With the advent of Reusable Launch Vehicles (RLVs) on the European launch market, the German Aerospace Centre (DLR)’s In-Air Capturing (IAC) recovery technique promises a competitive reduction in launch costs. Central to this IAC procedure is the Aerodynamically Controlled Capturing Device (ACCD), which is deployed from a Towing Aircraft (TA) by means of a tether. It intercepts a returning RLV, and provides a structural interface between the TA and RLV during tow-back. Once in the vicinity of a landing site, the ACCD releases the towing connection, and the RLV lands autonomously. Whereas the aerodynamic shell of the ACCD has been defined and studied in the past, an in-depth exploration of its electromechanical design space has not been performed previously. This work proposes a systems engineering approach to fill the corresponding design gap, based on the definition of key requirements and risks characterizing the ACCD’s functionalities, and its operational environment.

First, the behaviour of the ACCD as part of a larger towing system is analysed. For this, a two-dimensional (2D), quasi-steady-state Towing Model is developed, enabling a swift exploration of the ACCD’s design space, as well as a preliminary estimation of extreme tow-back loads. This Towing Model shows that the ACCD’s towing position relative to the TA is highly sensitivity to the ACCD’s Centre of Gravity (CoG); in order to stay clear of the TA’s disturbing wake zone, a maximum allowable CoG position of 1 m behind the vehicle’s nose is defined. Additionally, a significant pitch control authority is demonstrated, indicating substantial manoeuvrability. Finally, a study of extreme tow-back conditions results in the definition of a 300 kN axial towing design load - representing a 72% increase compared to previous estimates.

Next, the ACCD’s relative navigation system is studied, used to estimate the position and attitude of the RLV. A comparison between state-of-the-art systems is made, and the so-called VisNav solution is proposed - with a sensor attached to the RLV, and active beacons mounted on the ACCD. In order to identify a representative configuration for this system, a geometric VisNav Model is developed, analysing the visibility of beacons from the sensor’s Point of View (PoV). Based on comparative studies performed with this model, a design with three asymmetric rings of twelve beacons is proposed. Combined with Global Navigation Satellite System (GNSS) and Inertial Navigation System (INS) data, this configuration enables relative position and attitude estimation with an accuracy of 20 cm and 1°.

A representative electromechanical design for the ACCD is then proposed, based on results from the foregoing design space exploration. On the one hand, the selection of Commercial-off-the-Shelf (COTS) avionics is covered, as well as a preliminary dimensioning of the vehicle’s power system. Here, the ACCD’s actuators are identified to be the main drivers of its electronic design in terms of Size, Weight and Power (SWaP) footprint. Additionally, the feasibility of a battery-powered design is demonstrated. On the other hand, a Computer-Aided Design (CAD) study of mechanical subassemblies is performed, which are further analysed in terms of extreme operational loads. Most noticeably, a separation of the docking and release systems is proposed, as combining both into a single mechanism is deemed infeasible, due to the continuous presence of significant towing loads. Based on the presented design, a Bill of Materials (BoM) for the ACCD is established, yielding a total functional mass of 159.14 kg. Because of the substantial impact of the vehicle’s wings and docking system on its overall CoG, an additional 16.3 kg trim mass is required to reach the desired CoG position. Compared to previous studies, the final estimate for the ACCD’s total mass has increased by 33%.

Concluding the study, compliance with the vast majority of the postulated design requirements is demonstrated, while a number of future improvements and complementary studies are identified. These include an aerodynamic redesign and shortening of the ACCD’s shell, to eliminate the need for a trim mass. Additionally, prototyping of this unique vehicle is highly recommended, in order to further study its behaviour, and validate the theoretical design proposed in this work.

The prediction of high-frequency combustion instabilities is of relevance for liquid propellant rocket engines, to prevent risks or damages caused by overpressure and increased heat transfer to the walls. The main prediction method in rocket engines is through CFD simulations, which are computationally expensive, and require a significant set up process. This document presents a low-order tool capable of providing predictions in significantly less time than CFD simulations, with little to no setup time required by the user. The tool has the capability to predict stability, as well as computing the acoustic resonant frequencies of the system with an accuracy comparable to higher order models. The tool employs what is known as a "network model" and linear acoustics. The model includes new developments that improve its prediction capabilities with respect to the existing network model approaches present in literature. This includes acoustic elements capable of describing any isentropic flow, as well as distributed flame models that asses the stability of the combustion chambers of rocket engines. The document describes these new models in detail as well as their derivations. The new models are validated using experimental data from various rocket engines and load points tested at DLR Lampoldshausen. The validation shows that the acoustic resonant frequencies are predicted with an average error less than 5%. Furthermore, no single validation case presents an error over 10%. The tool is capable of predicting stability based on two parameters, the time lag and gain between perturbations in pressure and the flame response. Given these parameters the tool adequately predicts stability of the engines used for validation.

Beam-powered thermal propulsion, utilizing an external energy beam to heat a propellant, is proposed as a promising alternative to traditional chemical and electrical propulsion systems. This research aims to contribute to this technology’s numerical and experimental development, offering a comprehensive analysis of the receiver-absorber cavity
(RAC) performance and its optimization.
The work begins with a detailed background study of the principles of beam-powered propulsion. This sets the stage for the subsequent chapters, which investigate the numerical modelling and experimental testing of the thruster.
The numerical part of this thesis focuses on developing accurate prediction tools for the total absorbed beam power by the RAC. This information is added to a previously
developed RAC performance prediction tool and the development of a new tool considering spatial information is started. The optimization process regarding efficiency is
presented with an example case study with optimal values above 80% with a margin to improve.
The experimental component involves the design, manufacturing, and testing of a prototype beam-powered thermal thruster. Leak testing, test bench calibration and a
cold flow test are conducted. Detailed test procedures and setups for these are presented, alongside their results. The experimental findings are compared with the numerical predictions to assess the accuracy and reliability of the models.
The thesis concludes with evaluating the research questions, and discussing the successes and challenges encountered. Recommendations for future work are provided,
aiming to guide further advancements in beam-powered thermal propulsion technology. The combination of numerical and experimental insights gained through this research contributes to the feasibility and optimization of beam-powered thermal propulsion systems.

A laser-sustained plasma (LSP) generator capable of operating as a laser-thermal thruster was designed and tested. The apparatus was powered by a 3-kW 1070-nm fiber-optic laser and was operated with argon gas. The laser absorption, radiated spectrum, and change in static pressure were recorded, and high-speed video footage of the LSP was acquired. Of special interest is the heat-deposition efficiency into the working gas resulting from LSP ignition, which was estimated by tracking the pressure change in the test section. Laser transmission through the plasma was also measured to quantify two energy loss mechanisms: incomplete laser absorption and radiation of heat to the walls of the test section. Minimizing these losses would be critical in the realization of a practical and efficient laser-thermal propulsion (LTP) thruster, as most of the laser energy should be deposited as heat in the propellant to maximize thruster specific impulse.

The electric-pump cycle is a rocket engine configuration that uses an electric motor to power the pumps instead of a turbine. This offers several expected advantages such as simpler design, lower development costs, and easier restartability, but comes with reduced performance compared to conventional cycles. Previous research has primarily compared the electric-pump cycle to the gas generator cycle and has been limited to direct comparison. To extend the previous research a Rocket Cycle Analysis Tool was developed called RoCAT. It models the last-named cycles as well as the open expander cycle for a broad scope of thrusts, burn times, and chamber pressures, and several propellants. In addition, RoCAT optimizes several inputs for each cycles individually for a fairer comparison. Besides analyzing the electric-pump cycle's current performance, this research also estimates its performance in the future based on historic trends in its key technologies like the battery and electric motor.

High-frequency combustion instabilities represent the main technical risk faced by design engineers when developing new chemical rocket propulsion systems. Such instabilities are driven by the coupling between the flames’ heat release rate and the combustion chamber’s acoustic field. As such, they can only be assessed during detailed design phases, making design iterations costly. Despite constant efforts being invested in developing numerical tools for their analysis, their reliable prediction remains a major issue. Furthermore, with the advent of the New Space industry, a need to cost-effectively analyze the stability of Hydrocarbon-based systems is stressed.

As a result, this work contributes towards the development of a reliable numerical framework to predict the onset of high-frequency combustion instabilities in liquid rocket engines by ameliorating their acoustic modeling. More specifically, the performance of a novel Linearized Navier-Stokes (LNS) solver in COMSOL is benchmarked against its established Helmholtz solver and validated against the data of two hot-fires of DLR’s LOX/CH4 LUMEN engine. Following a step-wise increase in modeling complexity and by investigating how the background flow description influences the results, sub-5% eigenfrequency errors were obtained. The results also showed that the moving background flow in rocket engines significantly reduces the eigenfrequencies, and that its influence can be accounted for by prescribing varying flow properties to the acoustic solvers. Besides achieving the most accurate eigenfrequency results for a LOX/CH4 system, a more representative velocity disturbance distribution is also attained with the LNS simulation, allowing future studies to better account for the influence of velocity fluctuations on the heat release rate driving the instabilities.

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.

Contrasting experience with Reusable Launch vehicles—Space Shuttle and Falcon 9— has established that potential of reusable launch vehicles to achieve low launch costs is driven by the design choices made. Methalox propellant is one such design choice that has been touted to power future missions and potentially replace traditional propellants — Hydrolox and Kerolox. The high density of methalox compared to hydrolox and improved specific impulse compared to kerolox, potentially make methalox an ideal propellant choice. To justify any new design choice, cost-based analysis is essential, especially given the persisting issue of high launch cost. Additionally, it is essential to benchmark the performance of any new design choice with existing practices, especially when they are expected to replace current practices. For this, a cost based comparative analysis of methalox based launchers with hydrolox and kerolox launchers is performed using a tool capable of launch vehicle design and cost analysis. Rather than designing a tool from scratch, existing First Stage Recovery Tool (FRT), which was developed by M. Rozemeijer to modify and cost existing expendable launchers to include reusability, was extended to include a launch vehicle design module, which was previously lacking. A Multidisciplinary Design Analysis and Optimization (MDAO) methodology was applied for the design module. The design module was developed by verified and validated models implemented from literature for hydrolox, kerolox and methalox propellants and an optimization scheme to minimize for Gross Lift-Off Mass (GLOM). This design module in conjunction with the FRT enables design and costing of expendable and reusable launcher configurations. For the current study, two missions were considered—15600 kg payload to Low Earth Orbit (LEO) and a 5000 kg payload to Geostationary Transfer Orbit (GTO)—for different propellant combination-based launchers. Additionally, different launcher configurations— expendable, reusable via non-propulsive recovery and reusable via propulsive recovery—were considered, enabling cost comparison of propellants for different scenarios. Results indicate that methalox based launchers are cost-effective solution when compared to hydrolox, regardless of mission type or launcher configuration considered in the current study. Compared to kerolox, only a marginal cost benefit can be achieved, for the case of expendable configuration and in combination with kerolox. For reusable configurations, purely methalox shows potential to achieve costs within 10% of kerolox. Sensitivity analysis showed the potential to reduce this gap by including the lower refurbishment requirement of methalox, given low soot formation possibility. Furthermore, it showed the need for a better engine model for methalox, to refine comparison between methalox and kerolox. The tool, however, is not complete and should be extended to include reliability assessment, especially for methalox systems, which are not flight proven unlike hydrolox and kerolox. There also remains issues with the accuracy and uncertainty in certain models, which make the current version suitable only for comparative studies.

This thesis details the methods used to build KOTUS (Kit Optimization Tool for Upper Stages) and the application of the tool to the future evolved Ariane 6 upper stage evolution, the Multi-function Upper Stage Express (MUSE), in order to gauge the impact of selected mission-specific kit elements on its payload performance for a range of Earth-bound missions.
Cryogenic bi-propellants such as the liquid hydrogen and oxygen combination are among the most potent propellants, typically sporting specific impulses in excess of 450 s in vacuum. However, the non-propulsive loss of cryogenic propellants sustained in missions that feature multiple boost phases can quickly become prohibitive. These losses limit the efficacy of cryogenic stages for a wide array of mission profiles.
A cryogenic stage is typically optimized for the characteristic challenges imposed by the most prominent mission in the portfolio offered by the launch service provider. The payload performance of this stage is then inevitably suboptimal for the other missions in the portfolio. This performance penalty can be addressed by introducing complementary or supplementary hardware elements into the stage architecture. A new configuration of the stage is then created using a combination of such elements, known as a kit. KOTUS addresses a need to determine the optimal kit for every mission profile at a conceptual level in an expedient manner. This can help to steer the design in the optimal direction from the onset. The tool is capable of performing this analysis for an arbitrarily defined mission, stage, and liquid bi-propellant combination.
In this work, the processes that are relevant to the payload performance analysis are discussed, along with the methods used in the modeling efforts of these processes. This includes the absorbed heat fluxes through insulation, attitude control and propellant orientation, propellant conditioning and pressurization, and feed line management. The agglomeration of these processes allow the modeling of the performance of a baseline stage when embarking on any mission. A set of elements was subsequently curated that can be equipped either individually or in combination with others to this baseline stage. The elements chosen are primarily focused on enhancing the operational performance of the stage by reducing non-propulsive losses. The integrated element set includes various insulation options, boost pumps, feed line heat exchangers, vapor-cooled shields, and thermodynamic venting systems. The modeling efforts for each of these elements and their individual impact on the upper stage relative to the baseline configuration is discussed at length. Sensitivity studies of uncertain input parameters and correction factors were conducted, as well as sub-model validations, to provide an indication of the tool’s credibility.
Results showed that for medium duration missions with one or more restarts, payload performance enhancements upwards of 100 kg are possible when a deployable sun shield is added to the baseline configuration. The degree to which the sun shield is able to limit absorbed heat flows allowed it to drastically reduce non-propulsive losses. For single boost missions, only the limiting of thermal residuals was demonstrated to be relevant to enhancing the payload performance of the stage. This was shown to be best accomplished through the use of a feed line heat exchanger in the LH2 feed line and a turbine-driven boost pump in the LOX feed line, the combination of which is estimated to enable an extra payload mass of 103 kg. In addition, it was shown that the default pressurization setup of the reference stage (i.e., the heterogeneous pressurization of the LOX tank and autogenous pressurization of the LH2 tank) was already optimal, and that no tank or feed line sizing changes would be beneficial.
The potential for the deployment of hardware kits has been demonstrated. Even short duration, single boost missions were shown to be able to benefit from tailor-made kit configurations, though the potential for payload performance enhancement is significantly larger for medium to long duration missions, particularly those featuring multiple engine restarts. A number of future work recommendations are listed in the conclusion, chief among them to verify the assumptions made in this work and to improve the determination of thermal residuals through enhanced knowledge of incident heat fluxes and spatial energy distributions in the propellant.

The renaissance of reusable space launch vehicles that is being witnessed in the current space industry demands a new cost estimating approach, which takes into account not only the time dependency in the lifecycle brought upon by reusability, but also the complementary considerations of minimizing cost and reliability. To achieve this end, a methodology was developed which allows to estimate both cost and reliability, and in a trade-space test various configurations at the development and operating phases, in order to achieve an optimum configuration which minimizes cost, maximizes reliability, and fulfills the design requirements. The development, manufacturing and operations costs of reusable launch vehicles were obtained using the Theoretical First Unit equivalence method. This includes the particular cost features of reusability, including the recovery hardware, the retrieval of the reusable equipment and its refurbishment. Furthermore, a model for the failure cost was developed, including: Flight replacement, Insurance Penalties, Failure Investigation, Modifications and Downtime. A reliability model is implemented in parallel to the cost model, which allows to obtain a flight-dependent reliability estimate, taking into account the possible reliability increase methods, such as redundancy, derating, engine-out capability, and extensive testing. This is achieved using publicly available test and operational data and using non-parametric (Kaplan-Meier) as well as parametric (Exponential and Weibull Maximum Likelihood Estimates) techniques in order to obtain life-cycle reliability. These two models coalesce into a trade-space where it is possible to tune the launcher configuration in order to meet its requirements, while minimizing cost and/or maximizing reliability. At the design level, the features of different configurations and testing programs can be compared, in order to achieve the lowest cost possible while attaining the necessary reliability targets. On the other hand, at the operation level, the reliability figures can be used in conjunction with the cost per flight in order to obtain the expected cost of failure and value throughout the life-cycle of the launch vehicle. This allows the designer to forecast the potential financial losses throughout the operating life of the vehicle, and look to minimize these losses while securing proper funding to cover them, which is especially critical in the first years of operations. Two use-cases are presented. Firstly, the model is applied to the Falcon 9 vehicle. Verification of the cost figures shows that the error of the methodology is under 30%. A projection of the life-cycle costs is obtained, as well as the expected cost of failure and expected value. Recommendations are given in order to mitigate losses and advise the planning of the life-cycle of the fleet and its reuses. It was found that at an initial stage, where cost reduction from learning and reliability growth haven’t yet taken hold, flights have a negative expected value, meaning that the launch provider can expect to lose money with these flights. Insurance options are investigated in order to mitigate these losses. Secondly, a study is done based on the Ariane 6 publicly available data, in order to study the impact of varying the number of engines in its first stage as well as engine commonality with the upper stage. It was found that for an expendable vehicle such as the Ariane 6, a single-engine configuration provides the most value, as the manufacturing costs of producing several engines per flight does not compensate eventual lower development costs. On the other hand, it was found that for a reusable vehicle, the stricter the reliability requirements, the more a multi-engine configuration is advantageous. This is due to the lower development costs combined with the replacement of the high cost of manufacturing engines with refurbishment costs, with an optimum for a configuration of 4-5 engines.

The Delfi program at the Delft University of Technology aims to develop nano satellites between 1 and 10 kg to provide high level education and to develop a platform for novel space technologies. Miniaturized satellites require miniaturized propulsion systems to provide orbital adjustments for functions like station keeping, extending mission duration and precise formation flying. Current research into micro-propulsion systems with aims to improve performance and efficiency is focused on Micro-electromechanical systems (MEMS), solar thermal and Commercial off-the-shelf (COTS) resistojet thrusters. The COTS design and prototype by Versteeg was the first leak resistant micro-thruster produced and successful experiments were performed using hot and cold nitrogen as propellant. The usage of liquid water as propellant for a Vaporizing Liquid Micro-resistojet (VLM) has been suggested as it is easy to store, has a high storage density, decent performance and is relatively safe to use. However as of writing this thesis, experimental tests using liquid water as propellant have been unsuccessful. In this thesis thrust tests with nitrogen have been repeated and show similar results to the data obtained by Versteeg verifying results of both experiments. In order for liquid water tests to be performed in the absence of a liquid mass flow sensor, a syringe pump was used to expel a constant volumetric flow. The liquid mass flow has been calculated by calibrating the syringe pump combined with the measured chamber pressure. Using liquid water, it was found that the chamber pressure increased over the course of the thrust experiment due to the feed system being volume based instead of driven by pressure. Since the test bench is dependent on the center of gravity of the thruster, thermal expansion significantly effects the measured thrust. Using nitrogen this effect can accurately be corrected for, while for water the effect is more difficult to correct due to the unpredictability of the liquid water remaining between the last propellant valve and the thruster chamber. This led to a decrease in accuracy of the measured thrust using liquid water compared to nitrogen, where the error grew from ∼1.5% to ∼13%. Thrust tests with water have been successfully performed at a chamber pressure of 1.02 bar and temperature of 300 °C for 15 minutes producing 8.0-8.3 mN of thrust with a specific impulse of 94-100 seconds with an accuracy of 14%. Several areas of improvement have been listed for both thruster design and experimental setup.

The existing literature on Laser-Thermal Propulsion (LTP) and associated concepts such as Directed-Energy Propulsion (DEP) and Laser-Sustained Plasma (LSP) is reviewed in preparation for an experimental thesis project on LTP.

The space community is currently focusing on defining system architectures able to perform multiple interplanetary missions to support deep space exploration. In particular, placing orbital propellant depots in strategic locations in space would allow to increase the useful payload mass at launch. Despite all the efforts that are being invested in finding effective propellant solutions, cryogenics are still widely used within the space industry. In particular, the hydrolox combination is the most used as it returns the best Δ𝑣 performance thanks to its high specific impulse. However, this class of propellants faces an issue that could jeopardize all the mass savings: boil-off. This consists in propellant evaporation due to heat penetration in the tank structure, and it could become so severe for long storage times (in the order of months) that the additional propellant mass used to make up for the boil-off losses would nullify all the benefits that come from the high impulse of these mixtures.

The approach taken in this work consists in developing a propellant depot sizing model that allows determining the effect of different thermal control design options, thermal environments, and depot configurations for varying mission durations. The model performs a multi-nodal thermal analysis to estimate boil-off rates and allows for taking into account different thermal control methods, which include Multi-Layer Insulation with uniform and variable density distributions and Vapor Cooled Shields.

Using the developed tool, the analysis of different depot architectures is performed. In particular, Low Earth and Geostationary orbits are investigated as locations for a depot, with the possibility to extend the results to the Earth-Moon Lagrange point L1. As for the depot, both single and multi-tank configurations are studied.
Results showed that variable density MLI (VDMLI) improves the effectiveness of the thermal insulation compared to the uniform density case: the most mass efficient designs in terms of insulation plus boil-off mass are achieved with less total number of layers when VDMLI is used. The use of a vapor cooled shield (VCS) additionally reduces the propellant boil-off. However, for the VCS to become advantageous in terms of mass, a minimum storage duration is needed: this corresponds to two months for the LEO depot, and at least six for the GEO depot. This is due to the mass penalty introduced by the addition of the metal shield into the system.
Lastly, it is shown that multi-tank depots have increased boil-off rates than single tank depots because of reflected thermal fluxes among the depot elements. However, some propellant savings come from placing the cryogenic tank in the shadow of the other tanks of the depot station.

The Europa Lander is a mission by NASA to land scientific instruments on the Jovian Moon by 2030. The current design of the landing system envisions fixed landing legs with the lander being lowered down at very low velocities (≈ 0.1 m/s) by use of a Skycrane. This latter system not only adds mass, but also requires complicated moving parts and a high reliance on control systems to work properly. This thesis wants to propose an alternative design to the Europa Lander Landing System that uses sturdier legs that can withstand landing at higher velocities than originally envisioned. This eliminates the need to use a Skycrane system. In order to pack them in a more efficient way while coasting, the landing legs are of the deployable type.
In order to track the deployment dynamics of the landing system a 2-dimensional Matlab deployment model based on double pendulum motion with damping is created. Furthermore, in order to prove the landability with a prescribed landing case, a landing model is created using FEA software Abaqus. The landing model is used to make sure that stresses, strains and G forces applied to the rest of the lander don’t exceed prescribed values as dictated by requirements. The implementation of both models is verified by comparing models found in literature with the ones made for the purpose of this thesis. Good correlation is shown with less than 10% of difference between literature results and results for this thesis.
The result through the modelling found a specific geometry with parameters that include, but are not limited to: length of leg elements, joint damping and maximum deployment angles. The 2-dimensional geometry features output by the Matlab deployment model are fed into the Abaqus landing model in order to assess the landing performance with a single leg landing. The landing case has been chosen as a worst case scenario. 2 landing leg designs, with the same geometry, but one made of metallic parts and the other made of composite materials are compared in the landing model. In this model it was found that both designs pass the stress (or strain for the composite leg design) and G forces
requirements when a shock absorber is attached to the leg and when they land on deformable Lunar like soil. On the other hand, the landing legs don’t pass G forces requirements when landing on stiff soil.
Through the modelling, it was also possible to give preliminary design guidelines over the subsystems making the landing system. Therefore, system engineering practices have been used in order to initiate the design of some of the subsystems such as the landing damping, the leg joints and the locking mechanism to lock the legs at the desired angle and make sure they don’t retract. Simplifications have been made in order to model some of the landing and deployment aspects of the system for this thesis. Matters related to the adaptability and the detailed design of each of the subsystems have been left as future work.
In the end, a comparison between the design developed in this thesis and the current Europa Lander Landing System design is made difficult by the lack of technical data on the latter related to mass. Nevertheless, a design that uses the decelerations of the engine module before touchdown is developed and the suggested material for the leg elements is CFRP. The design, though, didn’t pass mass fraction requirements when computing the mass of the extra subsystems. Even so, locations where potential gains through mass optimization can be performed for future studies have been pointed out.

In the space industry miniaturization has been a trend for decades, with satellite masses ranging from 0.1-10kg for pico- and nano-satellites. Propulsion subsystems for these satellites have not kept up with this trend performance-wise. They are often limited by safety regulations requiring inert propellants and
power budget constraints of the satellite.

Vaporizing Liquid Microthruster (VLM) are a type of resistojet considered for pico- and nano-satellites and are currently in development at Space Engineering (SE) department of the Technical University of Delft (TUD). VLM expel an electrically heated (inert) propellant, which is stored in the liquid phase.

Currently, VLM are inefficient and design tools are needed to find more optimal designs. In this research such a design tool is developed, and applied to thrusters with water as a propellant. A distinction with earlier research on VLM is that specific attention was paid to the field of microchannel flow boiling, so that knowledge from this discipline could be applied in the heat transfer model of the VLM heating channels.

The design tool was successfully developed, and is extendible with novel heat transfer and pressure drop relations when required for future research. When applied to 15 mN thrust range, the outcome is that the chamber temperature can be increased without increasing the total power consumption significantly. Heating chamber design also tends to be more optimal if more and smaller channels are used than in earlier published VLM designs.

The optimal wall temperatures were about 40 K higher than the chamber temperature. The optimal channel width and wall thickness were constrained by structural limitations, at 20 and 50 micron, respectively. The optimal number of channels increased with thrust from 3-4 channels at 1 mN to 8-9 at 5 mN.

The aim of this investigation is to improve the first stage of mission analysis on trajectories in the Circular Restricted Three-Body Problem (CR3BP). This will be achieved by developing a versatile tool that can optimize any transfer in the CR3BP in terms of ΔV and time of flight. Hence, the output will not be
a single solution, but a Pareto front with multiple non-dominated solutions that range in duration and propellant consumption, such that the user can easily identify the one that better suits his/her particular mission requirements.
The tool considers three types of arcs: direct transfers, manifold trajectories and flyby arcs which can be combined in any way to connect the departure and destination orbits. These can be defined as Keplerian orbits or CR3BP periodic solutions, including Lyapunov, Halo orbits or NRHOs. This is done in a very intuitive way thanks to the careful selection of design variables. Therefore the developed tool can be used by designers that do not have a large amount of experience solving trajectory optimization problems in the CR3BP.
In order to demonstrate the capabilities of the tool, two specific transfers were optimized. In both cases, the Pareto front is obtained with the combination of optimal solutions that include direct transfers, and manifold and flyby arcs. First, a LEO to L2 Halo orbit with Az = 2000 km was selected to compare the results obtained by the software with those from literature. The tool not only obtained more results than the previous research, but the solutions found were also improved in terms of ΔV and time of flight. The second transfer was a GTO to lunar Gateway NRHO trajectory, which was chosen to show the applications of this study to problems of high scientific and industrial interest in present times, as well as to prove the versatility of the algorithm. Again, the results are better than the ones found in literature, if only by the number of different solutions that are obtained. Moreover, these results can easily be exported into a higherfidelity
software such as ASTOS.

Dawn Aerospace is start up company that is developing a reusable sub-orbital spaceplane. The MK-II rocket powered spaceplane uses High Test Peroxide for its main engine and reaction control system. Catalyst beds are required to decompose High Test Peroxide (HTP) to steam and Oxygen. To investigate the decomposition of HTP in a pellet catalyst bed a numerical model was developed which is capable of predicting High Test Peroxide flow decomposition in catalyst beds.

The model developed is based on a two-phase flow through the catalyst bed that is chemically reacting by undergoing decomposition. The flow is assumed homogeneous, adiabatic, and continuous resulting in steady state model outputs. Fast equilibrium adsorption and finite rate desorption of Hydrogen Peroxide on to the catalyst surface is assumed in a single step. The decomposition is modelled using the Arrhenius relations. Additionally, the flow pressure losses are computed using an extended Ergun relation for two phase flowthrough packed beds. The model can predict the flow temperature distributions up to 98% reaction advancement. The model was verified by comparing it to existing model results, this showed good agreements in 7 of the 8 cases. The results showed consistently that the model predicted 98% of the decomposition progression is achieved in a shorter catalyst bed length than the catalyst bed lengths in an existing model (4 to 14% shorter). The flow pressure losses of the models showed deviations compared to the existing models ranging from 5.55% to 12.18%. The differences in the flow temperature and flow pressures are discussed and related to the governing theories of the models.

To validate the model, experiments were performed. The test data were used in model validation and a good agreement was found between the test data and the simulated results. The simulated flow temperature and the measured temperature deviated from 0.41% to 1.93% of the adiabatic flow temperature. The overall pressures showed good agreement with the experimental average, with pressure drop deviations less than 1% of the operating pressure. In all experiment cases the flow distribution predicted was within the measured pressure roughness.

Many simulations were performed and compared to investigate the impact of varying design choices and operational conditions. In all cases the flow development along the catalyst bed length was assessed. In comparison to the simulations results it was found that an increase of operational parameters such as pressure promoted gas phase decomposition resulting in shorter bed length requirements. Similarly, the influence of feed High Test Peroxide concentration was evaluated and finding showed increasing concentration to shorten the required bed length. The analysis of the simulated conditions showed that the increase of mass flow inversely impacts the required catalyst bed length. This model serves as tool to identify the operational parameters and their impacts in catalyst beds. The Model outcomes demonstrating the effects of operational parameters and design choices on catalyst bed helps in understanding their interactions and can aid in catalyst bed design and optimisations processes.

A current trend, in recent decades, in spacecraft engineering is the miniaturization of satellites and their subsystems. One of these subsystems is for propulsion, which is important for orienting satellites and performing orbital manoeuvres. Currently one such micropropulsion system, being developed at the Space Engineering department at the Delft University of Technology, is called the ``Vaporizing Liquid Microthruster". These thrusters are manufactured using Deep Reactive Ion Etching on silicon wafers. They operate by resistively heating the propellant, vaporizing and expanding it throughout the nozzle to create thrust. Currently, several experimental investigations have been conducted to formulate models that can predict the performance of these microthrusters, on a wide range of operating conditions, without the use of expensive testing. These models can then be validated comparing them to experimental data. Thrust tests were performed with nitrogen, for three different thrusters in ambient, cold conditions, for operating pressures of 2-5bars. From these tests, the discharge coefficient was predicted within 10\% of the experimental values. So the model's prediction of the discharge coefficient was valid. It was also immediately clear that the thruster was not designed for ambient conditions due its low specific impulse values of 7.6-22.7s, resulting in incomplete flow expansion in the nozzle. The current 1st generation thruster interface used at the department is limited to 150ºC due to its teflon material. Therefore, an alternate design using a printable circuit board was investigated to reach at least 600K, where the thruster would be glued to it using a thermally resistant, ceramic based glue (Thermeez 7020). Unfortunately, the design was limited to 150ºC due to the low adhesive quality of the glue, which meant that the thruster inlet could not be sealed off and a less suitable replacement glue had to be used. Heated thrust tests were performed for two thrusters up to around 150ºC in ambient conditions. For this increase in temperature , throat Reynolds numbers decreased from 2915 to 779, while the discharge coefficient increased from 0.71 to near 1. However, a less suitable mass flow sensor had to be used, so the validation of the model based on the hot thrust tests was inconclusive for discharge coefficient and the Isp efficiency. Overall, the goals for this experimental research study were partially reached. Thrust tests were conducted, using nitrogen as propellant, for cold conditions and up to 150ºC. The experiments could not substantially validate the prediction of the Isp efficiency, however the model is accurate to within 10% of the experimental values for the discharge coefficient. Experimental data from literature also could not validate the Isp efficiencies of the model, unless the fitting constant of the thrust coefficient relation was altered. It is recommended to focus on experimenting with the use of water as a propellant and make minor improvements to the thruster design, to obtain accurate performance data at higher temperatures.

The supersonic split line (SSSL) nozzle has been considered for years in the use of solid rocket motors. The SSSL nozzle is an attractive alternative to the more conventional submerged movable (SM) nozzle. The SSSL nozzle can save up to 43% on nozzle mass and 25% on the thrust vector control (TVC) system mass. The SSSL nozzle also has the benefit of the amplification factor, which is a result of the SSSL nozzle turning the flow in the supersonic region. The shock waves formed deflect the flow more than the mechanical deflection. Only limited research is available on the SSSL nozzle. This thesis aims to expand on the knowledge of the SSSL nozzle in cooperation with Nammo Raufoss. The aspects that are researched in detail are as following: The effect on the amplification factor when moving the split and changing the expansion ratio of the nozzle. Furthermore, to determine how well the SSSL nozzle compares to a more traditional SM nozzle, a comparison between the SSSL and the SM nozzle is performed at equal performance. This comparison has a focus on the mass differences. The performances of the SM nozzle and the SSSL nozzle are determined by simulations with computational fluid dynamics (CFD). From the CFD simulations, the amplification factor of the SSSL nozzle was determined. In order to compare the SSSL nozzle and the SM nozzle a mass model is developed. The model uses the calculated performance from the CFD simulations to calculate the mass difference between the two nozzle and TVC systems. The results show that the amplification factor increases when moving the split further down stream. Furthermore, when increasing the expansion ratio the amplification factor decreases. For the SSSL nozzle with a total expansion ratio of 12 and the split location at an expansion ratio of 1.75, the maximum amplification factor is 1.57. Additionally, a linear relationship between the amplification factor and the ratio of the expansion ratio at the split to the total expansion ratio of the nozzle is observed. The SSSL nozzle experiences increasing thrust and Isp losses while vectoring when the split is located further downstream, while receiving only a slight increase in the amplification factor. The results from the mass model show that the SSSL nozzle is on average 37% lighter than the SM nozzle for a vectoring duty cycle of 0.25. To conclude, for an increasing ratio between the expansion ratio at the split to the total expansion ratio of the nozzle, the amplification factor increases. Additionally, the SSSL nozzle is the lighter option when only the aerodynamic effects are considered. It was also seen that the temperature and heat flux around the split area was high, to overcome this the SSSL nozzle might need additional thermal protection which could potentially lead to a higher mass. It was observed that the largest mass difference between the SSSL and SM nozzle system was found when the split of the SSSL nozzle is located at an expansion ratio that gives the highest possible Isp.

Electric propulsion is generally characterized by higher specific impulse than traditional chemical rockets, which gives them propellant saving benefits. These have been applied in deep-space missions and micro-satellites. A Pulsed Plasma Thruster consists of two electrodes around a block of teflon. A pulsed discharge ablates the teflon and ionizes the vapor. This plasma forms a current bridge between the electrodes which is accelerated outwards by the self-induced magnetic field, producing thrust.

The current research project has developed and used a 1D model to investigate the current bridge with a focus on the plasma electrode interaction. The influence of electron temperature and density, cathode temperature, voltage, geometry and propellant type on the current density and magnetic field has been explored. The influence of thermionic electron emissions on the current-voltage characteristic has been demonstrated.