To address the threat posed by micrometeoroid impacts to Earth-orbiting satellites, this study conducts an analysis of the near-Earth micrometeoroid environment. Using data from the Meteor Radar of China's Meridian Project, we inverted the mass, velocity, and flux distributions of near-Earth micrometeoroids and compared with the classical Gru & uml;n model. Analysis of the selected meteor radar data indicates that the meteoroid flux is concentrated in the northwest direction (azimuth 270 degrees-360 degrees) and exhibits distinct seasonal fluctuations, with summer activity intensity reaching 2.8 times winter levels. The velocities are concentrated in the 20-40 km s-1, while the heights are concentrated in the 75-125 km. The least-squares polynomial curve fitting method was applied to model the diurnal flux-density relationships and velocity versus number density distributions. Results reveal the improved fitting accuracy at higher polynomial orders, although the discrepancies emerged for velocities exceeding 40 km s-1. Comparisons between the inverted fluxes and model predictions indicate there is good agreement for larger meteoroids, but significant deviation for smaller masses, potentially attributable to the limited observational samples. The radar directly detects meteor trail echoes in the similar to 70-110 km region, while the meteoroid states/ fluxes discussed at higher altitudes are model-inferred quantities extrapolated from the detected layer and should be interpreted as preliminary. (c) 2026 Published by Elsevier B.V. on behalf of COSPAR.

The rapid development of low‐Earth orbit (LEO) satellites brings increased attention to spacecraft collisions, space debris, orbital decay, and satellite reentry. Neutral density and associated drag force on the satellite orbits elevate space risks, significantly determined by space weather disturbances, particularly geomagnetic storms. On 3 September 2024, the Australian Binar‐2, 3, and 4 satellites were deployed, but their actual lifetimes were only around 2 months, much shorter than designed, only 20% that of the nearly identical Binar 1 satellite launched in 2022. For the first time, we analyze the premature reentries of the Binar‐2, 3, and 4 satellites to unveil the severe space weather impact on their lifetimes, especially the influence of accuracy of medium‐ and long‐term space weather prediction on satellite lifetime designs. Our findings reveal that the premature reentries of Binar‐2, 3, and 4 satellites were caused by enhanced neutral density due to much higher solar and geomagnetic activities than predicted. The actual satellite lifetimes align well with estimations based on observed space parameters, whereas large deviations occur when using predicted parameters. The widely used predictions for medium‐ and long‐term space weather underestimated the F10.7 index and sunspot numbers during Solar Cycle 25 (SC25), especially during the peak time, leading to discrepancies in the designed satellite lifetimes. Our results illustrate the importance of medium‐to long‐term space weather forecasting for the lifespan of LEO satellites.

AbstractSatellite drag coefficients are crucial for determining the neutral mass densities that affect spacecraft operations in the thermosphere. Many studies typically utilize a constant drag coefficient of 2.2 to calculate the neutral density. However, due to the variability of space environment, uncertainties in the drag coefficient can lead to significant systematic discrepancies in neutral density measurements. Satellite drag coefficient may fluctuate in the thermosphere under various geomagnetic activities and altitudes. For the first time, we calculate the spherical satellite drag coefficient using data from the “Orbital Atmospheric Density Detection Experimental Satellite,” referred to as the QX satellite. Our findings reveal that the drag coefficient can be estimated by thermospheric temperature and density, which are dependent on geomagnetic activity and altitude. At an altitude of ∼510 km, drag coefficients are adjusted to around 2.425, instead of the constant value of 2.2. Furthermore, the drag coefficient may decrease due to the significant influence of increasing geomagnetic activity, such as geomagnetic storms, on thermospheric density and temperature. These estimates of the drag coefficient can also be used to reduce discrepancies when deducing the ballistic coefficient. Consequently, using the estimated drag coefficient can accurately determine the QX‐derived neutral density, which agrees well with the density from Swarm‐B satellite.

On November 15, 2021, Russian dead satellite Cosmos 1408 broke up, which attracted widespread attention from media and researchers both domestic and foreign. This paper analyzed this event. According to TLEs released by US Space Surveillance Network, the orbit change of the parent body was analyzed. By using the orbital elements of breakup debris, the information such as breakup time and intensity was preliminarily determined. Finally, two spacecraft on LEO were selected to carry out the collision risk assessment. The impact of Cosmos 1408 breakup debris (CBD) on the international space station (ISS), and the China space station (CSS) was analyzed. The results show that the impact risk of the CBD on the ISS and CSS could not be ignored.

In recent years,the large number of deployments of low-orbit giant constellation satellites have a significant impact on the operational safety of low-orbit satellites in orbit,and it is critical to de-tect orbital anomalies of constellation satellites in time.As a result,this study picks the Starlink satel-lite constellation as the research objects for detecting orbit anomaly and presents an improved orbit anomaly detection approach-segmentation optimization.Based on the concept of dynamic optimization,the approach improves the orbit anomaly detection method by using satellite orbit semi-major axis data as an analysis parameter,transforming it from threshold screening to optimization search.First,by as-suming the segmentation points and randomly distributing them across the entire semi-major axis data series,the data in the time window are randomly segmented.Based on the variance of the segmented da-ta,each segment's loss function is built.The iterative function are then designed using the loss function for optimization iterations.In order to determine the optimal segmentation method,the random segmen-tation is finally optimized with the goal of minimizing the sum of the total loss functions.This research finds that the anomaly detection effect is the best for the semi-major axis data by the differential pro-cessing after evaluating a range of data.The segmentation optimization approach has various sensitivi-ties to different data after removing the noisy data.In conclusion,this work uses the TLE(Two-line ele-ment)data and the ephemeris data from the Starlink satellite for example verification,which was launched on 28 February 2023.The method's effectiveness in detecting orbital anomaly of constellation satellites is demonstrated,which are simple and efficient.

Interplanetary meteoroids and space debris can impact satellites orbiting the Earth or spacecraft traveling to the Moon. Targeting China Space Station（CSS）, 7 satellites selected from the constellation of Beidou Navigation Satellite System Phase Ⅲ（BDS-3）, and 3 spacecraft orbiting the Moon, we have adopted in the paper the Meteoroid Engineering Model 3, Divine-Staubach meteoroid environment model, and Jenniskens-Mc Bride meteoroid steam model to analyze the meteoroid environment with the mass range of 10 –6 ～10 g. Orbital Debris Engineering Model 3.1 space debris model is used to analyze the orbital debris environment faced by these satellites.The flux of space debris with a size larger than 100 μm is compared with that of the meteoroids. The results show that the space debris flux encountered by China Space Station is much higher than that of the meteoroids with sizes in the above range. And quite the opposite, the meteoroids flux impacting the 7 satellites from the BDS-3 is higher.Upon adopting the double-layer Whipple protection measure, the catastrophic collision flux of these satellites encountering meteoroids is about 10 –6 times of that without protection, or even less, implying that the Whipple protection effectively guarantees the safety of the satellites in orbit. Besides, it is also found that the flux of the highdensity meteoroid population encountered by each satellite is greater than that of the low-density population,whereas the impact velocity is lower for each satellite. These results can aid the orbit selection and the protection design for satellites and spacecraft.

根据热层物理、经验模型原理和代码分析，研究模型构建方法。进而剖析国内热层大气探测和热层模型构建现状，提出存在的困难和未来的发展思路。以GOST模型为基础，分析模型的工作机制、地磁扰动期大气密度预报误差来源和密切相关的模型系数，推导模型密度对相关模型系数的偏导数矩阵。利用天基实测密度，有针对性地构建磁暴期大气模型(DAM)。并通过独立于建模的实测密度数据，验证DAM模型性能。统计发现，地磁活动指数Ap介于100～132时GOST, MSIS00和DAM模型的相对误差均值依次为64.32%，–176.72%，–14.83%。Ap指数80～132时，相对误差均值对应为77.44%，–136.74%，–14.14%，DAM模型性能较GOST和MSIS00均有明显提升。证明通过搭建大气模型框架和实测密度数据估计模型参数的建模方法是可行和有效的。

The Space Radiation Threat Sensor is one of a payload of CX-12(01), an imaging satellite operating in medium earth orbit satellite. A key component of the payload is the space particle detector, which is used for detecting space protons, electrons and a-particles. In this paper, the conceptual design of the detector was presented and energy deposition of particles was simulated by Geant4 Monte Carlo tool. Results show that the system meets the detection requirements and can successfully identify protons in 6~500 MeV, electrons in 0.5~10 MeV, and a-particles in 25~150 MeV.

1 "星链"事件始末 据SpaceX公司官网报道,世界标准时间2022年 2 月 3 日 18:13,SpaceX 公 司 使 用 猎 鹰 -9 (Falcon-9)火箭从佛罗里达州肯尼迪航天中心的LC-39A发射场向近地轨道发射了49颗"星链"卫星,卫星顺利进入近地点210km、远地点340km的轨道.按照预定流程,卫星入轨后会停泊在近地点210km的极低轨道,进行一系列在轨测试,确认卫星状态良好后,才会启动电推进器进行升轨.然而,卫星在2月4日遭遇地磁暴,据SpaceX公司报道,相比以往发射,本次任务大气阻力增加了50％.虽然SpaceX公司命令卫星进入安全模式,转为侧面迎风的姿态飞行,以最大限度减少阻力,但该策略并未奏效,低空增加的阻力使卫星离开安全模式,导致多颗卫星无法顺利升轨.

The wide application of satellite constellations in the field of space-based global communications and remote sensing has led to a substantial increase in small-satellite launch plans, a sharp increase in the density of space objects in low-Earth orbit (LEO), and a reduction in available orbit and frequency resources. This will further aggravate the trend of deterioration of the space debris environment. Taking the Starlink constellation as an example, this paper describes the influence of the constellation from the environmental debris flux of the satellite, the evaluation of the number of evasion maneuvers, the change of risk level, the success rate of post mission disposal (PMD) and the growth rate of space objects. The simulation results show that the collision risk of the Starlink constellation is related to the orbital parameters, and the higher success rate of post-mission disposal (PMD) can reduce the collision risk of the constellation. The large constellations increases the growth rate of space objects, and even if all the satellites are disposed of after the mission, the impact of constellations on the space environment can not be offset.

In predicting the collision of space debris, the propagated orbital uncertainty may not follow a Gaussian distribution if the initial orbital uncertainty is large or the propagation time is long. In this paper, a Gaussian mixture uncertainty propagation method developed by (Psiaki et al., 2015) is used to calculate the collision probability. The initial Gaussian distribution is fitted by the weighted Gaussian mixture components. The linear matrix inequality is optimized to prevent the covariance matrix of Gaussian mixture components from being too small, and an appropriate number of Gaussian mixture components is used to approximate the initial orbital covariance. At the same time, this paper provides a method to calculate the collision probability of two objects in which a Gaussian mixture is used to represent the distribution of orbital uncertainty. The linear method and the unscented Kalman filter (UKF) method for propagating the Gaussian covariance are analysed. The results of numerical simulations show that compared with the linear covariance propagation method, UKF method, and high-precision Monte Carlo covariance propagation method for space objects with a large initial orbital uncertainty, the Gaussian mixture method can be effectively applied to capture the non-Gaussian characteristics of the predicted non-linear orbital dynamic uncertainty. Compared with the univariate splitting method, the advantage of this Gaussian mixture method is that it does not need to search for the most nonlinear direction. The accuracy of the collision probability calculation is improved from 1.460 x 10(-3) to 1.663 x 10(-3). A comparison of the computational burden between the Gaussian mixture algorithm and Vittaldev's algo-rithm to achieve the same results is presented. The calculation burden of the Gaussian mixture method is approximately 3 times that of the univariate Gaussian method. (C) 2021 COSPAR. Published by Elsevier B.V.

In this paper, optical target monitoring orbit determination using the Gooding algorithm based on the 400 km altitude space platform was researched. The measurement error was set to 3″ and 6″ respectively for 800 km, 1500 km altitude low Earth orbits and 36 000 km altitude geosynchronous orbit to determine the initial orbit and precise orbit. Simulation results show that the initial orbit determination accuracy of the 4~15 min arc is about 10 km, and the 1~2 min arc is about 100 km when the measurement data error is between 3″ and 6″. The error of 15 min initial determination arc is 100 m. As the arc is less than 10 min, the improved error accuracy of the orbit is in the order of km scale. As the measurement data error is 3″, the initial orbit determination accuracy of the 15~20 min arc is about tens of kilometers, and that of the 8~10 min arc is 100 km. The improved orbit has an error of km. When the measurement data error is 6″, the accuracy of the initial orbit determination of the 20 min arc is at the magnitude of 10 km, the accuracy of the initial orbit determination of the 8~15 min arc is at the magnitude of 100 km, and the error accuracy of the improved orbit is at the magnitude of 10 km.

Collision warning and avoidance is the main method of the satellite risk mitigation to the catalog debris. The warning precision and confidence are the main problems of the current collision warning work. This paper first put forward the problems on the collision warning work which include the data, model and error cope methods, and then introduced the collision warning work flow and development in NSSC (National Space Science Center, Chinese Academy of Sciences). Some refined work were introduced from two aspects: the first is to improve orbit prediction precision, such as the TLE precision improvement, drag coefficient analysis, the space environment effect to the prediction accuracy; The other is to refine the collision probability computation, such as the covariance analysis. These studies could improve the warning confidence to some extent.

目的 基于硬质落球,分析探测临近空间大气密度和水平风场的载荷需求和探测性能.方法 仿真计算自由下落探测段受力情况,建立运动方程.以假定释放条件,仿真落球的探测过程,给出各作用力量级.结果 面质比的变化不能改变气动力范围.相同加计灵敏度下,增大面质比可提高探测高度.150 km以下的探测中可忽略光压力,探测到水平风场米级变化的加计灵敏度需求为1×10-5m/s2.结论 面质比的变化不能改变气动力范围.相同加计灵敏度下,增大面质比可提高探测高度.

Currently, more than 19,000 pieces of debris with a diameter exceeding 10 cm have been catalogued around the Earth. Consequently, collision warning systems are of great significance, especially for low Earth orbit (LEO) space assets. For an encounter between two objects, the US Joint Space Operations Center (JSpOC) can compute and report a value commonly referred to as a collision probability, which can be used by a satellite owner/operator as a warning and may motivate a personal investigation of the collision. Other countries, including China, have also established collision warning programmes to protect their spacecrafts. The collision probability is an accepted and calibrated parameter for evaluating the collision between two space objects. (Foster and Estes 1992), (Patera 2001&2005), (Chan 1997&2004&2003), and (Alfano 2005) have developed methods for calculating the collision probability. Foster derived the collision probability by applying polar coordinates in the encounter plane, which is perpendicular to the relative velocity of the two objects; hence, in this approach, the probability density function (PDF) is expressed as a polar parameter. This collision probability model is currently applied by NASA to assess the on-orbit collision risk for the International Space Station. Chan used series expansion to analytically approximate the collision PDF; in this technique, the three-dimensional Gaussian probability density function is transformed into a two-dimensional PDF by projecting the Gaussian covariance onto the encounter plane. This model is currently applied in the Analytical Graphics, Inc., Satellite Tool Kit. A mathematically equivalent expression was developed by Patera to represent the collision PDF, and this method is implemented within the Collision Vision Tool by the Aerospace Corporation. Alfano’s method employs series expansion to represent the collision PDF. (Alfano 2009) applied a simplified Monte Carlo process based on a two-body analytical propagation to assess the satellite collision probability computations with different methods. Alfano provided 12 test cases involving linear and nonlinear relative motion for satellites in geosynchronous orbit (GEO) and LEO, as well as nonlinear relative motion for highly eccentric orbit (HEO), and compared the accuracy of each method with that of the Monte Carlo method. In addition to the methods mentioned above, many scholars have employed the Monte Carlo method to calculate the collision probability. (Sabol, Binz et al. 2011) used a special perturbation-based Monte Carlo method to investigate approaches for estimating the probability of collision between two satellites. Comparisons were made against an analytical method and a two-body Monte Carlo method for LEO and GEO satellites. (de Vries and Phillion 2010) calculated the collision

在太阳活动高低年的地磁平静/扰动环境下，利用不同热层大气模式J77，DTM78，MSIS00，JB2008和CHAMP加速度计反演密度，分析有无先验信息条件下的轨道预报误差.结果表明无先验信息的精密轨道预报中，热层模式的性能可能被弹道系数等参数偏差干扰，此时预报误差不能作为模式性能的评价标准.先验信息对轨道预报精度提升非常明显，尤其是地磁扰动期先进热层模式性能得以展现，轨道预报误差为无先验信息情况下的10%～25%.目前热层模式的主要缺陷存在于地磁扰动期.各模式之间的差异是：JB2008模式可以通过线性和单一频率周期项补偿，而J77及DTM78等模式还存在更多频率的误差.本文对不同情况下精密轨道预报的研究结果可为空间碎片碰撞预警等工程实践提供参考.

分析Jacchia70（J70）热层模式原理、美国空军高精度卫星拖曳模型（HASDM）的修正方法及选取球面调和函数的原因.推导模式密度对球谐系数（SH）的偏导数，给出利用模式密度泰勒展开进行线性化处理、迭代求解球谐系数的具体过程.针对2003年10月29日大磁暴事件，基于CHAMP和GRACE A/B卫星加速度计实测密度，进行修正方法的性能评估.统计相对误差的均值及标准差变化：改进前分别为-81.7%，74.4%；改进后分别为-5.9%，53.1%，验证了改进算法的有效性.从热层上下边界温度角度，详细分析了热层模式动态修正原理，研究结果为类HASDM修正模式的工程应用提供了理论基础.

In the process of space object cataloging, because the observation data of new object is sparse, the orbit determination error is large and the orbit correlation success rate of fixed threshold method is low. This paper aims at associate tracks that cannot be associated with the object database through the orbit covariance propagation and the Mahalanobis distance dynamic association method. The simulation process of UCTs correlation includes the calculation of orbital ephemeris, the simulation of observation data, orbit determination and covariance propagation, and the dynamic correlation based on Mahalanobis distance. Covariance propagation is the main process of UCTs association. In this paper, a covariance generation and propagation algorithm based on linear and unscented Kalman filter (UKF) is established. Jacobian transform and UKF method were applied to space object orbital covariance propagation. The model of covariance propagation is applied to the project of uncorrelated track dynamic association. Through the model of covariance propagation and the uncorrelated track dynamic association method, a statistical measure called k is computed to measure how closely the two state correlate. 20 of 104 satellites which launched by India are simulated as UCTs to correlate, when the threshold of k is set as 200, all of the 200 satellites can be correlated. But as the interval between two tracks increases, the value of k decrease. With the same interval, the value of k of different satellite is different.

In this paper,ant colony optimization algorithm is used to study the path optimization in the removal progress of multi-debris.The modified steepest descent method is used to optimize the time of removing each piece of debris reasonably which further reduces the aggregate requirements of velocity increment.The comparisons with the order of orbit height,inclination or RAAN (Right Ascension of the Ascending Node) indicate that using the order after the optimization of the ant colony algorithm can greatly save the velocity increment.Three groups of debris produced in Chinese activities are chosen for optimization.Results show that the optimized order may be different from three kinds of order mentioned above,and the optimized order can save more velocity increment in the same task time.In addition,the task time also has an impact on the best removal order of debris.

碰撞能量转移方案的基本思路是通过准弹性碰撞将碎片携带的部分能量转移到航天器,在完成碎片降轨任务的同时,推动航天器完成轨道转移,移除更高轨道的碎片.进一步,采用气动力辅助变轨方式调整航天器轨道参数,促使航天器与目标碎片多次碰撞,从而减小碰撞速度差.首先介绍了利用碰撞能量转移方案移除碎片的航天器的结构组成、工作机理和关键技术,然后对影响对心碰撞和非对心碰撞的因素进行分析研究,找出对航天器携带弹性捕网的性能要求低、适合多次碰撞的区域.