. The measurements of SPEXone provides highly relevant information on aerosol type for source apportionment. This raises the question whether SPEXone can serve as a stand-alone air quality monitoring system, thereby responding to the increasing awareness of the importance of monitoring air quality worldwide. Thus, this thesis focuses on the applicability of the SPEXone instrument regarding air quality monitoring by investigating plume polarization detection based on two relevant use cases. In this study, the plumes were simulated according to Gaussian plume behaviour in combination with a collection of observation scenarios. SPEXone measurements were simulated by using a radiative transfer model and instrument noise model to simulate the degree of linear polarization at the satellite sensor. This study found that for very optimistic emission rates, plume detection was possible for about 90 percent out of 144 measurements and between 79 and 67 percent out of 144 measurements at a wavelength of 670 nm. However, lowering the emission rate to realistic values showed that, apart from some very rare cases, no detection was possible. This leads to the conclusion that polarization signatures from the emitted aerosols in the simulated plumes are too weak in order to be detected above the polarization signature of the background. Nonetheless, due to the complexity of aerosol modelling, not all contributing factors were given the same level of detail, i.e. varying microphysical properties and the exact influence of the vertical height profile of the plume. Future research needs to focus on these aspects in order to yield a conclusive result of the applicability of the SPEXone instrument regarding air quality monitoring. Additionally, a different approach to characterize the background noise may increase the probability for plume detection.