Micropropulsion is universally considered to be a key technology enabling nano- and pico- satellites to perform more complex missions. However, past research has shown that nozzle efficiencies at the microscale are far inferior to their macro scale counterparts. These low efficiencies can be attributed to the relatively high viscous losses associated with this microscale. This thesis conducted a three-dimensional numerical study to investigate the impact of the nozzle geometry on the viscous losses. Furthermore, the designed micronozzles are fabricated and the ground work is laid for experimental testing of these nozzles. Results of the numerical study show that by application of a new double depth micro aerospike design nozzle efficiencies can be improved by as much as 41.2%.

Resonance ignition devices utilize the gas dynamic phenomenon to heat propellant gases to their auto-ignition temperature in order to initiate combustion. These devices are a promising alternative to current rocket engine ignition systems. In this thesis a python code was developed to generate a resonance ignition system design. The design was tested and parameters characterized using ANSYS Fluent. This resulted in a detailed design of a GOX-Ethanol resonance ignition device that is successful in heating the resonance gas to above its auto-ignition temperature. This project presents the python and Fluent tools developed to conduct this study, analyses the effect of different input parameters on the device, and presents a working detailed design.

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.

Reusable rocket stages is a trend in launch vehicles that have been observed for some time now. Even reusable upper stages have been seen in the Space Shuttle, Kistler K-1 and now the SpaceX Starship. This thesis is on the development of a genetic optimisation algorithm to determine the economic feasibility of upper stage reusability. Genetic optimisation has been chosen due to the mixed-integer nonlinear nature of the Reuse Index proposed by United Launch Alliance. The optimisation algorithm uses a modified version of ParSim v3 made by the Parachute Research Group (PRG) of Delft Aerospace rocket Engineering (DARE). The thesis gives an overview of the various dynamic, mass and cost models used in the tool together with the verification and validation of these models. Based upon the conclusions of the various runs performed recommendations are given on if and when upper stages can be reused.

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.

the aim of this thesis is to regard all aspects concerning mass optimization with respect to small satellite structures and deployers, specifically aimed at the DelfiPQ PocketQube mission.

The influence of using different chemistry model and different chemical kinetics schemes on modeling combustion inside a rocket engine igniter are investigated. An igniter for a LOX/CH4 rocket engine is designed and a CFD model is developed which is used to model the designed igniter. Verification and validation methods are discussed.

In the search for understanding the formation and evolution of a solar system, terrestrial planets & any exoplanetary systems, the understanding of the evolution & formation of giant planet's play a crucial role. The primary goal is to study the atmospheres, their physical & chemical processes governed by gases, clouds, and hazes. This research focuses on developing a new technique to characterize aerosol particles using a nephelometer. The new forward model is a two-part model, in the first part, an atmosphere radiative model for Saturn & Venus is developed and the polarization internal field is evaluated using Monte-Carlo technique. The second part of the model is to demonstrate the feasibility of the new instrument technique developed in first part using space system engineering approach. The research concludes that the forward model developed could be used to evaluate the internal polarization field of any planetary atmosphere. The feasibility study of new instrument technique for Saturn’s atmosphere demonstrates a possibility to derive the vertical structure up to 2[bar] pressure & for Venus’s atmosphere in all layers.

The heat fluxes in (liquid) rocket engines can go up to high values and they need active cooling to prevent the chamber wall from failing. Transpiration cooling is identified as a way to achieve lower wall temperatures than commonly used regenerative cooling and regenerative cooling with additional film cooling (referred to as film cooling from now on). This can lead to higher performance of the engine. However, transpiration cooled engines are not in use today and this is mainly attributed to problems with the required porous wall materials. Additive manufacturing (AM) is a promising solution for these material problems. Better cooling is especially useful for Inconel additively manufactured rocket engines as such engines experience higher wall temperatures due to the low thermal conductivity of the material. This thesis has two purposes. Firstly, an analysis was performed to see if transpiration cooling actually performs better than regenerative and film cooling in total engine performance. Secondly, it was investigated if AM can be used to create the porous walls required for transpiration cooling. To compare the cooling techniques, a simplified model for each cooling technique was developed. These models were verified and validated using data from literature. A new (transcritical) film cooling model was created by combining three existing film cooling models. The wall temperatures obtained from the three cooling methods were compared by applying them to a reference liquid rocket engine with either Inconel or copper as wall material. Subsequently, the losses in specific impulse and dry mass were determined. Then, a delta-v calculation for each cooling method was made to objectively compare them. The conclusions on the comparison of the cooling techniques are that the regenerative cooled engine reaches wall temperatures above the material limits. Therefore, it is not a feasible to use this cooling method for the reference engine. Film and transpiration cooling can both achieve temperatures below the limit. Transpiration cooling requires less coolant than film cooling to achieve the same temperature. This will result in lower losses in specific impulse compared to film cooling. However, to achieve these low coolant mass flows, thick chamber walls are required to achieve the specified pressure drop over the wall. When comparing transpiration cooling to film cooling on the total delta-v achieved, it is found that the Inconel chambers perform better than the copper ones. However, it depends on the pore size if transpiration cooled engines outperform film cooled ones. A smaller pore size is better. Additionally, it was found that the pore sizes producible with AM are an order of magnitude larger than required. Therefore, experiments were performed on the pressure drop over AM porous walls with different geometries. With these experiments, it was found that a new porous wall geometry made using AM techniques has a lower pressure drop than the geometry used in the calculations, being 1.32 lower. A geometry designed to achieve an as large as possible pressure drop increased the pressure drop 24.9 times. However, this geometry does not achieve a uniform coolant injection required for transpiration cooling, so further research is required.

This thesis includes an experimental study about the performance of a third generation MEMS-VLM chip. The influence of the throat Reynolds number on the discharge factor and the nozzle quality was investigated. This was done by measuring the mass flow and thrust force for a chamber pressure range of 2-5 [bar] using nitrogen gas as propellant. It was found that increasing the Reynolds number from 600 to 2000 was beneficial since the nozzle quality increased from 0.19 to 0.60 and the discharge factor from 0.77 to 0.91. The found nozzle quality and discharge factor were lower than values found in literature. It was hypothesized that this was because the 2D design of the MEMS-VLM chip is more sensitive for boundary layer formation.

The conducted study investigates the potential of a newly released multi-phase solver to simulate atomization in liquid rocket injectors. The "VOF-to-DPM" solver was used to simulate primary and secondary atomization in an air-blast atomizer with a coaxial injector-like geometry. The solver uses a hybrid Eulerian/Eulerian-Lagrangian formulation with a geometric transition criteria between the two models. The conducted study assumed isothermal, non-reacting flow at room temperature. The primary focus was predicting Sauter Mean Diameter and droplet velocity data at a sampling plane downstream of the injection site. The results showed that the solver is able to produce the expected data and to predict trends similar to those found in experimental measurements. The accuracy of the produced droplet diameters was roughly a factor 2 off compared to experiment. This is attributed to mesh resolution. Measurements were obtained via a cooperative agreement between TU Delft and The University of Sydney. It was concluded that the solver has the potential to predict atomization at a reasonable computational cost, but further study is needed to confirm its full capabilities.

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.

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

The 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 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 order to make access to space cheaper, the first stage of a launcher can be recovered. For this research, different options are investigated to make the cost per flight cheaper. Two vehicles are investigated going to two different orbits. The results show that recovering the complete stage is better for vehicles that have a lower separation velocity. When the separation velocity is higher, recovering only the engine is better. The best results are obtained when using a Ringsail parachute and a Hypersonic Inflatable Aerodynamic Decelerator. The design for these systems is optimised using a surrogate optimisation algorithm that reduces the computation time. The set target of 30% cost savings could not be achieved, but cost savings of at least 20% were achieved for both vehicles.

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.

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.

Even though hundreds of CubeSats have been launched, few have launched with a micropropulsion module on board. Propulsion would allow for an extended lifetime and better mission capabilities, thus greatly increasing the attractiveness of CubeSats. TU Delft is working on a type of thruster suitable for CubeSats, called the Vaporizing Liquid Microthruster (VLM). For micro-propulsion it is viable to generate sufficient thrust by using an external heat source to vaporize a liquid and expanding the vapour via a nozzle. The usage of such micropropulsion modules is still inhibited by issues in the nozzle and heating chamber. One issue is the multiphase flow occurring inside the micronozzle, which lowers the thrust and specific impulse performance of the thruster as well as the stability due to the explosive boiling phenomena. Previous research indicates that the multiphase flow in the nozzle most likely occurs inside the heating chamber and flows into the nozzle. Little is known on how exactly this multiphase flow looks like, ranging from annular flow to dispersed droplets, nor on how it is generated within the chamber. It is indicated that the boiling may lead to large pressure oscillations inside the chamber, which contributes to the instability of the thruster. These pressure oscillations occur due to multichannel interactions from nucleate boiling in microchannels, as well as explosive boiling. In this thesis, the lattice Boltzmann method (LBM) is used to simulate the complex phase change multiphase flow occurring inside the microchannels of the vaporizing liquid micro-thruster. This method has recently gained a lot of attraction in simulating microscale fluid problems. Research shows that LBM applied to multiphase flows can be an order of magnitude faster than conversional Navier-Stokes solvers. This thesis utilizes the Bhatnagar–Gross–Krook (BGK) collision operator in combination with the pseudopotential method to simulate the multiphase flow. A modified Guo forcing scheme is used to ensure thermodynamic consistency. The water is modelled using the non-ideal Peng-Robinson equation of state. The thermal problem is solved using the double distribution function method, in combination with a semi-hybrid source term solved via a multiple relaxation time (MRT) collision operator. The phase change term is derived from the local balance of entropy, using the equation of state to calculate the latent heat. The open-source lattice Boltzmann method solver OpenLB is extended to be capable of simulating phase change flow. Verification tests are performed to show that the code extension is correctly implemented. The code was released to the OpenLB, making it the first open-source lattice Boltzmann method framework capable of simulation phase change flow. In the analysis of the tool developed, the pool boiling simulations showed how nucleation, bubble formation, and departure occurs. Correct behaviour are shown in the thermodynamic consistency, Laplace pressure and wall wettability. The effect of the spurious currents on the thermal solution are analysed in more depth than found in literature. The procedure for the multichannel simulation verification and validation is given, including a novel analysis method. The novel analysis method can estimate the multiphase flow type occurring inside real thrusters without any additional equipment required in a test setup. The required experiment data are the specific impulse values at various massflow rates, and the massflow rate at which multiphase flow starts to occur. The simulations are strongly limited by the BGK collision operator, which is apparent by the large spurious currents at high density ratio multiphase flow. The spurious currents can cause simulation instabilities, but are mitigated by using a thicker diffuse interface. This mitigation results in a larger metastable phase change region, which inhibits the nucleation process. Thus, nucleation inside the microchannels did not occur. However, using a MRT collision operator should reduce the spurious currents without needing to increase the diffuse interface thickness, allowing for nucleation to take place in the microchannels.