Spacecraft and launcher development has recently focused on new design concepts employing intelligent propulsion systems, enabled by advanced AI-based paradigms for operations and condition monitoring (OCM) and structural health monitoring (SHM). The main challenge, however, remains providing abundant sensing data points to ensure reliable OCM and SHM processes for effective onboard systems control. This paper presents a case study on a smart spacecraft propellant tank prototype manufactured through carbon fiber filament winding and additive manufacturing of carbon fiber-reinforced polymer, using a Distributed Optical Fiber Sensor (DOFS) system. The optical fiber sensors (OFS) embedding technique is discussed, highlighting methods to optimize temperature isolation from strain variation effects. Composite structure post-processing considerations are also addressed for compatibility with acrylate-coated fibers. Thermal test results, using a high-backscattering OFS interrogated by a Luna ODiSI-6000 OFDR system, are presented.

AbstractThis study investigates the feasibility of using space debris as a supplemental resource for Lunar infrastructure, with a particular focus on the mission design and energy requirements of debris transfer operations. While recycling methods themselves remain at a conceptual stage, this work establishes a technical baseline for how orbital debris—specifically upper stages in GTO could be captured and transported for Lunar processing. The analysis highlights the central challenge of orbital transfer alignment under long-term perturbations and evaluates multiple capture and transfer scenarios, comparing them against direct material delivery missions. Both chemical and electric propulsion architectures are assessed, demonstrating potential energy savings of up to 30 % per kilogram of material, with further reductions when rideshare configurations are employed. By quantifying the mission energy expenditure, this study clarifies the role that efficient transfer design can play in making debris recycling a viable supplement to In-Situ Resource Utilization and reducing reliance on costly terrestrial launch. The results are intended to inform future research on processing methods by first establishing the transfer architectures under which recycling missions could realistically operate.

Designing high-performance and aerospace-grade heat exchangers requires detailed characterization of the as-manufactured geometry, including cross-sectional area and surface texture, to reduce uncertainties in performance prediction and issues regarding subsequent system integration. This paper presents experimental testing and analysis of microchannels fabricated using the laser powder directed energy deposition (LP-DED) additive manufacturing (AM) process. Research has shown that as-built surfaces result in differential pressure higher than what is predicted with current correlations and surface enhancements may be required for heat exchangers built using AM to meet the desired pressure drop specifications. Various surface enhancement techniques including abrasive flow machining (AFM), chemical milling (CM), and chemical mechanical polishing (CMP), were applied to the internal surfaces of the channels to tailor flow dynamics and induce variations in pressure drop. Based on experimental flow testing, channels processed with surface enhancements provide a tenfold reduction in differential pressure compared to the as-built channels. After testing, the samples were destructively sectioned to obtain geometric and detailed surface texture information. This characterization helped to inform a new prediction method for determining hydraulic diameter and equivalent sand grain roughness, thus reducing the uncertainty of predicted friction factors. The new correlation allows to estimate friction factor and pressure drop with a deviation from the experimental data that is within 20% of their value. The identification of the mechanisms at the basis of the formation of surface texture allowed to categorize distinct aspects related to friction factor ranges: roughness peaks, peak smoothing/reduction, minimized roughness, and combined waviness and valley reduction.

Green SWaP (Green Solar-to-propellant Water Propulsion) is a project funded by the European Innovation Council (EIC) Pathfinder program which aims at developing the core technologies for a new class of in-space water propulsion. Water offers unparalleled handling and storage advantages both on Earth and in orbit, and through in-situ resource utilization on the Moon and other bodies, can become a renewable source of propellant. Moreover, water provides dual-use benefits in future human outposts by serving as radiation shielding and as a working fluid in life support systems. Green SWaP pioneers the direct onboard conversion of water into hydrogen peroxide and gaseous hydrogen using solar energy, yielding a propellant combination with superior storability compared to conventional water electrolysis systems. The project aims to increase the technology readiness levels of all key subsystems, including microgravity water conversion systems, concentration of hydrogen peroxide to rocket grade levels, and the safe storage of hydrogen in an inflatable tank. These propellants are then employed in two innovative thrusters: a 1 N hydrogen Solar Thermal Thruster (STT) for precise attitude control and a 200 N bipropellant engine that uses High-Test Peroxide (HTP) and hydrogen for main propulsion. By integrating these building blocks, Green SWaP lays the foundation for renewable, self-sustaining mobility in space, extends water-based propulsion to higher thrust regimes, and enables new mission architectures leveraging in-situ resources.

The EU-EIC Pathfinder project Green SWaP (Green Solar-to-Propellant Water Propulsion) develops a sustainable in-space mobility architecture that directly converts water into hydrogen (H₂) and hydrogen peroxide (H₂O₂) using solar energy. This approach enables a reusable propulsion system combining a 200 N chemical thruster for primary manoeuvres with 1 N solar-thermal thrusters (Isp ∼500 s) for attitude control. By harvesting energy in orbit and producing propellant onboard, the system enhances operational safety, supports water circularity in space, and reduces dependence on Earth-supplied resources. Such a capability extends spacecraft lifetime, enables in-orbit refueling and in-situ resource utilization (ISRU), and broadens the feasibility of reusable orbital stages. To evaluate this potential, a dedicated mission analysis was performed after a selection process, focusing on a reusable kick-stage concept as a case study. The results provide preliminary sizing of key enabling technologies, such as bi-modal propulsion, inflatable hydrogen storage, and solar-to-fuel conversion, and demonstrate the transformative impact of Green SWaP on sustainable space logistics and future mission architectures.

In the original version of the book, on page xi, one of the authors listed for Chapter 2 is “R. Schnmehl”, which should be “R. Schmehl”. On page 21, the same correction needs to be made twice, in the listed authors at the top of the page and also in the footnotes. “Schnmehl” should become “Schmehl” in both the cases. This has now been rectified and the author’s name has been corrected. The correction to this book have been updated with the changes.

With a growing trend in the miniaturisation of satellites, there is an increasing need to develop micro-propulsion systems for these satellites. Since scaling down conventional propulsion systems is challenging and not always possible, new concepts need to be developed. These concepts, although often based on already known systems and principles, require significant modifications to make them meet the requirements of miniaturized propulsion. One such concept is the micro-resistojet, using an electrical resistance to increase the propellant temperature. At Delft University of Technology, two thrusters based on this concept were developed: the Vaporized Liquid Micro-resistojet (VLM) and the Low-Pressure Micro-resistojet (LPM), which were specifically designed for being demonstrated on-board a PocketQube satellite. To determine the operating regime of the thrusters for this demonstration, there is a need to develop simplified analytical models that accurately predict their performance without significant computational expenses. Although there were previous attempts to model these thrusters, they did not provide a complete representation of their performance. For the VLM thruster, the focus of the model presented in this paper was on coupling the heating chamber and the nozzle, to obtain a more accurate value for the mass flow rate through the thruster. The heating chamber section was discretized into finite one-dimensional cells and convective heat transfer equations were used to model parameters such as density, pressure, wall temperature and heat transfer coefficient. The nozzle was modelled based on ideal rocket theory corrected with adequate loss factors. The mass flow rate was calculated iteratively by coupling the two sections until it reached convergence. For the LPM thruster, the focus was on including an accommodation coefficient to account for heat transfer efficiency between thruster walls and propellant. Rarefied gas dynamics equations were used to calculate performance parameters due to the low-pressure conditions within the thruster. The models proved to produce realistic results when compared to available numerical and experimental values, although still with some limitations in modelling heat transfer, which could not be fully overcome yet due to the lack of available data for validation. Optimal operating points were determined for both thrusters by maximizing an objective function based on performance parameters such as thrust-to-power ratio, specific impulse, and mass flow rate. Constraints included thrust, power, and temperature requirements, which led to different optimal points for the thrusters under varying operational conditions.

The 16U4SBSP mission concept is based on using a swarm of CubeSats to perform a scaled demonstration of Space-Based Solar Power (SBSP) from Earth orbit. In this demonstration mission, seven identical spacecraft of 16U format are used to provide wireless energy in the kW-scale using Radio-Frequency (RF) Wireless Power Transfer (WPT), and the spacecraft in the swarm are designed to be suitable to both space-to-ground or space-to-space WPT applications. The main objective of the mission is to validate the general concept of providing SBSP using a swarm of satellites instead of a monolithic configuration, as well as some of the involved miniaturized technologies, in view of full-scale missions which could serve users in remote areas with low power requirements or support emergency operations in blackout zones affected by unpredicted hazards (e.g. natural disasters). More in general, the mission would represent a low-cost precursor towards MW-GW scale SBSP to supply clean and affordable energy from space to large areas on the Earth surface. A pre-Phase A study of the mission, funded by the European Space Agency (ESA) through the Sysnova campaign “Innovative Missions Concepts enabled by Swarms of CubeSats”, has led to encouraging results on the feasibility of the mission concept.

This paper presents in detail the final outcome of the pre-Phase A design effort for the 16U4SBSP spacecraft. The trade-off studies conducted to select all sub-systems and components are presented and their final outcomes detailed and justified, together with the technical budgets and the main areas of attention for the spacecraft design. Particularly critical for the success of the mission are the choices related to: the power transmission payload (DC-RF converter, transmitting antenna and heat dissipation system); the ADCS subsystem and in particular the sensors required to provide sufficient accuracy in the knowledge of the 3-axis attitude (both absolute and relative to the other spacecraft in the swarm); the relative navigation system, based on inter-satellite link between the spacecraft in the swarm and on a beacon link to the receiving station on ground, for efficient beaming coordination; the main propulsion system for continuous formation flying control through the whole mission lifetime; the electric power system, based on orientable solar arrays by means of a SADA mechanism and a set of batteries with sufficient capacity for beaming the required amount of power while in eclipse conditions.

This research evaluates Laser Powder Directed Energy Deposition (LP-DED) for producing fine feature internal microchannels. This study is focused on enhancing and characterising the surfaces of microchannels produced using techniques such as abrasive flow machining, chemical milling, chemical mechanical polishing, electrochemical machining, and thermal energy method to modify internal surfaces of microchannels made from NASA HR-1 Fe-Ni-Cr alloy. Flow testing for discharge coefficient measurement is conducted on processed microchannel samples, followed by characterisation through optical microscopy, Scanning Electron Microscopy (SEM), and Computed Tomography. Findings reveal variations in surfaces due to powder adherence, melt pool undulations, and polishing mechanisms. The study emphasises the significance of removing material equivalent to the mean powder diameter to reduce surface roughness and impact the discharge coefficient. The research proposes a ratio for planarising roughness and waviness peak height and density, offering insights for tailored surface adjustments in specific applications requiring reduced flow resistance. Highlights Internal microchannels with thin-walls were fabricated using the laser powder directed energy deposition process. Various surface enhancements and polishing processes were developed to modify the surface texture of the LP-DED channels. Flow testing was conducted to determine the discharge coefficient. Post-test characterisation was completed to obtain cross sectional area, perimeter, surface texture, and general surface condition to analyse results. Ratio of roughness and waviness peak and density (Spk/Spd and Wp/WPc) is proposed as a relevant surface characterisation parameter. Tailored surface modifications for specific end-use applications.

Lunar meteoroid impacts have caused in the past a substantial change in the lunar surface. With no atmospheric shield, the Moon is subject to many impacts from meteoroids, ranging from a few grams to a few kilograms. The high impact rate on the lunar surface has important implications for future human and robotic assets that will inhabit the Moon for significant periods of time. Therefore, a better understanding of the meteoroid population in the cislunar environment is required for future exploration of the Moon. Moreover, refining current meteoroid models is of paramount importance for many applications, including planetary science investigations. Studying meteoroid impacts can help deepening the understanding of the spatial distribution of near-Earth objects in the Solar System. The ability to predict impacts is therefore critical to many applications, both related to engineering aspects of space exploration, and to more scientific investigations regarding evolutional processes in the Solar System. The Lunar Meteoroid Impacts Observer (LUMIO) is a CubeSat mission to observe, quantify, and characterise lunar meteoroid impacts, by detecting their impact ashes on the far-side of the Moon. This complements the information available from Earth-based observatories, which are bounded to the lunar near-side, with the goal of synthesising a global recognition of the lunar meteoroid environment. LUMIO envisages a 12U CubeSat form-factor placed in a halo orbit at Earth-Moon L2. The detections are performed using the LUMIO-Cam, an optical instrument capable of detecting light ashes in the visible spectrum (450-950 nm). LUMIO has successfully passed the PDR and is currently moving towards Phase C. We present the latest results on the modelling of the meteoroid environment in the Earth-Moon system, including an estimate of LUMIO's potential impact on our existing knowledge of meteoroids, supported by high-fidelity simulation data. An overview of the present-day LUMIO CubeSat design is also given, with a focus on the latest developments involving both the ongoing/planned scientific activities and the development of the payload.

This study proposes the concept of recycling space debris as a novel means of supplying material resources for the establishment of a permanent Lunar presence while simultaneously cleaning up Earth's orbital environment. Upon the creation of a space debris dataset and characterizing debris objects as resources and reserves, spent Ariane 5 upper stages in GTO are identified as prime candidates for recycling. However, orbital transfer alignment poses a critical challenge due to orbit perturbations over time. Mission scenarios, including debris capture, transfer and Lunar processing, are analyzed, with global mission energy expenditure used to compare them to direct material delivery missions. Both chemical and electric propulsion transfer architectures are highlighted as enabling feasible and efficient recycling mission scenarios, with potential energy savings of up to 30% per kg of material. The significant reduction in launch mass as a direct consequence of capturing the mission payload in orbit allows for the inclusion of rideshare configurations, increasing efficiency to over 60% less energy investment per kg.

Modern spacecraft and launch vehicle design is more oriented towards reducing system-level design and assembly complexities. In order to maintain high overall system performance while reducing these complexities, the use of smart materials and smart structural components is a well-known practice and is currently of rising interest to space systems' designers. The paper discusses a concept of smart space structures, in particular, a carbon fiber composites structure embedded with Optical Fiber Sensors (OFS) for spacecraft and launch vehicle applications. This study highlights the operational requirements for such tank and the smart features enabled by the optical fiber sensors. For the latter aspect, a quantitative comparison between Fiber Bragg Grating sensors (FBGs) and Distributed Optical Fiber Sensors (DOFS) based on Optical Frequency Domain Reflectometry (OFDR) is presented to state their core performance parameters, such as the sensitivity, sensing range, dynamic measurement capability, and spatial resolution. The increased performance and reliability in harsh environments associated with fiber optic sensors come with a reduction in size, mass, and power consumption compared to the conventional electronic sensors. Optical fiber sensors embedded in carbon fiber structures have proven their capability in providing accurate real-time measurements of temperature and monitoring structural integrity while detecting precisely possible points of rupture and failure as discussed and demonstrated in the literature review. The applications of fiber optic sensing in smart propellant tanks may extend to detecting fluid leakage, also providing increased precision in propellant gauging through temperature mapping, and can be used in on-ground qualification, pre-flight testing, as well as in-orbit operation, condition, and structural health monitoring. The article presents a statement for an optimal FOS embedding approach in composite pressure vessels and discusses the related placement and orientation method for the fiber optic sensors, coupled with a one component simplified analytical stress-strain transfer model deriving the stress component along the maximum principal direction (i.e., σ_MaxPrincipal). The novel approach is believed to serve the optimal employment of embedded FOS in composite structures, e.g., pressure vessels and light-weight structures in spacecraft, among other applications. The simplified model is believed to pave the way for effective data interpretation and processing, utilizing the available limited computational resources on-board the spacecraft.

The 16U4SBSP mission concept is based on using a swarm of CubeSats to perform a scaled demonstration of SpaceBased Solar Power (SBSP) from Earth orbit. In this demonstration mission, seven identical spacecraft of 16U format are used to provide wireless energy in the kW-scale using Radio-Frequency (RF) Wireless Power Transfer (WPT), and the spacecraft in the swarm are designed to be suitable to both space-to-ground or space-to-space WPT applications. The main objective of the mission is to validate the general concept of providing SBSP using a swarm of satellites instead of a monolithic configuration, as well as some of the involved miniaturized technologies, in view of full-scale missions which could serve users in remote areas with low power requirements or support emergency operations in blackout zones affected by unpredicted hazards (e.g. natural disasters). More in general, the mission would represent a low-cost precursor towards MW-GW scale SBSP to supply clean and affordable energy from space to large areas on the Earth surface. A pre-Phase A study of the mission, funded by the European Space Agency (ESA) through the Sysnova campaign “Innovative Missions Concepts enabled by Swarms of CubeSats”, has led to encouraging results on the feasibility of the mission concept. This paper summarizes the main results of the 16U4SBSP pre-Phase A study, including the mission design and the possible way forward to the following steps in mission implementation (Phase A and later). A summary of the spacecraft system design and mission analysis is presented, as well as a short description of the mission CONOPS (concept of operations). Particular focus is given to the available options and objectives of the SBSP demonstration, and to the proposed solutions for RF power beaming, formation flying maintenance and end-of-life operations for compliance to the new ESA regulations on space debris mitigation.

Establishing a permanent lunar base has gained increasing attention since it offers opportunities for international cooperation and the commercialization of space, forming the foundation and testing ground for a human existence independent from Earth. Essential to future missions beyond cislunar space is the exploration and in situ processing of the Moon's resources, especially the sustainable production of energetic resources and propellants. Utilizing in situ generated propellants can dramatically reduce transportation costs by removing the need to source propellants from Earth. Resources on the Moon are limited, and the extraction of available resources are energy-intensive processes demanding advanced techniques and technologies. Consequently, one of the biggest challenges lies in developing process architectures with a positive energy balance, for which comprehensive analyses are still missing. The focus currently lies on the extraction of water ice from lunar regolith and the production of hydrogen and oxygen through water electrolysis. However, alternative fuel and process options may reduce the energy cost while providing equivalent energetic revenue. In the scope of this research, the infrastructure and technologies required for extraction, refining, and storing are assumed to exist in cislunar space; therefore, only the operating cost is considered. Exergy analyses of in situ extraction methods are conducted to investigate whether the required energetic budget allows sustainable implementation. The analysis includes extraction methods and propellant options to reveal the extent to which alternatives to hydrogen are feasible. Exergy analyses determine thermodynamic losses of energy flows giving the ground for process optimization. The exergy destructed represents the margin of improvement within the process architecture and thus reflects the process's thermodynamic and economic value while allowing a more distinct examination of energy use. Assuming the availability of water and carbon dioxide ice in permanently shadowed regions, the analysis shows that choosing methane instead of hydrogen in combination with oxygen as propellants can reduce the required exergy input by up to a third. An example mission allows to directly compare the operating cost of the extraction processes for the different propellant options. The mission entails a spacecraft propelled by a liquid bipropellant engine utilizing the extracted propellant and transporting a payload of the same propellant to a depot located in lunar near-rectilinear halo orbit (NRHO). Although abundant in space, the results suggest that hydrogen may not be the only or even energetically cost-effective resource for developing cislunar and Martian space infrastructures. Likewise, sustainable extraction of propellants suitable for current and future propulsion systems will foster humanity's reach further into the solar system.

The 16U4SBSP mission aims to demonstrate Space-Based Solar Power (SBSP) using a CubeSat (CS) swarm from Earth orbit. This mission employs seven 16U CSs to deliver 1 kW-scale wireless energy via Radio-Frequency (RF) beaming, adaptable for space-to-ground and space-to-space applications. The goal is to validate SBSP provision using a satellite swarm and to explore miniaturized technologies for future large-scale missions. A pre-Phase 0 study funded by the European Space Agency (ESA) through the Sysnova campaign has shown encouraging feasibility results. This paper presents a study on the formation flying and orbital dynamics of a CS mission, using a model that includes Earth’s gravitational perturbations, solar radiation pressure (SRP), atmospheric drag, and lunar and solar gravity. The swarm configuration consists of seven CSs, with one at the center and six in a hexagonal arrangement. The Concept of Operations (CONOPS) is divided into three phases: deployment and acquisition, maintenance, and disposal. CSs are deployed at 30-second intervals, followed by a one-day Launch and Early Orbit Phase (LEOP) for subsystem checks. A 1000-meter formation is initially established, then reduced to 100 meters for the first half of the mission and 10 meters for the second half, maintained by a bang-bang limit-cycle controller. A disposal strategy compliant with ESA’s Space Debris Mitigation Requirements is outlined. The analysis characterizes propellant consumption at various altitudes, proposes optimal initial conditions and launch dates, and performs a trade-off analysis, resulting in a detailed mission characterization and baseline definition. The work presented in the paper proves the feasibility of the 16U4SBSP mission, which would supply clean energy from space through wireless power transfer.

In order for off-Earth top surface structures built from regolith to protect astronauts from radiation, they need to be several metres thick. In a feasibility study, funded by the European Space Agency, Technical University Delft (TUD aka TU Delft) explored the possibility of building in empty lava tubes to create rhizomatic subsurface habitats. With this approach natural protection from radiation is achieved as well as thermal insulation because the temperature is more stable underground. It involves a swarm of autonomous mobile robots that survey the areas and mine for materials such as regolith in order to create cement-based concrete reproducible on Mars through in-situ resource utilisation (ISRU). The concrete is 3D printed by means of additive Design-to-Robotic-Production (D2RP) methods developed at TUD for on-Earth applications with the 3D printing system of industrial partner, Vertico. The printed components are assembled using a Human--Robot Interaction (HRI) supported approach. The 3D printed and HRI-supported assembled structures are structurally optimised porous material systems with increased insulation properties. In order to regulate the indoor pressurised environment a Life Support System (LSS) is integrated, which in this study is only conceptually developed. The habitat and the D2RP production system are powered by an automated kite power system and solar panels developed at TUD. The long-term goal is to develop an autarkic, automated and HRI-supported D2RP system for building autarkic habitats from locally obtained materials.

Interplanetary CubeSat missions are currently becoming more popular, with a significant number of recently planned missions. The context of this paper is a Mars mission, starting from a parking orbit around Earth: the adoption of a chemical propulsion system for the Earth-Mars transfer phase is investigated, considering the recent technological developments for CubeSats. A trade-off of propulsion system type and propellant results in the choice of a mono-propellant system adopting the HAN-based propellant AF-M315E (ASCENT). The main challenge for the propulsion system is to fit inside a CubeSat standardised volume, which can range up to 24 U, for which the implementation of a suitable COTS micro-pump is considered. Finally, the complete architecture and design of the propulsion system is presented. This work demonstrates the feasibility of adopting full chemical propulsion for an interplanetary CubeSat mission, with consequent advantages in terms of transfer time and required power, at the cost of relatively small mass and volume left for the other subsystems. Even better results can be expected for interplanetary missions requiring slightly lower AV budgets, such as Near Earth Objects exploration or asteroid fly-by missions.

This manuscript aims to present and evaluate the applicability of combining optical line-of-sight (LoS) navigation with crosslink radiometric navigation for deep-space cruising distributed space systems. To do so, a set of four distributed space systems architectures is presented, and for each of those, the applicability of the combination is evaluated, comparing it to the baseline solutions, which are based on only optical navigation. The comparison is done by studying the performance in a circular heliocentric orbit in seven different time intervals (ranging from 2024 to 2032) and exploiting the observation of all the pairs of planets from Mercury to Saturn. The distance between spacecraft is kept around 200 km. Later, a NEA mission test case is generated in order to explore the applicability to a more realistic case. This analysis shows that the technique can also cope with a variable inter-satellite distance, and the best performance is obtained when the spacecraft get closer to each other.

Metal additive manufacturing (AM) is being used for mission-critical applications in both developmental and production components, driven by economic and technical benefits. Laser powder directed energy deposition (LP-DED) allows manufacturing of thin wall geometric features for various components at diameters larger than 2 m. The characterization of geometric capabilities and limitations is critical for establishing guidelines for end users of the technology. Within this study, several samples of enclosed vertical tracks were fabricated and characterized using LP-DED, with 1 mm-thick walls and varying inclination angles up to 45° using the NASA HR-1 alloy (Fe-Ni-Cr). The wall thickness, melt pool, and surface texture, inclusive of waviness and roughness, were evaluated and results presented. The experimental results indicate that the wall thickness increases exponentially above 30°. The surface texture was shown to be dependent on 1) excess powder adherence, 2) melt pool irregularities causing material droop, and 3) excess material. The experiment revealed that the mean roughness reduces with increasing wall angle for the downskin surface. The upskin roughness reaches a maximum peak at 20° and slowly reduces as powder adheres within the valleys. Both the downskin and upskin surface textures are dominated by irregular waviness generated by the melt pool.

Line-of-Sight (LoS) navigation is an optical navigation technique that exploits the direction to visible celestial bodies, obtained from an onboard imaging system, to estimate the position and velocity of a spacecraft. The directions are fed to an estimation filter, where they are matched with the actual position of the observed bodies, retrieved from onboard stored ephemerides. As LoS navigation represents a really promising option for the next-generation deep-space spacecraft, the objective of this work is to provide new insights into the performance. First, the information matrix is analyzed to show the influence of the geometry between the spacecraft and the observed planet(s). Then, a Monte Carlo approach is used to investigate the influence of measurement error (ranging from 0.1 to 100 arcsec), and tracking frequency (ranging from four observations per day to one observation every two days). The effect on navigation performance is quantified by two indicators. The first is the 3D position and velocity Root-Mean-Square-Errors, computed once the estimation is considered to be steady-state. The second is the convergence time, which quantifies the required time for the estimation to reach the steady-state behaviour. The simulation is based on a set of four planets, which do not follow the common heliocentric dynamics but rotate around the Sun with the same (distance-independent) angular velocity of the spacecraft. This approach allows the separation of scenario-dependent behaviours from navigation intrinsic properties, as the same relative geometry between observer and observed objects is maintained during the whole simulation. The results provide a useful guide for the next-generation autonomous navigation system, for both the definition of hardware requirements and the design of an appropriate navigation strategy. Considerations are then applied to Near-Earth Asteroid fly-by mission scenarios for the definition of the navigation strategy and hardware requirements. It is shown the importance of relative angles between the spacecraft and the planets. In the single-planet observation scenario, when the angle between the position vectors of the spacecraft and planet approaches a null value, the estimation error decreases. In the double-planets observation scenario, when the separation angle between the two LoS directions gets close to 90°, the estimation error decreases. The main influence on the performance is driven by the measurement error, which with current technologies is shown to be able to provide a position error in the order of a few hundred kilometers, while with a lower measurement error (0.1 arcsec) it would be possible to have a position error below 100 km. Finally, it is demonstrated that tracking frequency plays a secondary role in the performance, and only influences tangibly the convergence time.