Micro-nano Satellites Research Articles

Design and On-Orbit Performance of Ku-Band Phased-Array Synthetic-Aperture Radar Payload System

The current emphasis in the advancement of space-based synthetic-aperture radar (SAR) is on lightweight payloads under 100 kg with resolutions surpassing 1 m. This focus is directed toward meeting the launch criteria for multiple satellites on a single rocket and cutting costs. This article discusses the creation and progress of a Ku-band SAR payload for the Taijing-4(03) satellite, launched on 23 January 2024 and accompanied by four other satellites. The SAR payload design was customized to meet the demands of a micro-nano satellite platform, resulting in a lightweight, flat design weighing less than 80 kg, seamlessly integrated with the plate-shaped satellite platform. The article also introduces a beam optimization strategy for the phased array SAR antenna, significantly boosting the SAR system’s performance. The SAR payload provides various operating modes like slide-spot, strip, Scan 1, Scan 2, and others, with a maximum achievable resolution exceeding 1 m. Extensive in-orbit testing of the payload produced numerous high-quality SAR images with potential uses in emergency disaster mitigation, safeguarding ecosystems, monitoring forests, managing crops, tracking sea ice, and more.

A high-precision and near-zero residue laser-controlled gel propellant

Micro-nano satellites, with their advantages of small size, low power consumption, short development cycles, and capability for formation flying and networking, play a significant role in scientific research, national defense, and commercial applications. However, currently developed micro-nano satellites globally possess very limited maneuverability and lack effective attitude and orbit propulsion control. Solid propellants can offer some advantages such as high energy density and low cost, and they are extensively used in micro-and nano satellite propulsion systems. However, most of these propellants exhibit poor controllability and produce massive amounts of residual by-products. Laser-controlled gel propellants have been confirmed to be the most effective solution to address these issues in this study. The present study exhibits the preparation of ADN-based laser-controlled gel propellants using the freeze–thaw method and their controllable combustion characteristics under laser irradiation. The results show that the ignition delay time of gel propellants decreases from 93.0 ms to 34.6 ms, the extinction delay time is stable at around 2 ms, and the average burning rate increases from 0.6 mm/s to 1.5 mm/s under continuous laser irradiation with the power density ranging from 0.6 to 1.3 W/mm2. The average utilization rate of gel propellant is as high as 99.2 %. Upon testing, the residual solid matter was identified as unreacted propellant that had not been irradiated by the laser. Laser-controlled gel propellants represent a new type of energetic material that can be modulated to control its ignition, extinction, and burning rate. These capabilities overcome the challenges associated with the inability of traditional chemical propellants to extinguish and restart, thereby fulfilling the propulsion system requirements of micro-nano satellites.

Axial emission characteristics of an ionic liquid electrospray thruster with a circular emitter

Ionic Liquid Electrospray Thrusters (ILETs) are well suited for micro-nano satellite applications due to their small size, low power consumption, and high specific impulse. However, the limited thrust of a single-emitter ILET restricts its use in space missions. To optimize the performance of ILETs and make them suitable for a wider range of space missions, we designed a Circular-emitter ILET (CILET) to convert a one-dimensional (point) emission into a two-dimensional (line) emission. The CILET can self-organize multiple Taylor cones simultaneously. The cones were photographed and the axial emission currents were measured under different voltage and pressure difference conditions with a CILET experimental system. The emission can be divided into two stable states and one unstable state based on the flow and current characteristics. The current in Stable state I increases non-linearly with the voltage, while that in Stable state II is nearly linear with respect to the voltage. The number of cones increases with the voltage in stable states, while the cones become short and crowded under high-voltage conditions. The variation law of the number of cones can be explained with the self-organization theory. The variation in the current exhibits a good correlation with the number of cones. This study demonstrates the feasibility of circular emitters and experimentally indicates that the emission current is improved by approximately two orders of magnitude compared to that of a single capillary.

Implementation and Verification of a Micro-Jet-Vane System of a Solid Rocket Motor for a Micro-Nano Satellite

To achieve rapid vector maneuvering of a space micro-nano satellite, a micro-sized solid rocket motor was utilized as its propulsion system, and a micro-jet-vane-thrust-vector control system was devised. Computational fluid dynamics (CFD) numerical simulations were conducted on the designed micro-vane structure at various deflection angles to ascertain the lateral force and flow field characteristics. The motor’s combustion temperature is 1380 K. Therefore, materials such as 45 steel, alumina ceramics, and tungsten–molybdenum alloy were chosen for the jet vanes to carry out ground-based-motor-jet-ablation experiments and measure the ablation amount. Concurrently, experimental data, including lateral force, were gathered. The tests demonstrated that despite 45 steel having a higher melting point than the combustion temperature significant ablation still occurred. Alumina ceramics exhibited defects and experienced ablation and fragmentation post-test. In contrast, tungsten–molybdenum alloy, being a refractory metal, showed minimal ablation after testing, making it an ideal material for micro-jet vanes. At a 20° deflection of the jet vanes, the lateral force calculated via numerical simulation was 3.76 N, whereas the lateral force obtained from the test was approximately 3.8 N, resulting in an error within 1% and validating the numerical simulation’s validity and accuracy. The jet vanes can generate a maximum steering angle of 8°, thus ensuring the micro-nano satellite’s swift vector maneuvering at large angles.

Design of a focusing system for micro-nano satellite remote sensing camera based on thermal control technology

Remote sensing cameras mounted on micro-nano satellite mare susceptible to optical system defocusing during launch and operation due to environmental influences. Traditional focusing system encounter issues such as large volume and mass, making them unsuitable for the design requirements of micro-nano satellite. Proposes a thermal control-based focusing system for compact systems with maintained imaging quality. The design is developed according to the characteristics of a compact coaxial four-mirror remote sensing camera optical system and the overall structure of the system. The focusing system utilizes structural thermal expansion to alter the distance between the primary and secondary mirrors, achieving the purpose of adjusting the system’s focal length. Finite element simulations were conducted to analyze the modal and thermal deformations of the overall structure, validating the rationality of the system design. Experimental verification of the thermal control focusing system demonstrated its adjustment accuracy and range, with clear imaging obtained after system focusing. The results indicate that when the temperature control accuracy reaches ±0.5 °C, the focal length adjustment accuracy can reach 0.0052 mm. This article proposed thermal control focusing system can adjust the camera’s focal length in orbit, exhibiting advantages such as a small volume, high stiffness, and high precision.

Simulation of liquid cone formation on the tip apex of indium field emission electric propulsion thrusters

Field emission electric propulsion (FEEP) thrusters possess excellent characteristics, such as high specific impulse, low power requirements, compact size and precise pointing capabilities, making them ideal propulsion devices for micro-nano satellites. However, the detection of certain aspects, such as the evolution process of the liquid cone and the physical quantities at the cone apex, proves challenging due to the minute size of the needle tip and the vacuum environment in which they operate. Consequently, this paper introduces a computational fluid dynamics (CFD) model to gain insight into the formation process of the liquid cone on the tip apex of indium FEEP. The CFD model is based on electrohydrodynamic (EHD) equations and the volume of fluid (VOF) method. The entire cone formation process can be divided into three stages, and the time-dependent characteristics of the physical quantities at the cone apex are investigated. The influences of film thickness, apex radius size and applied voltage are compared. The results indicate a gradual increase in the values of electrostatic stress and surface tension stress at the cone apex over an initial period, followed by a rapid escalation within a short duration. Apex configurations featuring a small radius, thick film and high voltage exhibit a propensity for liquid cone formation, and the cone growth time decreases as the film thickness increases. Moreover, some unstable behavior is observed during the cone formation process.

Design and testing of a Mini-RF plasma thruster with permanent magnets

Compared with traditional electric propulsion, RF plasma thruster have attracted much attention due to their characteristics of no electrodes, long life, and high ionization rate. In recent years, the development of micro-nano satellites has put forward requirements for the miniaturization, modularization, and integration of satellite thrusters, and the miniaturization of RF plasma thrusters has a broad prospect. In this paper, a mini-RF plasma thruster with a discharge chamber inner diameter of 10 mm is designed, and a magnetic field is generated around it by two annular samarium cobalt permanent magnets. The discharge state and plasma parameters of the mini-RF thruster are tested using optical emission spectrometry (OES) and target thrust stand. In the experiment, we changed the flow rate of argon gas, the presence or absence of a magnetic field, the power of the RF source and analyzed the effects of these factors on the mode transition, plasma density, electron temperature, and thrust. According to the experimental results, it is found that the magnetic field (maximum strength 0.14T) helps to increase the plasma density at low power but delay the jump power threshold of the CCP-ICP.

Preliminary design of lunar high-resolution remote sensing micro-nano satellite

The development and utilization of lunar resources and the construction of lunar scientific research stations are important parts of China’s future lunar exploration plan. Aiming at the problems of China’s micro-nano satellites with low radiation resistance, lack of autonomous Earth-Moon transfer capability, and limited optical system volume and quality, a preliminary design scheme of a lunar remote sensing micro-nano satellite with autonomous Earth-Moon transfer capability is proposed. Using the analytical orbit average technology and transcription and collocation method to calculate the satellite’s Earth-Moon transfer orbit, the satellite can use the Hall thruster with large specific impulse and low thrust to realize the autonomous Earth-Moon transfer. The passive + active method is used to optimize the radiation resistance of the system, and the total radiation dose of some single machines can reach more than 25krad. A compact high-resolution multispectral camera is designed, and imaging simulation and verification are carried out. In the end, the satellite can achieve high-resolution imaging of the moon with a resolution of 0.2 m at an orbital height of 100 km to achieve the purpose of fine exploration of information such as lunar surface topography and geological structure.

Microwave ignition characteristics and distinction of typical nanothermites under different electromagnetic radiation

As a commonly used agent in micro-nano satellites and other small pyrotechnics, it is crucial to understand the response of nanothermite to microwave due to the growing interest in space and the increasingly complex electromagnetic environment. Microwaves can penetrate typical energetic materials and interact with them spatially and holistically. There are two ways of microwave radiation to the samples: microwave drill (contact) and single mode resonator (non-contact). To date, all previous studies were carried out in a single environment, and only for a single thermite system. In this work, the ignition performance of three representative nanothermite systems (Al/CuO, Al/MnO2, Al/Fe3O4) under two microwave ignition devices (custom-made microwave drill and single mode resonator) was first systematically studied, which contain both magnetic and non-magnetic, as well as high and low reactivity. The results showed that the relationship between ignition delay time and power basically accorded with exponential function. The ignition delay time decreased rapidly with the increase of power and finally stabilized within a fixed range. Compare to the metal oxide, pure Al was insensitive to microwave. However, the mechanism of the two ignition processes was different. The ignition sequence of the nanothermites was the same as metal oxides, indicating that the metal oxide was the decisive factor and driver of nanothermite ignition under microwave drill. For the ignition in resonator, the nanothermite behaved like bulk metal, reflecting most of the electromagnetic energy away. The ignition sequence was consistent with the aluminum content in the composites, implying nAl was the determining factor. To further analyze the ignition mechanism, the hot wire and laser sensitivity of nanothermites were also compared.

Random Error Analysis of Launch and Injection Positions for Distributed Multi Micro-nano Satellite System

Accurate satellite orbit calculation has always been a focus in the aerospace field. Especially, the error distribution of the orbit involving various internal and external factors in the satellite launch process has an important influence on the effectiveness of the satellite cluster. To qualitatively and quantitatively analyze the cooperative efficiency of multi-satellite clusters involving various random errors, this paper analyzes the random error sources in the launch process of micro-nano satellites and studies the distribution law of these satellite orbit position errors. Based on the parameter estimation and the central limit theorem of probability theory, the expression of the error function for the satellite’s orbit position is also constructed. In addition, the probability distribution model of the satellites in a certain range is designed, and the relationship between the number of satellites and the overall errors is proposed. Finally, the result is verified by the simulation experiment, which provides the rationale behind the position and altitude adjustment and constellation optimization of micro-nano satellites.

Hardware Acceleration of Satellite Remote Sensing Image Object Detection Based on Channel Pruning

Real-time detection of satellite remote sensing images is one of the key technologies in the field of remote sensing, which requires not only high-efficiency algorithms, but also low-power and high-performance hardware deployment platforms. At present, the image processing hardware acceleration platform mainly uses an image processing unit (GPU), but the GPU has the problem of large power consumption, and it is difficult to apply to micro-nano satellites and other devices with limited volume, weight, computing power, and power consumption. At the same time, the deep learning algorithm model has the problem of too many parameters, and it is difficult to directly deploy it on embedded devices. In order to solve the above problems, we propose a YOLOv4-MobileNetv3 field programmable gate array (FPGA) deployment scheme based on channel layer pruning. Experiments show that the acceleration strategy proposed by us can reduce the number of model parameters by 91.11%, and on the aerial remote sensing dataset DIOR, the average accuracy of the design scheme in this paper reaches 82.61%, the FPS reaches 48.14, and the average power consumption is 7.2 W, which is 317.88% FPS higher than the CPU and reduces the power consumption by 81.91%. Compared to the GPU, it reduces power consumption by 91.85% and improves FPS by 8.50%. Compared with CPUs and GPUs, our proposed lightweight algorithm model is more energy-efficient and more real-time, and is suitable for application in spaceborne remote sensing image processing systems.

A random error estimation method for satellites being launched into orbit based on ellipsoid theory

Accurate and fast random error estimation is essential to improve satellite control. In this paper, the random error of the satellite being launched into orbit is explored. According to the uncertainty matrix of six satellite orbit parameters associated with the project, the uncertainty of the orbit position in the three axes of the geocentric inertial frame of reference is calculated via the propagation rate of the covariance matrix. Then, the random error probability density of the satellite position is constructed for normal distribution. The surface of the error with the same probability density is approximated as an ellipsoid, the size is deduced along with the ellipsoid's shape and direction, and the satellite's spatial position probability is obtained. Monte Carlo method and STK software were employed to simulate low-earth orbit micro-nano satellite launching into the same orbital position. It is confirmed that the results obtained via the theoretical model are in line with practice. Therefore, the ellipsoid theory proposed in this paper can be used to describe the satellite random position error.

Thermal control system design and analysis for a micro-nano satellite stays on target satellite

Aiming at the future satellites on-orbit service requirements, this paper researched the thermal control system for a micro-nano satellite. The micro-nano satellite can be attached on target satellite. The extreme temperature conditions of the micro-nano satellite has been determined by analyzing the complex space heat environment and power consumption, the thermal control system of micro-nano satellite has been designed, the finite element model (FEM) of the satellite is established, and the temperature fields about the extreme temperature conditions are calculated. The simulation results proves the feasibility of the thermal control system.

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.

Agile Micro-nano Satellite Attitude Estimation with Lost Measurements

Aiming at the problem of lost measurement in attitude determination process of the agile micro-nano satellite, a novel algorithm is proposed employing the Bernoulli process to describe the loss process. Firstly,the model of gyro and star sensors as well as the satellite attitude estimation system simplified with gyro is given. Since the missing observations, the modified EKF algorithm is derived by combining a Bernoulli random variable and a real-time measurement noise correction algorithm. Then the convergence of the improved algorihtm is analyzed, and the bottom boundary condition for the observation arrival rate is presented. Finally, the improved method is compared with the conventional one through digital simulation, and the experimental results show that the preciseness of the privious one is one or two orders improved compared with the latter.

Multi-axis vibration test technology of satellite based on vector-fixture’s design and applications

Aiming at the problems that the three-direction uniaxial sequential vibration test cannot effectively simulate the real launch environment of the satellite, and the high cost of the multi-axis shaker and the limitation of engineering application, the multi-axis vibration test scheme based on vector-fixture is proposed. Taking the vibration magnitude of uniaxial vibration test on the satellite mounting surface as equivalent reference, the appropriate vector direction is determined to carry out vector-fixture design. By analyzing the multi-axial vibration test data of the real structure of a satellite, it can be seen that: The “satellite-fixture” mounting surface can reach the set vibration magnitude at low frequencies, and undertest occurs at high frequency due to structural resonance and other factors; By optimizing the control strategy, the degree of over-test and under-test in the three directions can be balanced. According to the mission characteristics and development cycle of micro-nano satellites, this paper provides an efficient vibration test scheme.

Friction properties of the single-crystal Si in collision sliding contacts under different lubrication conditions

With the development of nanotechnology, micro-nano electromechanical systems are widely used in various high-precision instruments, e.g., micro-nano satellites, pico-satellites, in which single-crystal Si is always employed. In outer space, there will be random collisions in space mechanisms caused by unexpected and irregular small disturbances especially under the microgravity environment. In this paper, a molecular dynamics model is used to simulate the motion of the mechanism based on the motion characteristics of random collisions happened within space mechanisms, and irregular collisions are represented by sinusoidal vibrations of the indenter in Y and Z directions. The friction properties of the collision sliding contacts under the lubrication of gold films with different layers are investigated. The results show that a nine-layer gold film presents good vibration absorption and lubrication performance. Then, rectangular textures are designed and made on the gold film. The textured surface can significantly reduce the friction forces at high vibration frequencies. In addition, the friction properties of a molybdenum disulfide (MoS2) film and the combined lubrication behaviors of molybdenum disulfide films and gold films are studied. The molybdenum disulfide film shows excellent friction properties at different vibration frequencies, and the combination of molybdenum disulfide and gold films can significantly reduce the average friction force within a specific vibration frequency range.

On-Orbit Radiometric Calibration of Hyperspectral Sensors on Board Micro-Nano Satellite Constellation Based on RadCalNet Data

The stability and accuracy of the on-orbit radiometric calibration of hyperspectral sensors are prerequisites for the quantitative application of satellite hyperspectral data. The Zhuhai-1 micro-nano satellite constellation is composed of eight hyperspectral satellite missions. The Orbita Hyperspectral Sensor (OHS) on board each satellite has a gradient filter spectroscopic design. When observing the Earth, eight integration stages can be set for each band according to different lighting conditions. Due to high manufacturing costs, OHSs are not equipped with on-board calibration devices. Therefore, it is very difficult to accurately calibrate OHSs for all of the integration stages. On the other hand, it is extremely important to ensure radiometric consistency between different OHSs within the Zhuhai-1 micro-nano satellite constellation. To carry out the rapid radiometric calibration of the Zhuhai-1 constellation, an on-orbit radiometric calibration model considering all of the integration stages related to hyperspectral sensors was built based on the BOA reflectance and atmosphere parameters published by the Committee on Earth Observation Satellites (CEOS) radiometric calibration network (RadCalNet). The RadCalNet product was used to derive the TOA radiance base in the Second Simulation of the Satellite Signal in the Solar Spectrum (6S) radiative transfer (RT) model. In this paper, we analyzed the radiometric stability of the same sensor and the consistency of different calibration results regarding four RadCalNet sites, and the on-orbit radiometric performance evaluation of OHSs was also carried out. The data retrieved from OHSs regarding hyperspectral surface reflectance were preliminarily validated using site-synchronous surface reflectance measurements.

Research on Design, Simulation, and Experiment of Separation Mechanism for Micro-Nano Satellites

The separation mechanism is a critical device that transports and releases satellites during launch and on-orbit. However, as satellites become smaller and more compact, the traditional belt-locking device or pyrotechnic release and separation device cannot meet the micro-nano satellite’s separation requirements. A novel separation mechanism kinematic system was designed, analyzed, and experimentally verified to achieve non-interference and non-pyrotechnic separation of the satellites from the launch vehicle while maintaining the initial separation attitude. First, an overall structural strategy for the kinematics system was proposed based on the MF (product code of the satellite) satellite’s structural properties. The structural characteristics of the separation mechanism were also confirmed by the principle of energy conservation. Then, a finite element model and a dynamic model of the kinematic coupling system between the MF satellite and the separation mechanism during launch and deployment were constructed, along with an analysis of the mechanical characteristics and a kinematic simulation. Finally, the mechanical characteristics and dependability of the separation mechanism were verified using a ground vibration test and a separation test. The prototype test results demonstrate that the separation process is essentially identical to the numerical simulation results and that the MF satellite can be deployed successfully without interference, with the MF satellite’s initial separation attitude fully satisfying the designed technical indexes.

Micro-nano Satellite Phasing Maneuver Method under Microthrust Constraint

The micro-thruster of micro-nano satellite cannot provide a large impulse for a single time to satisfy the requirements of the general phasing maneuver because of the constraints of thrust and the stability of attitude control system. A multi impulses phasing method is proposed based on the four impulses phasing maneuver, which satisfies the constraint of the working time of the micro thruster. In this method, a large impulse is divided into multiple small impulses, and the impulses can be re-optimized under the constraint of phasing time, which can realize the phasing maneuver of micro-nano satellites using microthrusters. The simulation results show that the error of phasing time is less than 30 s, and the phasing accuracy in two-body orbit dynamic is less than 0.2 km. Considering the perturbation and the error of the micro thruster, the phasing accuracy is less than 5 km, which meets the time limit and fuel consumption limit in phasing maneuver of the micro-nano satellite.