The accurate measurement of inter-satellite distances is fundamental to the successful operation of distributed spacecraft missions, facilitating diverse applications ranging from scientific exploration to navigation and communication. This study comprehensively overviews applications requiring inter-satellite tracking by analyzing 44 multi-spacecraft missions. These missions are divided into seven categories based on their use of inter-satellite distance measurements. Each category necessitates varying levels of accuracy, prompting the utilization of distinct tracking methods. The analysis reveals that missions near Earth typically rely on Global Navigation Satellite Systems measurements, achieving millimeter-level accuracy, while lunar missions opt for radio ranging for centimeter-level accuracy. Inter-satellite Laser Ranging Interferometry emerges as the preferred method for missions demanding exceptionally high accuracy (nanometer to picometer range), such as those dedicated to gravitational wave detection and gravimetry. Notably, the analysis identifies a burgeoning trend towards the adoption of Inter-satellite Laser Transponder Ranging, capable of achieving sub-millimeter accuracy. Furthermore, this work proposes a novel concept: integrating inter-satellite ranging with Laser Communication Terminal (LCTs), either to substitute for existing tracking methods or to enhance measurement accuracy within established frameworks. However, its full potential rests upon the successful adaptation of existing LCTs for ranging functionality without compromising data rates. Future research will play a critical role in quantifying achievable ranging performance by characterizing systematic errors within LCTs.@en

The Da Vinci Satellite is a 2U CubeSatellite that is being developed with the goal to inspire and enthuse the youth to learn more about technology and space travel. The team at Delft University of Technology consists of over 80 Bachelor and Master students from disciplines such as Aerospace Engineering, Computer Science, Electrical Engineering and Applied Physics. There are also multiple Precision Engineers part of the team, that have joined from the Leiden Instrumenten Maker School (LIS) in the Netherlands. Through working on the satellite and educational modules, the multidisciplinary team focuses on demystifying space and making it a fun and engaging subject for children in primary schools and high schools. That is why the satellite harbours two novel custom-made payloads on board; the Dice Payload and the BitFlip Payload.

The Dice Payload consists of a small chamber with five small aluminium dice of different colours, which will be used by primary school children. In collaboration with the LIS, a special mechanism has been designed to ‘roll the dice’ in microgravity and clamp them such that a picture of the numbers can be taken with the Earth as a backdrop. After the design, manufacturing, and assembly of the parts, the payload underwent a series of tests. These tests have included multiple 0g flight tests and a vibration test. Through the extensive testing, there have been iterative design changes to improve the payload’s overall performance and design.

The second payload of the Da Vinci Satellite is the BitFlip Payload. This novel payload recently has been tested in a proton accelerator facility at the Paul Scherrer Institute. This subsystem is a stack of PCB’s with SRAM that has been designed for high school students. High school students can send a picture of themselves to the satellite where the data will be stored on the SRAMs. Because of the radiation environment in LEO and the susceptibility of the memory, bitflips will occur. These changes in the information from a 1 to a 0 (or the other way around), will result in the information of the picture being changed. When the picture has been compressed using a compression algorithm such as GIF, JPEG, or PNG interesting effects can occur. Once the student will receive the altered picture, they will be able to compare it with the original and learn about space, radiation, compression algorithms, and electronics.@en

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.@en

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.@en

In recent years, the trend towards satellite miniaturization has led to a considerable rise in picosatellite missions in Low Earth Orbit (LEO). Due to their size, identifying and tracking small objects poses significant challenges. While the detectability of individual picosatellites has been proven, the potential implications of clusters flying in formation remain unexplored. Therefore, the Delft University of Technology is starting a pioneering mission involving multiple picosatellites to improve Space Situational Awareness (SSA) by demonstrating the capabilities and limits of both inspace and ground-based tracking means. This paper outlines a high-level mission analysis investigating the feasibility of this formation-flying mission: in the proposed model, the autonomy of each satellite is enhanced by integrating a Global Navigation Satellite System (GNSS) receiver, which enables independent orbital determination and facilitates the validation of satellite position data against other tracking systems. The mission concept involves deploying a cluster of two identical 3P PocketQubes, launched as a single spacecraft into a near-circular orbit. Following deployment, they will be separated using springs, considering factors such as relative velocity, direction, and angle to carefully study the release process. Additionally, their relative distance is controlled using differential drag, i.e. adjusting the satellite drag area by deploying solar panels at varying angles. Through the integrated use of STK and MATLAB, two mission control sequences are defined: the former, characterized by satellite propagation stopping conditions based solely on relative distance, and the latter, in which they are based on both relative distance and relative velocity. Their comparison reveals the latter as the most effective strategy: despite the challenge of controlling satellite distance during close passes (conjunctions), the optimal sequence prioritizes maximizing time in close proximity. This approach reduces the average relative velocity and minimizes the duration of the high drag configuration, resulting in a significant extension of the mission lifetime. The simulations, along with a power budget, support the definition of both mission and GNSS receiver payload requirements. Finally, a suitable candidate is selected from among miniaturized GNSS space receivers and tested through multiple hardware-in-the-loop simulations, using a GNSS signal simulator. These simulations aim at verifying the accuracy of receiver positioning measurements and assessing power consumption. The results of this paper represent the cornerstone of a disruptive mission, providing insights into the future development of satellite control optimization strategies to minimize collision risk. Furthermore, the remarkable payload test results, while reliable, underscore the need to improve testing systems to reduce position errors and achieve higher tracking accuracy for LEO picosatellites equipped with GNSS receivers. @en

The need for higher data rates has transitioned the satellites towards laser communication, opening up new opportunities for inter-satellite distance measurements. The stringent requirements for space missions like gravimetry, formation flying, collision avoidance, and precise orbit determination require highly precise distance knowledge, which can be offered by lasers. The past and present missions depend on radio ranging and laser interferometers to achieve up to centimeter-order and picometer-order precision, respectively. However, these methods either require additional hardware or interfere with the communication data rates as the power available is shared between communication and ranging operations. Therefore, this research explores the potential of laser communication terminals in measuring the inter-satellite distance. Numerical simulations are performed to analyze the effect of the precision of inter-satellite distance measurements on atmospheric density estimation. The analysis shows that such range measurements can only improve atmospheric density estimates if the uncertainty in the drag coefficient can be reduced below the current range of 3-5%.@en

The Delft University of Technology has been working on Delfi–PQ, a 3P PocketQube developed by Aerospace Engineering students during their education. The satellite, while being only 50x50x178 mm and having a mass of 545 g, shares the same problems and requirements as larger satellites. Delfi–PQ was launched on January 13th, 2022, and stayed operational till it decayed on January 9th, 2024. This paper presents the design concept, development, and cost of Delfi–PQ to help other teams in their development. This paper will provide a detailed overview of the hardware cost of Delfi–PQ. This cost breakdown will be from multiple angles; system costs, components, structural pieces, and the cost of a specific function in a subsystem. Such a detailed breakdown can be used for future satellite cost modeling and create a foundation for other institutions for their satellite projects.@en

Current monitoring systems to detect sporadic E use ground-based setups, ionosondes, and the network of GNSS satellites in order to assess the phenomenon of sporadic E. This paper aims to monitor sporadic E using a miniature space-based platform in an atypical way. The setup consists of multiple radio-amateur beacon systems aboard satellites, each having a specific modulation and transmission scheme. This Radio Amateur Beacon System for the Investigation of the Ionosphere (RABSII) is coupled to a GNSS receiver, revealing the location of the platform. Multiple beacon data streams are sequentially sent from a satellite platform towards the Earth. By receiving and comparing the Signal-to-Noise ratios of these streams using a dedicated ground-based radio-amateur network of receiving stations, the presence of sporadic E can be determined, and a location-based model can be built. The advantage of this miniaturized, low-power, low-cost instrument is its ability to be put on any satellite platform in the future in order to map sporadic E.@en

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.@en

The last years saw the diffusion of nano, pico and femto satellite missions launched by multiple entities thanks to the launch cost reduction and the electronics miniaturization. Such missions usually present limited capabilities in terms of precise orbit determination and extremely small radar and optical cross-sections. Often these missions carry one or more laser retro-reflectors for precise orbit determination but precise orbital measurements cannot be found in the literature. Miniaturized GNSS receivers are also often carried out but due to the experimental nature of such missions, the reliability and time span of such measurements is limited, leaving radar tracking as the only reliable tracking method. Due to the size of such satellites, the signal-to-noise ratio of such radar measurements is typically low and satellite identification (when launched on ride-share launches with a hundred or more other satellites) proves difficult and time-consuming. Being these very small satellites at the edge of the radar detection capabilities and not providing independent orbit determination means, their position uncertainty could be quite significant, leading to an increased orbit collision perceived risk. With this paper, we present a dedicated small satellite formation, made by multiple nano and pico satellites to evaluate the space surveillance network tracking capabilities and limits. The formation is made by a 3U CubeSat to be deployed as part of a rideshare launch. The satellite would be equipped with multiple means to track it, including a GNSS receiver, a set of multiple laser retro-reflectors, and LEDs for optical, laser, and radar tracking, allowing to characterize also different detection means in terms of capabilities. Such a satellite is made of two independent smaller satellites that can be un-docked in orbit upon command, reducing the satellite size and cross-section. This would push the detection limit for the space surveillance networks starting from an already acquired object and with limited clutter around it. Independent laser and GNSS tracking would allow ground measurement validation and validate position estimations. Further pico-satellites would be deployed by each sub-satellite to further push the detection limits and validate up to which size objects are trackable (still optically, radar and GNSS), thanks to miniaturized GNSS receivers already flown by several other missions. Sub-satellite separation is implemented upon command to ensure the process can be followed and executed at lower altitudes to limit the orbital lifetime of eventually hard-to-track small objects that could worsen the space debris problem. Ground characterization (in terms of optical and radar properties) will be performed, also including polarimetric measurements used to identify the separate satellites. All these technologies together would contribute to creating a unique tool to estimate the tracking capabilities of multiple instruments, specifically tailored for very small objects, the hardest to track, as compared to other characterization activities performed on much bigger objects.@en

Spaceborne Global Navigation Satellite Systems (GNSS) receivers have become ubiquitous sensors for spacecraft navigation, especially in Low Earth Orbits (LEOs), often also supporting science endeavors or as acting dedicated science payloads. Due to the large number of space-capable GNSS receiver models available, spacecraft designers, as well as scientists, may find it difficult to have or gain an overview of suitable state-of-the-art models for their purposes and constraints. Based on a literature review that included more than 90 different receiver models, this paper aims to provide an overview of space-capable GNSS receivers that have a heritage in space missions. It analyses trends from the collected data and provides an outlook on miniaturized GNSS receiver models, which have a high potential of being used in future space missions.@en

The Da Vinci Satellite project is a non-profit initiative started at the Delft University of Technology to inspire and enthuse the youth to learn more about technology and space travel. The team does this by focussing on demystifying space and making it a fun and engaging subject. The non-profit student team is divided into different sub teams, two of which are the technical team and the educational team. The technical team has been building a 2U CubeSat with two payloads that have been designed to support educational packages for children from primary schools and high schools. The educational team works to make these educational modules available for schools all around the world such that children have the opportunity to interact directly with space via The Da Vinci Satellite. @en

The Delft University of Technology has been working on Delfi-PQ, a 3 P P ocketQube d eveloped b y Aerospace Engineering students during their education. The satellite, while being only 50x50x178 mm and having a mass of 545 g, shares the same problems and requirements of bigger satellites. This paper presents the design concept, development, and testing of Delfi-PQ to help other teams in their development. All the combined information will help to generate a big picture for institutions to start their own small satellite mission.@en

In recent years, there has been a growing interest in lunar missions, particularly with the growing role of small satellites facilitated by piggyback launch opportunities. Typically, ground-based radiometric tracking is the workhorse to establish the necessary navigation solution in these missions, however, this could be expensive, while small satellites development is expected to be at low cost. To address this challenge, autonomous navigation presents a potential solution: this study explores the satellite-to-satellite tracking-based autonomous on-board orbit determination method for a satellite formation in cislunar space. Several factors affect the performance of orbit determination, and one critical aspect is the timing of tracking windows. Basically, it is crucial to determine when to collect the most useful observations to optimize the outcome of the navigation filter. In some cases, there might be operational constraints such as inter-satellite distance due to the limited onboard power for ranging. This study investigates particle swarm optimization-based satellite-to-satellite tracking window planning. The findings of this work demonstrate that particle swarm optimization offers a near-optimal solution for tracking windows, taking into account constraints arising from the spacecraft itself or from other design choices. In summary, particle swarm optimization provides near-optimal tracking windows by minimizing the overall orbit determination error. The results presented have the potential to enhance the design of satellite formations performing autonomous on-board orbit determination and contribute to cost- effective mission planning solutions.@en

This chapter provides an overview of the command and data handling system (CDHS) in small satellites and CubeSats. The chapter presents first analysis of radiation effects, specifically targeted at this subsystem, to justify components and architecture choices. Improvements in radiation testing strategies are also presented, specifically for small satellites. State-of-the-art components are then presented, providing an overview of the current market and the most common architectures. An overview of past and current missions is also presented, providing a clear mapping of the presented state-of-the-art components and architectures to guide future designs. High-level design considerations are also presented to help the reader follow some of the current trends in the sector. This chapter, overall, aims at presenting the most common approaches for the CDHS system and comparing this with traditional satellites, showing where the main differences lay with component selection and testing strategies being the fundamental points driving the architecture choices.

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The Earth-Moon system is constantly bombarded by meteoroids of different size and impact speed. Observation of the impacts on the Moon can enable thorough characterization of the Lunar meteoroid flux, which is similar to that of the Earth. While Earth-based Lunar observations are restricted by weather, geometric and illumination conditions, a Lunar-based observation campaign can improve the detection rate and, when observing the Lunar far side, complement in both space and time the observations taken from Earth. The Lunar Meteoroid Impact Observer (LUMIO), one of the two winning concepts of the ESA SysNova Lunar CubeSats for Exploration challenge, is a mission designed to observe, quantify, and characterize the micro-meteoroid impacts on the Lunar far side. It is based on a 12U CubeSat that carries the LUMIO-Cam, a custom-designed optical instrument capable of detecting light flashes in the visible spectrum. The spacecraft is placed on a halo orbit about the Earth–Moon L2 point, where permanent full-disk observation of the Lunar far side can be performed with excellent quality, given the absence of Earth background noise. After passing Phase 0 and an independent feasibility study in the ESA Concurrent Design Facility, the mission has successfully completed its Phase A in March 2021. Although the Phase 0 design of the LUMIO spacecraft was assessed as feasible by the ESA CDF study, a number of critical issues were identified, which have been tackled by the Phase A design. The paper presents the outcome of this Phase A design effort for the LUMIO spacecraft. Particularly relevant changes or updates in the spacecraft design include: a consolidated design of the LUMIO-Cam, with longer baffle for straylight protection; a set of ADCS sensors and actuators with increased redundancy; a combination of Direct-to-Earth communication and inter-satellite link with a mothership in Lunar orbit; use of Earth ranging to complement and validate the current innovative autonomous navigation strategy based on optical observations of the Moon by means of the LUMIO-Cam; re-assessment of the COTS components selection for the power and propulsion systems.

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In this study, the application of the Linked Autonomous Interplanetary Satellite Orbit Navigation (LiAISON) method for cislunar satellite formations is investigated considering range only and range-rate only measurements. The LiAISON method provides an autonomous orbit determination solution solely using satellite-to-satellite measurements such as range and/or range-rate. This paper presents a comparison between range only and range-rate only measurements in satellite formations at cislunar space including the Earth-Moon L!, L" and Lunar orbits by presenting the results of Monte Carlo simulations and observability analysis. The results show that range observations in general provide better state estimations than range-rate observations for cislunar satellite formations in the autonomous navigation applications. However, range-rate only measurements could be an alternative to range-only measurements if range measurements are not precise and high precise range-rate measurements could be collected on-board. It has been found that range only measurements could be good enough to meet the orbit determination requirements for certain small satellite missions and allow to simplify the communication system design and reduce power usage. @en

The Lunar Meteoroid Impacts Observer (LUMIO) is a CubeSat mission to observe, quantify, and characterize the meteoroid impacts on the lunar farside by detecting their flashes. This complements the knowledge gathered by Earth-based observations of the lunar nearside, thus synthesizing global information on the lunar meteoroid environment and contributing to the lunar situational awareness. The goal of LUMIO is to advance our current knowledge of meteoroid models in the solar system. In this work, we present the methodology devised to predict the scientific contribution of LUMIO. Our approach relies on combined modeling and simulation of payload, orbit, and environment. The analyses carried out have been used to drive the design of the LUMIO mission and its payload, the LUMIO-Cam. A payload radiometric model is derived and exploited to assess the quality of the scientific measurements. A dedicated study about straylight rejection is carried out to assess how straylight noise affects LUMIO-Cam measurements. Our results indicate that a 150 mm baffle grants good performance when the Sun angle is between 20° and 90°. Furthermore, the present-day LUMIO mission has the potential to detect more than 6000 impact flashes during the activity peak of the Geminids in 2024 in the range of the equivalent impact kinetic energy at Earth of [10 ⁻⁶,10 ⁻¹]kton TNT Equivalent. Compared to previous programmes, LUMIO could refine information and fill the knowledge gap about the meteoroid population in the ranges of the equivalent impact kinetic energy at Earth of [10 ⁻⁶,10 ⁻⁴]kton TNT Equivalent and [10 ⁻⁴,10 ⁻¹]kton TNT Equivalent, respectively.

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Recent advances in space technology provide an opportunity for small satellites to be launched in cislunar space. However, tracking these small satellites still depends on ground-based operations. Autonomous navigation could be a possible solution considering the challenges presented by costly ground operations and limited onboard power available for small satellites. There have been various studies on autonomous navigation methods for cislunar missions. One of them, LiAISON, provides an autonomous orbit determination solution solely using inter-satellite measurements. This study aims at providing a detailed performance analysis of crosslink radiometric measurements based on autonomous orbit determination for cislunar small satellite formations considering the effects of measurement type, measurement accuracy, bias, formation geometry, and network topology. This study shows that range observations provide better state estimation performance than range-rate observations for the autonomous navigation system in cislunar space. Line-of-sight angle measurements derived from radiometric measurements do not improve the overall system performance. In addition, less precise crosslink measurement methods could be an option for formations in highly observable orbital configurations. It was found that measurement biases and measurements with high intervals reduce the overall system performance. In case there are more than two spacecraft in the formation, the navigation system in the mesh topology provides a better overall state estimation than the centralized topology.

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This study presents an autonomous orbit determination system based on crosslink radiometric measurements applied to a future lunar CubeSat mission to clearly highlight its advantages with respect to existing ground-based navigation strategies. This work is based on the Linked Autonomous Interplanetary Satellite Orbit Navigation (LiAISON) method which provides an autonomous navigation solution solely using satellite-to-satellite measurements, such as range and/or range-rate, to estimate absolute spacecraft states when at least one of the involved spacecraft has an orbit with a unique size, shape, and orientation. The lunar vicinity is a perfect candidate for this type of application due to the asymmetrical gravity field: the selected lunar mission, an Earth-Moon L2 (EML2) Halo orbiter, has an inter-satellite link between a lunar elliptical frozen orbiter. Simulation results show that, even in case of high-measurement errors (in the order of 100 m, 1σ), the navigation filter estimates the true states of spacecraft at EML2 with an error in the order of 500 m for position, and 2 mm/s for velocity, respectively and the elliptical lunar frozen orbiter states can be estimated in the order of 100 m for position and 1 cm/s for velocity, respectively. This study shows that range-only measurements provide better state estimation than range-rate-only measurements for this specific situation. Different bias handling strategies are also investigated. It has been found that even a less accurate ranging method, such as data-aided ranging, provides a sufficient orbit determination solution. This would simplify the communication system design for the selected CubeSat mission. The most observable states are found to be position states of the lunar orbiter via the observability analysis. In addition, the best tracking windows are also investigated for the selected mission scenario.@en