This thesis addresses the optimization of on-demand satellite servicing in Sun-Synchronous Orbits through dynamic routing and refueling strategies. Traditional satellite operations, limited by finite onboard fuel, lead to premature mission termination. This research focuses on optimizing a refueling infrastructure comprising fuel stations, a service satellite, and client satellites, specifically for SunSynchronous Orbits. The algorithm is designed to efficiently manage fuel resources by determining the most effective servicing routes to minimize mission time and fuel consumption. A multi-objective optimization framework is developed to model these operations, utilizing a custom algorithm that refines transfer trajectories through high-precision numerical simulations that account for perturbations. Robustness analysis demonstrates that the algorithm can effectively handle uncertainties, consistently converging to the same optimal solutions across varying conditions. The developed algorithm efficiently identifies optimal servicing paths and is scalable to more complex missions involving multiple satellites and refueling stations.

Final Report Design Synthesis Exercise Spring 2024. AE3200: Design Synthesis Exercise. DSE Group 22

The design of multi-target rendezvous trajectories, which see a spacecraft approaching a sequence of objects in orbit as efficiently (by some metric) as possible, is a challenging problem of critical importance for Active Debris Removal (ADR), On-Orbit Servicing (OOS) and cis-Lunar logistics more widely. This thesis investigates two primary challenges in space Vehicle Routing Problems (VRPs): the application of Neural Combinatorial Optimization (NCO) methods for ADR missions and the integration of verifiable trajectory optimization techniques for OTV payload deployment.
The first research focus assesses the efficacy of NCO methods in designing multi-target rendezvous trajectories for ADR missions. An Attention-based routing policy, comprising a Graph Attention Network and a Pointer Network, was developed and trained using Reinforcement Learning (RL) algorithms, including REINFORCE, Advantage Actor-Critic (A2C), and Proximal Policy Optimization (PPO). Through hyperparameter analysis utilizing ANOVA, embedding dimension and the number of encoder layers were identified as critical factors influencing model performance. The trained policy was evaluated on scenarios involving 10, 30, and 50 transfers based on the Iridium 33 debris cloud. In missions with 10 transfers, the NCO policy achieved a mean optimality gap of 32%, outperforming the Dynamic RAAN Walk (DRW) heuristic in both mission cost and runtime. However, performance degraded in more complex scenarios with 30 and 50 transfers, indicating limited generalization beyond the training conditions. Grid search hyperparameter optimization revealed that while model performance improves with increased complexity, gains are marginal, and larger training datasets enhance convergence speed with only slight improvements in final performance. These findings demonstrate that NCO methods are effective for ADR missions with a limited number of targets but face scalability and generalization challenges in more complex scenarios.
The second research focus involves the design and optimization of multi-rendezvous trajectories for the UARX Space OSSIE mission using a modular framework that integrates Heuristic Combinatorial Optimization (HCO) with Sequential Convex Programming (SCP). This framework successfully determined optimal target sequences and generated near fuel-optimal trajectories for OSSIE, a translational and mass-dynamic payload delivery platform.
An Attention-based routing policy trained with RL was integrated into the combinatorial optimization process, enhancing the efficiency of mission planning. Applied to the OSSIE mission, the framework effectively explored the mission design space, optimizing 5000 mission scenarios and affirming the vehicle’s capability to fulfill advertised services. The modularity of the framework ensures adaptability to mission-specific constraints and facilitates future extensions, such as the incorporation of low-thrust propulsion profiles. Overall, this thesis confirms that NCO methods are applicable and effective for specific instances of space VRPs, particularly in optimizing ADR missions with a limited number of targets and in near-static mission scenarios where RAAN convergence is not required. The integration of verifiable trajectory optimization techniques with advanced routing policies presents a viable approach for efficient and adaptable mission planning. However, scalability and generalization remain challenges that necessitate further research. Recommendations include refining NCO model architectures to enhance scalability and generalization, exploring hybrid approaches that combine NCO with traditional heuristics, and developing automated machine learning frameworks to optimize model performance and robustness.
The project successfully achieved its primary objectives: developing and implementing heuristic and neural combinatorial optimization solvers for space VRPs, designing a modular trajectory optimization framework, and conducting comprehensive mission analyses for the OSSIE OTV. In doing so it has increased the mission design capabilities for space logistics missions at SENER Aerospace & Defence, as well as provided a strong foundation for future research and development aimed at addressing the increasing complexities of space operations.

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 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.

RLV development can be considered as the modern step towards mission design due to financial and strategic decisions. In the past, reusability has been addressed however the level of maturity of the technology, both in terms of hardware and software was not yet reached. There are several aspects to developing a RLV, and these can be categorized into optimization of the LV, optimization of the trajectory, and cost analysis. TO be able to determine the feasibility of the mission, it is not just necessary to develop a suitable configuration, but to also determine the physical feasibility of the trajectory. Several methods exist of which convex optimization is selected. This class of algorithms have risen in popularity in the regime of powered descent guidance, and present a desirable trade-off between performance and computational cost. An already existing algorithm, DESCENDO, for a two-staged vehicle CALLISTO purposed for a mission to a geo-synchronous orbit, is taken as reference. The algorithm is rewritten in YALMIP allowing it to perform more efficiently by saving computation time through creation of multiple controllers based on a discretized burn schedule. A potential candidate for reusability in the future is selected, which is a VEGA variant, considered as a two-staged vehicle with set requirements on the mission and configuration. Through closed-loop simulations, the feasibility of RTLS for a particular mission of this VEGA variant can be studied. The disciplines involved in the study include the launch vehicle optimization, engine sizing, preliminary ascent & descent, and 3-DoF simulations. Previous research at TU Delft on RLV has included the work Rozenmeijer, Vandamme, Van Kesteren, Miranda, and Contant, graduate students of the TU Delft Aerospace Engineering faculty. The work relied on the usage of the TUDAT C++ software environment and based its feasibility or reusability of operations through a cost-analysis. A shift in direction is taken away from cost-analysis to examine at a greater detail the physical feasibility of the trajectory for a nominal candidate RLV. This is done by examining the influence of simulator to guidance algorithm dynamics and guidance algorithm parameters. Moreover, a nominal payload class between 100 and 500 kg is selected to determine the configuration of the vehicle ideal for this mission. To be able to determine feasibility of RTLS, three metrics are considered, which are the final landing velocity, final landing position, and maximumdynamic pressure. The study performs higher fidelity analysis only on the return phase, and as such the starting conditions for descent are determined through a preliminary design process by considering a drag-less ascent. This returns a starting altitude of around 26-30 km, with similar values for starting downrange position, and varying conditions of initial mass and velocity. For the preliminary descent, it is found that the metric of dynamic pressure does not reach more than around 60% of the limit imposed by the VEGA-C, which is similar for other VEGA variants. This coincides with research done with the CALLISTO vehicle. All in all, the best cases for these metrics and one included as an overall best case where candidates for the 400 kg payload class. The selection criteria for best cases of the preliminary descent involved the dynamic pressure, final velocity, and required propellant mass for descent. Moreover, the vehicle optimization results showed that this contained the most variation of vehicle characteristics, and as such it was deemed as a desirable class to work with for its flexibility in design. The best case burnt mass result was selected as the best case velocity required extensive propellant mass to burn for only a less than 7 m/s difference in result, which would not be indicative of what the convex algorithm could achieve due to dynamics involved in the preliminary experiment. This nominal candidate is then tested for various variations of controller tuning parameters combinations, burn schedules, and simulator/guidance frequencies. The results showed a clear desirable region for the final time of just between 300 and 310 seconds for return, favouring shorter burns. The solution envelope for the burn schedule showed gaps in zones of feasibility as well as optimality, suggesting some performance issues with the algorithm due to it failing to find solutions. Nevertheless this envelope is well defined and several feasible solutions existed. The influence of parameter tuning and simulator frequency was studied. It was determined that no set of guidance parameters could give an advantage over the other, but that some values did favour feasibility more. This is somewhat in contrast to the selection of frequencies, as despite the fact there was also a large difference for higher frequency ratios between the Q3 to Q4 and Q0 to Q2, there is a noticeably trend that higher ratios are favoured. Moreover, a local optimal ratio of frequency of 10-1 was also found, and being the same ratio used for the other experiments as well as the CALLISTO study, provides more evidence that this effect is intended. The uncertainties studied are for the initial state variations, errors in reading of position and velocities, and process time delays. It was noted that almost all the errors in the initial state variations shared similar distributions for ranges of values of the metrics. The overall majority returned feasible as well as the large minority of this returned optimal. Of little to no significance was the processing time delay, of which the overwhelming majority returned optimal results, and the rest where outliers whose process time factor where beyond the 3σ limit imposed in the creation of normal random variables. The largest errors arose from the real-time uncertainty in the velocities and position, modelled after pseudo-range errors. Although results showed a high density in the feasible and optimal regions for metrics of final time and position, there was also a high density past the feasible regions. It can be considered that the feasibility of RTLS operations for such a VEGA vehicle is restricted by such errors as expected, but nevertheless results are promising in what can be the main steps to lead to an error analysis study by the use of state estimation techniques and testing various modifications made to the SOCP problem to improve performance.

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.

A flight termination system (FTS) is one of the most important safety features on a launch vehicle, as it allows for the vehicle's flight to be terminated in a controlled way when predetermined safety criteria or boundaries are exceeded. Currently, T-Minus Engineering B.V. aims to develop a custom flight termination system for the Dart payload stage of their DART XL sounding rocket. A sounding rocket is by definition sub-orbital, so an important safety constraint is its impact location. Based on this impact location, an advice can be formed for a range safety officer (RSO) to base flight termination decisions on. Therefore, the objective of this thesis is creating an algorithm to predict the instantaneous impact point (IIP) of the DART XL's ballistic payload stage in near real-time during a launch, considering impact probability area boundaries and requirements for the accuracy and update rate.

To meet this objective, two separate models were created. The first was a pre-flight nominal trajectory model aimed at evaluating the nominal vehicle trajectory and thereby producing inputs and verification tools to assist the development of the second model. The second model, the in-flight IIP prediction model, is intended to run during a DART XL launch to continuously predict the ballistic trajectory and IIP of the Dart stage in the post-separation phase. An important element of developing both models is determining the model definitions. The final environment model selection consists of a US76 atmospheric model, a central gravity model including the J2 effect, a local wind model and a rotating, ellipsoidal Earth model. Furthermore, using 3-degrees of freedom (3-DoF) for the IIP prediction model resulted in IIP predictions close to those predicted using a nominal trajectory model of more degrees of freedom.
Additionally, an Euler integrator with a step-size of 0.1 s was used for the IIP prediction model. With this step-size, the IIP prediction model successfully met the computation time requirement of an update rate below 1 s. For the accuracy requirement, the IIP prediction model shall only include parameters that influence the prediction in the order of 1 km. This requirement was successfully met using the step-size of 0.1 s, with a distance to a verification IIP of less than 600 m.
Lastly, the nominal impact point distribution was determined by performing Monte Carlo simulations of the pre-flight nominal trajectory model. The resulting IIP distribution was used to determine the 3-σ impact probability area, which serves as a measure of the uncertainty of the nominal impact point.

It should be noted that uncertainties remain in this nominal prediction larger than the desired 1 km accuracy radius. However, it is unlikely that the uncertainty of the nominal IIP prediction can be decreased below a radius of 1 km due to the inherent unpredictability of launch conditions. The overall conclusion is therefore that the thesis work has successfully developed a model for the near real-time instantaneous impact point prediction of the DART XL sounding rocket's Dart payload stage.

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 small satellite market is rapidly growing and becomes more competitive by the day. In order to reduce launch costs, the concept of reusable launch vehicles wins popularity. Only few companies have managed to successfully recover and re-use their launch vehicle, of which SpaceX is the only one using a powered descent technique. This study focuses on applying that same technique on a small satellite launcher: Rocket Labs Electron. A 3-DOF trajectory optimization is performed to find the fuel optimal ascent and descent trajectories of the Electron. The original design of the Electron is adhered to as much as possible. The optimization is performed for a variety of Single- and Multi objective optimizers and different settings. Although this is only a feasibility study in which many assumptions are made, the results look promising for further research to be done.

Dawn Aerospace is developing a horizontal take-off and landing two stage to orbit partially-reusable launcher concept. The re-usable first stage spaceplane operates on a return to launch site trajectory and integrates into the existing airspace, flying as an UAV. For any re-entry vehicle the design needs to account for aerothermodynamic behaviour around the vehicle, to ensure the structure can survive the re-entry temperatures. The unique mission of the Mk-III means the thermal design considerations are unique and provide a new engineering challenge and research topic.

This thesis investigates the aerothermodynamic behaviour of the Mk-III flow and structure. A loosely coupled model was created for this purpose, which couples engineering methods to predict the aerothermodynamics and the thermal behaviour of the structure. The coupling is done by transferring the external skin temperature and convective heat flux between the two simulations.

The primary research question for this thesis is “What thermal protection systems have potential to be implemented on the Dawn Aerospace Mk-III spaceplane for a range of different design trajectories.” Two thermal protection systems (TPS) and material choices have been identified as potential solutions. The first is a fully titanium structure, which can handle the temperature experienced by the Mk-III for all trajectories at every point along the vehicle. The second is a combined titanium and BMI CF structure with an insulation layer on the BMI CF. The titanium is required for the temperatures experienced on the vehicle’s leading edges, while the BMI CF has been identified as suitable in areas away from the leading edge if protected by an insulation TPS. A benefit this option produces is that the insulation layer can be changed in thickness to lighten the vehicle for lower design trajectories, therefore creating different vehicles for different trajectories. However, combining a metal with a composite could pose manufacturing issues such as cost for different manufacturing processes and joining problems. This thesis could not properly trade-off between these two options due to it being outside the scope of this thesis and due to limitations in detailed structural knowledge.

These two material choices were chosen from four materials analysed in the thesis and four different TPS. Other TPS were not suitable at decreasing the structural temperature for the Mk-III mission or had an unjustifiable weight penalty. The other material choices would have been suitable in certain situations but were heavier than the current proposed options and might have required a TPS.

This thesis provides valuable insight in what suitable material and TPS choices could be for the Mk-III mission. It also shows that for any future material and TPS choice to be made, a comprehensive structural analysis is required. Mass is a key trade-off parameter between the two proposed solutions and a better structural analysis is required for any further trade-off.

In this thesis, the achievable performance and limitations of a Synthetic Aperture Radar (SAR) seeker on an Anti-Ship Ballistic Missile (ASBM) for a set of parameters has been investigated. This was achieved by modelling the dynamics of an ASBM and the constraints related to the acquisition requirements of a maritime surface target in the "Generation and Improvement Algorithm for Nonlinear Trajectories" (GIANT) optimisation tool provided by TNO. The angular velocities of the ASBM were set as the control variables for the optimal control problem where the objectives were set to the minimisation and maximisation of the exposure time and vertical end velocity of the missile and the minimisation of the required time for creating one SAR image by the SAR seeker.

Three experiments were carried out to investigate the influences for different initial conditions and resolution constraints on the optimised trajectory. In the first experiment, the trajectory was optimised for initial reentry conditions that are representative for typical ASBMs. The target was placed at the location that would match the uncorrected ballistic flight impact location of the ASBM. This allowed to compare and visualise the optimised manoeuvres of the ASBM relative to its ballistic flight trajectory. In the second experiment, the target was placed at the same location, but the initial re-entry conditions were optimised by GIANT, that provided the optimal trajectory. In the last experiment, the constraints related to the resolution of the target were lowered to analyse the maximum performance of the system.

The experiments show that a larger initial squint angle of the re-entry position of the ASBM improves the overall performance. This is because a larger squint angle allowed the missile to obtain a larger seeker look angle during the SAR phase while performing minimal manoeuvres during the re-entry phase. This permitted the dwell time of the SAR seeker and the exposure time of the missile to be as short as possible, while the missile’s vertical end velocity could be maximised. It was also shown that a ground range and crossrange resolution of 1.60 m could be achieved for an average dwell time of 0.1170 s. For better resolutions, the bandwidth of the SAR seeker appeared to be the limiting factor of the system for the chosen set of parameters.

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.

In previous research, the necessity of non-rigid body simulators was stated to be a requirement for long, slender bodies. However, it remains unclear for which vehicle design parameters non-rigid body effects become destabilising. A set of 2+N linearly independent equations of motions can be combined in a single matrix to determine the motion of a flexible launch vehicle with N slosh masses. The flexibility of the vehicle is modelled using the method of assumed modes. The sloshing motion is modelled by assuming that the slosh mass behaves like a pendulum. To couple the different motions, a constraint matrix is used. The model is developed in Matlab Simulink. Using non-quiescent starting conditions with no external forces acting on the vehicle except gravity, it was found that the energy of the system remained constant. The flight data from the Stratos III launch vehicle was obtained and compared to the simulated non-rigid body model data. It was found that slosh motion can result in destabilisation when this vehicle experiences a sudden decrease in acceleration. The destabilisation that occurred could not be observed when rigid body equations of motion were used. Based on a sensitivity analysis over a range of vehicle parameters, it is recommended for length over diameter (L/D) ratios above 20 to consider the use of non-rigid body models. From this ratio, the maximum and total flexibility strongly increase. For L/D ratios below 20, flexibility is negligible and sloshing becomes the main perturbing factor. It was found that a PID controller that is designed using a rigid body model is still able to stabilise the non-rigid body during flight. Oscillations occurred in both the rigid and non-rigid body model when a single set of gains was used. For both models, the oscillations were removed by introducing gain scheduling.

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.

For decades the space domain has been used for military purposes. When enemy forces establish counter-space capabilities, this could become a weakness. Responsive Space is a concept that focuses on replenishing lost capabilities and/or granting additional capabilities from space assets quickly. The continuous miniaturization of technologies has enabled small satellites to fulfill many desired functions. To launch such satellites, it might be possible to modify a missile into a small launch vehicle that can be launched from a naval ship using a shipborne missile canister launching system, the Mark 41 Vertical Launching System, thereby giving the military autonomous and quick-response space launch capability. Modifying a missile that has already been integrated in the VLS instead of developing a fully new launch vehicle could have the benefit that integration cost could be lower. Also, a fully new vehicle would require more development, which can be costly and time consuming. It is investigated what the maximum payload is of a launch vehicle developed from an already integrated missile that is launched from the Mark 41 Vertical Launching System and whether it is capable of launching microsatellites. The Block IIA version of the Standard Missile-3 anti-ballistic missile is determined to be the most capable missile to modify. A multidisciplinary design optimization tool has been created that maximizes the payload mass and the perigee altitude of the orbit of the new launch vehicle. The results show that it is possible to launch a microsatellite to various Low-Earth orbits from different launch locations. For example, for an equatorial orbit, it is possible to launch a payload mass of up to 17.8 kg to a near-circular orbit with a perigee altitude of 200 km.

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%.

Oceans cover a significant part of the Earth's surface. The coupling between the upper ocean and the atmosphere is very complicated with defied theoretical understanding, while it is essential for climate studies, weather prediction, and marine ecosystems. With the advent of spaceborne Synthetic Aperture Radar (SAR) systems, surface signatures of ocean and atmospheric processes have been revealed. As winds blowing over the ocean excite the wind waves, all undulations of the ocean surface are assumed as waves in this study. The primary sources for ocean surface signatures in SAR images are waves that are created by the exertion of the local wind stress. Wind waves cause changes in the backscattered power due to three mechanisms: specular reflections, Bragg scattering, and a contribution from wave breaking. A statistical multi-static normalized radar cross-section (NRCS) background model in terms of the directional wave spectrum is developed, considering both Bragg and non-Bragg mechanisms for various polarization states. As the qualitative comparison between optical and SAR data reveals a significant correlation in sea surface signatures, a synthetic attempt is made to estimate the SAR signals from optical signatures. This is realized with the transformation of the wave spectrum in a nonuniform medium, as a consequence of surface currents, and varying near-surface wind fields. A comparison between modeled NRCS and observations is presented. This modulated NRCS model advances the quantitative interpretation of the upper ocean dynamics from satellite measurements.