On-Orbit Mission Overview of the Low Power Hall Thruster Propulsion System aboard Venμs Satellite

On-Orbit Mission Overview of the Low Power Hall Thruster Propulsion System aboard Venμs Satellite | Research Square window.SnipcartSettings = { analytics: { enabled: false } }; (function() { var accessVector = localStorage.getItem('access_vector') || ''; window.dataLayer = window.dataLayer || []; if (accessVector) { window.dataLayer.push({ user: { profile: { profileInfo: { snid: accessVector } } } }); } })(); (function(w,d,s,l,i){w[l]=w[l]||[];w[l].push({'gtm.start':new Date().getTime(),event:'gtm.js'});var f=d.getElementsByTagName(s)[0],j=d.createElement(s),dl=l!='dataLayer'?'&l='+l:'';j.async=true;j.src='https://www.googletagmanager.com/gtm.js?id='+i+dl;f.parentNode.insertBefore(j,f);})(window,document,'script','dataLayer','GTM-K279D39R'); Browse Preprints In Review Journals COVID-19 Preprints AJE Video Bytes Research Tools Research Promotion AJE Professional Editing AJE Rubriq About Preprint Platform In Review Editorial Policies Our Team Advisory Board Help Center Sign In Submit a Preprint Cite Share Download PDF Research Article On-Orbit Mission Overview of the Low Power Hall Thruster Propulsion System aboard Venμs Satellite Daniel Katz-Franco, Dan Lev, Boaz Shoor, Amoz Davidson, Raul Zimmerman, and 5 more This is a preprint; it has not been peer reviewed by a journal. https://doi.org/ 10.21203/rs.3.rs-4022336/v1 This work is licensed under a CC BY 4.0 License Status: Under Review Version 1 posted 7 You are reading this latest preprint version Abstract Venµs is a satellite launched in 2017, for super-spectral Earth imaging and Electric Propulsion System (EPS) demonstration. In this paper we overview EPS design and operation throughout all five mission phases, from open/closed-loop orbit control, through orbit descent (720→410 km), orbit maintenance under high drag environment, to orbit raising (410→560 km). The EPS consisted of two throttleable Hall thrusters, PPU, Propellant Management Assembly (PMA), and a 9 liter propellant tank carrying 16 kg of Xenon. Both thrusters operated in the 300–550 W power range, generated a combined total impulse of 158.1 kN-sec and consumed all propellant. Two methods are described to compute the remaining propellant mass – “Bookkeeping” and “PTV” methods. The advantages and disadvantages of each method are discussed in light of the Venµs mission. Thruster performance was measured on-orbit using the “Orbit Determination” method and compared to laboratory experiments conducted on the ground with identical thrusters. The measured performance on-orbit was found to be on average lower by up to 5% than the performance measured on the ground. Lastly, we present several repeating events in which the propulsion system suffered from sudden beam-outs or ignition difficulties. We present the methods used to construct a solution and implement it. Hall Effect Thruster Electric Propulsion System Technological Mission In Orbit Test Figures Figure 1 Figure 2 Figure 3 Figure 4 Figure 5 Figure 6 Figure 7 Figure 8 Figure 9 Figure 10 Figure 11 Figure 12 I. Introduction The past two decades have seen a strikingly emerging presence of spacecraft equipped with electric propulsion systems, and specifically Hall thruster-based systems. Various past satellite missions utilizing Hall thruster technologies reported on their on-orbit experience and even attempted to assess propulsion system performance in space. These missions are for Hall thruster technologies ranging in power from less than 100 W [ 1 ] up to 4.5 kW [ 2 ]. Those missions reported propulsion system cumulative operation duration from several minutes [ 3 ] to thousands of hours and ignition cycles [ 4 ]. Most space missions, after the spacecraft is injected into initial orbit, perform a preliminary phase that is called In-Orbit Testing (IOT) where all systems are tested for proper operation, after the launch phase. The Electric Propulsion System (EPS) runs cascading housekeeping and health checks with the goal of assessing its health [ 1 – 7 ]. These include basic PPU electrical checks, pressure and temperature sensors health checks, initial valve activation, and cathode heater startup. IOT is usually performed over the course of several earth revolutions. In the case of the DEIMOS-2 mission the IOT phase led to a re-calibration of the cathode mass flow rate [ 7 ]. Throughout the mission there is a need to monitor propellant remaining level in order to track the cumulative impulse used and remaining budget. This is conducted by using two methods – Bookkeeping and Pressure Temperature Volume (PTV) [ 2 , 3 , 8 ]. The bookkeeping method, as its name implies, is based on the integration of the mass flow rate over time to compute the mass used to the point in the mission. As with any integrative method, bookkeeping may accumulate any errors in the intended mass flow rate. Its advantage is that it does not rely on sensor measurements. The PTV method computes the remaining mass using the propellant tank temperature and pressure sensor measurements along with the known internal volume of the tank. This measurement can be performed in real time at any point during the mission and does not require historical mission information. None of the propulsion systems reporting on its mission experience found significant discrepancies between the two methods. To assess propulsion system performance on orbit the Orbit Determination (OD) method is used. This method uses data of the spacecraft orbit change due to propulsion system activity. All EP systems estimating thrust on orbit used this method. A thorough description of the OD method is given in refs. 9 and 10. On-ground and on-orbit performance data for a variety of Hall thruster-based missions is presented in Table 1 . It can be seen in the table that some Hall thruster-based systems exhibited an improved performance in space while others had better performance on the ground. On average, it can be determined that the performance of these thrusters is approximately the same on orbit as on the ground. It should be noted that even gridded ion engines reported on-orbit thruster performance similar to that measured on the ground [ 12 , 13 ]. Table 1 On-orbit vs. on-ground thruster performance data for different Hall thruster-based propulsion systems. Thruster Name Anode Power [W] On Ground Thrust [mN] On Orbit Thrust [mN] Discrepancy (on-orbit to on-ground) ExoMG[ 1 ] ~ 50 1.8, 2.4 2.2, 3.01 <+25% HEPS-200 [ 7 ] 200–400 ~ 11 ~ 11 Approx. the same ACE[ 3 ] 320 20.6 22.4 ± 5.5 Within uncertainty SPT-100[ 5 ] 1350 83 83.12 Approx. the same SPT-100[ 11 ] 1350 80 ± 3 83 ± 3 <+3.7% SPT-100[ 12 ] 1350 84.05 ± 1.3 78.5 ± 6 <-7% PPS-1350-G[ 6 ] 1350 -2% SPT-140[ 2 ] 4000 226–236 228–234 Mostly higher In addition to performance data and regular mission operations some missions reported unexpected events. SMART-1 exhibited several unexpected shutdowns during different phases of the mission [ 6 ]. These were resolved by uploading revised software versions which reduced the number of unexpected occurrences. DEIMOS-2 experienced ignition and operation issues from the first few thruster operations. These were tracked back to erroneous cathode flow setting, resulting in a much lower cathode flow rate than required [ 7 ]. This issue was resolved by recalibration of the cathode mass flow line. The SPT-100 on-board Inmarsat-4F1 exhibited noisy discharge current during initial operations [ 5 ]. These were attributed to thruster outgassing even after propulsion system venting followed by gas purging. Saleh et al. investigated different anomalies and unexpected events with electric propulsion systems on orbit [ 14 ]. He concluded that the mean time between minor anomalies, such that do not impede system operation, is on average 116 days between all Hall thruster-based missions. Saleh also concluded that most of these are the so-called infant-minor-anomalies that are associated with the first several operations of the propulsion system. The VENµS Program In the recent two decades, Rafael Ltd. has engaged in electric propulsion systems research and development. To date, most efforts have been focused on low power Hall thruster-based propulsion systems [ 15 ]. One such propulsion system was developed for the Venµs mission. Vegetation and Environment monitoring on a New Micro Satellite, or VENµS, is a joint program of both the Israeli and French space agencies [ 16 ]. Within the frame of the program, two missions have been executed: a scientific mission and a technological mission. The mission commenced with the launch of the VENµS satellite in August 2017 on-board the Vega launcher, with the In Orbit Test (IOT) phase. The space mission was planned for 4½ years. During this period, the satellite conducted a combination of scientific and technological mission phases. The Venµs mission consisted of five consecutive phases named VM1 through VM5 (Venµs Mission X). VM1 consisted of 2.5 years at 720 km altitude, during which Venµs performed super-spectral imaging with a two-day revisit time, and the EPS performed 29 characterization experiments and controlled the orbit. The next phase, VM2, was devoted to a descending orbit transfer, solely performed by the EPS, to 410 km altitude and 1.2° inclination. Then, in VM3, Venµs imaged again with a two-days revisit time, this time with an enhanced resolution due to proximity – while the EPS constantly corrected the orbit in this higher drag environment. The next phase, VM4, consisted of a second orbit transfer, also performed by the EPS, using the remaining Xe propellant to reach 560 km. The next phase, VM5, was solely devoted to imaging in a one-day repeating orbit. Venµs has depleted all of its Xenon propellant – and the orbit is controlled by the remaining chemical propellant. All mission orbits were sun-synchronous and earth repeating. The scientific mission has been super-spectral imaging, using a camera able to capture 12 narrow spectral bands (bandwidth between 25 to 40 nm). Its goal is to image and monitor agricultural growth, vegetation and water bodies, for environment studies. Since the super-spectral imaging relays mainly on light reflection, the imaging angles must be temporarily constant. Thus, sun-synchronous & earth repeating orbits were chosen for each Venµs phase. Hence the big challenge is controlling the orbit precisely, during all phases and altitudes. This delicate task was assigned to the technological mission and payload – the EPS, which aided by autonomous orbit control algorithms, strictly controlled the required orbit to satisfy the imaging requirements. . II. Technological Mission The ‘Technological Mission’ phases consisted of the operation and qualification of the electric propulsion system in space, orbit control and orbit transfer. As its technological payload, VENµS incorporated the Israeli Hall Effect Thruster (IHET), which was developed for this purpose. The thruster was designed to fulfill the requirements of the Venµs small satellite platform of 300 kg, which can produce a relatively low power. IHET’s anode power range is between 300 W and 550 W and it consumed a total of 16 kg of Xenon propellant throughout its mission. The IHET is part of Rafael’s 1st generation of Electric Propulsion Systems family, denoted as the R-400EPS. The two main objectives of the technological mission, space verification and mission validation, were achieved by a series of experiments that tested the EPS performance, followed by five mission phases in which the EPS changed the orbit and maintained it [ 17 ]. All Venµs technological mission phase was driven by the Technological Mission Module (TMM), which is a Rafael provided flight software component, residing in the on-board computer (OBC). The TMM was specifically designed for the Venµs technological mission goals. It has three major operation modes: (1) EPS experiment test, (2) Orbit transfer, and (3) Orbit control. The TMM operates autonomously and is fed by commands initiated by the Technological Mission Center (TMC) which is the ground-based facility managing the technological mission. The remarkable autonomous operation of the TMM consists of the ability to estimate the actual orbit on-board the satellite and generate thrust firing EPS commands to correct the orbit to whatever the TMC commanded it [ 9 ]. This article describes the operation of the EPS on-board the Venµs satellite mission. First, we present all EPS components and their respective role in the system. We then overview on-orbit mission activities which were divided into five phases. Subsequently, we present propellant consumption calculations using two separate methods, as well as the assessment of propulsion system performance using Orbit Determination (OD). Lastly, we discuss unexpected events such as unplanned thruster shutdowns or several failed ignitions attempts, and the methods used to overcome these while in space. III. VENµS EPS A. System The Venµs EPS (see Refs. [ 16 , 17 ]) is designed to support and operate two IHET-300 thrusters, one at a time. Besides the thrusters, the EPS main components are the Propellant Management Assembly (PMA), the Digital Xenon Flow Controller (DXFC), the Power Processing Unit (PPU), and 2 electrical Filter Units (FU). The propellant tank stores highly pressurized Xenon. The system architecture is presented in Fig. 1 , and the Venµs propulsion plate is shown in Fig. 2 . One notable feature of the EPS is the ability to throttle the anodic power of its thrusters, according to the instantaneous available power supplied by the bus platform in orbit. This proves to be an essential feature in a LEO satellite, since the power levels produced by the solar arrays are variable and change according to satellite’s true anomaly angle in each revolution. The power level is automatically computed by the TMM and commanded to the PPU which in turn regulates the propellant flow and operates the thrusters at the commanded power level (see Ref. [ 17 ]). A Thruster Selection Unit (TSU) is responsible for selecting the active thruster, as decided by the TMM and commanded by the PPU. B. IHET-300 Thruster The heart of the EPS is the Hall Effect Thruster (HET), codenamed IHET-300, shown in Fig. 3 . It operates on Xenon, which is ionized by electrons emitted from the cathode and accelerated in the form of plasma via a high electric field. Table 2 IHET-300 Main Characteristics @ 300W Thrust @ 300W (EOL) > 14.3 mN Specific Impulse @ 300W (EOL) > 1210 sec Power Operation Range 250W to 600W Operating Life > 1100 hours Number of Operation Cycles > 2000 Total Impulse > 135 kNs The main characteristics of IHET-300 are listed in Table 2 . This thruster, which employs low power, is specifically designed to be used onboard small and micro satellites. Its useful range of operation is between 250 to 600W. On Venµs platform we limit the operation power to 550W, due to platform power limitations. Thus, it can utilize the instantaneous available power from the satellite as explained above. The IHET-300 mass is about 1.5 kg and its dimensions are 170×120×90 mm. C. Propellant Management Assembly (PMA) The PMA is responsible for feeding Xenon gas, at controlled pressure and flowrate, to the selected thruster. It consists mainly of adapted and qualified COTS components. Ref. [ 5 ] describes the development and qualification process of most EPS components. A picture of the Digital Xenon Flow Controller and functional schematic are shown in Fig. 4 . The PMA consists of two (2) pressure regulators connected in series. The pressure regulation works in an open loop mode pre-calibrated on the ground, while the first regulator reduces the tank high pressure feed to nearly the operating pressure and the second regulator gives a more accurate output of 1.8 bara. This setup provides a very accurate pressure at the output of the pressure regulators. The Digital Xenon Flow Controller (DXFC) (shown in Fig. 4 ) uses a set of six (6) valves and flow restrictors arranged as a gas manifold to control the total Xenon flow through the anodes; therefore, controlling the discharge power and thrust levels. This flow controller setup is fully redundant in the case some of the valves fail to open. The DXFC controlled by the PPU would select a configuration of valves to maintain the required discharge power within a tolerance of 20W, the control loop checks the discharge power and if required it updates the DXFC valve configuration. Only one controller was installed on board the EPS while the flow was directed to each thruster by opening a series of valves downstream near the thrusters. D. PPU The PPU contains the diverse power supplies and control circuits for operation and management of primarily the thrusters and the whole EPS components. The Sequencer and Control Unit (SCU) is the control module of the PPU. Besides controlling the power supply and the ignition sequence, the PPU also implements communication with the satellite OBC, monitors all signals and prepares the EPS telemetry. The PPU block diagram is shown in Fig. 5 herein with its main interfaces to the EPS sub-assemblies. The PPU is thermally insulated from the satellite and its radiator on the bottom side is exposed to space for cooling by heat radiation. Figure 5 also shows the PPU flight model, while Table 3 shows its main characteristics. Table 3 PPU main characteristics Characteristic Value Mass 12kg Efficiency at max. power 91% Anodic power up to 600W @ 300 VDC Thermal control Autonomous, through bottom radiator Anode voltage PS 300 VDC Magnetic Field PS 0.6–2.6A, variable control for optimal performance Cathode Heater Current Selectable up to 11 ADC Cathode keeper Voltage 60 VDC Auxiliary PS 5 VDC, 16 VDC and 28 VDC Redundancy Full cold redundancy, reconfigurable in space Interface Power: Unregulated bus, 34 to 42 V Communications: Dual redundant CAN-Bus IV. On-Orbit Operation After the launch of Venµs satellite, a series of In-Orbit Tests (IOT) were performed. These tests checked and assured that the platform and its payloads were ready for the intended mission. The technological mission experiments started as planned in early 2018 and continued until the end of VM1. All other mission objectives were fully accomplished, and all propellant was depleted in 2022. A. Overview of electric propulsion system operation The IOT consists of a large series of trials in which an additional function is checked in turn until a full EPS activation is achieved, including uplink communication and download of telemetry. The success of the IOT is a prerequisite for the next EPS mission phases. Table 4 summarizes the main tests performed. Table 4 In-Orbit Testing (IOT) set of checks and tests Tests Name Test Description PPU Telephone Home PPU switched on for the first time, tested and parameters are set PPU Uplink/Downlink Test all power supplies, redundant paths and cross commands Pneumatics Check A series of tests in which valve is open and close. Test verification by examining the pressure transducers. System vent Low-pressure segment valves are opened and each branch is vented. Cathodes Outgassing Cathodes outgassing EPS first ignition Each thruster is ignited for the first time to check proper operation. EPS autonomous module A concluding end-to-end station keeping test for automatic orbit maintenance After successful completion of the IOT, the EPS was ready to commence its technological mission execution in each of the five Venµs phases, as shown in Table 5 . During VM1, the satellite performed each month a technological mission experiment. The goal of these experiments was to estimate the thrust performance of each Thruster Unit (TU). For this purpose, the TMM commanded the EPS a predefined series of firing pulses, that would change the orbit by increasing and decreasing the satellite’s Semi-Major Axis (SMA). The whole maneuver consisted of orbit extrication in three Open Loop (OL) segments and finished with orbit re-insertion in a Closed Loop (CL) segment, as shown in Fig. 6 . Each OL segment consists of three satellite revolutions, with a one long EPS activation at a constant power between 300 and 550W (meaning constant thrust) for each revolution. In the first segment, EPS activations raised the orbit by aligning the satellite's thruster orientation with the velocity vector. The second and third segments also contained EPS activations for three consecutive revolutions each, but the satellite attitude was commanded by the TMM so that the orbit Semi-Major Axis (SMA) will decrease. At the end of each segment, the EPS was left idle for one revolution, dedicated to orbit measurements and estimation. The overall design goal was to finish the three segments in an orbit lower than the initial one. Subsequently, the CL segment commences. The TMM in CL mode commands the EPS to fire and operate on each consecutive revolution for a duration and altitude instantaneously calculated, until the orbit reached the target value. The orbit data telemetry is then analyzed by the TMC, which estimates what was the force applied that resulted in these orbit changes. Then, all parasitic forces are compensated, resulting in the average thrust actually applied by the TU. During this VM1 phase the EPS was well characterized in terms of operational parameters (mass flow rates requirements, anode current and anode current noise measurements, system various temperatures, etc.), ignition sequence, downlink and uplink etc., and some parameters calibrations were made through the uplink to the satellite such us mass flow rate to the anode during the ignition. This phase ended on October 15th, 2020, with a total of 29 experiments, 380 ignitions, 86.05 operating hours and a total calculated impulse of 6.7kNs accumulated by the EPS. During this phase of characterization, the EPS was activated across all its discharge power range, which is 300W to 520W. The average firing duration was 20 minutes for each thruster including the startup ignition sequence. During VM2, the satellite was commanded to make an autonomous orbit transfer from a 720 km SSO to 410 km SSO. The maneuver started on November 4th, 2020. During this phase, the EPS thrusters were activated in activated in 90% of the revolutions in each day in which TU1 was operated for approx. 25 minutes and TU2 was operated for approx. 15 minutes. During the remaining revolutions telemetry was downloaded, and additional checks were performed on other satellite subsystems. At each revolution the EPS operated both thrusters to lower the altitude of the satellite. To do so, after one thruster completed its burn, the satellite was rotated about 180° to align the second thruster with the correct velocity vector direction. After several tens of revolutions two significant phenomena were observed, the first is that the maximum continuous discharge power that the EPS delivered was 420W. An investigation was performed on this matter which concluded that the cause of this phenomena was the thermal dissipation on the EPS mounting plate and the mass flow controller, affecting its inlet conditions, therefore reducing the effective mass flow rate. One possible solution to this matter was to use the redundancy of the DXFC to compensate the lower mass flow rate to the anodes. However, the implementation of this solution required some on-orbit testing to revalidate operation, resulting in delaying the scheduled maneuver. At this stage it was decided to limit and leave the maximum discharge power command to 420 W for this phase in order not to disturb the mission's maneuver schedule. The second phenomenon was that some beam-out events were reported via telemetry, mostly immediately after thruster ignition. Beam-out events are events in which the discharge spontaneously extinguishes. At this point, it was decided to stop the autonomous maneuver and keep activating the EPS manually using direct scripts which were uplinked to the satellite. To mitigate the beam-out events some tests were performed at Rafael's space propulsion facility resulting in an update of the magnetic field parameters and a change in the ignition routine to start the thrusters with no magnet current, that is also referred to as " Glow mode " ignition. This event proved the benefit of project risk management that was conducted during the development phase, which resulted in two redundant and optional modes of TU ignition and operation. The target orbit of 410 km was reached on September 5th, 2021, with a total of 4,077 ignitions, 1,382.5 operating hours and a total calculated impulse of 110.6 kNs. A total of 23 beam-out events were registered. The main objective of VM3 in terms of the EPS was to provide autonomous station keeping maneuvers which enable the scientific payload to image with a better resolution than in the previous orbit. This was achieved thanks to the shorter distance to Earth and at the cost of more frequent station keeping maneuvers. By performing this phase, we demonstrated, the feasibility of enabling enhanced imaging resolution without increasing the size and mass of the satellite’s camera. Venµs satellite maintained this orbit for a total duration of one month, while four station keeping maneuvers, each consisting of 9 thruster operations, were required to overcome the higher drag at this altitude. The autonomous module performed a total of 36 thruster ignitions throttling the discharge power between 300W and 480W, accumulating 11.8 thruster operating hours. By the end of this phase of the mission the EPS accumulated a total of 4,113 ignitions, 1394.5 operating hours and a total of calculated impulse of 111.6 kNs. After VM3 was completed, at altitude of 410 km, VM4 phase started, the EPS was commanded to raise the satellite's orbit to a 560 km SSO and ended at VM5 after a successful orbit insertion at this altitude to continue the scientific mission. The VM4 maneuver started on November 9th, 2021, using the same technique as in VM2, but this time it was manually commanded by the Ground Station, on open loop, and ended with the depletion of the Xenon propellant on March 5th, 2022. Overall, throughout the mission, and until all propellant was consumed, the EPS achieved a total of 6,311 successful ignitions accumulated 2105.1 operating hours, that is 158 kNs and a total throughput of 16 kg of xenon. Table 5 Summary of the EPS utilization at different mission phases. Mission Phase Description # Ignitions [#] # Operating Hours [hr] Impulse [kNs] Accumulated Propellant Usage [kg] TU1 TU2 TU1 TU2 TU1 TU2 PTV Bookkeeping IOT In Orbit Test 10 12 2 1.4 0.1 0.1 Too low to compute 0.02 VM1 EPS Characterization and orbit correction 140 218 41.4 41.2 3.3 3.2 0.68 0.63 VM2 Orbit Transfer (720 km→410 km) 1863 1834 844.5 452 66.5 37.5 11.57 10.41 VM3 Orbit Maintenance (410 km) 18 18 7.7 4.2 0.6 0.4 Insuff. data 10.49 VM4 Orbit Transfer (410 km→560 km) 920 940 258 330.2 17.8 21.2 14.74 14.34 VM5 End of EPS Mission 22 316 3.9 118.5 0.3 7.1 15.85 15.03 Total EOL 2973 3338 1157.5 947.5 88.6 69.5 6311 2105 158.1 B. Propellant consumption During the propellant loading campaign in 2017, the propellant tank was pressurized with 16.1 kg high purity Xenon. As of the end of mission, the propellant was entirely depleted. Subsequently and throughout all mission phases, calculations were made to determine the propellant used and remaining. Two conventional methods were used to determine the remaining propellant mass at any time during the mission: (1) PTV method and (2) Bookkeeping method [ 2 , 3 , 8 ]. For the first method, periodic telemetry provides the pressure and temperature data of the Xenon tank. Using this data, the pressure dependent tank volume and basic thermodynamic state tables, the Xenon propellant mass was calculated. This method can be used at any point during the mission and regardless of previous propellant mass estimations used earlier. On the other hand, the bookkeeping method is integrative. It integrates the mass flow rate injected by the propulsion system over the operation time to compute the amount of propellant expelled. This amount is then subtracted from the initial propellant mass prior to the respective operation. The estimation of the mass flow rate injected into the thruster and cathode were calibrated using mass flow meters before satellite integration. This method has the disadvantage of accumulating the computation error from each operation to the next. Additionally, this method relies on the initial calibration of the mass flow controller prior to installation on-board the satellite and may accrue any mass flow calibration errors. In addition to the two methods, we used a third method based on laboratory experiments on an identical thruster unit. We operated an identical IHET-300 Flight Model (FM) thruster to find the correlation between the thruster discharge power and the mass flow rate. The background pressure during the test was between 1.8×10 − 5 Torr and 2.9×10 − 5 Torr (N 2 corrected). We then used the ṁ→P d relation to assess the mass flow rate used during each thruster operation in space and based on the operating anode power. Figure 7 shows the propellant consumption with the cumulative impulse of the propulsion system. Propellant consumption is approximately linear, even though the activation time for each thruster varied throughout the mission. This does not come as a surprise as the slope of the trend shown in the figure is proportional to the inverse of the specific impulse. If the specific impulse is approximately constant so will be the slope. It can also be seen that all three methods produce approximately the same result, albeit an increasing deviation from each other as the mission progresses. This increasing deviation is most likely due to a cumulative error due to a slight underestimation of the expelled propellant during all activations. It is interesting to note that the error associated with the “Bookkeeping Method” increases with EPS activation. This is due to the integrative nature of this method, that relies on the previous assessment of the remaining mass when computing the propellant mass at a given time. Since the temperature and pressure of the DXFC has changed throughout the mission, it was necessary to correct the calculation of the mass flow controller according to the temperature and input pressure. The DXFC was calibrated on the ground at pressure and temperatures of 1.8 bara and 20°C. On orbit, the input pressure was measured by the low-pressure transducer located upstream of the DXFC. The value of the pressure varied during firing due to temperature changes. The temperature was measured using a thermistor located on the body of the DXFC. Since the DXFC is located close to the PPU it was subjected to heat flux that raised its temperature. Figure 8 shows DXFC temperature during a five-day long EPS activations cycle (tens of thruster activations) at VM2 (second mission phase). It can be seen from the figure that the DXFC temperature increases from ~ 22°C to 28°C during each activation cycle. The corresponding correction was made to assess the mass flow rate deviation from the calibrated values. In any case, the PPU operated the EPS in a closed-loop manner, while adjusting the mass flow rate to match the available/commanded power. The errors associated with the “PTV Method” are different in nature since this method relies on real-time pressure and temperature measurements. The challenge with the “PTV Method” for Xenon is that Xenon density is very sensitive to changes in pressure in the 60–80 bar range, at temperatures around 25°C. This can be observed in Fig. 9 where propellant mass is plotted against tank pressure at the measured tank temperatures (~ 25°C). Small variations in pressure may lead to large variations in measured propellant mass. For this reason, propellant mass uncertainty can go up to over 0.5 kg during the intermediate phases of the mission when the propellant mass is approximately half of the initial loaded mass. The “PTV Method” errors presented in Fig. 7 are based on measuring device error of 0.6 bar for the high-pressure gauge and 0.2°C for the temperature measurement by the propellant tank thermistor. C. On-orbit performance estimation Propulsion system performance was estimated on-orbit using Orbit Determination (OD). To use this method the orbit of the satellite was determined using GPS location before a satellite maneuver was performed. Subsequently, the propulsion system was operated, and the new orbit was determined after the maneuver was completed. This method was used in previous Hall thruster-based missions and is thoroughly described in refs. 9 and 10. Additionally, an identical Hall thruster FM was operated in a vacuum chamber at Rafael to measure its performance and compare to the on-orbit data. The identical FM was at its beginning of life and only operated during its acceptance test previously. To adequately compare thruster operation the lab test operation was conducted for 20 minutes, approximately the same duration as its on-orbit counterparts. Background pressure during lab operation was between 1.8×10 − 5 Torr and 2.9×10 − 5 Torr (N 2 corrected). Past laboratory acceptance tests with the units operating in space were deemed unreliable due to exceedingly high chamber pressure. The on-orbit and lab thrust data are presented in Fig. 10 . It can be observed in the figure that the thrust increases approximately linearly with discharge power for on-orbit and lab operations. It can also be observed that the thruster operating in lab environment measured higher thrust values than both on-orbit thrusters by approximately up to 5% on average. We can attribute this to the difference in background pressure, although the background pressure during the lab test is considered sufficient for the estimation of performance in space [ref. 18]. Also, ‘Thruster 2’ exhibited higher performance than ‘Thruster 1’ by about 5% yet within thrust assessment error bars. Another set of laboratory experiments were conducted to understand thruster operation differences at the Beginning of Life (BOL) and End of Life (EOL), as well when it is considered operating cold (after 20 minutes of operation) and when it is considered hot (> 2 hrs operation). To do this another laboratory thruster was used that completed the lifetime qualification test of 130 kNsec. The qualification thruster is considered at its EOL. Experiment results are presented in Fig. 11 . It can be seen in the figure that there is little discernable difference between “hot” and “cold” operations of the thrusters in lab, whether the thrusters are at BOL or EOL. It can also be seen that the EOL thruster performance is slightly lower than that of the BOL thruster. We believe that this behavior is attributed to the increased inner thruster volume due to ceramic channel erosion. The greater volume at the thruster EOL causes a reduced density within the discharge channel, that in turn, may lead to lower ionization, thus lower mass utilization efficiency, as previously observed with low power Hall thrusters [ 19 ]. Note that EOL thrust data on-orbit is unavailable since orbit determination was performed only at the beginning of the mission during the first phase (VM1). D. Mission challenges and solutions (On-Orbit Troubleshooting) As previously mentioned, some difficulties were encountered with EPS operation during In-Orbit Testing (IOT) and VM2 where the satellite descended from 720 km to 410 km. For each issue that arose the team in charge of the EPS worked together with the mission operation team to troubleshoot the issue in a prompt manner. Some issues only required tracking of the EPS parameters, and some required a more extensive investigation to reach an optimal solution. A short review of the main issues is presented in this section. Ignition difficulties during IOT Before the system was delivered to the satellite a long End-to-End firing campaign was performed in which the whole propulsion plate was installed on a dedicated mounting fixture installed in the vacuum test facility at Rafael. During this test all the EPS functions and sequences were tested. In particular, the anode mass flow rate parameters for ignition were determined. During IOT the same test was repeated in space. It was found that the discharge current was lower than expected after anode breakdown resulting sometimes in spontaneous beam-out after a few seconds of operation. Some tests were performed on orbit to determine the source of this issue. Thanks to the flexibility of the EPS and particularly the PPU, operational parameters were uploaded in each test to evaluate their effects on the EPS ignition ability. Some examples for parameters testing were magnet current during ignition, cathode mass flow rate, anode mass flow rate, valve timing sequences etc. The first solution was to initiate anode breakdown with no magnetic field. Although this solution succeeded in solving the ignition problem, it was discarded due to the high currents observed on the discharge current circuit which may cause a PPU shutdown trigger. After several trials the most acceptable solution was to increase the mass flow rate through the anode during ignition and reduce it to nominal values in the subsequent seconds. Limited Maximum Discharge Power During AIT and VM1 the thrusters were operated only for a short number of revolutions, and the maximum discharge power was 520 W. When VM2 started the EPS was operated intensively to execute the orbit change. Each thruster was operated on each revolution and for a longer duration per operation. It was observed that the maximum continuous discharge power that the EPS delivered was 420 W when the command was 480 W. When telemetry parameters were investigated it was found that the mass flow controller (DXFC) reached its maximum nominal output, meaning that the discharge power control loop was functioning as intended and the issue was most probably hardware related. The respective investigation concluded that the cause of this phenomenon is the thermal dissipation on the EPS mounting plate and the mass flow controller, affecting its thermal inlet conditions, therefore reducing the effective mass flow rate through the DXFC mass flow restrictors (see Fig. 8 ). A possible solution was to open the sixth DXFC valve (redundant valve) to compensate for the lower mass flow rate to the anodes. However, the implementation of this solution required some on orbit testing, resulting in delaying the scheduled maneuver at this stage. In coordination with the mission operation team, it was decided to limit the maximum discharge power command to 420 W for this phase in order not to disturb the mission's maneuver. Beam-out after ignition After VM2 started some beam-out events were reported via telemetry, mostly immediately after thruster ignition and mainly on Thruster Unit 1 (TU1). At this point, it was decided to stop the autonomous maneuver and keep activating the EPS manually using direct scripts which were uploaded to the satellite. To mitigate the beam-outs events some tests were performed at Rafael's space propulsion facility. An identical Flight Model (FM) IHET-300 thruster was fired to accumulate the total operating time and the total accumulated ignitions performed already with TU1 in space, to simulate as close as possible the thruster operation history. The ground testing at this step did not reveal any issues and the beam-out phenomena was not recreated during the lab test. At this point other tests were performed with the goal of treating the symptoms to facilitate uninterrupted thruster ignition and thruster operation. One possible option was to reduce the intensity of the magnetic field so to reduce discharge resistivity and to facilitate a spontaneous reignition in the case of a quick beam-out, while taking advantage of the fact that the keeper-emitter circuit could be kept on during thruster operation. A test of magnet current sensitivity was performed and thrust, and anode current oscillations were measured. Figure 12 presents thruster discharge current oscillations, main frequency and thrust as measured in the lab. The data shows that regardless of the magnetic field current, instabilities are below 15% of the discharge current and with frequencies in the ~ 10 kHz range. These frequencies are well above the cutoff frequency of the PPU filter. The thrust results show the decrease with magnetic coil current. This decrease can surpass 10% when the magnet current is below 1.5 A. This investigation affirmed that the magnet current can be decreased by 25% of its designed value with little consequences to thruster performance and current oscillations. Thanks to the investigation, updated magnet current parameters were uploaded to the satellite. In addition, it was decided to increase the cathode mass flow rate during TU ignition and TU steady state operation. Lastly, the ignition sequence was updated to ignite the thruster in "Glow mode", that is with no magnet current during anode circuit breakdown. To do so, and to reduce the risk of the high inrush currents on the discharge current circuit that may cause a PPU shutdown trigger, the four anodes power supply modules were turned on consequently to reduce the total output capacitance at the anode output line during circuit breakdown. V. Conclusions Venµs is a sophisticated mission where the propulsion system was responsible for a variety of complex tasks, from orbit maintenance to orbit transfer (up and down) and thrusters’ characterization. The technological mission operated the EPS to demonstrate LEO mission enhancement capability and validate the EPS in space. The EPS consisted of two throttleable Hall thruster, PPU, Propellant Management Assembly (PMA), and a 9-liter propellant tank carrying 16 kg of Xenon. The system was capable of throttling operation in the 300–600 W discharge power range. The EPS successfully passed its In-Orbit Testing (IOT) phase and proved to operate properly according to specifications and expectations. The subsequent mission phases included orbit maintenance demonstration, 720 km→410 km orbit transfer, orbit maintenance at high drag environment, and 410→560 km orbit transfer. Throughout all mission phases the EPS operated within the designated power range, yet most of the time under discharge power of 450 W. Two methods were used to compute the remaining propellant mass – “Bookkeeping” and “PTV” methods. Both methods proved successful and corresponding measurement errors were computed for each method. Thruster performance was measured on-orbit, using the “Orbit Determination” method and compared to laboratory experiments conducted on the ground with identical thrusters, one at its beginning of life, while the other after it past the lifetime test of 130 kN-sec. The measured performance on-orbit was found to be on average lower by up to 5% than the performance measured on the ground. Additional tests were conducted to assess performance variation between BOL and EOL thrusters, as well as a cold ( 120 mins) operation. Lastly, we presented several repeating events in which the propulsion system suffered from sudden beam-outs or ignition difficulties. The methods used to assess a possible solution and implement it were presented. Although designed as a fully redundant EPS, until the end of the technological mission there was no need to switch any redundant component or sub-system. All initial assigned components and branches – completed the mission successfully. Declarations Funding Declaration The Venus project was supported by the Israeli Space Agency (ISA). Author Contribution All authors took part in the design, development, qualification of the electric propulsion system.All authors were involved in the operation of the propulsion system on-orbit, along with conceiving solutions to the issues presented in the manuscript. Acknowledgements The Authors wish to thank the Israeli Space Agency for its long-term support of the program. Although this paper authors are the limited team of the operation phase only, we would like to thank the grand Venµs team that developed, manufactured, tested, integrated and operated the EPS and the technological mission, since its inception in 2004 and until now. In addition, special thanks to ASRI team in the Technion, which developed the TMC, for their commitment and accompaniment of operations. References Paul Lascombes M, Montès A, Fiorentino T, Gelu M, Fillastre, Antonio Gurciullo. Lessons learnt from operating the first Cubesat mission equipped with a Hall thruster. Proceedings of the 35th Annual Small Satellite Conference. Logan, UT, USA, 7–12 Aug 2021. SSC21-XI-01 Ian k, Johnson G, Santiago J, Li, Baldwin J 100,000hrs of On-Orbit Electric Propulsion and MAXAR’s First Electric Orbit Raising, AIAA 2020 – 0189. AIAA Scitech 2020 Forum. January 2020. Gill M, Martinez R, Zannos A, Bailey T, Cassidy M, Cooney J, Crawford A, Cuadra J, Delgado J, Fuller J, Haverty M, Hopkins M, Keyes E, Lee L, Massey D, McClellan K, Pittman J, Schmidt M, Shiyani D, Sorel T On-Orbit Data and Validation of Astra’s ACE Electric Propulsion System. Proceedings of the 36th Annual Small Satellite Conference. Logan, UT, USA, 6–11 Aug 2022. SSC22-S2-08 Jorge J, Delgado JA, Baldwin, Corey RL Space Systems Loral Electric Propulsion Subsystem: 10 Years of On-Orbit Operation. Proceedings of the 34th International Electric Propulsion Conference (IEPC), July 4–10, 2015, Hyogo-Kobe, Japan, IEPC-2015-04 Howard Gray S, Provost M, Glogowski and Alain Demaire. Inmarsat 4F1 Plasma Propulsion System Initial Flight Operations. Proceedings of the 29th International Electric Propulsion Conference (IEPC), Princeton University, Princeton, New Jersey, USA. October 31 – November 4, 2005, IEPC-2005-082 Milligan D, Gestal D, Camino O (2006) SMART-1 Electric Propulsion: An Operational Perspective. Proceedings of the 9th International Conference on Space Operations, Rome, Italy. June 19 – 23, AIAA 2006–5767 Mar Luengo A, Mazzoleni FA, Bravo P, Pisabarro F, Pirondini H, Lee Y-H, Jeong, Se-young Pyo S, Korea (2016) 16–20 May AIAA 2016–2566 Snyder JS, Johnson IK, August, Taylor Kerl and Jeff Baldwin (2021). Pre-Flight Assessment of Xenon Propellant Usage and Usage Uncertainty for the Psyche Mission, AIAA 2021–3427. AIAA Propulsion and Energy 2021 Forum Kim E-H, Kim Y-H, Park J-S, Koh D-W (2015) Yun-Hwang Jeong and Hyun-Woo Le. Orbit Evolution Analysis of DubaiSat-2 using Hall-effect Thruster. J Korean Soc Aeronaut Space Sci 43(4):377–386 Oliver Jia-Richards (2024) Optimized Measurement Timing for In-Space Thrust Inference, AIAA 2024–2071. AIAA SCITECH 2024 Forum Ronald L, Corey, Pidgeon DJ Electric Propulsion at Space Systems/Loral. Proceedings of the 31st International Electric Propulsion Conference (IEPC), University of Michigan, Ann Arbor, Michigan, September 20–24, 2009, USA, IEPC-2009-270 Manzella D, Jankovsky R, Elliott F, Mikellides I, Jongeward G, and Doug Allen (2001). Hall Thruster Plume Measurements On-Board the Russian Express Satellites. Proceedings of the 27th International Electric Propulsion Conference (IEPC), Pasadena, California, USA, October 14–19, IEPC-2001-211217 Jia-Richards O, Lafleur T (2023) Iodine Electric Propulsion System Thrust Validation: From Numerical Modeling to In-Space Testing. J Propul Power 39(6):896–904. https://doi.org/10.2514/1.B39198 Joseph H, Saleh F, Geng M, Ku M, Walker Electric Propulsion Reliability: Statistical Analysis of On-orbit Anomalies and Comparative Analysis of Electric versus Chemical Propulsion Failure Rates, Acta Astronautica , (139) July 2017, pp. 141–156 Herscovitz J, Zuckerman Z, Lev D Electric Propulsion Development at Rafael. Proceedings of the 34th International Electric Propulsion Conference (IEPC), July 4–10, 2015, Hyogo-Kobe, Japan, IEPC-2015-30 Herscovitz J, Karnieli A (2008) VENµS Program: Broad and New Horizons for Super-Spectral Imaging and Electric Propulsion Missions for a Small Satellite Proceedings of the AIAA/USU Conference on Small Satellites, Coming Attractions, SSC08-III-1, http://digitalcommons.usu.edu/smallsat/2008/all2008/14/ Herscovitz J, et-al (2017) VENµS – A Novel Technological Mission Using Electric Propulsion, Proceeding of the 35th International Electric Propulsion Conference, Georgia Institute of Technology, Atlanta, GA, USA, October 8–12 Walker M (2005) Effects of facility backpressure on the performance and plume of a Hall thruster Ahedo E, Gallardo J (2003) Scaling down Hall thrusters 28th Int. Electric Propulsion Conf. (Toulouse, France) Fairview Park OH Electric Rocket Propulsion Society) IEPC 2003 – 104 Additional Declarations No competing interests reported. Cite Share Download PDF Status: Under Review Version 1 posted Editorial decision: Revision requested 05 Apr, 2024 Reviews received at journal 27 Mar, 2024 Reviewers agreed at journal 17 Mar, 2024 Reviewers invited by journal 17 Mar, 2024 Submission checks completed at journal 12 Mar, 2024 Editor assigned by journal 12 Mar, 2024 First submitted to journal 06 Mar, 2024 You are reading this latest preprint version Research Square lets you share your work early, gain feedback from the community, and start making changes to your manuscript prior to peer review in a journal. As a division of Research Square Company, we’re committed to making research communication faster, fairer, and more useful. We do this by developing innovative software and high quality services for the global research community. Our growing team is made up of researchers and industry professionals working together to solve the most critical problems facing scientific publishing. Also discoverable on Platform About Our Team In Review Editorial Policies Advisory Board Help Center Resources Author Services Accessibility API Access RSS feed Manage Cookie Preferences © Research Square 2026 | ISSN 2693-5015 (online) Privacy Policy Terms of Service Do Not Sell My Personal Information {"props":{"pageProps":{"initialData":{"identity":"rs-4022336","acceptedTermsAndConditions":true,"allowDirectSubmit":false,"archivedVersions":[],"articleType":"Research Article","associatedPublications":[],"authors":[{"id":278802797,"identity":"87520352-41f2-437a-8e0a-0753cc99adc3","order_by":0,"name":"Daniel Katz-Franco","email":"","orcid":"","institution":"Rafael Advanced Defense Systems (Israel)","correspondingAuthor":false,"prefix":"","firstName":"Daniel","middleName":"","lastName":"Katz-Franco","suffix":""},{"id":278802798,"identity":"22336cef-e24c-45db-890f-c6ae8f696730","order_by":1,"name":"Dan Lev","email":"data:image/png;base64,iVBORw0KGgoAAAANSUhEUgAAAZAAAAAyAQMAAABI0h/eAAAABlBMVEX///8AAABVwtN+AAAACXBIWXMAAA7EAAAOxAGVKw4bAAAA1klEQVRIiWNgGAWjYDCCA2DSgoGBvQFIG1gQrUWCgYEHxDKQIEWLRAKUQQjwHT97gLmCQULO4Obzqxt+FEgw8Ld3J+DVInkmL4HxDIOEscHtnLKbPUCHSZw5uwGvFoMDOQaMDQwSiTNn56Td4AFqMZDIJaDl/BuwlvqZM8+k3fxDlJYbEFsS+CXYj90myhbJG28MDjYYSBj28+Sw3ZYxkOAh6Be+8zmGDxsqbOTZ2I8/u/nmj40cf3svfi0gcIDBAETxQEiCypEA+wNSVI+CUTAKRsEIAgAOekLbUuGB+gAAAABJRU5ErkJggg==","orcid":"","institution":"Georgia Institute of Technology","correspondingAuthor":true,"prefix":"","firstName":"Dan","middleName":"","lastName":"Lev","suffix":""},{"id":278802799,"identity":"9076c098-2fd1-4dfb-aab3-4b66e8bded73","order_by":2,"name":"Boaz Shoor","email":"","orcid":"","institution":"Rafael Advanced Defense Systems (Israel)","correspondingAuthor":false,"prefix":"","firstName":"Boaz","middleName":"","lastName":"Shoor","suffix":""},{"id":278802800,"identity":"97e4f2fc-933e-4fd2-b1e7-7c9dad7140a6","order_by":3,"name":"Amoz 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(Israel)","correspondingAuthor":false,"prefix":"","firstName":"Jacob","middleName":"","lastName":"Herscovitz","suffix":""}],"badges":[],"createdAt":"2024-03-06 23:00:05","currentVersionCode":1,"declarations":"","doi":"10.21203/rs.3.rs-4022336/v1","doiUrl":"https://doi.org/10.21203/rs.3.rs-4022336/v1","draftVersion":[],"editorialEvents":[],"editorialNote":"","failedWorkflow":false,"files":[{"id":52712582,"identity":"32c7f36d-f983-4119-8530-f2e9887784e2","added_by":"auto","created_at":"2024-03-14 20:10:55","extension":"png","order_by":1,"title":"Figure 1","display":"","copyAsset":false,"role":"figure","size":52210,"visible":true,"origin":"","legend":"\u003cp\u003eVenµs EPS schematic block diagram\u003c/p\u003e","description":"","filename":"1.png","url":"https://assets-eu.researchsquare.com/files/rs-4022336/v1/31463cf4a4f92d05559b5d73.png"},{"id":52712584,"identity":"21a3ea23-648b-4bbc-b88b-902e8f3ece27","added_by":"auto","created_at":"2024-03-14 20:10:56","extension":"png","order_by":2,"title":"Figure 2","display":"","copyAsset":false,"role":"figure","size":1438869,"visible":true,"origin":"","legend":"\u003cp\u003eVenµs propulsion plate\u003c/p\u003e","description":"","filename":"2.png","url":"https://assets-eu.researchsquare.com/files/rs-4022336/v1/b8adeb856d3d2d62e25c4846.png"},{"id":52713013,"identity":"ee11340a-fb38-43e1-a121-9210e6a4a192","added_by":"auto","created_at":"2024-03-14 20:18:56","extension":"jpg","order_by":3,"title":"Figure 3","display":"","copyAsset":false,"role":"figure","size":8877,"visible":true,"origin":"","legend":"\u003cp\u003eHET-300 Flight Model (FM)\u003c/p\u003e","description":"","filename":"3.jpg","url":"https://assets-eu.researchsquare.com/files/rs-4022336/v1/fd1e7723852abe827a4ce413.jpg"},{"id":52712583,"identity":"f2c086a4-d206-4533-885c-6e3a047b114e","added_by":"auto","created_at":"2024-03-14 20:10:56","extension":"png","order_by":4,"title":"Figure 4","display":"","copyAsset":false,"role":"figure","size":247021,"visible":true,"origin":"","legend":"\u003cp\u003eDigital Xenon Flow Controller (DXFC): Picture of the Flight Model (left) and block diagram (right)\u003c/p\u003e","description":"","filename":"4.png","url":"https://assets-eu.researchsquare.com/files/rs-4022336/v1/4ed6df2dcb7a79469e4be828.png"},{"id":52712585,"identity":"d5a6cbbc-1371-4fc4-96c0-5df127aa9452","added_by":"auto","created_at":"2024-03-14 20:10:56","extension":"png","order_by":5,"title":"Figure 5","display":"","copyAsset":false,"role":"figure","size":279737,"visible":true,"origin":"","legend":"\u003cp\u003eLeft: PPU block diagram; Right: Venµs PPU Flight Model (FM)\u003c/p\u003e","description":"","filename":"5.png","url":"https://assets-eu.researchsquare.com/files/rs-4022336/v1/2968a9ef9ffc8e2da9bf21a3.png"},{"id":52712587,"identity":"34b1790b-ec1b-4469-87fe-4ded706303be","added_by":"auto","created_at":"2024-03-14 20:10:56","extension":"png","order_by":6,"title":"Figure 6","display":"","copyAsset":false,"role":"figure","size":23953,"visible":true,"origin":"","legend":"\u003cp\u003eIllustration of a typical orbit change set of maneuvers test during the VM1 phase\u003c/p\u003e","description":"","filename":"6.png","url":"https://assets-eu.researchsquare.com/files/rs-4022336/v1/34129a6b06e6c817f5e11051.png"},{"id":52712580,"identity":"f757968b-bb8c-48d3-b783-e009f421f1a4","added_by":"auto","created_at":"2024-03-14 20:10:55","extension":"png","order_by":7,"title":"Figure 7","display":"","copyAsset":false,"role":"figure","size":48051,"visible":true,"origin":"","legend":"\u003cp\u003ePropellant consumption as a function of the cumulative impulse throughout the mission using three methods – Bookkeeping, PTV and Bookkeeping based on lab test with an identical thruster.\u003c/p\u003e","description":"","filename":"7.png","url":"https://assets-eu.researchsquare.com/files/rs-4022336/v1/7310c67a7b4ef0052338a164.png"},{"id":52712590,"identity":"8a444654-3a95-4434-a852-3b638d4b4904","added_by":"auto","created_at":"2024-03-14 20:10:56","extension":"png","order_by":8,"title":"Figure 8","display":"","copyAsset":false,"role":"figure","size":34581,"visible":true,"origin":"","legend":"\u003cp\u003eTemperature of the DXFC during VM2 (phase 1) EPS operation. Three operation cycles are captured.\u003c/p\u003e","description":"","filename":"8.png","url":"https://assets-eu.researchsquare.com/files/rs-4022336/v1/f504bb068e3e2ac8f75b71fe.png"},{"id":52712589,"identity":"aac19f88-c04a-42ff-b603-46fef7de352e","added_by":"auto","created_at":"2024-03-14 20:10:56","extension":"png","order_by":9,"title":"Figure 9","display":"","copyAsset":false,"role":"figure","size":35386,"visible":true,"origin":"","legend":"\u003cp\u003ePropellant mass as a function of tank pressure at the measured tank temperatures (~25°C). The horizontal errorbars represent the 2% pressure gauge reading uncertainty.\u003c/p\u003e","description":"","filename":"9.png","url":"https://assets-eu.researchsquare.com/files/rs-4022336/v1/cf954da39a80e06e0d7c92c8.png"},{"id":52713015,"identity":"9503e617-cff9-496d-b1b0-2d3fa6179807","added_by":"auto","created_at":"2024-03-14 20:18:56","extension":"png","order_by":10,"title":"Figure 10","display":"","copyAsset":false,"role":"figure","size":41698,"visible":true,"origin":"","legend":"\u003cp\u003eOn-orbit thrust estimation using the Orbit Determination (OD) method for maneuvers performed by each thruster and measured thrust in the lab using an identical thruster at the beginning of life after operation of 20 minutes.\u003c/p\u003e","description":"","filename":"10.png","url":"https://assets-eu.researchsquare.com/files/rs-4022336/v1/da7ba6c0dd85cf79e0f6fade.png"},{"id":52712591,"identity":"e9df534c-85d3-4e40-bb6c-c919fcd4b733","added_by":"auto","created_at":"2024-03-14 20:10:56","extension":"png","order_by":11,"title":"Figure 11","display":"","copyAsset":false,"role":"figure","size":29558,"visible":true,"origin":"","legend":"\u003cp\u003eThrust vs. discharge power as produced by two IHET-300 Hall thrusters in lab, at the Beginning of Life (BOL) and End of Life (EOL) after 20 mins and after 120 mins of operation.\u003c/p\u003e","description":"","filename":"11.png","url":"https://assets-eu.researchsquare.com/files/rs-4022336/v1/59802487276aaea1f4f20ce1.png"},{"id":52713014,"identity":"949c3245-3808-4d8a-93cb-44c1366d8112","added_by":"auto","created_at":"2024-03-14 20:18:56","extension":"png","order_by":12,"title":"Figure 12","display":"","copyAsset":false,"role":"figure","size":54830,"visible":true,"origin":"","legend":"\u003cp\u003eRelative discharge current instability STDV, instability frequency and thrust as a function of electromagnet current as measured in a lab test. The nominal magnet current is 2.5 A.\u003c/p\u003e","description":"","filename":"12.png","url":"https://assets-eu.researchsquare.com/files/rs-4022336/v1/3e674943cbca02c1404c0746.png"},{"id":52713276,"identity":"b7a296f2-5c77-4ec2-a343-7ccf0951c0e9","added_by":"auto","created_at":"2024-03-14 20:26:57","extension":"pdf","order_by":0,"title":"","display":"","copyAsset":false,"role":"manuscript-pdf","size":2396732,"visible":true,"origin":"","legend":"","description":"","filename":"manuscript.pdf","url":"https://assets-eu.researchsquare.com/files/rs-4022336/v1/8ebb2496-e125-47a4-b198-a48baa441827.pdf"}],"financialInterests":"No competing interests reported.","formattedTitle":"On-Orbit Mission Overview of the Low Power Hall Thruster Propulsion System aboard Venμs Satellite","fulltext":[{"header":"I. Introduction","content":"\u003cp\u003eThe past two decades have seen a strikingly emerging presence of spacecraft equipped with electric propulsion systems, and specifically Hall thruster-based systems. Various past satellite missions utilizing Hall thruster technologies reported on their on-orbit experience and even attempted to assess propulsion system performance in space. These missions are for Hall thruster technologies ranging in power from less than 100 W [\u003cspan class=\"CitationRef\"\u003e1\u003c/span\u003e] up to 4.5 kW [\u003cspan class=\"CitationRef\"\u003e2\u003c/span\u003e]. Those missions reported propulsion system cumulative operation duration from several minutes [\u003cspan class=\"CitationRef\"\u003e3\u003c/span\u003e] to thousands of hours and ignition cycles [\u003cspan class=\"CitationRef\"\u003e4\u003c/span\u003e].\u003c/p\u003e\n\u003cp\u003eMost space missions, after the spacecraft is injected into initial orbit, perform a preliminary phase that is called In-Orbit Testing (IOT) where all systems are tested for proper operation, after the launch phase. The Electric Propulsion System (EPS) runs cascading housekeeping and health checks with the goal of assessing its health [\u003cspan class=\"CitationRef\"\u003e1\u003c/span\u003e\u0026ndash;\u003cspan class=\"CitationRef\"\u003e7\u003c/span\u003e]. These include basic PPU electrical checks, pressure and temperature sensors health checks, initial valve activation, and cathode heater startup. IOT is usually performed over the course of several earth revolutions. In the case of the DEIMOS-2 mission the IOT phase led to a re-calibration of the cathode mass flow rate [\u003cspan class=\"CitationRef\"\u003e7\u003c/span\u003e].\u003c/p\u003e\n\u003cp\u003eThroughout the mission there is a need to monitor propellant remaining level in order to track the cumulative impulse used and remaining budget. This is conducted by using two methods \u0026ndash; Bookkeeping and Pressure Temperature Volume (PTV) [\u003cspan class=\"CitationRef\"\u003e2\u003c/span\u003e, \u003cspan class=\"CitationRef\"\u003e3\u003c/span\u003e, \u003cspan class=\"CitationRef\"\u003e8\u003c/span\u003e]. The bookkeeping method, as its name implies, is based on the integration of the mass flow rate over time to compute the mass used to the point in the mission. As with any integrative method, bookkeeping may accumulate any errors in the intended mass flow rate. Its advantage is that it does not rely on sensor measurements. The PTV method computes the remaining mass using the propellant tank temperature and pressure sensor measurements along with the known internal volume of the tank. This measurement can be performed in real time at any point during the mission and does not require historical mission information. None of the propulsion systems reporting on its mission experience found significant discrepancies between the two methods.\u003c/p\u003e\n\u003cp\u003eTo assess propulsion system performance on orbit the Orbit Determination (OD) method is used. This method uses data of the spacecraft orbit change due to propulsion system activity. All EP systems estimating thrust on orbit used this method. A thorough description of the OD method is given in refs. 9 and 10. On-ground and on-orbit performance data for a variety of Hall thruster-based missions is presented in Table\u0026nbsp;\u003cspan class=\"InternalRef\"\u003e1\u003c/span\u003e. It can be seen in the table that some Hall thruster-based systems exhibited an improved performance in space while others had better performance on the ground. On average, it can be determined that the performance of these thrusters is approximately the same on orbit as on the ground. It should be noted that even gridded ion engines reported on-orbit thruster performance similar to that measured on the ground [\u003cspan class=\"CitationRef\"\u003e12\u003c/span\u003e, \u003cspan class=\"CitationRef\"\u003e13\u003c/span\u003e].\u003c/p\u003e\n\u003cdiv class=\"gridtable\"\u003e\n\u003cdiv class=\"colspec\" align=\"left\"\u003e\u0026nbsp;\u003c/div\u003e\n\u003cdiv class=\"colspec\" align=\"left\"\u003e\u0026nbsp;\u003c/div\u003e\n\u003cdiv class=\"colspec\" align=\"left\"\u003e\u0026nbsp;\u003c/div\u003e\n\u003ctable id=\"Tab1\" border=\"1\"\u003e\u003ccaption\u003e\n\u003cdiv class=\"CaptionNumber\"\u003eTable 1\u003c/div\u003e\n\u003cdiv class=\"CaptionContent\"\u003e\n\u003cp\u003eOn-orbit vs. on-ground thruster performance data for different Hall thruster-based propulsion systems.\u003c/p\u003e\n\u003c/div\u003e\n\u003c/caption\u003e\n\u003cthead\u003e\n\u003ctr\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eThruster Name\u003c/p\u003e\n\u003c/th\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eAnode Power\u003c/p\u003e\n\u003cp\u003e[W]\u003c/p\u003e\n\u003c/th\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eOn Ground Thrust\u003c/p\u003e\n\u003cp\u003e[mN]\u003c/p\u003e\n\u003c/th\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eOn Orbit Thrust\u003c/p\u003e\n\u003cp\u003e[mN]\u003c/p\u003e\n\u003c/th\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eDiscrepancy\u003c/p\u003e\n\u003cp\u003e(on-orbit to on-ground)\u003c/p\u003e\n\u003c/th\u003e\n\u003c/tr\u003e\n\u003c/thead\u003e\n\u003ctbody\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eExoMG[\u003cspan class=\"CitationRef\"\u003e1\u003c/span\u003e]\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e~\u0026thinsp;50\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e1.8, 2.4\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e2.2, 3.01\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e\u0026lt;+25%\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eHEPS-200 [\u003cspan class=\"CitationRef\"\u003e7\u003c/span\u003e]\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e200\u0026ndash;400\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e~\u0026thinsp;11\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e~\u0026thinsp;11\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eApprox. the same\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eACE[\u003cspan class=\"CitationRef\"\u003e3\u003c/span\u003e]\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e320\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e20.6\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e22.4\u0026thinsp;\u0026plusmn;\u0026thinsp;5.5\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eWithin uncertainty\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eSPT-100[\u003cspan class=\"CitationRef\"\u003e5\u003c/span\u003e]\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e1350\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e83\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e83.12\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eApprox. the same\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eSPT-100[\u003cspan class=\"CitationRef\"\u003e11\u003c/span\u003e]\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e1350\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e80\u0026thinsp;\u0026plusmn;\u0026thinsp;3\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e83\u0026thinsp;\u0026plusmn;\u0026thinsp;3\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e\u0026lt;+3.7%\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eSPT-100[\u003cspan class=\"CitationRef\"\u003e12\u003c/span\u003e]\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e1350\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e84.05\u0026thinsp;\u0026plusmn;\u0026thinsp;1.3\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e78.5\u0026thinsp;\u0026plusmn;\u0026thinsp;6\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e\u0026lt;-7%\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003ePPS-1350-G[\u003cspan class=\"CitationRef\"\u003e6\u003c/span\u003e]\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e1350\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\u0026nbsp;\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\u0026nbsp;\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e-2%\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eSPT-140[\u003cspan class=\"CitationRef\"\u003e2\u003c/span\u003e]\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e4000\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e226\u0026ndash;236\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e228\u0026ndash;234\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eMostly higher\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003c/tbody\u003e\n\u003c/table\u003e\n\u003c/div\u003e\n\u003cp\u003eIn addition to performance data and regular mission operations some missions reported unexpected events. SMART-1 exhibited several unexpected shutdowns during different phases of the mission [\u003cspan class=\"CitationRef\"\u003e6\u003c/span\u003e]. These were resolved by uploading revised software versions which reduced the number of unexpected occurrences. DEIMOS-2 experienced ignition and operation issues from the first few thruster operations. These were tracked back to erroneous cathode flow setting, resulting in a much lower cathode flow rate than required [\u003cspan class=\"CitationRef\"\u003e7\u003c/span\u003e]. This issue was resolved by recalibration of the cathode mass flow line. The SPT-100 on-board Inmarsat-4F1 exhibited noisy discharge current during initial operations [\u003cspan class=\"CitationRef\"\u003e5\u003c/span\u003e]. These were attributed to thruster outgassing even after propulsion system venting followed by gas purging. Saleh et al. investigated different anomalies and unexpected events with electric propulsion systems on orbit [\u003cspan class=\"CitationRef\"\u003e14\u003c/span\u003e]. He concluded that the mean time between minor anomalies, such that do not impede system operation, is on average 116 days between all Hall thruster-based missions. Saleh also concluded that most of these are the so-called infant-minor-anomalies that are associated with the first several operations of the propulsion system.\u003c/p\u003e\n\u003cp\u003e\u003cstrong\u003eThe VEN\u0026micro;S Program\u003c/strong\u003e\u003c/p\u003e\n\u003cp\u003eIn the recent two decades, Rafael Ltd. has engaged in electric propulsion systems research and development. To date, most efforts have been focused on low power Hall thruster-based propulsion systems [\u003cspan class=\"CitationRef\"\u003e15\u003c/span\u003e]. One such propulsion system was developed for the Ven\u0026micro;s mission. Vegetation and Environment monitoring on a New Micro Satellite, or VEN\u0026micro;S, is a joint program of both the Israeli and French space agencies [\u003cspan class=\"CitationRef\"\u003e16\u003c/span\u003e]. Within the frame of the program, two missions have been executed: a scientific mission and a technological mission.\u003c/p\u003e\n\u003cp\u003eThe mission commenced with the launch of the VEN\u0026micro;S satellite in August 2017 on-board the Vega launcher, with the In Orbit Test (IOT) phase. The space mission was planned for 4\u0026frac12; years. During this period, the satellite conducted a combination of scientific and technological mission phases.\u003c/p\u003e\n\u003cp\u003eThe Ven\u0026micro;s mission consisted of five consecutive phases named VM1 through VM5 (Ven\u0026micro;s Mission X). VM1 consisted of 2.5 years at 720 km altitude, during which Ven\u0026micro;s performed super-spectral imaging with a two-day revisit time, and the EPS performed 29 characterization experiments and controlled the orbit. The next phase, VM2, was devoted to a descending orbit transfer, solely performed by the EPS, to 410 km altitude and 1.2\u0026deg; inclination. Then, in VM3, Ven\u0026micro;s imaged again with a two-days revisit time, this time with an enhanced resolution due to proximity \u0026ndash; while the EPS constantly corrected the orbit in this higher drag environment. The next phase, VM4, consisted of a second orbit transfer, also performed by the EPS, using the remaining Xe propellant to reach 560 km. The next phase, VM5, was solely devoted to imaging in a one-day repeating orbit. Ven\u0026micro;s has depleted all of its Xenon propellant \u0026ndash; and the orbit is controlled by the remaining chemical propellant. All mission orbits were sun-synchronous and earth repeating.\u003c/p\u003e\n\u003cp\u003eThe scientific mission has been super-spectral imaging, using a camera able to capture 12 narrow spectral bands (bandwidth between 25 to 40 nm). Its goal is to image and monitor agricultural growth, vegetation and water bodies, for environment studies. Since the super-spectral imaging relays mainly on light reflection, the imaging angles must be temporarily constant. Thus, sun-synchronous \u0026amp; earth repeating orbits were chosen for each Ven\u0026micro;s phase. Hence the big challenge is controlling the orbit precisely, during all phases and altitudes. This delicate task was assigned to the technological mission and payload \u0026ndash; the EPS, which aided by autonomous orbit control algorithms, strictly controlled the required orbit to satisfy the imaging requirements.\u003c/p\u003e\n\u003cp\u003e.\u003c/p\u003e"},{"header":"II. Technological Mission","content":"\u003cp\u003eThe \u0026lsquo;Technological Mission\u0026rsquo; phases consisted of the operation and qualification of the electric propulsion system in space, orbit control and orbit transfer. As its technological payload, VEN\u0026micro;S incorporated the Israeli Hall Effect Thruster (IHET), which was developed for this purpose. The thruster was designed to fulfill the requirements of the Ven\u0026micro;s small satellite platform of 300 kg, which can produce a relatively low power. IHET\u0026rsquo;s anode power range is between 300 W and 550 W and it consumed a total of 16 kg of Xenon propellant throughout its mission. The IHET is part of Rafael\u0026rsquo;s 1st generation of Electric Propulsion Systems family, denoted as the R-400EPS.\u003c/p\u003e \u003cp\u003eThe two main objectives of the technological mission, space verification and mission validation, were achieved by a series of experiments that tested the EPS performance, followed by five mission phases in which the EPS changed the orbit and maintained it [\u003cspan citationid=\"CR17\" class=\"CitationRef\"\u003e17\u003c/span\u003e].\u003c/p\u003e \u003cp\u003eAll Ven\u0026micro;s technological mission phase was driven by the Technological Mission Module (TMM), which is a Rafael provided flight software component, residing in the on-board computer (OBC). The TMM was specifically designed for the Ven\u0026micro;s technological mission goals. It has three major operation modes: (1) EPS experiment test, (2) Orbit transfer, and (3) Orbit control. The TMM operates autonomously and is fed by commands initiated by the Technological Mission Center (TMC) which is the ground-based facility managing the technological mission. The remarkable autonomous operation of the TMM consists of the ability to estimate the actual orbit on-board the satellite and generate thrust firing EPS commands to correct the orbit to whatever the TMC commanded it [\u003cspan citationid=\"CR9\" class=\"CitationRef\"\u003e9\u003c/span\u003e].\u003c/p\u003e \u003cp\u003eThis article describes the operation of the EPS on-board the Ven\u0026micro;s satellite mission. First, we present all EPS components and their respective role in the system. We then overview on-orbit mission activities which were divided into five phases. Subsequently, we present propellant consumption calculations using two separate methods, as well as the assessment of propulsion system performance using Orbit Determination (OD). Lastly, we discuss unexpected events such as unplanned thruster shutdowns or several failed ignitions attempts, and the methods used to overcome these while in space.\u003c/p\u003e"},{"header":"III. VENµS EPS","content":"\u003cp\u003eA. \u003cb\u003eSystem\u003c/b\u003e\u003c/p\u003e \u003cp\u003eThe Ven\u0026micro;s EPS (see Refs. [\u003cspan citationid=\"CR16\" class=\"CitationRef\"\u003e16\u003c/span\u003e, \u003cspan citationid=\"CR17\" class=\"CitationRef\"\u003e17\u003c/span\u003e]) is designed to support and operate two IHET-300 thrusters, one at a time. Besides the thrusters, the EPS main components are the Propellant Management Assembly (PMA), the Digital Xenon Flow Controller (DXFC), the Power Processing Unit (PPU), and 2 electrical Filter Units (FU). The propellant tank stores highly pressurized Xenon. The system architecture is presented in Fig.\u0026nbsp;\u003cspan refid=\"Fig1\" class=\"InternalRef\"\u003e1\u003c/span\u003e, and the Ven\u0026micro;s propulsion plate is shown in Fig.\u0026nbsp;\u003cspan refid=\"Fig2\" class=\"InternalRef\"\u003e2\u003c/span\u003e.\u003c/p\u003e \u003cp\u003e \u003c/p\u003e \u003cp\u003eOne notable feature of the EPS is the ability to throttle the anodic power of its thrusters, according to the instantaneous available power supplied by the bus platform in orbit. This proves to be an essential feature in a LEO satellite, since the power levels produced by the solar arrays are variable and change according to satellite\u0026rsquo;s true anomaly angle in each revolution. The power level is automatically computed by the TMM and commanded to the PPU which in turn regulates the propellant flow and operates the thrusters at the commanded power level (see Ref. [\u003cspan citationid=\"CR17\" class=\"CitationRef\"\u003e17\u003c/span\u003e]).\u003c/p\u003e \u003cp\u003eA Thruster Selection Unit (TSU) is responsible for selecting the active thruster, as decided by the TMM and commanded by the PPU.\u003c/p\u003e \u003cp\u003e \u003c/p\u003e \u003cp\u003eB. \u003cb\u003eIHET-300 Thruster\u003c/b\u003e\u003c/p\u003e \u003cp\u003e \u003c/p\u003e \u003cp\u003eThe heart of the EPS is the Hall Effect Thruster (HET), codenamed IHET-300, shown in Fig.\u0026nbsp;\u003cspan refid=\"Fig3\" class=\"InternalRef\"\u003e3\u003c/span\u003e. It operates on Xenon, which is ionized by electrons emitted from the cathode and accelerated in the form of plasma via a high electric field.\u003c/p\u003e \u003cp\u003e \u003cdiv class=\"gridtable\"\u003e\u003ctable float=\"Yes\" id=\"Tab2\" border=\"1\"\u003e \u003ccaption language=\"En\"\u003e \u003cdiv class=\"CaptionNumber\"\u003eTable 2\u003c/div\u003e \u003cdiv class=\"CaptionContent\"\u003e \u003cp\u003eIHET-300 Main Characteristics @ 300W\u003c/p\u003e \u003c/div\u003e \u003c/caption\u003e \u003ccolgroup cols=\"2\"\u003e \u003cdiv align=\"left\" class=\"colspec\" colname=\"c1\" colnum=\"1\"\u003e\u003c/div\u003e \u003cdiv align=\"left\" class=\"colspec\" colname=\"c2\" colnum=\"2\"\u003e\u003c/div\u003e \u003cthead\u003e \u003ctr\u003e \u003cth align=\"left\" colname=\"c1\"\u003e \u003cp\u003eThrust @ 300W (EOL)\u003c/p\u003e \u003c/th\u003e \u003cth align=\"left\" colname=\"c2\"\u003e \u003cp\u003e\u0026gt;\u0026thinsp;14.3 mN\u003c/p\u003e \u003c/th\u003e \u003c/tr\u003e \u003c/thead\u003e \u003ctbody\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eSpecific Impulse @ 300W (EOL)\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003e\u0026gt;\u0026thinsp;1210 sec\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003ePower Operation Range\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003e250W to 600W\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eOperating Life\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003e\u0026gt;\u0026thinsp;1100 hours\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eNumber of Operation Cycles\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003e\u0026gt;\u0026thinsp;2000\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eTotal Impulse\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003e\u0026gt;\u0026thinsp;135 kNs\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003c/tbody\u003e \u003c/colgroup\u003e \u003c/table\u003e\u003c/div\u003e \u003c/p\u003e \u003cp\u003eThe main characteristics of IHET-300 are listed in Table\u0026nbsp;\u003cspan refid=\"Tab2\" class=\"InternalRef\"\u003e2\u003c/span\u003e. This thruster, which employs low power, is specifically designed to be used onboard small and micro satellites. Its useful range of operation is between 250 to 600W. On Ven\u0026micro;s platform we limit the operation power to 550W, due to platform power limitations. Thus, it can utilize the instantaneous available power from the satellite as explained above.\u003c/p\u003e \u003cp\u003eThe IHET-300 mass is about 1.5 kg and its dimensions are 170\u0026times;120\u0026times;90 mm.\u003c/p\u003e \u003cp\u003eC. \u003cb\u003ePropellant Management Assembly (PMA)\u003c/b\u003e\u003c/p\u003e \u003cp\u003eThe PMA is responsible for feeding Xenon gas, at controlled pressure and flowrate, to the selected thruster.\u003c/p\u003e \u003cp\u003eIt consists mainly of adapted and qualified COTS components. Ref. [\u003cspan citationid=\"CR5\" class=\"CitationRef\"\u003e5\u003c/span\u003e] describes the development and qualification process of most EPS components. A picture of the Digital Xenon Flow Controller and functional schematic are shown in Fig.\u0026nbsp;\u003cspan refid=\"Fig4\" class=\"InternalRef\"\u003e4\u003c/span\u003e.\u003c/p\u003e \u003cp\u003eThe PMA consists of two (2) pressure regulators connected in series. The pressure regulation works in an open loop mode pre-calibrated on the ground, while the first regulator reduces the tank high pressure feed to nearly the operating pressure and the second regulator gives a more accurate output of 1.8 bara. This setup provides a very accurate pressure at the output of the pressure regulators.\u003c/p\u003e \u003cp\u003eThe Digital Xenon Flow Controller (DXFC) (shown in Fig.\u0026nbsp;\u003cspan refid=\"Fig4\" class=\"InternalRef\"\u003e4\u003c/span\u003e) uses a set of six (6) valves and flow restrictors arranged as a gas manifold to control the total Xenon flow through the anodes; therefore, controlling the discharge power and thrust levels. This flow controller setup is fully redundant in the case some of the valves fail to open. The DXFC controlled by the PPU would select a configuration of valves to maintain the required discharge power within a tolerance of 20W, the control loop checks the discharge power and if required it updates the DXFC valve configuration. Only one controller was installed on board the EPS while the flow was directed to each thruster by opening a series of valves downstream near the thrusters.\u003c/p\u003e \u003cp\u003e \u003c/p\u003e \u003cp\u003eD. \u003cb\u003ePPU\u003c/b\u003e\u003c/p\u003e \u003cp\u003eThe PPU contains the diverse power supplies and control circuits for operation and management of primarily the thrusters and the whole EPS components. The Sequencer and Control Unit (SCU) is the control module of the PPU. Besides controlling the power supply and the ignition sequence, the PPU also implements communication with the satellite OBC, monitors all signals and prepares the EPS telemetry.\u003c/p\u003e \u003cp\u003eThe PPU block diagram is shown in Fig.\u0026nbsp;\u003cspan refid=\"Fig5\" class=\"InternalRef\"\u003e5\u003c/span\u003e herein with its main interfaces to the EPS sub-assemblies. The PPU is thermally insulated from the satellite and its radiator on the bottom side is exposed to space for cooling by heat radiation.\u003c/p\u003e \u003cp\u003e \u003c/p\u003e \u003cp\u003eFigure \u003cspan refid=\"Fig5\" class=\"InternalRef\"\u003e5\u003c/span\u003e also shows the PPU flight model, while Table\u0026nbsp;\u003cspan refid=\"Tab3\" class=\"InternalRef\"\u003e3\u003c/span\u003e shows its main characteristics.\u003c/p\u003e \u003cp\u003e \u003cdiv class=\"gridtable\"\u003e\u003ctable float=\"Yes\" id=\"Tab3\" border=\"1\"\u003e \u003ccaption language=\"En\"\u003e \u003cdiv class=\"CaptionNumber\"\u003eTable 3\u003c/div\u003e \u003cdiv class=\"CaptionContent\"\u003e \u003cp\u003ePPU main characteristics\u003c/p\u003e \u003c/div\u003e \u003c/caption\u003e \u003ccolgroup cols=\"2\"\u003e \u003cdiv align=\"left\" class=\"colspec\" colname=\"c1\" colnum=\"1\"\u003e\u003c/div\u003e \u003cdiv align=\"left\" class=\"colspec\" colname=\"c2\" colnum=\"2\"\u003e\u003c/div\u003e \u003cthead\u003e \u003ctr\u003e \u003cth align=\"left\" colname=\"c1\"\u003e \u003cp\u003eCharacteristic\u003c/p\u003e \u003c/th\u003e \u003cth align=\"left\" colname=\"c2\"\u003e \u003cp\u003eValue\u003c/p\u003e \u003c/th\u003e \u003c/tr\u003e \u003c/thead\u003e \u003ctbody\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eMass\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003e12kg\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eEfficiency at max. power\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003e91%\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eAnodic power\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003eup to 600W @ 300 VDC\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eThermal control\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003eAutonomous, through bottom radiator\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eAnode voltage PS\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003e300 VDC\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eMagnetic Field PS\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003e0.6\u0026ndash;2.6A, variable control for optimal performance\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eCathode Heater Current\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003eSelectable up to 11 ADC\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eCathode keeper Voltage\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003e60 VDC\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eAuxiliary PS\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003e5 VDC, 16 VDC and 28 VDC\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eRedundancy\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003eFull cold redundancy, reconfigurable in space\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003ctr\u003e \u003ctd align=\"left\" colname=\"c1\"\u003e \u003cp\u003eInterface\u003c/p\u003e \u003c/td\u003e \u003ctd align=\"left\" colname=\"c2\"\u003e \u003cp\u003ePower: Unregulated bus, 34 to 42 V\u003c/p\u003e \u003cp\u003eCommunications: Dual redundant CAN-Bus\u003c/p\u003e \u003c/td\u003e \u003c/tr\u003e \u003c/tbody\u003e \u003c/colgroup\u003e \u003c/table\u003e\u003c/div\u003e \u003c/p\u003e"},{"header":"IV. On-Orbit Operation","content":"\u003cp\u003eAfter the launch of Ven\u0026micro;s satellite, a series of In-Orbit Tests (IOT) were performed. These tests checked and assured that the platform and its payloads were ready for the intended mission. The technological mission experiments started as planned in early 2018 and continued until the end of VM1. All other mission objectives were fully accomplished, and all propellant was depleted in 2022.\u003c/p\u003e\n\u003cp\u003eA. \u003cstrong\u003eOverview of electric propulsion system operation\u003c/strong\u003e\u003c/p\u003e\n\u003cp\u003eThe IOT consists of a large series of trials in which an additional function is checked in turn until a full EPS activation is achieved, including uplink communication and download of telemetry. The success of the IOT is a prerequisite for the next EPS mission phases. Table\u0026nbsp;\u003cspan class=\"InternalRef\"\u003e4\u003c/span\u003e summarizes the main tests performed.\u003c/p\u003e\n\u003cdiv class=\"gridtable\"\u003e\n\u003cdiv class=\"colspec\" align=\"left\"\u003e\u0026nbsp;\u003c/div\u003e\n\u003cdiv class=\"colspec\" align=\"left\"\u003e\u0026nbsp;\u003c/div\u003e\n\u003ctable id=\"Tab4\" border=\"1\"\u003e\u003ccaption\u003e\n\u003cdiv class=\"CaptionNumber\"\u003eTable 4\u003c/div\u003e\n\u003cdiv class=\"CaptionContent\"\u003e\n\u003cp\u003eIn-Orbit Testing (IOT) set of checks and tests\u003c/p\u003e\n\u003c/div\u003e\n\u003c/caption\u003e\n\u003cthead\u003e\n\u003ctr\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eTests Name\u003c/p\u003e\n\u003c/th\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eTest Description\u003c/p\u003e\n\u003c/th\u003e\n\u003c/tr\u003e\n\u003c/thead\u003e\n\u003ctbody\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003ePPU Telephone Home\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003ePPU switched on for the first time, tested and parameters are set\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003ePPU Uplink/Downlink\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eTest all power supplies, redundant paths and cross commands\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003ePneumatics Check\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eA series of tests in which valve is open and close. Test verification by examining the pressure transducers.\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eSystem vent\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eLow-pressure segment valves are opened and each branch is vented.\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eCathodes Outgassing\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eCathodes outgassing\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eEPS first ignition\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eEach thruster is ignited for the first time to check proper operation.\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eEPS autonomous module\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eA concluding end-to-end station keeping test for automatic orbit maintenance\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003c/tbody\u003e\n\u003c/table\u003e\n\u003c/div\u003e\n\u003cp\u003eAfter successful completion of the IOT, the EPS was ready to commence its technological mission execution in each of the five Ven\u0026micro;s phases, as shown in Table\u0026nbsp;\u003cspan class=\"InternalRef\"\u003e5\u003c/span\u003e.\u003c/p\u003e\n\u003cp\u003eDuring VM1, the satellite performed each month a technological mission experiment. The goal of these experiments was to estimate the thrust performance of each Thruster Unit (TU). For this purpose, the TMM commanded the EPS a predefined series of firing pulses, that would change the orbit by increasing and decreasing the satellite\u0026rsquo;s Semi-Major Axis (SMA). The whole maneuver consisted of orbit extrication in three Open Loop (OL) segments and finished with orbit re-insertion in a Closed Loop (CL) segment, as shown in Fig.\u0026nbsp;\u003cspan class=\"InternalRef\"\u003e6\u003c/span\u003e.\u003c/p\u003e\n\u003cp\u003e\u0026nbsp;\u003c/p\u003e\n\u003cp\u003eEach OL segment consists of three satellite revolutions, with a one long EPS activation at a constant power between 300 and 550W (meaning constant thrust) for each revolution. In the first segment, EPS activations raised the orbit by aligning the satellite's thruster orientation with the velocity vector. The second and third segments also contained EPS activations for three consecutive revolutions each, but the satellite attitude was commanded by the TMM so that the orbit Semi-Major Axis (SMA) will decrease. At the end of each segment, the EPS was left idle for one revolution, dedicated to orbit measurements and estimation. The overall design goal was to finish the three segments in an orbit lower than the initial one. Subsequently, the CL segment commences. The TMM in CL mode commands the EPS to fire and operate on each consecutive revolution for a duration and altitude instantaneously calculated, until the orbit reached the target value. The orbit data telemetry is then analyzed by the TMC, which estimates what was the force applied that resulted in these orbit changes. Then, all parasitic forces are compensated, resulting in the average thrust actually applied by the TU.\u003c/p\u003e\n\u003cp\u003eDuring this VM1 phase the EPS was well characterized in terms of operational parameters (mass flow rates requirements, anode current and anode current noise measurements, system various temperatures, etc.), ignition sequence, downlink and uplink etc., and some parameters calibrations were made through the uplink to the satellite such us mass flow rate to the anode during the ignition. This phase ended on October 15th, 2020, with a total of 29 experiments, 380 ignitions, 86.05 operating hours and a total calculated impulse of 6.7kNs accumulated by the EPS. During this phase of characterization, the EPS was activated across all its discharge power range, which is 300W to 520W. The average firing duration was 20 minutes for each thruster including the startup ignition sequence.\u003c/p\u003e\n\u003cp\u003eDuring VM2, the satellite was commanded to make an autonomous orbit transfer from a 720 km SSO to 410 km SSO. The maneuver started on November 4th, 2020. During this phase, the EPS thrusters were activated in activated in 90% of the revolutions in each day in which TU1 was operated for approx. 25 minutes and TU2 was operated for approx. 15 minutes. During the remaining revolutions telemetry was downloaded, and additional checks were performed on other satellite subsystems. At each revolution the EPS operated both thrusters to lower the altitude of the satellite. To do so, after one thruster completed its burn, the satellite was rotated about 180\u0026deg; to align the second thruster with the correct velocity vector direction. After several tens of revolutions two significant phenomena were observed, the first is that the maximum continuous discharge power that the EPS delivered was 420W. An investigation was performed on this matter which concluded that the cause of this phenomena was the thermal dissipation on the EPS mounting plate and the mass flow controller, affecting its inlet conditions, therefore reducing the effective mass flow rate. One possible solution to this matter was to use the redundancy of the DXFC to compensate the lower mass flow rate to the anodes. However, the implementation of this solution required some on-orbit testing to revalidate operation, resulting in delaying the scheduled maneuver. At this stage it was decided to limit and leave the maximum discharge power command to 420 W for this phase in order not to disturb the mission's maneuver schedule.\u003c/p\u003e\n\u003cp\u003eThe second phenomenon was that some beam-out events were reported via telemetry, mostly immediately after thruster ignition. Beam-out events are events in which the discharge spontaneously extinguishes. At this point, it was decided to stop the autonomous maneuver and keep activating the EPS manually using direct scripts which were uplinked to the satellite. To mitigate the beam-out events some tests were performed at Rafael's space propulsion facility resulting in an update of the magnetic field parameters and a change in the ignition routine to start the thrusters with no magnet current, that is also referred to as \"\u003cem\u003eGlow mode\u003c/em\u003e\" ignition. This event proved the benefit of project risk management that was conducted during the development phase, which resulted in two redundant and optional modes of TU ignition and operation. The target orbit of 410 km was reached on September 5th, 2021, with a total of 4,077 ignitions, 1,382.5 operating hours and a total calculated impulse of 110.6 kNs. A total of 23 beam-out events were registered.\u003c/p\u003e\n\u003cp\u003eThe main objective of VM3 in terms of the EPS was to provide autonomous station keeping maneuvers which enable the scientific payload to image with a better resolution than in the previous orbit. This was achieved thanks to the shorter distance to Earth and at the cost of more frequent station keeping maneuvers. By performing this phase, we demonstrated, the feasibility of enabling enhanced imaging resolution without increasing the size and mass of the satellite\u0026rsquo;s camera. Ven\u0026micro;s satellite maintained this orbit for a total duration of one month, while four station keeping maneuvers, each consisting of 9 thruster operations, were required to overcome the higher drag at this altitude. The autonomous module performed a total of 36 thruster ignitions throttling the discharge power between 300W and 480W, accumulating 11.8 thruster operating hours. By the end of this phase of the mission the EPS accumulated a total of 4,113 ignitions, 1394.5 operating hours and a total of calculated impulse of 111.6 kNs.\u003c/p\u003e\n\u003cp\u003eAfter VM3 was completed, at altitude of 410 km, VM4 phase started, the EPS was commanded to raise the satellite's orbit to a 560 km SSO and ended at VM5 after a successful orbit insertion at this altitude to continue the scientific mission. The VM4 maneuver started on November 9th, 2021, using the same technique as in VM2, but this time it was manually commanded by the Ground Station, on open loop, and ended with the depletion of the Xenon propellant on March 5th, 2022.\u003c/p\u003e\n\u003cp\u003eOverall, throughout the mission, and until all propellant was consumed, the EPS achieved a total of 6,311 successful ignitions accumulated 2105.1 operating hours, that is 158 kNs and a total throughput of 16 kg of xenon.\u003c/p\u003e\n\u003cdiv class=\"gridtable\"\u003e\n\u003cdiv class=\"colspec\" align=\"left\"\u003e\u0026nbsp;\u0026nbsp;\u003c/div\u003e\n\u003ctable id=\"Tab5\" border=\"1\"\u003e\u003ccaption\u003e\n\u003cdiv class=\"CaptionNumber\"\u003eTable 5\u003c/div\u003e\n\u003cdiv class=\"CaptionContent\"\u003e\n\u003cp\u003eSummary of the EPS utilization at different mission phases.\u003c/p\u003e\n\u003c/div\u003e\n\u003c/caption\u003e\n\u003cthead\u003e\n\u003ctr\u003e\n\u003cth rowspan=\"2\" align=\"left\"\u003e\n\u003cp\u003eMission Phase\u003c/p\u003e\n\u003c/th\u003e\n\u003cth rowspan=\"2\" align=\"left\"\u003e\n\u003cp\u003eDescription\u003c/p\u003e\n\u003c/th\u003e\n\u003cth colspan=\"2\" align=\"left\"\u003e\n\u003cp\u003e# Ignitions [#]\u003c/p\u003e\n\u003c/th\u003e\n\u003cth colspan=\"2\" align=\"left\"\u003e\n\u003cp\u003e# Operating Hours [hr]\u003c/p\u003e\n\u003c/th\u003e\n\u003cth colspan=\"2\" align=\"left\"\u003e\n\u003cp\u003eImpulse [kNs]\u003c/p\u003e\n\u003c/th\u003e\n\u003cth colspan=\"2\" align=\"left\"\u003e\n\u003cp\u003eAccumulated Propellant Usage [kg]\u003c/p\u003e\n\u003c/th\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eTU1\u003c/p\u003e\n\u003c/th\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eTU2\u003c/p\u003e\n\u003c/th\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eTU1\u003c/p\u003e\n\u003c/th\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eTU2\u003c/p\u003e\n\u003c/th\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eTU1\u003c/p\u003e\n\u003c/th\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eTU2\u003c/p\u003e\n\u003c/th\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003ePTV\u003c/p\u003e\n\u003c/th\u003e\n\u003cth align=\"left\"\u003e\n\u003cp\u003eBookkeeping\u003c/p\u003e\n\u003c/th\u003e\n\u003c/tr\u003e\n\u003c/thead\u003e\n\u003ctbody\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eIOT\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eIn Orbit Test\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e10\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e12\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e2\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e1.4\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e0.1\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e0.1\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eToo low to compute\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e0.02\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eVM1\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eEPS Characterization and orbit correction\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e140\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e218\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e41.4\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e41.2\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e3.3\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e3.2\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e0.68\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e0.63\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eVM2\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eOrbit Transfer\u003c/p\u003e\n\u003cp\u003e(720 km\u0026rarr;410 km)\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e1863\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e1834\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e844.5\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e452\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e66.5\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e37.5\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e11.57\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e10.41\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eVM3\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eOrbit Maintenance (410 km)\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e18\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e18\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e7.7\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e4.2\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e0.6\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e0.4\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eInsuff. data\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e10.49\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eVM4\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eOrbit Transfer\u003c/p\u003e\n\u003cp\u003e(410 km\u0026rarr;560 km)\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e920\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e940\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e258\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e330.2\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e17.8\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e21.2\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e14.74\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e14.34\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eVM5\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003eEnd of EPS Mission\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e22\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e316\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e3.9\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e118.5\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e0.3\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e7.1\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e15.85\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e15.03\u003c/p\u003e\n\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd rowspan=\"2\" align=\"left\"\u003e\n\u003cp\u003eTotal\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd rowspan=\"2\" align=\"left\"\u003e\n\u003cp\u003eEOL\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e2973\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e3338\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e1157.5\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e947.5\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e88.6\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\n\u003cp\u003e69.5\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\u0026nbsp;\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\u0026nbsp;\u003c/td\u003e\n\u003c/tr\u003e\n\u003ctr\u003e\n\u003ctd colspan=\"2\" align=\"left\"\u003e\n\u003cp\u003e6311\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd colspan=\"2\" align=\"left\"\u003e\n\u003cp\u003e2105\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd colspan=\"2\" align=\"left\"\u003e\n\u003cp\u003e158.1\u003c/p\u003e\n\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\u0026nbsp;\u003c/td\u003e\n\u003ctd align=\"left\"\u003e\u0026nbsp;\u003c/td\u003e\n\u003c/tr\u003e\n\u003c/tbody\u003e\n\u003c/table\u003e\n\u003c/div\u003e\n\u003ch2\u003eB. Propellant consumption\u003c/h2\u003e\n\u003cp\u003eDuring the propellant loading campaign in 2017, the propellant tank was pressurized with 16.1 kg high purity Xenon. As of the end of mission, the propellant was entirely depleted. Subsequently and throughout all mission phases, calculations were made to determine the propellant used and remaining. Two conventional methods were used to determine the remaining propellant mass at any time during the mission: (1) PTV method and (2) Bookkeeping method [\u003cspan class=\"CitationRef\"\u003e2\u003c/span\u003e, \u003cspan class=\"CitationRef\"\u003e3\u003c/span\u003e, \u003cspan class=\"CitationRef\"\u003e8\u003c/span\u003e]. For the first method, periodic telemetry provides the pressure and temperature data of the Xenon tank. Using this data, the pressure dependent tank volume and basic thermodynamic state tables, the Xenon propellant mass was calculated. This method can be used at any point during the mission and regardless of previous propellant mass estimations used earlier. On the other hand, the bookkeeping method is integrative. It integrates the mass flow rate injected by the propulsion system over the operation time to compute the amount of propellant expelled. This amount is then subtracted from the initial propellant mass prior to the respective operation. The estimation of the mass flow rate injected into the thruster and cathode were calibrated using mass flow meters before satellite integration. This method has the disadvantage of accumulating the computation error from each operation to the next. Additionally, this method relies on the initial calibration of the mass flow controller prior to installation on-board the satellite and may accrue any mass flow calibration errors.\u003c/p\u003e\n\u003cp\u003eIn addition to the two methods, we used a third method based on laboratory experiments on an identical thruster unit. We operated an identical IHET-300 Flight Model (FM) thruster to find the correlation between the thruster discharge power and the mass flow rate. The background pressure during the test was between 1.8\u0026times;10\u003csup\u003e\u0026minus;\u0026thinsp;5\u003c/sup\u003e Torr and 2.9\u0026times;10\u003csup\u003e\u0026minus;\u0026thinsp;5\u003c/sup\u003e Torr (N\u003csub\u003e2\u003c/sub\u003e corrected). We then used the ṁ\u0026rarr;P\u003csub\u003ed\u003c/sub\u003e relation to assess the mass flow rate used during each thruster operation in space and based on the operating anode power.\u003c/p\u003e\n\u003cp\u003eFigure \u003cspan class=\"InternalRef\"\u003e7\u003c/span\u003e shows the propellant consumption with the cumulative impulse of the propulsion system. Propellant consumption is approximately linear, even though the activation time for each thruster varied throughout the mission. This does not come as a surprise as the slope of the trend shown in the figure is proportional to the inverse of the specific impulse. If the specific impulse is approximately constant so will be the slope. It can also be seen that all three methods produce approximately the same result, albeit an increasing deviation from each other as the mission progresses. This increasing deviation is most likely due to a cumulative error due to a slight underestimation of the expelled propellant during all activations.\u003c/p\u003e\n\u003cp\u003eIt is interesting to note that the error associated with the \u0026ldquo;Bookkeeping Method\u0026rdquo; increases with EPS activation. This is due to the integrative nature of this method, that relies on the previous assessment of the remaining mass when computing the propellant mass at a given time. Since the temperature and pressure of the DXFC has changed throughout the mission, it was necessary to correct the calculation of the mass flow controller according to the temperature and input pressure. The DXFC was calibrated on the ground at pressure and temperatures of 1.8 bara and 20\u0026deg;C. On orbit, the input pressure was measured by the low-pressure transducer located upstream of the DXFC. The value of the pressure varied during firing due to temperature changes. The temperature was measured using a thermistor located on the body of the DXFC. Since the DXFC is located close to the PPU it was subjected to heat flux that raised its temperature. Figure\u0026nbsp;\u003cspan class=\"InternalRef\"\u003e8\u003c/span\u003e shows DXFC temperature during a five-day long EPS activations cycle (tens of thruster activations) at VM2 (second mission phase). It can be seen from the figure that the DXFC temperature increases from ~\u0026thinsp;22\u0026deg;C to 28\u0026deg;C during each activation cycle. The corresponding correction was made to assess the mass flow rate deviation from the calibrated values. In any case, the PPU operated the EPS in a closed-loop manner, while adjusting the mass flow rate to match the available/commanded power.\u003c/p\u003e\n\u003cp\u003eThe errors associated with the \u0026ldquo;PTV Method\u0026rdquo; are different in nature since this method relies on real-time pressure and temperature measurements. The challenge with the \u0026ldquo;PTV Method\u0026rdquo; for Xenon is that Xenon density is very sensitive to changes in pressure in the 60\u0026ndash;80 bar range, at temperatures around 25\u0026deg;C. This can be observed in Fig.\u0026nbsp;\u003cspan class=\"InternalRef\"\u003e9\u003c/span\u003e where propellant mass is plotted against tank pressure at the measured tank temperatures (~\u0026thinsp;25\u0026deg;C). Small variations in pressure may lead to large variations in measured propellant mass. For this reason, propellant mass uncertainty can go up to over 0.5 kg during the intermediate phases of the mission when the propellant mass is approximately half of the initial loaded mass. The \u0026ldquo;PTV Method\u0026rdquo; errors presented in Fig.\u0026nbsp;\u003cspan class=\"InternalRef\"\u003e7\u003c/span\u003e are based on measuring device error of 0.6 bar for the high-pressure gauge and 0.2\u0026deg;C for the temperature measurement by the propellant tank thermistor.\u003c/p\u003e\n\u003cp\u003eC. \u003cstrong\u003eOn-orbit performance estimation\u003c/strong\u003e\u003c/p\u003e\n\u003cp\u003ePropulsion system performance was estimated on-orbit using Orbit Determination (OD). To use this method the orbit of the satellite was determined using GPS location before a satellite maneuver was performed. Subsequently, the propulsion system was operated, and the new orbit was determined after the maneuver was completed. This method was used in previous Hall thruster-based missions and is thoroughly described in refs. 9 and 10. Additionally, an identical Hall thruster FM was operated in a vacuum chamber at Rafael to measure its performance and compare to the on-orbit data. The identical FM was at its beginning of life and only operated during its acceptance test previously. To adequately compare thruster operation the lab test operation was conducted for 20 minutes, approximately the same duration as its on-orbit counterparts. Background pressure during lab operation was between 1.8\u0026times;10\u003csup\u003e\u0026minus;\u0026thinsp;5\u003c/sup\u003e Torr and 2.9\u0026times;10\u003csup\u003e\u0026minus;\u0026thinsp;5\u003c/sup\u003e Torr (N\u003csub\u003e2\u003c/sub\u003e corrected). Past laboratory acceptance tests with the units operating in space were deemed unreliable due to exceedingly high chamber pressure.\u003c/p\u003e\n\u003cp\u003eThe on-orbit and lab thrust data are presented in Fig.\u0026nbsp;\u003cspan class=\"InternalRef\"\u003e10\u003c/span\u003e. It can be observed in the figure that the thrust increases approximately linearly with discharge power for on-orbit and lab operations. It can also be observed that the thruster operating in lab environment measured higher thrust values than both on-orbit thrusters by approximately up to 5% on average. We can attribute this to the difference in background pressure, although the background pressure during the lab test is considered sufficient for the estimation of performance in space [ref. 18]. Also, \u0026lsquo;Thruster 2\u0026rsquo; exhibited higher performance than \u0026lsquo;Thruster 1\u0026rsquo; by about 5% yet within thrust assessment error bars.\u003c/p\u003e\n\u003cp\u003eAnother set of laboratory experiments were conducted to understand thruster operation differences at the Beginning of Life (BOL) and End of Life (EOL), as well when it is considered operating cold (after 20 minutes of operation) and when it is considered hot (\u0026gt;\u0026thinsp;2 hrs operation). To do this another laboratory thruster was used that completed the lifetime qualification test of 130 kNsec. The qualification thruster is considered at its EOL. Experiment results are presented in Fig.\u0026nbsp;\u003cspan class=\"InternalRef\"\u003e11\u003c/span\u003e. It can be seen in the figure that there is little discernable difference between \u0026ldquo;hot\u0026rdquo; and \u0026ldquo;cold\u0026rdquo; operations of the thrusters in lab, whether the thrusters are at BOL or EOL. It can also be seen that the EOL thruster performance is slightly lower than that of the BOL thruster. We believe that this behavior is attributed to the increased inner thruster volume due to ceramic channel erosion. The greater volume at the thruster EOL causes a reduced density within the discharge channel, that in turn, may lead to lower ionization, thus lower mass utilization efficiency, as previously observed with low power Hall thrusters [\u003cspan class=\"CitationRef\"\u003e19\u003c/span\u003e].\u003c/p\u003e\n\u003cp\u003eNote that EOL thrust data on-orbit is unavailable since orbit determination was performed only at the beginning of the mission during the first phase (VM1).\u003c/p\u003e\n\u003cp\u003eD. \u003cstrong\u003eMission challenges and solutions (On-Orbit Troubleshooting)\u003c/strong\u003e\u003c/p\u003e\n\u003cp\u003eAs previously mentioned, some difficulties were encountered with EPS operation during In-Orbit Testing (IOT) and VM2 where the satellite descended from 720 km to 410 km. For each issue that arose the team in charge of the EPS worked together with the mission operation team to troubleshoot the issue in a prompt manner. Some issues only required tracking of the EPS parameters, and some required a more extensive investigation to reach an optimal solution. A short review of the main issues is presented in this section.\u003c/p\u003e\n\u003cp\u003e\u003cem\u003eIgnition difficulties during IOT\u003c/em\u003e\u003c/p\u003e\n\u003cp\u003eBefore the system was delivered to the satellite a long End-to-End firing campaign was performed in which the whole propulsion plate was installed on a dedicated mounting fixture installed in the vacuum test facility at Rafael. During this test all the EPS functions and sequences were tested. In particular, the anode mass flow rate parameters for ignition were determined.\u003c/p\u003e\n\u003cp\u003eDuring IOT the same test was repeated in space. It was found that the discharge current was lower than expected after anode breakdown resulting sometimes in spontaneous beam-out after a few seconds of operation. Some tests were performed on orbit to determine the source of this issue. Thanks to the flexibility of the EPS and particularly the PPU, operational parameters were uploaded in each test to evaluate their effects on the EPS ignition ability. Some examples for parameters testing were magnet current during ignition, cathode mass flow rate, anode mass flow rate, valve timing sequences etc. The first solution was to initiate anode breakdown with no magnetic field. Although this solution succeeded in solving the ignition problem, it was discarded due to the high currents observed on the discharge current circuit which may cause a PPU shutdown trigger. After several trials the most acceptable solution was to increase the mass flow rate through the anode during ignition and reduce it to nominal values in the subsequent seconds.\u003c/p\u003e\n\u003cp\u003e\u003cem\u003eLimited Maximum Discharge Power\u003c/em\u003e\u003c/p\u003e\n\u003cp\u003eDuring AIT and VM1 the thrusters were operated only for a short number of revolutions, and the maximum discharge power was 520 W.\u003c/p\u003e\n\u003cp\u003eWhen VM2 started the EPS was operated intensively to execute the orbit change. Each thruster was operated on each revolution and for a longer duration per operation. It was observed that the maximum continuous discharge power that the EPS delivered was 420 W when the command was 480 W. When telemetry parameters were investigated it was found that the mass flow controller (DXFC) reached its maximum nominal output, meaning that the discharge power control loop was functioning as intended and the issue was most probably hardware related. The respective investigation concluded that the cause of this phenomenon is the thermal dissipation on the EPS mounting plate and the mass flow controller, affecting its thermal inlet conditions, therefore reducing the effective mass flow rate through the DXFC mass flow restrictors (see Fig.\u0026nbsp;\u003cspan class=\"InternalRef\"\u003e8\u003c/span\u003e). A possible solution was to open the sixth DXFC valve (redundant valve) to compensate for the lower mass flow rate to the anodes. However, the implementation of this solution required some on orbit testing, resulting in delaying the scheduled maneuver at this stage. In coordination with the mission operation team, it was decided to limit the maximum discharge power command to 420 W for this phase in order not to disturb the mission's maneuver.\u003c/p\u003e\n\u003cp\u003e\u003cem\u003eBeam-out after ignition\u003c/em\u003e\u003c/p\u003e\n\u003cp\u003eAfter VM2 started some beam-out events were reported via telemetry, mostly immediately after thruster ignition and mainly on Thruster Unit 1 (TU1). At this point, it was decided to stop the autonomous maneuver and keep activating the EPS manually using direct scripts which were uploaded to the satellite. To mitigate the beam-outs events some tests were performed at Rafael's space propulsion facility. An identical Flight Model (FM) IHET-300 thruster was fired to accumulate the total operating time and the total accumulated ignitions performed already with TU1 in space, to simulate as close as possible the thruster operation history.\u003c/p\u003e\n\u003cp\u003eThe ground testing at this step did not reveal any issues and the beam-out phenomena was not recreated during the lab test. At this point other tests were performed with the goal of treating the symptoms to facilitate uninterrupted thruster ignition and thruster operation. One possible option was to reduce the intensity of the magnetic field so to reduce discharge resistivity and to facilitate a spontaneous reignition in the case of a quick beam-out, while taking advantage of the fact that the keeper-emitter circuit could be kept on during thruster operation. A test of magnet current sensitivity was performed and thrust, and anode current oscillations were measured. Figure\u0026nbsp;\u003cspan class=\"InternalRef\"\u003e12\u003c/span\u003e presents thruster discharge current oscillations, main frequency and thrust as measured in the lab. The data shows that regardless of the magnetic field current, instabilities are below 15% of the discharge current and with frequencies in the ~\u0026thinsp;10 kHz range. These frequencies are well above the cutoff frequency of the PPU filter. The thrust results show the decrease with magnetic coil current. This decrease can surpass 10% when the magnet current is below 1.5 A. This investigation affirmed that the magnet current can be decreased by 25% of its designed value with little consequences to thruster performance and current oscillations.\u003c/p\u003e\n\u003cp\u003eThanks to the investigation, updated magnet current parameters were uploaded to the satellite. In addition, it was decided to increase the cathode mass flow rate during TU ignition and TU steady state operation. Lastly, the ignition sequence was updated to ignite the thruster in \"Glow mode\", that is with no magnet current during anode circuit breakdown. To do so, and to reduce the risk of the high inrush currents on the discharge current circuit that may cause a PPU shutdown trigger, the four anodes power supply modules were turned on consequently to reduce the total output capacitance at the anode output line during circuit breakdown.\u003c/p\u003e\n\u003cp\u003e\u0026nbsp;\u003c/p\u003e"},{"header":"V. Conclusions","content":"\u003cp\u003eVen\u0026micro;s is a sophisticated mission where the propulsion system was responsible for a variety of complex tasks, from orbit maintenance to orbit transfer (up and down) and thrusters\u0026rsquo; characterization. The technological mission operated the EPS to demonstrate LEO mission enhancement capability and validate the EPS in space.\u003c/p\u003e \u003cp\u003eThe EPS consisted of two throttleable Hall thruster, PPU, Propellant Management Assembly (PMA), and a 9-liter propellant tank carrying 16 kg of Xenon. The system was capable of throttling operation in the 300\u0026ndash;600 W discharge power range.\u003c/p\u003e \u003cp\u003eThe EPS successfully passed its In-Orbit Testing (IOT) phase and proved to operate properly according to specifications and expectations. The subsequent mission phases included orbit maintenance demonstration, 720 km\u0026rarr;410 km orbit transfer, orbit maintenance at high drag environment, and 410\u0026rarr;560 km orbit transfer. Throughout all mission phases the EPS operated within the designated power range, yet most of the time under discharge power of 450 W.\u003c/p\u003e \u003cp\u003eTwo methods were used to compute the remaining propellant mass \u0026ndash; \u0026ldquo;Bookkeeping\u0026rdquo; and \u0026ldquo;PTV\u0026rdquo; methods. Both methods proved successful and corresponding measurement errors were computed for each method.\u003c/p\u003e \u003cp\u003eThruster performance was measured on-orbit, using the \u0026ldquo;Orbit Determination\u0026rdquo; method and compared to laboratory experiments conducted on the ground with identical thrusters, one at its beginning of life, while the other after it past the lifetime test of 130 kN-sec. The measured performance on-orbit was found to be on average lower by up to 5% than the performance measured on the ground. Additional tests were conducted to assess performance variation between BOL and EOL thrusters, as well as a cold (\u0026lt;\u0026thinsp;20 mins) and hot (\u0026gt;\u0026thinsp;120 mins) operation.\u003c/p\u003e \u003cp\u003eLastly, we presented several repeating events in which the propulsion system suffered from sudden beam-outs or ignition difficulties. The methods used to assess a possible solution and implement it were presented.\u003c/p\u003e \u003cp\u003eAlthough designed as a fully redundant EPS, until the end of the technological mission there was no need to switch any redundant component or sub-system. All initial assigned components and branches \u0026ndash; completed the mission successfully.\u003c/p\u003e"},{"header":"Declarations","content":"\u003ch2\u003eFunding Declaration\u003c/h2\u003e\n\u003cp\u003eThe Venus project was supported by the Israeli Space Agency (ISA).\u003c/p\u003e\u003ch2\u003eAuthor Contribution\u003c/h2\u003e\u003cp\u003eAll authors took part in the design, development, qualification of the electric propulsion system.All authors were involved in the operation of the propulsion system on-orbit, along with conceiving solutions to the issues presented in the manuscript.\u003c/p\u003e\u003ch2\u003eAcknowledgements\u003c/h2\u003e \u003cp\u003eThe Authors wish to thank the Israeli Space Agency for its long-term support of the program. Although this paper authors are the limited team of the operation phase only, we would like to thank the grand Ven\u0026micro;s team that developed, manufactured, tested, integrated and operated the EPS and the technological mission, since its inception in 2004 and until now. In addition, special thanks to ASRI team in the Technion, which developed the TMC, for their commitment and accompaniment of operations.\u003c/p\u003e"},{"header":"References","content":"\u003col\u003e\u003cli\u003e\u003cspan\u003ePaul Lascombes M, Mont\u0026egrave;s A, Fiorentino T, Gelu M, Fillastre, Antonio Gurciullo. Lessons learnt from operating the first Cubesat mission equipped with a Hall thruster. Proceedings of the 35th Annual Small Satellite Conference. Logan, UT, USA, 7\u0026ndash;12 Aug 2021. 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Electric Propulsion Conf. (Toulouse, France)\u003c/span\u003e\u003c/li\u003e \u003cli\u003e\u003cspan\u003eFairview Park OH Electric Rocket Propulsion Society) IEPC 2003\u0026thinsp;\u0026ndash;\u0026thinsp;104\u003c/span\u003e\u003c/li\u003e\u003c/ol\u003e"}],"fulltextSource":"","fullText":"","funders":[],"hasAdminPriorityOnWorkflow":false,"hasManuscriptDocX":true,"hasOptedInToPreprint":true,"hasPassedJournalQc":"","hasAnyPriority":false,"hideJournal":false,"highlight":"","institution":"","isAcceptedByJournal":true,"isAuthorSuppliedPdf":false,"isDeskRejected":"","isHiddenFromSearch":false,"isInQc":false,"isInWorkflow":false,"isPdf":false,"isPdfUpToDate":true,"isWithdrawnOrRetracted":false,"journal":{"display":true,"email":"[email protected]","identity":"journal-of-electric-propulsion","isNatureJournal":false,"hasQc":true,"allowDirectSubmit":false,"externalIdentity":"joeprop","sideBox":"Learn more about [Journal of Electric Propulsion](https://www.springer.com/journal/44205)","snPcode":"44205","submissionUrl":"https://submission.nature.com/new-submission/44205/3","title":"Journal of Electric Propulsion","twitterHandle":"","acdcEnabled":true,"dfaEnabled":true,"editorialSystem":"stoa","reportingPortfolio":"Springer Hybrid","inReviewEnabled":true,"inReviewRevisionsEnabled":true},"keywords":"Hall Effect Thruster, Electric Propulsion System, Technological Mission, In Orbit Test","lastPublishedDoi":"10.21203/rs.3.rs-4022336/v1","lastPublishedDoiUrl":"https://doi.org/10.21203/rs.3.rs-4022336/v1","license":{"name":"CC BY 4.0","url":"https://creativecommons.org/licenses/by/4.0/"},"manuscriptAbstract":"\u003cp\u003eVen\u0026micro;s is a satellite launched in 2017, for super-spectral Earth imaging and Electric Propulsion System (EPS) demonstration. In this paper we overview EPS design and operation throughout all five mission phases, from open/closed-loop orbit control, through orbit descent (720\u0026rarr;410 km), orbit maintenance under high drag environment, to orbit raising (410\u0026rarr;560 km). The EPS consisted of two throttleable Hall thrusters, PPU, Propellant Management Assembly (PMA), and a 9 liter propellant tank carrying 16 kg of Xenon. Both thrusters operated in the 300\u0026ndash;550 W power range, generated a combined total impulse of 158.1 kN-sec and consumed all propellant. Two methods are described to compute the remaining propellant mass \u0026ndash; \u0026ldquo;Bookkeeping\u0026rdquo; and \u0026ldquo;PTV\u0026rdquo; methods. The advantages and disadvantages of each method are discussed in light of the Ven\u0026micro;s mission. Thruster performance was measured on-orbit using the \u0026ldquo;Orbit Determination\u0026rdquo; method and compared to laboratory experiments conducted on the ground with identical thrusters. The measured performance on-orbit was found to be on average lower by up to 5% than the performance measured on the ground. Lastly, we present several repeating events in which the propulsion system suffered from sudden beam-outs or ignition difficulties. We present the methods used to construct a solution and implement it.\u003c/p\u003e","manuscriptTitle":"On-Orbit Mission Overview of the Low Power Hall Thruster Propulsion System aboard Venμs Satellite","msid":"","msnumber":"","nonDraftVersions":[{"code":1,"date":"2024-03-14 20:10:51","doi":"10.21203/rs.3.rs-4022336/v1","editorialEvents":[{"type":"communityComments","content":0},{"type":"decision","content":"Revision requested","date":"2024-04-05T19:52:46+00:00","index":"","fulltext":""},{"type":"editorInvitedReview","content":"","date":"2024-03-27T19:58:34+00:00","index":"hide","fulltext":""},{"type":"reviewerAgreed","content":"30b93689-6105-47f3-976b-6421592ddb53_SNPRID","date":"2024-03-17T22:29:30+00:00","index":"hide","fulltext":""},{"type":"reviewersInvited","content":"","date":"2024-03-17T16:39:47+00:00","index":"","fulltext":""},{"type":"checksComplete","content":"","date":"2024-03-12T09:29:44+00:00","index":"","fulltext":""},{"type":"editorAssigned","content":"","date":"2024-03-12T09:29:44+00:00","index":"","fulltext":""},{"type":"submitted","content":"Journal of Electric Propulsion","date":"2024-03-06T22:56:22+00:00","index":"","fulltext":""}],"status":"published","journal":{"display":true,"email":"[email protected]","identity":"journal-of-electric-propulsion","isNatureJournal":false,"hasQc":true,"allowDirectSubmit":false,"externalIdentity":"joeprop","sideBox":"Learn more about [Journal of Electric Propulsion](https://www.springer.com/journal/44205)","snPcode":"44205","submissionUrl":"https://submission.nature.com/new-submission/44205/3","title":"Journal of Electric Propulsion","twitterHandle":"","acdcEnabled":true,"dfaEnabled":true,"editorialSystem":"stoa","reportingPortfolio":"Springer Hybrid","inReviewEnabled":true,"inReviewRevisionsEnabled":true}}],"origin":"","ownerIdentity":"6098cf5d-d376-442d-9d68-8f46ca2a9565","owner":[],"postedDate":"March 14th, 2024","published":true,"recentEditorialEvents":[],"rejectedJournal":[],"revision":"","amendment":"","status":"under-review","subjectAreas":[],"tags":[],"updatedAt":"2024-07-01T16:18:51+00:00","versionOfRecord":[],"versionCreatedAt":"2024-03-14 20:10:51","video":"","vorDoi":"","vorDoiUrl":"","workflowStages":[]},"version":"v1","identity":"rs-4022336","journalConfig":"researchsquare"},"__N_SSP":true},"page":"/article/[identity]/[[...version]]","query":{"redirect":"/article/rs-4022336","identity":"rs-4022336","version":["v1"]},"buildId":"qtupq5eGEP_6zYnWcrvyt","isFallback":false,"isExperimentalCompile":false,"dynamicIds":[84888],"gssp":true,"scriptLoader":[]}