CubeSat Research Articles

Performance Evaluation of a silicon-based 6U Cubesat detector for soft $\gamma$-ray astronomy

Performance Evaluation of a silicon-based 6U Cubesat detector for soft $\gamma$-ray astronomy

Advancing CubeSats Capabilities: Ground-Based Calibration of Uvsq-Sat NG Satellite’s NIR Spectrometer and Determination of the Extraterrestrial Solar Spectrum

Uvsq-Sat NG is a French 6U CubeSat (10 × 20 × 30 cm) of the International Satellite Program in Research and Education (INSPIRE) designed primarily for observing greenhouse gases (GHG) such as CO2 and CH4, measuring the Earth’s radiation budget (ERB), and monitoring solar spectral irradiance (SSI) at the top-of-atmosphere (TOA). It epitomizes an advancement in CubeSat technology, showcasing its enhanced capabilities for comprehensive Earth observation. Scheduled for launch in 2025, the satellite carries a compact and miniaturized near-infrared (NIR) spectrometer capable of performing observations in both nadir and solar directions within the wavelength range of 1100 to 2000 nm, with a spectral resolution of 7 nm and a 0.15° field of view. This study outlines the preflight calibration process of the Uvsq-Sat NG NIR spectrometer (UNIS), with a focus on the spectral response function and the absolute calibration of the instrument. The absolute scale of the UNIS spectrometer was accurately calibrated with a quartz-halogen lamp featuring a coiled-coil tungsten filament, certified by the National Institute of Standards and Technology (NIST) as a standard of spectral irradiance. Furthermore, this study details the ground-based measurements of direct SSI through atmospheric NIR windows conducted with the UNIS spectrometer. The measurements were obtained at the Pommier site (45.54°N, 0.83°W) in Charentes–Maritimes (France) on 9 May 2024. The objective of these measurements was to verify the absolute calibration of the UNIS spectrometer conducted in the laboratory and to provide an extraterrestrial solar spectrum using the Langley-plot technique. By extrapolating the data to AirMass Zero (AM0), we obtained high-precision results that show excellent agreement with SOLAR-HRS and TSIS-1 HSRS solar spectra. At 1.6 μm, the SSI was determined to be 238.59 ± 3.39 mW.m−2.nm−1 (k = 2). These results demonstrate the accuracy and reliability of the UNIS spectrometer for both SSI observations and GHG measurements, providing a solid foundation for future orbital data collection and analysis.

Modeling and control of perturbation torques and mass distribution impact on a tethered system for a 12U CubeSat in sun-synchronous orbit

Tethered CubeSat systems are characterized by momentum exchange and gravity gradient stabilization, and encounter stability challenges, especially with shorter tethers in highly inclined and eccentric orbits. This study investigates the libration dynamics of a 12U CubeSat in its stowed configuration, which separates into two satellites connected via tether, throughout its deployment, station-keeping, and retrieval phases. Two deployed configurations, symmetrical 6U-6U, and asymmetrical 8U-4U connected by a non-conductive tether with a maximum length of 100m are analyzed. The study accounts for perturbations including Earth’s oblateness, atmospheric drag, solar radiation pressure, and lunisolar gravitation, modeling the tether as a rigid, extendable rod and the satellites as lumped masses. The translational and rotational dynamics are decoupled, assuming the system’s center of mass follows an unperturbed Keplerian Sun-synchronous orbit at altitudes between 400 and 600km. A tension control strategy based on Lyapunov’s direct method and supplemented by external actuation torques is explored. Results show high orbit eccentricity significantly affects the maximum in-plane libration angle. The 6U-6U system experiences smaller disturbance torques but greater tether tension variations than the 8U-4U system. Both configurations exhibit larger in-plane oscillations compared to near-zero out-of-plane oscillations, nonetheless, eclipse passages exacerbate the out-of-plane libration of the 8U-4U system. Relative stability is maintained during deployment, however, retrieval is chaotic, with notable oscillations in libration angles and tether tension. The tension control strategy effectively dampens oscillations during retrieval but loses effectiveness as tether tension approaches zero at retrieval’s conclusion. External actuation torques, ranging from 1-2mNm for deployment to 15mNm for retrieval, complement tension control.

Trajectory & maneuver design of the NEA Scout solar sail mission

The Near-Earth Asteroid Scout (NEA Scout) mission aimed to deliver a 6U CubeSat to a slow flyby of a Near Earth Asteroid using a solar sail as the primary propulsion system. NEA Scout was launched as a secondary payload aboard Artemis I, on November 16, 2022, but communication with the spacecraft was not achieved. This paper describes the target asteroid selection and the innovative trajectory and maneuver design for the NEA Scout mission, including cold gas maneuvers, multiple lunar flybys, solar sail thrusting, and design margins for solar sailing. The resulting trajectory design tools, solutions, and analyses are potentially applicable to future solar sail missions.

Implementation of a 6U CubeSat Electrical Power System Digital Twin

This paper presents the design of a digital twin for a 6U CubeSat electrical power system, including the solar arrays, solar array regulators, battery, power distribution unit, and load subsystems. The digital twin is validated by comparing its real-time outputs with those of the physical system. Experimental tests confirm its feasibility, showing that the digital twin’s real-time outputs closely match those of the physical system. Additionally, the digital twin can be used for control-hardware-in-the-loop and power-hardware-in-the-loop tests, allowing the real-time integration of simulated subsystems with hardware. This capability facilitates testing of new subsystems and optimization during the project’s development phases. Additionally, to demonstrate the advanced capabilities of this model, the digital twin is used to simulate the CubeSat electrical power system behavior in real time throughout a complete orbital cycle in low Earth orbit conditions. This simulation provides valuable insights into the CubeSat operation by capturing the transient and steady-state responses of the EPS components under real orbital conditions. The results obtained indicate that the digital twin significantly enhances the testing and optimization process of new subsystems during the development phases of the project. Moreover, the capabilities of the digital twin can be further augmented by incorporating real-time telemetry data from the CubeSat, resulting in a highly accurate replication of the satellite’s in-orbit behavior. This approach is crucial for identifying and diagnosing failures or malfunctions in the electrical power system, ensuring the robust and reliable operation of the CubeSat.

Optical payload design for downlink quantum key distribution and keyless communication using CubeSats

Quantum key distribution is costly and, at the moment, offers low performance in space applications. Other more recent protocols could offer a potential practical solution to this problem. In this work, a preliminary optical payload design using commercial off-the-shelf elements for a quantum communication downlink in a 3U CubeSat is proposed. It is shown that this quantum state emitter allows the establishment of two types of quantum communication between the satellite and the ground station: quantum key distribution and quantum keyless private communication. Numerical simulations are provided that show the feasibility of the scheme for both protocols as well as their performance. For the simplified BB84, a maximum secret key rate of about 80 kHz and minimum QBER of slightly more than 0.07% is found, at the zenith, while for quantum private keyless communication, a 700 MHz private rate is achieved. This design serves as a platform for the implementation of novel quantum communication protocols that can improve the performance of quantum communications in space.

Using the principle of gravitational stabilization and orientation in the design of small spacecraft

Currently, most small spacecraft (MCAS) are becoming increasingly relevant in the modern space industry. To successfully complete the tasks assigned to them, the ICS must be oriented in some way in space relative to the Earth. The orientation and stabilization system (SOS) is responsible for this task, which is necessary for monitoring and controlling the position of the MC in space. At the moment, the tasks for the MCA basically do not require complex turning maneuvers and high orientation accuracy, therefore passive and combined SOS have become widespread for them. Such systems have a long service life, are characterized by simplicity, high reliability and low weight. One of the typical systems used in modern space technology is gravity systems. The principle of gravitational SOS is based on the use of gravity acting on a body and moments of inertia relative to three mutually orthogonal axes. This article proposes a 3U CubeSat ICA project with a gravitational orientation system. The design of such a satellite requires a gravitational stabilization device, which is necessary to deploy the ICA after separating it from the launch vehicle, as well as to create a restoring moment. The gravitational device is supposed to be placed between the rigidly bonded 2U MC and the third U MCA. The advantage of this design is the fact that it becomes possible to place more payload on the MC without overloading it with various devices for the orientation and stabilization system.

Ground-Based Characterisation of a Compact Instrument for Gamma-ray Burst Detection on a CubeSat Platform

Gamma-ray bursts (GRBs) are intense and short-lived cosmic explosions. Miniaturised CubeSat-compatible instruments for the study of GRBs are being developed to help bridge the gap in large missions and assist in achieving full sky coverage. CubeSats are small, compact satellites conforming to a design standard and have transformed the space industry. They are relatively low-cost and are developed on fast timescales, which has provided unparalleled access to space. This paper focuses on GMOD, the gamma-ray module, onboard the 2U CubeSat EIRSAT-1, launched on December 1st 2023. GMOD is a scintillation-based instrument with a cerium bromide crystal coupled to an array of sixteen silicon photomultipliers, designed for the detection of GRBs. The characterisation of GMOD in the spacecraft, along with the validation of an updated spacecraft MEGAlib model is presented and this approach can be followed by other CubeSats with similar science goals. The energy resolution of the flight model is 7.07% at 662 keV and the effective area peaks in the tens to hundreds of keV, making it a suitable instrument for the detection of GRBs. An investigation into the instrument’s angular response is also detailed. The results from this characterisation campaign are a benchmark for the instrument’s performance pre-launch and will be used to compare with the detector’s performance in orbit.

SPRISS: Scalable and precise resistive impact sensor for smallsats, architecture description and tests

Space debris are a tangible risk for satellites in Earth orbits. Indeed, in the last decades fragmentation events have generated a large number of uncontrolled objects that have made the debris population grow. Modelling the space debris environment is becoming a fundamental task to evaluate the vulnerability of operational satellites, the probability of accidental collisions with uncontrolled objects, and the evolution of the debris population. With this aim, remote and in-situ measurements provide valuable data to tune the space debris population models and improve their reliability. While large satellites can be observed and tracked from ground, the sub-millimeter debris population requires in-situ measurements. In this context, in-orbit impact sensors are a key technology to obtain information about the sub-mm space debris environment. In this framework, a small-scale impact sensor sized to be integrated in a 2U CubeSat is being developed at the University of Padova. The sensor consists of a multitude of thin, conductive stripes arranged on a thin film of non-conductive material. When a debris hits the sensor, one or more stripes are severed, and the impact is detected. Moreover, the sensor design ensures low power consumption, making it feasible for CubeSat space missions. This work presents the latest outcomes obtained from the development of the sensor. Specifically, structural analyses are performed to assess that the sensor can withstand the launch loads, as well as thermal analyses to confirm its endurance capability with in-orbit temperatures. The number of expected impacts during the mission is predicted through orbital propagation, using state-of-the-art debris environment modelling tools. Finally, the paper presents a functional shooting test and vibration tests, executed onto a development model of the sensor, which verify its functionality and validate a Technology Readiness Level (TRL) of 4.

Passive thermal control design and flight operation results of nano moon lander OMOTENASHI

OMOTENASHI is a 6U CubeSat launched by the first NASA SLS rocket. Its mission was to demonstrate that a CubeSat can make a semi-hard landing on the Moon. Tight resource constraints make it difficult to mount a heater and a radiator in a CubeSat, so passive thermal control is a prominent design feature of OMOTENASHI that resolves the technical difficulty. In November 2019, a thermal vacuum test was conducted. This paper discusses the thermal design and the test results. OMOTENASHI was launched by SLS rocket on November 16, 2022. After the SLS rocket separation, OMOTENASHI's attitude control malfunctioned so it could not generate power from its solar cells. Eventually, communications were lost and have not been restored. Brief telemetry data confirmed that the temperature monitor functioned correctly for 30 min. This paper reports the flight results of the thermal status during the first ground station pass. Also, a comparison between the analysis model and telemetry data is shown. Finally, based on its estimated trajectory, the future thermal profile of the spacecraft and the result of the search operation are discussed.

HERMES: Gamma-ray burst and gravitational wave counterpart hunter

Gamma-ray bursts (GRBs) bridge relativistic astrophysics and multimessenger astronomy. Space--based gamma - and X-ray wide-field detectors have proven essential for detecting and localizing the highly variable GRB prompt emission, which is also a counterpart of gravitational wave events. We studied the capability of detecting long and short GRBs with the High Energy Rapid Modular Ensemble of Satellites (HERMES) Pathfinder (HP) and SpIRIT, namely a swarm of six 3U CubeSats to be launched in early 2025, and a 6U CubeSat launched on December 1 2023. We also studied the capabilities of two advanced configurations of swarms of more than eight satellites with improved detector performances (HERMES Constellations). The HERMES detectors, sensitive down to sim \,2-3\,keV, will be able to detect faint and soft GRBs, which comprise X-ray flashes and high-redshift bursts. By combining state-of-the-art long- and short-GRB population models with a description of the single module performance, we estimate that HP will detect sim $ long GRBs ($3.4^ $ short GRBs per year. The larger HERMES Constellations under study can detect between sim \,1300 and sim \,3000 long GRBs per year and between sim \,160 and sim \,400 short GRBs per year, depending on the chosen configuration, with a rate of long GRBs above $z>6$ of between 30 and 75 per year. Finally, we explored the capability of HERMES to detect short GRBs as electromagnetic counterparts of binary neutron star (BNS) mergers detected as gravitational signals by current and future ground--based interferometers. Under the assumption that the GRB jets are structured, we estimate that HP can provide up to sim 1 (14) $ joint detections during the fifth LIGO-Virgo-KAGRA observing run ( Einstein Telescope single triangle 10 km arm configuration). These numbers become sim \,4 (100) $, respectively, for the HERMES Constellation configuration.

MeSat Mission: Exploring Martian Environment with THz Radiometer Payload and Optimal Trajectory

Space exploration presents vast prospects for scientific, industrial, and economic progress. This paper introduces the MeSat mission as a pioneering approach to Mars exploration. The MeSat aims to deepen our understanding of Martian conditions and resources by employing an optimized Earth-to-Mars trajectory, enabling a comprehensive study of the Martian atmosphere and surface. The mission comprises a cargo microsatellite hosting three 6U CubeSats and two 3U CubeSats, deployed into four separate Mars orbits to form a constellation. Each CubeSat carries distinct payloads: a THz radiometer for Martian water vapor atmospheric observation, a high-resolution surface camera, a high-tech spectrometer, and a Fourier transform spectrometer (FTS) for wind speed readings. This paper includes the majority of the key parameters; however, we focus our discussion more on two aspects of this pioneering mission: the first aspect contains the proposal of four distinct payloads for the study of Mars’ atmosphere and the second aspect proposes an optimal mission design algorithm that analyzes a fuel-efficient low-thrust trajectory from Earth to Mars. Regarding the payloads, the THz radiometer requires a specific design; hence, we explain this payload in more depth; the rest of the payloads, we suggest utilizing commercially available elements for the cost-effective manufacture of a whole system. For mission trajectory optimization, the study employs a dual-step hybrid optimization algorithm (PSO-homotopy) to analyze fuel-efficient low-thrust trajectories from Earth to Mars, incorporating the ephemeris dynamics model to account for gravitational perturbations in the entire Solar System. In practical mission design, crucial factors like hyperbolic excess velocity, diverse opportunities for Earth launch and Mars rendezvous, varied propulsion systems, and time of flight (TOF) play vital roles in trajectory optimization. In summary, for the MeSat mission, we propose a comprehensive Mars environmental mission design. We consider all aspects of the mission from trajectory design to engineering detail design, since we would like to inspire future Mars missions with a complete report.

Feasibility Analysis of a CubeSat Mission for Space Rider Observation and Docking

AbstractIn the last few years, the number of orbiting satellites has increased exponentially, in particular due to the development of the New Space Economy. Even if this phenomenon makes the space more accessible, bringing a great contribution to the scientific, economic and technological fields, on the other hand it contributes to the overpopulation of the space background. Therefore, it is necessary to develop new techniques to manage the space environment, such as in orbit servicing, which is a procedure that aims to refuel and repair satellites to extend their operational life. A first step to reach this goal is to inspect closely the object of interest to study its features. In this framework, the Space Rider Observer Cube (SROC) mission is being developed. SROC is a payload that will be deployed by Space Rider (SR), an uncrewed and reusable robotic spacecraft designed by ESA (European Space Agency). SROC is a 12U CubeSat, whose goal is to carry out inspection manoeuvres around SR, then re-enter on board using a safe docking system to come back to Earth. The feasibility of a mission similar to SROC has been simulated during a university class, starting from the definition of the system requirements with particular focus on the analysis of the payloads and subsystems, to ensure the achievement of the mission goals. In particular, the CubeSat is equipped with an imaging payload to capture high resolution images of Space Rider surface and a docking mechanism. Then, the design of the orbit and the simulation of the effects of the space environment on the CubeSat have been studied using GMAT, SYSTEMA, MATLAB and other numerical tools. The results of the study are useful for future missions, aiming to inspect orbiting objects, such as operative satellites for in orbit servicing, space debris and dead satellites to study their geometries and plan their removal.

The Deep Space Radiation Probe: Development of a first lunar science payload for space environment studies and capacity building

Regions outside of Low Earth Orbit (LEO, altitudes above approximately 1000 km) are classified as “deep space”, including Medium Earth Orbit (MEO), geostationary orbit (GEO), as well as cislunar and lunar space. The deep space environment poses many challenges for human and robotic exploration, including stronger ionizing radiation fluxes, more extreme temperature variations, as well as limited data downlink volume. With the growth of the rideshare and hosted payload model aboard government and commercial lunar missions, developing the capacity to design and implement payloads and other space avionics for this environment is of increased importance this decade. Utilizing one of the growing number of rideshare opportunities offered by commercial lunar mission providers, National Central University (NCU) has been working on the rapid development of Taiwan’s first scientific payload for lunar lander use, with launch aboard the HAKUTO-R Mission 2 (M2) lander from ispace, inc. scheduled not earlier than Q4 2024. This Deep Space Radiation Probe (DSRP) will provide measurements of radiation dose, dose rate, and single event upset (SEU) rate during the Earth-Moon transit, in lunar orbit, as well as on the lunar surface. DSRP utilizes elements of the on-board computer (OBC) developed and flight qualified aboard the NCU-developed IDEASSat 3U CubeSat mission in 2021, and was developed by a student team, in consultation with experienced engineers from the lunar lander team. In this paper, we will report on the objectives, concept of operation, design, and implementation of the DSRP project. We will also describe the steps taken to facilitate parallel development of the DSRP payload and the HAKUTO-R M2 lander, as well as lessons learned during the design, implementation, and qualification process. The radiation data provided by DSRP will be beneficial for the development of future deep space spacecraft avionics, as well as crewed missions, and will also serve to build the capacity for deep space spacecraft and payload development at NCU.

3U CubeSat-Based Hyperspectral Remote Sensing by Offner Imaging Hyperspectrometer with Radially-Fastened Primary Elements.

This paper presents findings from a spaceborne Earth observation experiment utilizing a novel, ultra-compact hyperspectral imaging camera aboard a 3U CubeSat. Leveraging the Offner optical scheme, the camera's hyperspectrometer captures hyperspectral images of terrestrial regions with a 200 m spatial resolution and 12 nanometer spectral resolution across a 400 to 1000 nanometer wavelength range, covering 150 channels in the visible and near-infrared spectrums. The hyperspectrometer is specifically designed for deployment on a 3U CubeSat nanosatellite platform, featuring a robust all-metal cylindrical body of the hyperspectrometer, and a coaxial arrangement of the optical elements ensures optimal compactness and vibration stability. The performance of the imaging hyperspectrometer was rigorously evaluated through numerical simulations prior to construction. Analysis of hyperspectral data acquired over a year-long orbital operation demonstrates the 3U CubeSat's ability to produce various vegetation indices, including the normalized difference vegetation index (NDVI). A comparative study with the European Space Agency's Sentinel-2 L2A data shows a strong agreement at critical points, confirming the 3U CubeSat's suitability for hyperspectral imaging in the visible and near-infrared spectrums. Notably, the ISOI 3U CubeSat can generate unique index images beyond the reach of Sentinel-2 L2A, underscoring its potential for advancing remote sensing applications.

On-Board Image Enhancement on Remote Sensing Payload

CubeSats are designed to optimize applications within the strict constraints of space and power. This paper presents an On-Board Image Enhancement technique for remote sensing payloads, focusing on achieving Auto White Balance (AWB) with limited resources and enhancing the capabilities of small/microsatellites. The study introduces hardware-based techniques, including histogram adjustment, De-Bayer processing, and AWB, all tailored to minimize hardware resource consumption on CubeSats. The integrated 1U CubeSat system comprises a sensor board, an Image Data Processor (IDP) unit, and onboard computing, with a total power consumption estimated at 2.2 W. This system facilitates image capture at a resolution of 1920 × 1200 and utilizes the proposed algorithm for image enhancement on remote sensing payloads to improve the quality of images captured in low-light environments, thereby demonstrating significant advancements in satellite image processing and object-detection capabilities.

Agile Stage-Gate Approach for Design, Integration, and Testing of a 1U CubeSat

Life cycle stages are very important for the aerospace industry. Many models have emerged for handling the processes within and across the development of new products. Developing CubeSat-based missions has shortened the required time and has reduced expenses. However, the lack of strategic planning in the design, integration, and testing of product development models has been highlighted as one of the key issues contributing to failures. The objective of this study is to propose a new hybrid model for the physical development of a product using an Agile Stage-Gate methodology focused on a 1U CubeSat (AztechSat-1). This study aims to explain the full process throughout the project timeline from conceptualization to execution. The benefits of such a model include ensuring adaptive responses to not only improve technical integration but also allow the successful validation and verification of a nanosatellite. Our theoretical approach articulates an in-depth understanding of Agile Stage-Gate methodology through experience obtained from experts and team members. Our analysis supports the expected benefits of the iterative process at every stage. Through this approach, product development could benefit from reduced times and better innovations. Nevertheless, there are also drawbacks to this method. The requirements of greater human effort, more frequent demonstrations, and a constant review process have negative impacts. Additionally, particular modifications must be made for each area of research. For educational purposes, the initial results seem to be encouraging.

Mission analysis of a 1U CubeSat post-mission disposal using a thin-film vacuum arc thruster

We present the operation of the CubeSat De-orbiting All-Printed Propulsion System (Cube-de-ALPS), a thin-film Vacuum Arc Thruster being developed at the University of Southampton in collaboration with the European Space Agency to provide robust de-orbiting capability to sub-3U CubeSats. It comprises a flexible substrate on which coplanar arrays of vacuum arc micro-thrusters (micro-VAT) are printed alongside small supporting electronic subsystems. In particular, we focus on the application of a Cube-de-ALPS End-Of-Life disposal to an under-actuated 1U CubeSat with uncontrolled spin. In this scenario, a Faraday cup will provide coarse angle-of-attack estimates and trigger a single micro-VAT to ignite every time it points towards the forward velocity vector. Orbital lifetime estimates for different fuel materials, configurations, and operational modes are estimated using simplified dynamics and analytical thrust averaging. These results are compared to high-fidelity numerical simulations including full six degrees of freedom coupled attitude and orbital dynamics to confirm the viability of the concept and confirm that Cube-de-ALPS can de-orbit CubeSats from 1400 km.

First Results From REPTile‐2 Measurements Onboard CIRBE

AbstractCIRBE (Colorado Inner Radiation Belt Experiment), a 3U CubeSat, was launched on 15 April 2023 into a sun synchronous orbit (97.4° inclination and 509 km altitude). The sole science payload onboard is REPTile‐2 (Relativistic Electron and Proton Telescope integrated little experiment—2), an advanced version of REPTile which operated in space between 2012 and 2014. REPTile‐2 has 60 channels for electrons (0.25–6 MeV) and 60 channels for protons (6.5–100 MeV). It has been working well, capturing detailed dynamics of the radiation belt electrons, including several orders of magnitude enhancements of the outer belt electrons after an intense magnetic storm, multiple “wisps”‐ an electron precipitation phenomenon associated with human‐made very low frequency (VLF) waves in the inner belt, and “drift echoes” of 0.25–1.4 MeV electrons across the entire inner belt and part of the outer belt. These new observations provide opportunities to test the understanding of the physical mechanisms responsible for these features.

Modelling an Ion Thruster for a Small Spacecraft in Very Low Earth Orbit

A gridded ion thruster running on two different propellants (Xenon and Iodine) and sited within a 3U CubeSat satellite, is modelled in a Low-Earth-Orbit (altitude 400 km), using a Particle-in-Cell approach. Charged exchange collisions cause a plume backflow, which results in erosion or contamination of external spacecraft surfaces. In particular, this research is motivated by evaluating the risks to solar panel arrays from plume backflow. At low altitudes there is a wide range of ambient species able to interact electrostatically with the ion plume. In particular one must consider the plasma potential of the incoming particle flux at the solid boundaries as well as a sheath comparison and the temperature of the solar panel surfaces. This enables evaluation of the effect of temperature on the charged exchange region. A Particle-in-Cell is used, with a hybrid approach where electrons are treated as a fluid and the propellant as kinetic particles. Our results compare surface ablation for two propellants, focusing on solar panel arrays which are considered to be especially vulnerable to the backflow of ions from the plume expansion. It is found that the flux of incoming particles increases for lower satellite surface temperatures. Furthermore, Xenon results in having a lower overall effect on the sensitivity of the particle flux in comparison to Iodine.