Micropropulsion System Research Articles

Numerical investigation of micropropulsion systems for CubeSats: Gas species and geometrical effects on nozzle performance

In the present investigation, a numerical investigation is conducted to investigate cold-gas micropropulsion devices for CubeSat orbital control. Different micronozzle geometry configurations and gas species are analyzed to assess their influence on the flowfield structure, surface properties, and microthruster performance parameters. Due to the small length scales, the Direct Simulation Monte Carlo (DSMC) method is used to simulate rarefied argon and nitrogen in rectangular and curved micronozzles. The results indicate that nitrogen gas leads to a 23% increase in specific impulse at a relatively insignificant thrust cost. On the other hand, the use of a curved geometry increases the specific impulse and thrust generated by 23% and 35%, respectively. The results indicate that using lighter gases for cold gas microthrusters increases the micropropulsion efficiency. In addition, curved geometries significantly improve the overall performance of such devices, in contrast to rectangular geometries.

Multimodal electrospray thruster for small spacecraft: design and experimental characterization

Electrospray thrusters are a promising electric micropropulsion technology which could be used to meet the propulsion needs of nanosatellites, or for fine attitude control of larger spacecraft. Multimodal propulsion is the integration of two or more propulsion modes into a system which utilizes a common propellant. Indeed, spacecraft mission simulations and models have shown that this type of multimode propulsion capacity is exciting because of the flexibility and adaptability it provides mission designers and planners. A single spacecraft would have potential to execute drastically different mission profiles, and the exact mission could even be determined post-launch. The current paper investigates a micro-propulsion system which combines a droplet and ion mode electrospray emitter into a unified multimodal system (using an ionic liquid as the common propellant for both systems). The high relative thrust droplet mode emitter was fabricated from P3 borosilicate glass while the high efficiency ion mode emitter, Carbon Xerogel dense porous substrate, was fabricated in-house. To characterize the multimodal thruster, a full beam and time-of-flight (ToF) experimental setup were developed at the RMC Advanced Propulsion and Plasma Exploration Laboratory (RAPPEL) and experiments were conducted using a custom vacuum chamber. The ion mode emitter, with a beam comprised purely of ions had an onset voltage around 1400 V with an estimated thrust performance of 0.14 μN\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\mu N$$\\end{document} and specific impulse of 4040 s. For droplet mode, with a mixed beam comprised of around 17%\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\%$$\\end{document} droplets and 83%\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\%$$\\end{document} ions, an onset voltage of 1375 V with an estimated performance of thrust at 14 μN\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\mu N$$\\end{document} and specific impulse of 140 s were measured. The prototype thruster demonstrates how various electrospray emitters could be combined into a multimodal system to provide flexibility and adaptability in providing effective thrust for small satellites.

Propulsion Technologies for CubeSats: Review

This paper explores the wide-ranging topography of micro-propulsion systems that have been flown in different small satellite missions. CubeSats, known for their compact size and affordability, have gained popularity in the realm of space exploration. However, their limited propulsion capabilities have often been a constraint in achieving certain mission objectives. In response to this challenge, space propulsion experts have developed a wide spectrum of miniaturized propulsion systems tailored to CubeSats, each offering distinct advantages. This literature review provides a comprehensive analysis of these micro-propulsion systems, categorizing them into distinct families based on their primary energy sources. The review provides informative graphs illustrating propulsion performance metrics, serving as beneficial resources for mission planners and satellite designers when selecting the most suitable propulsion system for a particular mission requirement.

Flow pattern and pressure drop analysis of two-phase segmented flow driven by a thermomagnetic pump

The two-phase segmented flow is a flow pattern in which two immiscible fluids alternate in a flow channel with distinct segments. In this study, a thermomagnetic pump operated by heat and an external magnetic field was used to drive the flow and the temperature-sensitive magnetic fluid (TSMF) was chosen as the continuous-phase fluid to generate a two-phase segmented flow. The dispersed-phase fluid was selected as air or deionized water driven by a syringe pump. The two-phase segmented flow experiments were conducted to investigate the relation between segment length and the dispersed-continuous phase flow rate ratio at the air flow rate from 0.90 ml/h to 4.49 ml/h for air-TSMF experiment and water flow rate from 0.36 ml/h to 0.72 ml/h for water-TSMF experiment. The ratio was proportional to the bubble/droplet length and inversely proportional to the slug length. The air/water length per channel width can be up to 3.28 and 1.79, respectively, at the highest air/water flow rate in the current experiment. Moreover, the pressure drop was measured to increase with the dispersed-continuous phase flow rate ratio. The maximum pressure drop was observed as 428.2 Pa in the water-TSMF experiment with the highest water flow rate and the maximum pressure drop was 314.7 Pa for the air-TSMF experiment. These findings provide valuable insights for controlling the segment length in the two-phase segmented flow driven by a thermomagnetic pump and it could be used as precise control of liquid/gas fuel delivery in a micro-propulsion system in small satellites.

A Review on Micro-Combustion Flame Dynamics and Micro-Propulsion Systems

This work presents a state-of-the-art review of micro-combustion flame dynamics and micro propulsion systems. In the initial section, we focus in on the different challenges of micro-combustion, investigating the typical length and time scales involved in micro-combustion and some critical phenomena such as flammability limits and the quenching diameter.We present an extensive collection of studies on the principal types of micro-flame dynamics, including flashback, blow-off, steady versus non-steady flames, mild combustion, stable flames, flames with repetitive extinction, and ignition and pulsatory flame burst. In the final part of this review, we focus on micropropulsion systems, their performance metrics, conventional manufacturing methods, and the advancements in Micro-Electro-Mechanical Systems manufacturing.

Shock induced variable density flows in the vacuum microchannel: I. medium laser fluence

Laser-matter interactions with metal target cause plasma explosion and shock accelerated variable density flow instabilities in the Semiconfined Configuration (SCC). Their study gives deeper insight into the flow instabilities present in all microchannel devices. Blast wave motion along the SCC microchanel causes the Kelvin–Helmholtz (KH) instability and formation of vortex filaments for the critical Reynolds number. Appearing in all shear layers—it affects the fluid transport efficiency. Shear layer acceleration causes a Raleigh-Taylor instability (RTI). Oriented bubble growth by discrete merging indicates anisotropic RTI mixing. Similar RTI flame instability appears in the conversion of chemical energy into electricity affecting microcombustion efficiency. Another case of anisotropic RTI is the flow boiling for cooling of chips and microelectronic devices. The RTI boiling which appears for the critical heat flux is based on rising surface vapor columns (oriented bubble growth) with liquid counterflow (spike prominences) for the critical wavelength at density interface. The RT bubble merging graph trees determine turbulent mixing which affects the heat transfer rates. Bottom-wall turbulent flow in the SCC microchannel causes streaks of the low momentum fluid and formation of hairpin vortex packets with lattice organization. This makes possible to quantify parameters responsible for the evolution of hairpin vortex packets in the microchannel devices. Appearing from the low to the high Reynolds numbers they affect the transport properties, control of the fluid motion, enhancement of mixing, or the separation of fluids. Fluid particle ejecta—thin supersonic jets - evolve into long needle-like jets which start spiraling, helical pairing and swirling in the field of thermal gradients. Such instabilities appear in the microcombustion flame instability and in the space micropropulsion systems. Oscillating and spiral flames appear in the presence of thermal gradient in the microchannel, due to the combined effects of thermal gradient fields and the mixture flow rates.

Validation of DSMC mass flow modeling for transsonic gas flows in micro-propulsion systems

Introduction: In the present work, an inflow model for the DSMC method is presented and validated. The approach is based on inflow mass flow rate and temperature and is particularly suitable for arbitrary nozzle flow cases with higher density, subsonic inflow conditions.Methods: The validation is performed on a nozzle test case and the results are compared with experimental and numerical results based on DSMC and Navier-Stokes methods. Calculation of inflow and outflow boundary conditions on an analytical and numerical basis is presented.Results: Results for axial and radial density, temperature, and pressure are in good agreement and reasonable relationships are obtained.Discussion: Since only the inflow mass flow rate and temperature and the vacuum background pressure need to be known to apply the model, the calculation of the inflow velocity from analytical theory can be omitted, potentially eliminating possible sources of error resulting from theorybased calculations.

Development of thin-walled polymer glass beads reinforced 3D printed solid rocket motors

A polyamide 12(PA12) reinforced with glass beads (GBs) solid rocket motor (SRM) produced by 3D Printing was designed to achieve light weight of the motors. The pressure resistance and heat resistance of four motors with the thicknesses of 6, 2.5, 2 and 1.8 mm were recorded using a high-speed camera and an infrared thermometer. Porosity analysis was performed on four motors before the ignition experiment and the motor with the thicknesses of 2 mm after the ignition experiment. The results of the study show that the shell’s thickness of the motor can be as thin as 2 mm under the ignition condition of 12 Mpa and 2773 K. The overall and local porosity of the motor became larger after ablation. This work provides a new direction and practical references for the application of micropropulsion systems for satellites.

A review on solid propellant micro-thruster array based on MEMS technology

With the development of micro-spacecraft technology, micro-nano satellites have the advantages of small size, low power consumption, short development cycle, formation networking, etc., and can complete many complex space tasks at a lower cost. Micro-nano satellites require a micropropulsion system with the capability of performing precise total impulse and thrust to execute maneuvers, such as attitude control, orbital transfer, and gravitation compensation. In contrast to other micropropulsion systems, solid propellant microthrusters (SPM) arrays based on micro-electromechanical system (MEMS) technology possess a simple structure and quick response, which is a potential micropropulsion system. In recent years, many research groups have done a lot of research on SPM arrays. In this paper, the latest progress of SPM arrays is summarized from the aspects of structure design, propellant selection, bonding technology, ignition unit type and micro-thrust test, and some suggestions for the future development direction are given.

NO emission and enhanced thermal performances studies on Counter-flow Double-channel Hydrogen/Ammonia-fuelled microcombustors with Oval-shaped internal threads

Micro-combustion systems play a critical role in powering microreactors, micropropulsion and other micro-energy conversion systems. In this work, we have numerically investigated a double-channel counter-flow micro-combustor fuelled by premixed ammonia/hydrogen/oxygen parametric to investigate both NO emissions and thermal performance. The model is used to examine the effects of 1) the inlet velocity, vin, and 2) the inlet equivalence ratio, ϕ. The enhanced thermal performance is quantified by the increased mean wall temperature and the increased wall temperature uniformity. We find that thermal performance increases with increasing vin. The maximum level of NO is present atvin = 2 m/s. Meanwhile, the mean wall temperature benefits from ϕ = 1, and further increases in equivalence ratio can lessen the uniformity of the wall temperature. Furthermore, more fuel-rich ammonia combustion can lead to a lower NO emission and improve the thermal performance. To enhance ammonia combustion, further investigations are conducted by blending with different molar fractions of H2(XH2). It is found that mixing ammonia with more H2 can stabilize micro-combustion, and increase the temperature in the combustion field. Additionally, it makes the emission worse. The maximum mean wall temperature occurs atXH2 = 0.25, While the NO emission peaks atXH2 = 0.3. Moreover, the OH mole fraction can affect the formation of NO, positively.

DSMC Simulation of the Effect of Needle Valve Opening Ratio on the Rarefied Gas Flows inside a Micronozzle with a Large Length-to-Diameter Ratio

The cold gas micro-propulsion system can provide low noise and ultra-high accuracy thrust for satellite platforms for space gravitational wave detection, high-precision earth gravity field measurement. In this study, the effect of different needle valve opening ratios on the rarefied flow characteristics of a micro-nozzle in a cold gas micro-propulsion system was investigated based on DSMC method. The special feature of the currently studied micro-nozzle is that it has a section of micro-channel with a large length–diameter ratio up to 4.5. Due to the extremely small needle valve displacement of the nozzle (minimum needle valve displacement up to 1.7 μm), a finely structured mesh was used. The molecular particle and macro flow characteristics inside the micro-nozzle were calculated for the conditions of a needle valve opening ratio from 5% to 98%. The throttling effect of the throat has a significant effect on the rarefied flow in the micro-nozzle; especially under the tiny opening, this effect is more significant. The spatial distribution of continuous flow, transition flow, and free molecular flow in the micro-nozzle varies at different needle valve opening ratios. As the needle valve opening ratio increases, the continuous flow will gradually fill the microfluidic region.

MODELING of Rarefied Gas Flows Inside a Micro-Nozzle Based on the DSMC Method Coupled with a Modified Gas–Surface Interaction Model

In this study, we first considered the influence of micro-nozzle wall roughness structure on molecular collision and reflection behavior and established a modified CLL model. The DSMC method was used to simulate and analyze the flow of the micro-nozzle in the cold gas micro-propulsion system, and the deviation of simulation results before and after the improvement of CLL model were compared. Then, the rarefied flow characteristics under a small needle valve opening (less than 1%) were focused on the research, and the particle position, molecular number density, and spatial distribution of internal energy in the micro-nozzle were calculated. The spatial distributions of the flow mechanism in the micro-nozzle under different needle valve openings were compared and analyzed. It was found that when the needle valve opening is lower than 1%, the slip flow and transition flow regions move significantly upstream of the nozzle, the free molecular flow distribution region expands significantly, and the relationship between thrust force and needle valve opening is obviously different from that of medium and large needle valve openings. The effect of nitrogen temperature on the rarefied flow and thrust force is also discussed in this research. The numerical results showed that as gas temperature increases, the molecular internal energy, momentum, and molecular number density near the nozzle exit are enhanced. The thrust at small needle valve openings was significantly affected by the temperature of the working mass. The results of this study will provide key data for the design and development of cold gas micro-thrusters.

Direct thrust test and asymmetric performance of porous ionic liquid electrospray thruster

In order to meet the demand of CubeSats for low power and high-performance micro-propulsion system, a porous ionic liquid electrospray thruster prototype is developed in this study. 10 × 10 conical emitter arrays are fabricated on an area of 3.24 cm2 by computer numerical control machining technology. The propellant is 1-ethyl-3-methylimidazolium tetrafluoroborate. The overall dimension of the assembled prototype is 3 cm × 3 cm × 1 cm, with a total weight of about 15 g (with propellant). The performance of this prototype is tested under vacuum. The results show that it can work in the voltage range of ±2.0 kV to ±3.0 kV, and the maximum emission current and input power are about 355 μA and 1.12 W. Time of Flight (TOF) mass spectrometry results show that cationic monomers and dimers dominate the beam in positive mode, while a higher proportion of higher-order solvated ion clusters in negative mode. The maximum specific impulse is 2992 s in positive mode and 849 s in negative mode. The thrust is measured in two methods: one is calculated by TOF results and the other is directly measured by high-precision torsional thrust stand. The thrust (T) obtained by these two methods conforms to a certain scaling law with respect to the emission current (Iem) and the applied voltage (Vapp), following the scale of T ∼ IemVapp0.5, and the thrust range is from 2.1 μN to 42.6 μN. Many thruster performance parameters are significantly different in positive and negative modes. We speculate that due to the higher solvation energy of the anion, more solvated ion clusters are formed rather than pure ions under the same electric field. It may help to improve thruster performance if porous materials with smaller pore sizes are used as reservoirs. Although there are still many problems, most of the performance parameters of ILET-3 are good, which can theoretically meet the requirements of CubeSats for micro-propulsion system.

Additively manufactured vaporizing liquid microthruster with micro pin fins for enhanced heat transfer

Vaporizing liquid microthruster is a relatively new concept for evaporation based propulsive system. The research interest in vaporizing liquid microthruster mainly comes from its simplicity and the use of green propellant: water. This paper presents the study of a highly compact additively manufactured vaporizing liquid microthruster of 38 × 22 × 8 mm in dimension. The heat transfer enhancement effect of micro pin fins for vaporizing process has been investigated using modeling and computational fluid dynamics simulation, followed by experiments to validate the results. It has been shown that all micro pin fins (square, diamond, circular) with different array arrangements (staggered and inline) can enhance heat transfer. Square shaped pin fin exhibited the best heat enhancement performance with an increase in dimensionless total thermal resistance of 14.65% (as compared to blank vaporizing liquid microthruster) due to its large contacting area, while the circular shape was least effective. Meanwhile, the inline arrangement was found to induce a significant fluctuation in transverse flow velocity between the columns, which promotes heat convection. The vaporizing liquid microthruster prototype was produced using a specialty resin of good heat resistance up to 350 °C as confirmed by the thermogravimetric analysis. Thrust performance of the prototype was measured using a torsional thrust stand in a vacuum chamber. The vaporizing liquid microthruster with square micro pin fins and in inline configuration has produced the highest thrust of 740.2 μN at 2 μl/s flow rate, giving a specific impulse of 37.7s. The developed prototype could be implemented as micropropulsion system for nanosatellites, such as CubeSat and PocketQube.

Spectroscopic plasma plume study of a non-volatile liquid-fed pulsed plasma thruster

The development of small satellites weighing as little as 1 kg has accelerated along with the miniaturization of technology. However, there is still a lack of efficient micropropulsion systems at these scales. A possible solution is the pulsed plasma thruster, which has an inherently simple structure that is conducive towards miniaturization. As these thrusters were primarily developed in an era of larger satellites, there are systematic considerations that need to be taken into account. One of these is the propellant, where the conventional solid polytetrafluoroethylene (PTFE) propellant can be replaced with a non-volatile liquid perfluoropolyether (PFPE), enabling liquid-feeding approaches for propellant feeding. The physics behind the operation of a non-volatile liquid-fed pulsed plasma thruster still remains relatively unknown compared with the well-studied solid-fed archetype. We focus here on the use of emission spectroscopy to better understand the spatial distribution and temporal behavior of plasma components within a pulsed plasma thruster using non-volatile liquid PFPE. This is directly compared with solid PTFE to give us an understanding of the similarities and differences between the two propellant types. The overall results suggest that liquid PFPE may be a useful propellant in scenarios with a high thrust-to-power requirement and lower emphasis on electromagnetic acceleration. One such use-case is in coaxial or capillary discharge miniaturized pulsed plasma thrusters, where a significant portion of thrust is obtained from thermodynamic gas expansion rather than electromagnetic acceleration.

Method for On-Orbit Thrust Estimation for Microthrusters Based on Nanosatellites

AbstractThe development of nanosatellites is inseparable from micropropulsion systems, and a micropropulsion system cannot be separated from the evaluation of its micropropulsion performance. Altho...

The View of Micropropulsion Technology for China’s Advanced Small Platforms in Deep Space

In this paper, micropropulsion systems are analyzed in conjunction with the various mission requirements of China’s deep space exploration. As a great challenge facing the world, deep space exploration can be enabled only in a few countries with a success rate of around 50%. With the advancement of spacecraft and scientific instruments, it is now feasible to build small and low-cost spacecraft for a variety of deep space missions. As spacecraft become smaller, there is a need for proper micropropulsion systems. Examples of propulsion system selections for deep space exploration are discussed with a focus on products developed by Beijing Institute of Control Engineering (BICE). The requirements for propulsion systems are different in lunar/interplanetary exploration and gravitational wave detection. Chemical propulsion is selected for fast orbit transfer and electric propulsion for increasing scientific payloads. Cold gas propulsion and microelectric propulsion are good choices for space-based gravitational wave detection due to the capability of variable thrust output at the micro-Newton level. The paper also introduces the sub-1-U micropropulsion modules developed by BICE with satisfactory performance in flight tests, which are promising propulsion systems for small deep space platforms. A small probe with an electric sail propulsion system has been proposed for the future solar system boundary exploration of China. The electric sail serves as not only a propellant-free thruster but also a detector probing the properties of the space medium.

Performance measurement and evaluation of an ionic liquid electrospray thruster

As a novel micro-propulsion system for small satellites (from micro to nano), the ionic liquid electrospray propulsion system is a promising candidate. However, performance measurement and evaluation of the Ionic Liquid Electrospray Thruster (ILET) is one of the most challenging issues for practical application, due to the difficulties in the development of a prototype and direct measurements of micro-thrust and small flow rate. To address this issue, a Modular Ionic Liquid Electrospray Thruster (MILET) prototype is constructed, and a diagnostic system for thrust and mass flow rate is specially developed based on an analytical balance method. With the ionic liquid 1-ethyl-3-methylimidazolium tetrafluoroborate as the propellant, a series of experiments is carried out on the MILET prototype under a wide operating condition through changing the applied voltage to control the thrust. Under different applied voltages, the thrust and the mass flow rate of the propellant are directly measured. The propulsive performance parameters of the thruster, such as thrust, specific impulse, thrust-to-power ratio, thruster efficiency, etc., are comprehensively analyzed. Then, a performance comparison is made between the MILET and other representative ILETs. With a relatively low applied voltage ranging from 1550 V to 2000 V, the MILET achieves a quasi-constant specific impulse of 1263 s with the averaged thrust-to-power ratio of 65.2 μN/W and thruster efficiency of 40.7%. The performance of ILET is also compared with other typical electric propulsions. The results demonstrate that the ILET exhibits an excellent ability of minimalization with high specific impulse and thruster efficiency, which guarantees a great superiority in micro propulsions. Finally, the ways to further improve the performance of ILET are discussed, which further confirms the potential prospect of ILET. The present result helps to advance the development and application of ILET.

Molecular dynamics simulation of ionic liquid electrospray: Microscopic presentation of the effects of mixed ionic liquids

Micro-propulsion devices based on electrospray have advantages of high specific impulse, high precision and compact structure, and are able to meet the requirements of gravitational wave detection and other drag-free spacecrafts for micro-propulsion system. For electrospray thrusters, ionic liquid is the preferred propellant because of its low volatility, high electrical conductivity and thermal stability. Most of the studies presented today focus on single ionic liquid, while the properties of single ionic liquid are certain, which leads to a limitation of electrospray characteristics and hence propulsion performance. Mixing different kinds of ionic liquids broadens the properties of ionic liquids, and thus adjusts the characteristics of electrospray to meet the requirements of diversified space missions. In this paper, the effects of mixing two typical ionic liquids, 1-ethyltrimethylimidazole dicyanamide (EMIM-N(CN)2) and 1‑butyl‑trimethylimidazolium hexafluorophosphate (BMIM-PF6), on propellant properties and electrospray characteristics are investigated by molecular dynamics simulation and validated by comparing with experimental data. The result shows that the mixed ionic liquid changes the emission mode of electrospray thruster. Compared with BMIM-PF6, the specific impulse and thrust for the 50:50 (by mass) mixture increases by more than half of the difference between the two before mixing. In addition, the analysis shows that the range of extrusion voltage is widened by adjusting the mixing ratios to meet the mission requirements for certain specific impulse and thrust.

Ground performance tests and evaluation of RF ion microthrusters for Taiji-1 satellite

The “Taiji-1” satellite is a test satellite for the verification of those key technologies involved in the “Taiji Program in Space”, China’s space gravitational wave detection project; and the spacecraft drag-free control technology is one of its key technologies used to improve the microgravity level of spacecraft. Thus, a demand for a high-precision and continuously adjustable micronewton-level thrust has been proposed on the spacecraft micro-propulsion system. In allusion to such a task, a set of micronewton-level RF ion propulsion system was designed based on the principle of self-sustained discharge of RF plasma so as to conduct studies on the parameter optimization and engineering of RF ion thruster, and an engineering prototype of the micronewton-level continuously adjustable RF ion thruster was successfully developed to meet the design index requirements. The operating parameters have been solidified through further studies on the extreme operating conditions of the engineering prototype, and a series of ground simulation tests of space environment were successfully passed. The ground test and calibration experiment results of the micro-propulsion engineering prototype show that the engineering prototype has fully met the requirements of the Taiji-1 mission, thus having laid a solid foundation for the successful space verification of the key technologies for the “Taiji-1” satellite.