This thesis examines the thermal distribution and material behavior in in-space additive manufacturing (ISAM) using metal directed energy deposition (DED), addressing challenges of the orbital environment, including microgravity and vacuum. Experimental studies with a Scandium-modified Al-5183 alloy, conducted via Wire Are Additive Manufacturing, validated heat transfer coefficients and simulated ISAM thermal profiles. A finite element model, developed in Ansys Mechanical and calibrated with experimental data, accurately predicted overall thermal histories despite underestimating melt pool temperatures. Applied to orbital conditions, the model showed environmental healing effects were minimal for small components but significant for larger structures, supporting ISAM’s potential for space infrastructure. Material analysis revealed enhanced mechanical properties due to Scandium addition, with refined grains and Al3Sc precipitates. Integrating experiments, simulations, and characterization, this work advances ISAM thermal modeling, offering insights for future refinements in simulation accuracy and orbital testing.

This research investigated the application of electrical thrusters within the Attitude Determination and Control System (ADCS) of the European Space Agency’s LUMIO mission, a 12U CubeSat set to op- erate in a quasi-periodic halo orbit around the Earth-Moon L2 point. While CubeSats have become pivotal tools for lunar exploration, the use of electrical thrusters for ADCS, particularly in the lunar en- vironment, remains under-explored. This study aimed to evaluate the impact of replacing LUMIO’s current ADCS design by electrical thruster configurations on attitude control performance, robustness, and connectivity, bridging a critical gap in current literature.
A simulation framework was developed to assess spacecraft performance over a two-week period, ex- amining the LUMIO mission reaction wheel set-up, as well as four distinct thruster configurations. Key metrics included thrust output, angular velocity, and pointing accuracy, represented by the half-cone offset angle. In addition, power and energy requirements were investigated. Robustness tests eval- uated system adaptability to single thruster failures, solar array deployment, and high initial angular velocities. Physical validation was achieved by porting the control algorithm to an embedded STM32 Nucleo board, which was connected to a solenoid valve representing a dummy thruster. This setup allowed for real-time testing of execution accuracy, signal fidelity, and system integration.
The results demonstrated that the control algorithm consistently maintained high pointing accuracy, keeping the half-cone offset angle within the mission requirement of 0.18 [◦]. Reaction wheels and electrical thrusters effectively generated the required control torques, with negligible angular momen- tum build-up in the nominal simulation. Overdetermined thruster configurations showcased enhanced energy efficiency, reducing total energy consumption by approximately 29% compared to determinate setups. However, the study also highlighted challenges, such as the minimum impulse bit constraints of the Pocket Rocket thrusters, which limited precise low-thrust operations.
Robustness tests revealed that determinate configurations failed to accommodate single thruster fail- ures, while overdetermined systems adapted effectively, albeit with increased power demands. From the results, it is recommended for any space mission to be certain that the on-board computer is aware of any thruster fail occurring. The undeployed solar array configuration significantly reduced energy re- quirements, supporting its feasibility for early mission phases. The thruster configurations showed not to be able to realistically counteract de-tumbling manoeuvres, significantly degrading their feasibility in the actual LUMIO mission. Additionally, the embedded system validation confirmed accurate command signal generation and execution, bridging the gap between simulation and real-world implementation.
Although the electrical thruster configurations adhered to pointing accuracy requirements in nominal simulation conditions, they impose significantly higher power and energy demands and are unable to counteract high de-tumbling manoeuvrers compared to the current reaction wheel configuration in the LUMIO ADCS design. Addressing main engine parasitic torques and exploring alternative, higher- performance thrusters may offer a feasible solution for integrating electrical propulsion into lunar Cube- Sat ADCS. While the reaction wheel and non-electrical thruster combination remains the preferable option for now, this research demonstrates the potential for electrical thrusters in advancing CubeSat capabilities. Future studies should build upon this foundation by addressing current limitations, extend- ing simulation durations, and exploring innovative propulsion technologies to unlock their full potential in deep-space missions.

A promising design of a Space-Based Solar Power (SBSP) system for a lunar demonstration mission has been developed. SBSP technology can offer a continuous power supply using solar energy and wireless power transmission, addressing challenges like the 15-day lunar night. The proposed concept employs Laser Power Transmission (LPT) due to its compact design and high power density, delivering 3 kW from a Frozen Polar Low Lunar Orbit to a surface receiver. The satellite integrates deployable MegaFlex solar arrays, a custom LPT subsystem, and robust power management, thermal control, and communication systems. A system sizing model estimates a total mass of 4057.43 kg and a solar array output of 15.8 kW. The scalable and redundant architecture could support future operations up to 45 kW.

The high costs associated with launching satellites into orbit drives the market towards more compact satellites. These compact satellites require the development of miniaturized components, including a propulsion system. The Space Engineering (SpE) department of the Faculty of Aerospace engineering at the TU Delft is actively working on developing these miniaturized propulsion systems, including the Vaporizing Liquid Micro-resistojet (VLM). This thesis aims to "design a modular micro-resistojet propulsion system that is capable of delivering three distinct total impulse levels, based on a pre-existing VLM thruster developed by the TU Delft, with a focus on integration into a CubeSat configuration. The three systems developed in this thesis comprise a printed circuit board (PCB), a propellant tank with integrated propellant management devices (PMDs), piping, tank heaters, and a TunaCan add-on, utilizing water as the propellant and gaseous nitrogen as the pressurant. The PMD is a combination of a radial sponge, consisting of 70 sponge panels, and a 6-vane configuration. This vane configuration consists of 3 short and 3 long vanes to increase the expulsion efficiency without altering the ability to move propellant from the top of the tank towards the sponge and tank outlet. A preliminary thermal analysis on simplified versions of propulsion systems showed that a combination of tank heaters, each capable of providing 0.5 W, and a Zerlauts Z-93 white coating applied to the thruster side of the propulsion system are required to maintain a propellant temperature between 10 ºC and 50 ºC, whilst ensuring that the propellant within the piping does neither freeze nor boil. The final system designs achieve a thrust level of 1 mN and specific impulse of 130 s for a maximum input power of 7.9 W. A detailed thruster design is outside the scope of this thesis hence these performance values are calculated using ideal rocket theory. The three systems are capable of delivering a total impulse of 226 Ns, 465 Ns, and 700 Ns at a wet mass of 777 grams, 998 grams, and 1217 grams respectively. Several recommendations are proposed for further research, including the redesign of the thruster, enhancements to the tankage and PMD design, and improvements to the thermal and structural analyses.

Heat exchangers are essential for temperature regulation across various industries, from everyday applications to space exploration. Over time, these devices have evolved from simple clay vessels to complex structures made from advanced metal alloys and ceramics. In modern engineering, additive manufacturing (AM) has revolutionized heat exchanger production, enabling the creation of thin-walled structures with complex internal channels designed for efficient fluid flow. This is particularly critical in industries such as aerospace, power generation, and manufacturing, where components must be compact, lightweight, and capable of withstanding extreme temperatures and pressures. However, traditional AM methods like laser powder bed fusion (L-PBF) are constrained by size limitations, necessitating new manufacturing techniques to meet industry demands for compact large scale heat exchangers.

This research addresses the challenges of developing the Laser Powder Directed Energy Deposition (LP-DED) process for extreme-environment heat exchangers. Process and flow test experiments were conducted, along with comprehensive characterization of LP-DED-fabricated microchannels, which are thin-walled (1 mm) and capable of containing cryogenic or high-temperature pressurized fluids. The results from the research establishes LP-DED as a viable technology for heat exchanger fabrication by addressing challenges related to geometry, wall thickness, surface texture, and the fluid dynamics of these unique AM surfaces. Three key research questions guide the study:

1. How is the surface texture of heat exchanger microchannels affected by the LP-DED fabrication process?
2. What are the geometrical relationships and sensitivities affecting fluid flow performance if heat exchanger channels are manufactured with the LP-DED process?
3. What improvements can be made to control the surface texture of thin-wall LP-DED internal microchannels?

A detailed literature review identifies significant gaps in current thin-wall LP-DED manufacturing and internal surface enhancement techniques. An experimental study examines LP-DED process mechanics and build parameters, focusing on their influence of the thin-wall surface texture and the effect of build angles on both open and closed structures. This research establishes guidelines for Design for Additive Manufacturing (DfAM), addressing process limitations, surface texture, and wall thickness metrics for the LP-DED process.
This research introduces microchannels fabricated using LP-DED in various sizes, with their internal and external surface textures, wall thicknesses, and repeatability characterized. These microchannels are then processed internally using various surface enhancement techniques to provide variations of the surface finish. Two experimental studies were conducted, with comprehensive characterization performed of the internal channel surfaces to evaluate the variations in surface texture resulting from the enhancement processes and their relationship to flow resistance and friction factors.

Key innovations include the characterization of a new hydrogen-resistant alloy (NASA HR-1), which provides foundational data for heat exchanger design. The study also identifies surface texture mechanisms that affect fluid friction factors, resulting from distinct enhancement techniques such as peak smoothing, roughness minimization, waviness, and valley reduction. Additionally, friction factors and differential pressure in LP-DED-fabricated microchannels, both in as-built and surface-enhanced conditions, were investigated. Surface treatments such as abrasive flow machining, chemical milling, and chemical mechanical polishing were evaluated. The experimental results and comprehensive surface texture characterization led to the development of new correlations for calculating the hydraulic diameter of square channels and predicting sand grain roughness and friction factors. These correlations resulted in pressure drop predictions with deviations of less than 20% from experimental data, offering a 50% improvement over previous models.@en

In the renewed race for the Moon, there will be a high demand for transportation services to and from the lunar surface. The next generation of landers needs to have many capabilities, such as supporting longer exploration and global access, sustainability and versatility. ARCH-E, developed together with Airbus Defence and Space UK, fills this need as a reusable, crew/cargo-capable lander architecture. The lander's requirements and market are analysed, mission trajectory and landing are designed, followed by detailed subsystem design. The operations, risks, costs and sustainability considerations are assessed and highlighted. Finally, the design overview of a feasible lunar landing system is presented. Further design activities are recommended and a project timeline is laid out.

This study proposes the novel concept of recycling space debris as a means of supplying material resources for the establishment of a permanent Lunar presence while simultaneously cleaning up Earth’s orbital environment. This expands the scope of traditional debris mitigation efforts, which focus only on removing debris. As a first step, the study presents the creation of a space debris database. A novel classification is made to characterize space debris in terms of material resources and reserves usable for recycling. It is concluded that plentiful defunct upper stages currently drift in Geostationary Transfer Orbits (GTO), presenting substantial risk to other space assets. Their high relative metal content and elliptical orbits make these stages prime candidates for an efficient transfer to the Moon using an Orbital Transfer Vehicle. But as various perturbations have shifted the orbits of upper stages in GTO over the years, the orbital transfer alignment is found to be a critical complexity for a recycling mission. Several mission scenarios are analysed from debris capture to processing on the Lunar surface. The total energy expenditure across the entire mission is used as a novel tool to characterize and compare these scenarios to one another and to the alternative: a direct material delivery mission featuring ESA’s Argonaut lander. By harnessing the unique advantages of electric propulsion, this study shows that the concept of recycling defunct rocket stages is both feasible and viable. A direct, single-target recycling mission shows potential to require 30% lower energy investment per kg of raw material delivered compared to a standard lander mission. A novel thrust-arcing trajectory solves the alignment problem through flexible steering of the apogee to facilitate a Lunar encounter at minimal performance cost. Factoring in the potential to spread launch energy expenditure through a secondary client in a rideshare configuration results in an increase to over 60% less energy investment per kg. A preliminary analysis shows that even greater efficiencies could potentially be achieved through a continuous mission that returns multiple upper stage.
Ultimately, the novel concept of coupling space debris mitigation to advancing long-term Lunar exploration presented in this study shows significant potential and warrants further study. Moreover, a comprehensive analysis of the energy cycle highlights the potential for improved energy efficiency over conventional lander missions. Finally, the use of global energy expenditure as a tool for relative analysis shows potential as an adequate means of analysing space missions featuring distinct segments.

There is a growing demand for nano-satellite missions that require advancements in micro-propulsion capabilities to facilitate a wider range of orbital maneuvers. Among these capabilities, the ability to accurately control thrust would unlock new possibilities for nano-satellite applications, including missions such as space debris removal and orbit transfer.
Delft University of Technology is currently pioneering the development of an innovative green propellant-driven micro-propulsion system based on micro-electro-mechanical (MEMS) technologies for its PocketQube, known as Delfi-PQ, which features a compact form factor of 5x5x5 cm. While the thruster itself is in development, the interfacing and integration with other components are still ongoing.
Given the stringent mass, volume, and power limitations imposed by PocketQube satellite requirements, there is a pressing need for micro-scale components to realize a highly integrated propulsion system.
This thesis focuses on the design of a MEMS-based microvalve for proportional flow control in micro-resistojets. The design is conceptualized as comprising three components working in harmony: a valve seat with inlet and outlet, a flexible membrane, and a piezoelectric actuator. The valve seat with inlet and outlet, as well as the flexible membrane, utilize MEMS manufacturing techniques and are based on a silicon chip. And, a new design for piezoelectric actuators employing the d31 mode for contraction strokes is proposed.
The proposed preliminary design is a normally closed microvalve designed for a flow rate of 5g/hr, with the flexibility to accommodate higher flow rates if needed. It offers proportional flow control, ensuring precise regulation of fluid flow. Additionally, it promises a low power consumption of less than 1W and a low response time. Furthermore, this thesis provides a detailed outline of the MEMS fabrication process flow available at TU Delft’s Else Kooi Laboratory for the device. It also features a comprehensive test plan aimed
at assessing the feasibility of this design in future studies. Additionally, the thesis includes an elaborate risk analysis to help identify and mitigate potential risks during the manufacturing and testing phases.
The design shows promise and could pave the way for future developments of the microvalve within the department.

Progress in Laser powder bed fusion manufacturing has led to greater use of Additive Manufacturing in combustion chambers/nozzles. This allows for the creation of more intricate and efficient solutions, but generates the need for a better understanding of how surface roughness in micro-channels used for engine cooling relates to build angle and performance. The research presented in the framework of this thesis assesses the friction factors and internal surface roughness of 18 additively manufactured channels, through the means of flow testing and microscope inspection. Positive correlations between build angle and friction factors were found, and further investigation revealed relative roughness levels outside the applicability range of Moody charts. Comparison of open and closed contour channels confirmed roughness similarity, providing valuable input for the manufacturing of witness channels in cost/time sensitive applications. Finally, the relation between surface roughness parameters and sand-grain roughness were investigated, revealing their dependency on flow properties.

In line with the growing trend of satellite miniaturization, Delft University of Technology has developed advanced micropropulsion systems such as micro electro mechanical systems (MEMS) vaporizing liquid micro-resistojet (VLM) using water as propellant. These systems play a crucial role in enabling small satellites to achieve functions like maneuverability and attitude control. The objective of this research is to evaluate the performance of next-generation MEMS VLM by creating a numerical simulation model for the VLM heater microchannel and micronozzle and validating the model using existing experimental data. To this end, numerical simulations were conducted for the VLM microchannel and micronozzle using computational fluid dynamics (CFD) models. The microchannel simulation captured flow boiling characteristics, including bubble nucleation, growth, and heat transfer. Micronozzle simulations validated against experimental data showed deviations of less than 10% in key metrics like thrust and specific impulse and confirms the enhanced performance of next-generation designs.

The exponential increase in the number of microsatellite launches over the past decade has generated a research gap within the realm of satellite propulsion, needed to maintain the orbit of such satellites and hence increase their lifetime. The research towards the miniaturisation of propulsion systems has hence steadily gained traction within universities and other educational institutions. With this in mind, in collaboration with the Else Kooi Laboratory (EKL), the TU Delft is developing a novel resistojet thruster with integrated Titanium heaters based on Micro-Electromechanical Systems (MEMS) technology, called the Low Pressure Micro-Resistojet (LPM). Using their resistance, the newly fabricated thruster chips were characterised as temperature sensors for the range of 40-140 C. A new interface was designed and manufactured out of Teflon, allowing for the necessary mechanical and electrical connections to the thruster and adhering to Delfi-PQ system requirements. A maximum thrust of 0.37 mN was found during preliminary testing.

With the price of launching satellites remaining high, many companies and universities are turning towards smaller satellites as more affordable options. This requires the miniaturisation of their different components, including the propulsion subsystem. At TU Delft, Vaporising Liquid Microthrusters (VLM) are being developed, with the latest iteration being a new generation Micro-Electromechanical System (MEMS) design. In this work, the updated design of this microthruster is remanufactured and prepared for future operational validation, heat transfer and instabilities studies. This is done by simplifying the manufacturing process, creating a testing interface, and performing measurements and extensive experimental characterisation of the thruster and test set-up. The newly-created interface allows for the study of heat transfer efficiency and convection heat transfer coefficient (HTC) from the silicon wall to the propellant, as well as multiple ways to study the instabilities in the system. One interface + thruster assembly was shown to survive to at least 180 °C temperature and 6-7 bar relative pressure. On the manufacturing side, the fabrication process was streamlined by reducing the number of steps by 13%, most significantly by the use of a soft photolithography mask instead of a hard silicon dioxide one. The deviations in horizontal dimensions are in the same order of magnitude as the ones obtained by the previous student working on this project, and lower than other previous attempts at the Faculty of Aerospace Engineering. The etching process encountered some difficulties, leading to inconsistencies in the feature depth between wafers. Across wafers, etching rate deviations as large as 8% were observed. Within the final production wafer, the range of deviation from the design values on throat depth was from +3.5% to −16.5%, showing a measurable higher etching rate at the edges of the wafer than in the centre. Recommendations were given on which features’ depths are most important to be measured, in order to increase the confidence in the predicted values. With these results, future researchers are ready to use the produced thrusters and testing interfaces. It is recommended to continue with the experimental studies, especially by measuring the heat transfer efficiency, thrust levels, HTC, and by studying any instabilities that could appear.

Rocket Factory Augsburg is developing a launch vehicle that uses oxidiser rich staged combustion engines. The Helix engine is ignited with a single-shot hypergolic ampoule containing TEA-TEB. The goal for RFA is to be able to re-light the Helix engine in flight, making it possible to recover the first stage, and to extend the payload capabilities of the second stage. In order to achieve this, its single-shot ignition system has to be modified to allow for multiple in-flight ignitions. This thesis covers the modelling of the ignition fuel flow, a trade-off between existing re-ignition options for hypergolic ignited engines, and sizing of a re-ignition system for the current Helix engine.

The increasing need for miniaturization of radial pumps has been expressed across a variety of fields including rocket propulsion, which has driven research into their optimization. The literature provides design recommendations aimed at mitigating the adverse effects associated with the low Reynolds regimes of these small-scale pumps. This study applies some of these adjustments to a small, high-speed turbopump equipped on the Navier rocket engine. A sensitivity analysis is performed on the outlet total pressure and efficiency of the pump in order to gauge the effectiveness of the design changes. It is found that the impeller outlet width has the most on influence on the pressure, while the efficiency is most sensitive to the pump’s inlet diameter. A regression on the sample points generated for this analysis is then performed, becoming the basis for a surrogate model used in a simple optimization routine of the pump impeller. An optimized version of the Navier pump is obtained, with an outlet pressure fulfilling the engine’s requirements, and an increased overall efficiency.

The utilization of CubeSats as a cost-effective solution for satellite missions allows universities to engage in outer space research. Notably, the issue of increased tumbling rates during deployment is highlighted, attributed mainly to propulsion and guiding functions performed by existing contact actuators.
The proposed solution in this thesis employs viscous traction actuators, which manipulate a thin air film to generate both propulsion and guiding forces. The preliminary design of propulsion and guiding functions incorporates these actuators on the inner sides of the deployer case. The design requirements are established as a criterion to evaluate the redesigned propulsion and guiding functions, including minimal exit velocity, maximal tumbling rate, and maximum acceleration. Moreover, the inner envelope accommodates a 3U CubeSat, and estimation for air reservoir volume and pressure ensure adequate space for launch deployment.
Performance simulation is proposed to evaluate whether the design satisfies the design requirements. It consists of the simulation of the viscous traction actuator and the full model simulation of the deployer. For the simulation of the viscous traction actuator, the finite element method is used to calculate load capacity and viscous traction applied by the viscous traction actuator as input to the full model simulation of the deployer. The full model simulation of the deployer assesses CubeSat motion and propulsion force. The successful fulfillment of design requirements and safe deployment are confirmed.
While the preliminary design meets design requirements, a design method is introduced to cater to varied exit velocity and tumbling rate preferences. This approach adjust design parameters based on a propulsion force vs. time plot. Two distinct designs are presented based on this approach, catering to differing preferences.
The selected +EPC design, characterized by low tumbling rates, progresses to a detailed mechanical design encompassing deployer case, actuator, displacement sensor, and air-supplied system. The implementation of the viscous traction actuator follows established research, ensuring structural safety. Other aspects like displacement measurement, air reservoir selection are also discussed in this detailed design.

Future Mars exploration missions require safe and controlled landing on the planet’s surface. Conventional entry, descent and landing (EDL) technologies, such as parachutes and rigid aeroshells, face limitations in meeting increasing demands for heavier payloads, harsher entry conditions, and desired landing locations due to their limited deployment windows or geometry constraints set by current launchers. The stacked-toroid inflatable aerodynamic decelerator (IAD) has emerged as a promising EDL technology which has the potential to enable new and ambitious applications. Unlike conventional aeroshells, it utilizes flexible, high-temperature resistant materials that can be folded during orbital injection and transportation, and subsequently deployed before entering the Martian atmosphere. To address the complex interdependence of design variables and the multidisciplinary nature of stacked-toroid analysis, this research proposes a novel Multidisciplinary Design Analysis and Optimization (MDAO) framework. The framework integrates aerodynamics, aerothermodynamics, structural analysis, mass estimation, and Flexible-Thermal Protection System (F-TPS) sizing with trajectory simulations for a parametrized stacked-toroid. The major design variables are parameterized to trace model responses back to the design space. An additional novel contribution is the inclusion of a smaller torus on the IAD’s shoulder. For aerodynamics and aerothermodynamics modelling, the local-inclination panel method is implemented, separately addressing the continuum, rarefied, and transitional flow regimes using well-established analytical methods and bridging functions. The scalloping phenomenon of the deflected surface is accounted for through semi-empirical expressions and experimentally-fitted correlations, capturing the additional aerodynamic and aerothermal contribution. F-TPS sizing employs a 1D Finite Difference Method (FDM) with harmonic weighted averaging of material characteristics to accommodate abrupt thickness variations. Results from each discipline are compared to experimental, flight, and high-fidelity numerical data, showing consistent agreement under various conditions. All disciplines present mean percentage errors in the order of 10-20% which are deemed acceptable for early design stages. The framework is applied to the ESA MiniPINS study, to demonstrate its applicability to a novel EDL architecture for a penetrating probe. The proposed environment efficiently evaluates the stacked-toroid optimum design for minimum mass, weighing only 3.72 kg whilst complying with mission requirements and optimisation constraints. The design remains robust against variations in entry parameters and atmospheric density, requiring minor adjustments for the penetrator impact speed. The framework enables design space exploration, revealing trends favoring stacked-toroids with low inner torus radii and large numbers of tori to minimize aerothermal loads while ensuring sufficient aerobraking. The inflated radius remarkably results in a global variable that can further simplify the design space to a single input for near-optimum rapid evaluations. The feasible parameter ranges identified through the design space search expedite the evaluation and optimization of stacked-toroids’ multidisciplinary performance, aiding decision-making in early design stages of future Mars missions.

Quantifying the exergy requirement of propellant insertion into LEO can lead to insight into the feasibility of a lunar propellant-generating architecture. Spacecraft leaving from Earth can greatly reduce their lift-off mass if in-space refueling would possible. Exergy analyses quantify the available energy of a system and show where a reduction in usable energy occurs. Insight into the exergy destruction and input provides a key parameter into the scaling and design of processes and corresponding power systems. The present study aims to define an exergy environment in the lunar PSRs and then to analyze the exergy destruction related to the production of oxygen, ALICE, and hydrolox, in terms of both manufacturing and transportation using a two-stage fully reusable rocket. Extraction processes for ALICE and hydrolox were selected and analyzed w.r.t. the lunar environment to get an understanding of the exergy input. The behavior and exergy requirements of an LEO propellant depot was described. Two fully reusable two-stage rockets using ALICE and hydrolox were designed and compared based on their payload-to-propellant ratio for the oxygen, ALICE, and hydrolox payloads. The study found that the exergetic cost for the insertion of oxygen, hydrolox, and ALICE in LEO were 1.32 GJ/kg , 1.64 GJ/kg, and GJ/kg and 23.3 GJ/kg, 23.4 GJ/kg and 26.9 GJ/kg for the hydrolox and ALICE rocket, respectively.

CubeSats suffer from low reliability and have little to no Fault Detection, Isolation and Recovery mechanisms onboard. Advanced CubeSat missions such as the Lunar Meteroid Impact Observer (LUMIO), will use more complex attitude determination and control systems, increasing the need for advanced fault detection. Traditionally this requires model-based fault detection methods which are complicated, computation heavy, and highly sensitive to disturbances. Machine learning has proven proficient at fault detection in several non-space related applications, but training data including spacecraft faults is not available. In this research an especially lightweight unsupervised learning method for fault detection is designed for the LUMIO attitude determination and control
components. The result is a system capable of detecting artificially induced bias, calibration error, and drifting measurement faults on the scale of 0.1 mrad/s
in the IMU, with no false alarms being raised. The method was tested on simulated LUMIO telemetry from the IMU and reaction wheels as well as on real spacecraft telemetry from the OPS-SAT sun sensor, star tracker, reaction wheels and IMU. In both cases, excellent fault detection and false alarm performance was observed indicating the potential of this method for application in CubeSat AOCS fault detection and isolation.